FI74253B - VARNINGSSYSTEM FOER TAKTISKT FLYGPLAN. - Google Patents

Description

1 74253

Warning system for tactical aircraft

The present invention is directed to a system for assigning to a pilot of a tactical flight phase an altitude at which a corrective action must be initiated to prevent a fall below a predetermined altitude selected by the MDA or radio altitude "alarm" setting. The system also provides a warning of dangerous flight situations during non-tactical or routine parts of the flight, such as take-off, landing and cruising.

Ground-based warning systems are already known to provide warnings for various hazardous flight profiles. However, these systems 15 are generally designed for commercial aircraft and are designed to provide warnings primarily during the take-off and landing phases of a flight to prevent inadvertent off-road guided flight. Such systems are not designed to provide warnings during tactical movements, such as crashes, when an airplane is deliberately flown to the ground at high speed and at a steep angle, and thus provide inadequate or false warnings in such situations.

Thus, it is an object of the present invention to provide a warning system which overcomes many of the disadvantages of the prior art warning systems.

It is a further object of the present invention to provide a warning of a dangerous flight situation during tactical movements as well as during normal routine parts of the flight, such as take-off, landing and cruising.

It is a further object of the present invention to provide the pilot of a crashed aircraft with a warning of a corrective action that should be immediately initiated to prevent the aircraft from going below the desired minimum altitude.

2 74253

Tactical airplanes bring with them special problems in the design of ground-based warning systems that warn the pilot of an airplane of a dangerous flight situation. One such problem is due to the dual nature of the operation of the aircraft 5 in question. Typically, the airplane in question has a tactical or tuned mode that includes movements such as crashes and multiple low altitude movements, as well as a non-tactical or normal routine mode of operation that includes movements such as take-off, landing, and cruising. Because the movements in the tactical mode of operation of an airplane are substantially different from those typically in the non-tactical mode, a system designed for a non-tactical or commercial aircraft will provide adequate warnings during certain tactical movements while providing false warnings during other movements. Similarly, systems specifically designed to provide warnings during tactical movements do not provide adequate warnings during less demanding non-tactical movements of 20 flights.

The set objects are achieved and the above-mentioned drawbacks are avoided by the system according to the invention, which is characterized by the features set forth in the characterizing part of claim 1.

The system of the invention monitors the angle of inclination of the aircraft and the rate of change of pressure altitude in the tactical mode of operation and calculates the altitude at which the corrective action should be initiated to prevent the aircraft from falling below the desired mini-30 altitude selected by the MDA or radio altitude "alarm" setting. When the minimum height of the corrective action is reached, the warning system issues a special audible warning that is clearly distinguishable from non-tactical warnings, such as "ABORT", to indicate that the corrective action must be initiated immediately. Tactical warnings are only allowed during the tactical phases of the flight and certain non-tactical warnings are blocked during the tactical phases.

These and other objects and advantages of the present invention will become readily apparent from the following detailed description and accompanying drawings, in which: Figure 1 shows a logic block diagram of a warning system according to the invention;

Figure 2 shows the aircraft in a crash and illustrates the dynamics of the crash correction motion;

Fig. 3 is a diagram showing the loss of altitude of the aircraft 10 during a 4G correction motion due to a crash as a function of the rate of descent and the angle of inclination;

Fig. 4 is a diagram showing the loss of aircraft altitude during a 4G crash correction motion as a function of airspeed and longitudinal tilt angle; and

Figure 5 shows a block diagram of a proximity warning system according to the invention that can be used in both tactical and non-tactical movements.

Referring now to the drawings, and in particular to Figure 1, there is shown an embodiment of a tactical warning component for a proximity warning system according to the invention, generally designated 10. The system 10 of the invention is shown in Figure 1 in the form of a logic block diagram for gates, comparators 25 and the like; however, it is to be understood that the actual implementation of the logic may be different from that shown in Figure 1, with several analog and digital implementations possible. The signals used by the described system include the radio altitude, the rate of change of the pressure altitude, the airspeed and the longitudinal tilt angle of the aircraft, as well as signals indicating whether the armament is tuned and multiple eligibility signals. Depending on the type of aircraft on which the warning system is installed, the signals shown in Figure 1 can be obtained from private measuring devices such as altimeter 12, altitude rate circuit 14, radio altimeter 74253, vertical vortex 18, airspeed signal source 20, or ADC discrete circuit elements indicating whether the armament is tuned and whether the wheels have weight. Alternatively, these signals may be obtained from the digital data bus on certain newer aircraft.

The tactical warning part of the system according to the invention is designed to operate only during the tactical phase of the flight, during a bombing attack or a ground combat attack, in which case a situation of a sharp fall occurs. Thus, the system is allowed only when the conditions indicating that stage of operation occur. In the present embodiment, these situations are that the weight of the aircraft is off the wheels, that the airspeed is greater than a predetermined airspeed, such as 300 knots (562 km / h), and that the armament is tuned.

The enabling operation is performed by an AND gate 22 which receives signals indicating that the armament is tuned and that the weight is off the wheels and a signal comparator 24 connected to the airspeed signal source 20. The comparator 24 compares the signal from the airspeed signal source 20 to a reference signal which represents a predetermined airspeed, for example 300 knots (562 km / h) and indicates when the airplane airspeed exceeds 25 a predetermined airspeed, for example 300 knots (562 km / h). In addition, the AND gate 22 receives signals indicating that the radio altimeter is not blocked and that the pressure altimeter is not blocked to block the system when the meter is defective.

When all of the above conditions are met, the gate 22 allows a second AND gate 26 to be brought under the control of a comparator 28 which receives from the vertical vortex 8 a signal representing the aircraft angle of inclination, a signal representing the aircraft altitude change rate no-35 from the speed circuit 13, and a signal represents the minimum descent height set by the MDA setting. These input signals

J

5 Based on 74253, the comparator 28 determines the minimum correction altitude or warning altitude, Hw, at which the corrective action must be initiated to prevent the aircraft from falling below the minimum descent altitude. The comparator also receives a signal 5 representing the altitude of the aircraft from the ground altimeter 16 and provides a signal to the gate 26 in the event that the signal received from the altimeter 16 represents an altitude below the minimum correction altitude. When this signal is received, the gate 26 starts the operation of the warning generator, preferably the digital sound generator, thus causing the generator 30 to generate an audible warning. The audible warning shall be a special unambiguous warning, such as "ABORT", to indicate to the pilot exactly what action to take to avoid a dangerous flight situation. An audible warning is input either directly or indirectly to the transducer 32, such as a loudspeaker or earphone, to inform the instructor.

In addition to the speed and length of the aircraft, there are 20 other factors that determine the altitude at which a crash correction warning should be allowed to develop. These factors relate to aircraft performance and include the response time of the aircraft and the pilot and the value of G that the aircraft can develop during take-off. In a conventional repair motion, the flight path approximates the arc of the circle at the bottom of the take-off motion. The radius of the circle is determined by the value of G that the airplane can develop during the take-off motion, the radius of curvature of the curve being inversely proportional to the value of G developed.

Referring to Figure 2, an airplane 50 is shown which has just passed the minimum height of the repair warning, Hw, and begins its repair movement. Assuming that the correction is satisfactory, the aircraft turns horizontally to the minimum descent altitude level MDA and begins to ascend, as shown by dashed line 35. The correction path is a circle in its first approximation, and the radius R of the circle is determined by the value of G that the aircraft can develop during the correction motion.

6 74253

As the object travels in a circle, the normal acceleration, i.e., the radial, inward acceleration, is directly proportional to the square of the velocity of the object and inversely proportional to the radius of the circle traversed by the object. Mathematically defined, this ratio is: (D a = - n r where an is the inward normal acceleration to which the object is subjected during the revolution, V is the velocity of the object and R is the radius of curvature of the object's curve.

If the angle of descent of the aircraft 50, represented by the angle of descent or the approximation of the angle of inclination, is equal to Θ, the rate of descent of the aircraft 15 can easily be related to the speed of the aircraft using simple trigonometric functions:

Hb (2) V = -—F- K ’Sin Θ 20 where is the rate of descent of the airplane, more specifically the rate of change of pressure altitude. In the illustrated embodiment of the invention, the rate of change of pressure head is used in the calculation; however, the terrain approach rate or radio altitude change rate may be used in certain embodiments.

If the angle of inclination of the aircraft is Θ, then, according to geometrical principles, there is also the radius sen perpendicular to the flight path of the aircraft and the vertical angle kul (Figure 2). Thus, the distance between the center 52 of the circle 30 and the warning height Hw is equal to R cos Θ and the distance between the warning height Hw and the minimum descent height MDA is equal to R (l-cos0). The latter ratio allows the amount of altitude loss AH to be calculated between the warning altitude Hw and the minimum descent altitude 35 MDA as a function of the angle of inclination of the airplane and the radius R of the circle traveled by the airplane.

7 74253

Since the radius of the correction curve is a function of the airplane speed and the value of G that it can develop during take-off, the altitude loss between the warning altitude and the MDA can be calculated from the airplane speed and the value of G it can develop during take-off. In addition, since the rate of descent of the aircraft is a function of the speed of the aircraft and the angle of descent, the loss of altitude & H during the correction operation can be calculated as a function of the rate of descent and the value of G that it can generate ascending.

This is done as follows. Take the ratio (2), ne-lute it and make the appropriate trigonometric positions and obtain the following ratio: 2 Hh2 Hh2 Hk2 15 (3) V = -S- = - £ - = _ £ _

Sin29 (l-cos ^ 9) (l + cos9) (l-cos9) By rearranging the ratio (1) and placing the ratio (3) we get: 20 V2 H.2 (4) R = - = -S- an an ( l + cos9) (l-cos9)

Multiplying the ratio (4) by a factor (l-cos9) gives the height loss ΔΗ during the pull-up as follows:

Hb2 (5) H = - £ - an (l + cos9)

The latter ratio determines the loss of airplane altitude between the moment when the crash correction is initiated and the moment when the airplane reaches the horizontal flight portion; however, it does not take into account the reaction time of the airplane and the pilot and the minimum desired altitude has not been considered. These factors must also be taken into account in the equation determining the warning height Hw in order to ensure that the warning is given in sufficient time to enable the pilot to carry out the correction safely. The minimum descent height can be easily taken into account by adding the MDA setting to the height loss. The reaction time can be taken into account by adding to the equation a term equal to 5 times the reaction time of the pilot and the aircraft, for example two seconds, multiplied by the rate of descent. Thus, the warning height Hw is obtained from the following equation: H 2

(6) Hw = -h- + 2Hb + MDA

10 an (l + cos9) where the factor describing the Hb term represents the two-second reaction time of the pilot and the airplane.

The above equation can be used to determine the warning height Hw for several types of airplanes with several different performance characteristics. However, 4G pull-up is typically used during tactical movements because it does not expose the pilot to excessive strain. Thus, using 4G as the normal acceleration an, and solving the above equation for 4G correction and a reaction time of two seconds using numerical technique, the following result is obtained:

Hw = (Hb) 2 · (0,005175 + 0,0000236 · (θρ-11) + 0,0000321 * (θρ-40) + 2Hb + MDA

25 where:

Hw is the radio altitude warning height in feet Hb is the rate of change of pressure head in feet per second θρ is the longitudinal tilt angle in degrees 30 MDA is the desired minimum height in feet.

In the above equation, the terms (Θ -11) and (Θ -40) r r must not be negative. Thus, if the longitudinal tilt angle is less than 11 degrees or less than 40 degrees, respectively, the terms (θρ-11) and (θρ-40) are set to zero.

35 In addition, the angle of inclination of the airplane is used instead of the angle of inclination to represent the angle of descent, since the angle of inclination of the longitudinal inclination is easily obtained from a vertical vortex or the like and the angle of inclination is not; however, the actual pitch angle can be used in the calculation. The angle of inclination differs from the angle of inclination of the length by the angle of attack of the wings, which is typically only a few degrees. Thus, using a longitudinal tilt angle instead of a fall angle in the calculation provides a satisfactory approximation of the warning height. Similarly, a numerical approximation of the trigonometric equation determining the warning height is used because it allows the warning system of the invention to perform simple multiplications instead of much more complex trigonometric calculations, thus providing a faster response time.

The crash correction correction model for the crash of the warning system according to the invention for the correction of 4G is shown in Fig. 3. Fig. 3 shows the height loss ΔΗ as a function of the descent rate for several longitudinal tilt angles. As can be seen from Fig. 3, the height loss AH and thus the warning height Hw increases as the terrain approach speed increases and the longitudinal tilt angle increases. However, since the cosine function changes slowly at small angles, the curve of Figure 3 is relatively independent of the longitudinal tilt angle at smaller longitudinal tilt angles, i.e., at angles of inclination of less than 40 degrees. Thus, the same curve can be used to define a warning limit for longitudinal tilt angles of about 40 degrees and less.

In addition, as previously mentioned, are the rate of descent and the velocity of the air relative to the sinusoidal sine. Thus, the warning pattern can also be determined instead of the air velocity termeil-30 descent rate, as shown in Fig. 4, which shows the altitude loss ΔΗ as a function of air velocity and longitudinal tilt angle. Thus, in an alternative embodiment of the present invention, the air velocity signal may be fed to the comparator 35 instead of the descent rate 28, and the curves of Figure 4 may be used instead of the curves of Figure 3 to determine the altitude loss and warning altitude.

10 74253

Referring to Figure 5, a portion of a crash correction warning system utilized in conjunction with systems that generate dangerous flight situation warnings during multiple non-tactical flight situations is shown.

In the block diagram of Fig. 5, the same numbers denote the same components as in Fig. 1 and the comma-numbered numbers are used to indicate the non-tactical part of the warning system, which are analogous to the components of the tactical part of the system denoted by corresponding non-comma numbers. Thus, the non-tactical portion of the warning system 10 'is analogous to 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 'shown in Figure 5 also utilizes a plurality of signals representing a plurality of flight parameters of the aircraft, such as airspeed, radio altitude, pressure altitude change rate, radio altitude change rate, signal representing minimum descent altitude MDA, signal representing armament signal, arming signal , which represents the position of the landing gear as well as a plurality of qualification signals (not shown). As with system 10, system 10 'receives signals from either a plurality of individual meters 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 of a dangerous flight situation. Typical tasks performed by the system 10 'include those described in filed patent applications filed at the same time as the present application and invented by the same inventor as the present invention, the designations of which are "NEGATIVE CLIMB AFTER TAKE-OFF WARNING SYSTEM". "EXCESSIVE TERRAIN CLOSURE WARNING SYSTEM" ("System for ALERTING A PILOT OF A 35 DANGEROUS FLIGHT PROFILE DURING LOW LEVEL MANEUVERING"), "EXCESSIVE TERRAIN CLOSURE WARNING SYSTEM" ("EXCESSIVE TERRAIN CLOSURE WARNING SYSTEM") "EXCESSIVE DESCENT RATE WARNING SYSTEM FOR TACTICAL AIRDRAFT", and "WARNING SYSTEM FOR AIRCRAFT LANDING WITH LANDING GEAR UP" ("Above the landing gear"). warning system when the aircraft lands the country "). All of the above patent applications are the subject of the present invention and are incorporated herein by reference.

The system 10 'may include one or more of the warning functions set forth in the aforementioned applications or other warning functions and operates during the non-tactical phases of aircraft operation. The system 15 10 'controls the operation of the non-tactical warning generator 30' and causes the generator 30 'to generate a plurality of audible warnings input to the converter 32 or other suitable transducer to inform the pilot of a dangerous flight situation during the non-tactical phases of flight operations. Typical of those 20 warnings are "DON'T SINK", which indicates to the pilot that he is sinking after take-off, "TOO LOW", which indicates to the pilot that he is flying below the minimum descent altitude, "TERRAIN", which warns the pilot of excessive terrain. proximity during a low ta-25 flight, "SINK RATE," which indicates to the pilot that he is sinking too fast when approaching for landing, and "TOO LOW GEAR," which warns the pilot as he attempts to land on the landing gear at the top.

While all of these warnings are useful during the non-tactical phases of a flight, some of these warnings may be misdeveloped during the tactical phases of flight, thus causing harmful damage. In particular, the "TERRAIN" warning and the "SINK RATE" warning are likely to be developed during tactical movements because both the rate of change of ra-35 diode height and the rate of change of pressure altitude are high during tactical movements. Thus, those non-tactical states that are more likely to be misdone are prevented during tactical movements.

Thus, the system 10 'has an unarmed arming input that prevents the development of warnings, such as "TERRAIN" and "DON'T SINK" warnings, for example, and other warnings that could cause the development of adverse warnings during tactical movements. This signal can be easily obtained from the armament tuning switch 40 used to control the aircraft armament vi-10. This switch can be used to provide an ARMING NOT TUNED signal to the system 10 'to prevent any warnings that could be detrimental during tactical flight and to allow tactical warning system 10 during the tactical phases of flight.

i

Claims (8)

A system by which a pilot in a tactical flight phase aircraft is indicated at an altitude at which a corrective action to prevent submergence below a predetermined altitude must be started, characterized in that it comprises a comparison device (28), which on the basis of the aircraft's sinking speed (14) and angle of incidence (18) determine a minimum height at which corrective measures to prevent submergence below a predetermined altitude must be initiated so that the aircraft does not come too close to the ground, and a first warning device (30, 32); as on the basis of the altitude of the aircraft (16) and the comparator (28), the pilot of the aircraft warns when the corrective action must be started. 2. A system according to claim 1, characterized in that the comparator (28) responds to the barometric sinking velocity of the aircraft. 3. A system according to claim 1, characterized in that the comparator (28) responds to the aircraft's longitudinal angle of incidence. System according to claim 1, characterized in that the first warning device (30,32) responds to the height of the aircraft above the ground surface. The system according to claim 1, characterized in that it further comprises a second warning device (10 ', 30') which, on the basis of the aircraft flight conditions during non-tactical flight phases, generates a warning indicating a hazardous flight condition. 6. A system according to claim 5, characterized in that it comprises a device which, on the basis of the weapon's readiness signal (40), prevents the warning given by the second warning device where the aircraft is in tactical flight phase.