WIND SHEAR RECOVERY GUIDANCE SYSTEM WITH STALL PROTECTION

Abstract: A wind shear recovery system (10) for an aircraft utilizes inertially derived pitch information to provide the pilot with information (12) defining the optimum pitch angle for maximum climb during a wind shear condition. The system (10) utilizes a pitch reference modulator (22) that receives a stall warning discrete from a stall warning system (16) to reduce the commanded pitch guidance angle (12) upon the occurrence of a stall warning to reduce the possibility of stalling the aircraft during degraded performance conditions such as tail winds and engine-out conditions. The system utilizes inertially derived pitch information rather than air mass derived angle of attack information to avoid transients in this angle of attack vane signal that are caused by air turbulence encountered during wind shear, and that can cause pilot induced oscillations about the pitch axis.

Description: WIND SHEAR RECOVERY GUIDANCE SYSTEM WITH STALL PROTECTION 3ACKGROUND OF THE INVENTION This invention relates generally to aircraft wind shear recovery guidance systems, and more parti¬ cularly to wind shear recovery systems with stall protection. PRIOR ART Aircraft guidance systems for guiding a pilot during a wind shear condition are known. One such system is disclosed in United States Patent No. 4,347,572 to Berwick, Jr., et al. The Berwick, Jr., et al. system utilizes various inputs including air mass derived signals such as the angle of attack signal derived from an angle of attack vane to provide the pilot of an aircraft with a guidance signal defining the maximum climb rate obtainable without stalling the aircraft. However, while the system disclosed in the Berwick, Jr., et al. patent does provide guidance to a pilot, because the system utilizes air mass de¬ rived signals, it is responsive to turbulence often encountered in a wind shear condition and such turbu¬ lence can cause an oscillatory guidance signal, which, if responded to by the pilot, can cause pilot induced oscillation of the aircraft about its pitch axis. SUMMARY OF THE INVENTION Accordingly, it is an object of the present invention to provide a wind shear recovery guidance system that overcomes many of the disadvantages of the prior art systems. It is another object of the present invention to provide a wind shear recovery guidance system that utilizes inertially derived signals to avoid the ef¬ fects of turbulence on air mass derived systems. It is yet another object of the present invention to provide a wind shear recovery guidance system that utilizes a pitch angle reference modulator that modulates the optimum pitch angle as a function of a stall warning signal. Thus, in accordance with a preferred embodi¬ ment of the present invention, the pitch attitude of the aircraft is compared with a pitch reference signal to provide a pitch guidance command signal to the pilot that defines the optimum pitch angle required for maximum climb during a wind shear condition.' The pitch reference value is modulated by a stall warning discrete upon the occurrence of a stall warning to reduce the pitch guidance command signal during a stall warning to prevent the pilot from attaining a pitch attitude capable of causing a stall. Upon cessa¬ tion of the stall warning, the pitch guidance command signal is again increased toward the pitch reference value to permit the pilot to increase the pitch atti¬ tude to the reference value. 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 functional block diagram of the wind shear recovery guidance system as deployed in an aircraft; - 3-- FIG. 2 is a functional block diagram of the pitch guidance system utilized in the wind shear re¬ covery guidance system; and FIG. 3 is a logical flow diagram illustrating = the operation of the pitch angle reference modulator utilized in the pitch guidance system according to the invention. DETAILED DESCRIPTION Referring now to the drawing, with particular 0 attention to FIG. 1, there is illustrated a wind shear recovery guidance system 10 that provides a pitch guidance command signal to a pilot via an attitude director indicator 12 which may be an aural or visual indicator such as a gauge or a cathode ray tube. The 5 wind shear recovery guidance system 10 receives signals from a vertical gyro 14 that provides a signal repre¬ sentative of the pitch attitude of the aircraft, i.e., the angle of the longitudinal axis of the aircraft relative to the horizon. Other inputs to the wind 0 shear recovery guidance system 10 include a stall warning system 16, a flap position discrete 18 and other discretes 20 which may include signals represen¬ tative of the type of aircraft carrying the system and other aircraft parameters such as, for example, 5 landing gear position. The stall warning system 16 may be a conven¬ tional stall warning system, sometimes known as a "stick shaker" that monitors various aircraft para¬ meters such as angle of attack and airspeed and pro- 0 vides a warning to the pilot in the event of an impend¬ ing stall condition. The stall warning may be communi¬ cated to the pilot in a variety of ways, one way being a physical vibration of the aircraft's control, hence the terra "stick shaker". In addition, flap position 5 and other aircraft parameters affect the optimum pitch angle required for maximum climb of the aircraft, and hence, signals representative of flap position and aircraft type discretes are provided to the wind shear recovery guidance system 10 from the flap position discrete 18 and the aircraft type discretes 20. The aircraft type discretes define the type of aircraft that is carrying the system and the optimum pitch angles for maximum climb under various flight con¬ figurations. Typically, the optimum pitch angle for maximum climb is on the order of approximately 15 degrees for modern commercial aircraft. The guidance system 10 includes a pitch reference value generator 21 that responds to the aircraft type discretes and the flap position signal to provide a pitch reference value based on the type of aircraft and its flap posi¬ tion. The pitch reference value generator may take various forms such as, for example, a look-up table that has stored therein pitch reference values for various aircraft under various flap settings or a computation system that has stored therein a reference value for a particular aircraft and modifies this value as a function of flap setting. The wind shear recovery guidance system 10 is illustrated in greater detail in FIG. 2. The wind shear recovery guidance system 10 provides a pitch guidance command signal to the attitude director iπ- dicator 12 based on the pitch attitude of the aircraft, the pitch reference value which, as stated above, is determined by various parameters including the per¬ formance characteristics of the aircraft as defined ' by the aircraft type discretes 20 and aircraft con- figuration parameters such as flap position as obtained from the flap position discrete 18. The pitch reference value is operated on by a reference modulator 22 under the control of the stall warning discrete from the stall warning system 16. The operation of the reference modulator 22 will be described in greater detail in conjunction with FIG. 3, but briefly, the reference modulator 22 - _>-_ operates on the pitch reference value to reduce the pitch reference value upon the occurrence of a stall warning. The pitch reference value as modified by the stall warning discrete is applied to a comparator 24 that compares the present pitch attitude of the aircraft as obtained from the vertical gyro 14 with the modulated pitch reference value to provide a signal representative of the deviation of the present pitch attitude of the aircraft from the modulated pitch reference value. High frequency components of the pitch attitude signal from the vertical gyro 14 are passed through a high pass filter 26 (having a time constant of, for example, on the order of 2 seconds) and an amplifier 28. The amplified high frequency signals from the amplifier 23 are subtracted from the output of the comparator 24 by a summing junction 30 to provide a. damped pitch error signal to a command limiter 32. The command li iter 32 limits the ampli¬ tude of the signal from the summing junction 30 to a value avoid over controlling the aircraft, for example, to a range of zero to +20 degrees. The logical operation of the reference modu¬ lator 22 is illustrated in FIG. 3. The reference modulator monitors the pitch attitude signal from the vertical gyro 14 in conjunction with the pitch refer¬ ence value and the stall warning discrete to adjust the pitch guidance command signal as required relative to the pitch reference value and the stall warning discrete. As is illustrated in FIG. 3, the system first determines whether or not there is a stall warn¬ ing. If there is no stall warning, the system deter¬ mines whether the pitch guidance command signal is less than the reference value. If not, the output of the reference modulator is simply limited to the pitch reference value. If the pitch guidance command signal is less than the reference value, the pitch command is increased by a predetermined number of degrees. _ -- for example, two degrees. After a delay of a predeter¬ mined number of seconds, for example, two seconds, the process is repeated, beginning with a determination of whether a stall warning is present. 5 If a stall warning is present, then the pitch guidance command signal is reduced by a predeter¬ mined number of degrees, for example, two degrees, and after the predetermined time delay, the presence of a stall warning is monitored. If the stall warning -0 is still present, the pitch guidance command signal is again reduced, but if not, the pitch guidance com¬ mand signal is increased until it reaches the pitch reference value. Obviously, -any modifications and variations -5 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 de¬ scribed above. 20 What is claimed and desired to be secured by Letters Patent of the United States is: 25 30 35

Claims: CLAIMS: 1. A guidance system comprising: means for providing a pitch angle reference signal; means for providing a stall warning signal; and means responsive to said pitch angle refer¬ ence signal and said stall warning signal for providing a pitch guidance signal, said pitch guidance signal being determined by said pitch reference signal and modified by said stall warning signal. 2. A guidance system as recited in claim 1 wherein said pitch guidance signal providing means includes means for providing a signal equal to the pitch angle reference signal in the absence of a stall warning. 3. A guidance system as recited in claim 2 wherein said pitch guidance signal providing means includes means for reducing the value of the pitch guidance signal upon the occurrence of a stall warning signal. 4. A guidance system as recited in claim 1 wherein said pitch angle signal providing means in¬ cludes means responsive to a signal representative of the position of the flaps of the aircraft for altering the pitch. 5. A guidance system as recited in claim 4 wherein said pitch angle signal providing means in¬ cludes means responsive to data defining aircraft type. 6. A method for providing pitch guidance to the pilot of an aircraft comprising: monitoring the pitch angle of the aircraft relative to a reference pitch value and providing a guidance signal to the pilot in response to the differ¬ ence therebetween; and monitoring the occurrence of a stall warning and altering the value of the guidance signal upon the occurrence of a stall warning. 7. The method recited in claim 6 wherein the value of the guidance signal is reduced upon the occurrence of a stall warning.

Inventors: GLOVER J HOWARD, GLOVER, J., HOWARD
Applicants: SUNDSTRAND DATA CONTROL, SUNDSTRAND DATA CONTROL, INC.
Application year: 1989
Main classification: G05D1
IPC codes: B64D43, B64D45, G01P5, G05D1