GB1598005A

GB1598005A - Engine out control system for stol aircraft

Info

Publication number: GB1598005A
Application number: GB15702/78A
Authority: GB
Original assignee: Boeing Co
Current assignee: Boeing Co
Priority date: 1978-04-20
Filing date: 1978-04-20
Publication date: 1981-09-16

Description

(54) ENGINE OUT CONTROL SYSTEM FOR STOL AIRCRAFT (71) We, THE BOEING COMPANY, a corporation organized and existing under the Laws of the State of Delaware, of 7755 East Marginal Way South, Seattle, Washington, United States of America, do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to short takeoff and landing aircraft, more particularly this invention relates to an automatic control system for enabling a short takeoff and landing aircraft to safely execute takeoff and landing procedures with one engine malfunctional in that said engine does not provide the desired or scheduled amount of thrust.

In the majority of presently proposed short takeoff and landing (STOL) aircraft, the aircraft is equipped with deployable control surfaces such as specially configured trailing edge flaps which can be deployed during STOL procedures to substantially increase the aircraft lift coefficient. During other flight regimes, such as cruise, these control surfaces are generally deployed in a manner which essentially places the aircraft in a more or less conventional aerodynamic configuration. In one type of STOL aircraft, which is particularly suited to the practice of this invention, augmentation of the aerodynamic lift that is supplied by the flow of the ambient air over the wing is effected by a technique identified as upper surface blowing. In an aircraft utilizing upper surface blowing, the aircraft engines are mounted forwardly of and above the wings to discharge their exhaust stream chordwise across the upper airfoil surface of the wing. During normal flight, the exhaust stream is directed rearwardly to substantially generate forward thrust in a conventional manner. During STOL maneuvers, a type of trailing edge flaps, denoted herein as upper surface blown (USB) flaps, are employed to increase the chamber and chord of the wing and at the same time form a continuously curved, downward and rearward extension of the upper airfoil surface of the wing. When the USB flaps are so extended the exhaust stream traveling chordwise over the upper airfoil surface of the wing attaches itself by the Coanda effect to the downwardly and rearwardly curved surface to divert the ex haust stream downwardly and rearwardly. In this manner, a lift component, as well as a forward thrust component, is generated by the exhaust stream. The engine-generated lift component augments the conventional aero dynamic lift created by ambient airflow over the remaining portion of the wing to provide a STOL capability.

A serious problem is encountered when a STOL aircraft attempts to undertake a STOL landing or takeoff maneuver with one of the engines in malfunctional state, especially if such engine malfunction occurs while the air craft is engaged in the STOL maneuver. When this occurs, not only is the engine-generated forward thrust that is normally supplied by the malfunctional engine lost, but the lifting force produced by flow turning over the USB flap is also lost.

USB flaps are known for partially alleviat ing the loss of lift caused by a malfunct:onal engine by restoring some aerodynamic lift.

These USB flaps generally comprise apparatus for reconfiguring the USB flap to a configuration that corresponds to that of a conventional slotted flap arrangement such as those used on many commercial aircraft to produce mechanical lift. In particular, USB flaps have spanwise slots which can be opened during an engine malfunction condition such that ambient air can pass through the slots and produce mechanical lift to partially replace the lost engine-generated lift.

Although the USB flaps are operable to partially replace the loss of engine-generated lift, the use of such a flap alone often is not a satisfactory solution to the problem. First, it should be recognized that during a takeoif or landing procedure, the aircraft flight crew is performing under a substantial workload.

Accordingly, it is desirable to restrict additional procedures to a minimum to thereby ensure the safety of the aircraft. Secondly, it should be recognized that under some engine failure conditions, little time is available for the aircraft commander to react. For example, if an engine fails while executing a short distance takeoff procedure, the aircraft command der may have less than 10 seconds to execute the control actions that are necessary in order for the aircraft to clear the end of the run way.

In addition to the problems that arise due to the physical limitations of the aircraft crew, the use of a USB flap without taking other appropriate control action does not place the aircraft in a desirable aerodynamic configuration. In particular, even though a portion of the engine-generated lift is restored by the spanwise slots that are opened in the USB flap, the distribution of lift across the upper surface of the aircraft wing will not be symmetrical and the airplane will have a tendency to roll. Although the aircraft commander could actuate various control surfaces3 such as conventional spoilers or ailerons to reduce this rolling moment, such action not only increases crew workload but calls for judicious selection and operation of the proper control surafces.

In particular, less than optimal actuation of the control surfaces to alleviate the roll moment can cause a further increase in drag which can cause the aircraft to suffer a further loss in forward velocity or deteriorate flight path performance.

These and other problems are overcome in accordance with this invention by a control system which monitors engine operation and the position of both the USB and conventional flaps to detect whether an engine has failed and to detect whether the aircraft is executing a STOL maneuver. When an engine has failed and the aircraft executes a STOL procedure, the control system automatically activates the aircraft control surfaces to aerodynamically configure the aircraft so as to reduce both drag and rolling moment while increasing lift by opening spanwise slots in the USB flap located aft of the malfunctional engine.

In particular there is provided a method for automatically reconfiguring the control surfaces of a STOL aircraft having at least two engines when one of said engines is malfunctional and said aircraft executes a STOL takeoff or landing maneuver, said engines being mounted on left and right wings of said aircraft with each of said engines supplying an exhaust stream to an associated extendable upper surface blown flap, each of said upper surface blown flaps extending spanwise along the trailing edge of the aircraft wing at a position aft of said associated engine, each of said upper surface blown flaps including slot means for opening at least one spanwise slot in said upper surface blown flap in response to a predetermined signal, said STOL aircraft also having at least two conventional slotted flaps, at least one of said conventional slotted flaps extending spanwise along the trailing edge of the aircraft wing at a position outboard of an upper surface blown flap, said method com- prising the steps of: detecting that one of said engines is malfunctional; automatically extending the upper surface blown flap located aft of said malfunctional engine to an angle of extension which corresponds to the angle of extension of said conventional slotted flap located outboard of said upper surface blown flap being extended; supplying said predetermined control signal to open said spanwise slot in said upper surface blown flap located aft of said malfunctional engine; automatically retracting each of said conventional slotted flaps located on the wing of said aircraft not including said malfunctional engine, said conventional slotted flap being retracted to a position that at least partially counteracts the roll moment caused by said malfunctional engine.

An automatic control system for a STOL aircraft including first and second oppositely disposed wings projecting outwardly from a fuselage wherein at least one gas turbine engine is positioned on each of said first and second wings for supplying an exhaust stream to a lift augmentation flap extendable downwardly and rearwardly from the trailing edge of the wing upon which said engine is mounted, said lift augmentation flap including at least one spanwise slot and slot closure means for selectively opening or closing said spanwise slot, and wherein each of said first and second wings of said STOL aircraft further includes at least one slotted flap extendable downwardly and rearwardly from the trailing edge of said aircraft wings at a position in which said slotted flap is not supplied with an engine exhaust stream from said engine supplying said exhaust stream to said lift augmentation flap, said automatic control system activating said lift augmentation flaps and said slotted flaps when one engine is malfunctional, and comprising: engine failure detection means for detecting the malfunctional state of one of said gas turbine engines; and flight control means including; means responsive to said engine failure detection means for extending that lift augmentation flap normally supplied with an exhaust stream by said malfunctional engine to a position wherein the rearward and down ward extension of said lift augmentation flap corresponds to the rearward and downward extension of the slotted flap located on that wing of said aircraft including said malfunctional engine; means for activating said slot closure means to open said spanwise slot of said lift augmentation flap normally supplied with an exhaust stream by said malfunctional engine; and means for partially retracting the slotted flap of the wing not including said malfunctional engine to reduce roll moment associated drag caused by said malfunctional engine.

A better understanding of the present in vention can be derived by reading the ensuing specification in conjunction with the accompanying drawings wherein: FIGURE 1 is an isometric view of a twin engine STOL aircraft which can advantageously employ the present invention; FIGURES 2a and 2b are rear views of the aircraft of FIGURE 1, respectively depicting the deployment of the flight control surfaces during a normal, full power STOL landing procedure and the deployment of the flight control surfaces effected by this invention during a STOL landing with an inoperative engine; FIGURES 3a and 3b are rear views of the aircraft of FIGURE 1 respectively depicting deployment of the aircraft control surfaces during a normal, full powered takeoff procedure and the deployment of the aircraft control surfaces effected by this invention during a STOL takeoff procedure with an inoperative engine; FIGURE 4 is a graphic representation of the lift distribution across the upper surface of the aircraft wing which illustrates the asymmetric lift distribution normally present with an inoperative engine and illustrates the substantially more symmetrical lift distribution achieved by the operation of the present invention; FIGURE 5 is a block diagram illustrating an embodiment of the present invention; FIGURE 6 is a diagram depicting the flap trim utilized to reduce roll moment during a STOL landing procedure with one engine inoperative; and FIGURES 7, -8, 9 and 10 are logic diagrams which illustrate the operational arrangement of the flight control unit of FIGURE 5 to activate the aircraft flight control surfaces during STOL landing and takeoff maneuvers in which one engine is inoperative.

FIGURE 1 depicts a twin engine STOL aircraft 10 which can advantageously employ the control system of this invention. Although the aircraft of FIGURE 1 utilizes upper surface blowing to effect lift augmentation, it will be realized upon understanding the invention that the control system of this invention can be suitably embodied for use in various STOL aircraft of the powered lift variety.

In FIGURE 1, gas turbine engines 1Z and 14 are respectively mounted forwardly and above the wings 16 and 18 to direct exhaust gasses rearwardly across a plortion of the upper surfaces 20 and 22 of the wings 16 and 18 and across the upper surface of upper surface blown (USB) flaps 24 and 26 which are respectively mounted to the rear portion of the wings 16 and 18 at a position aft of the engines 12 and 14. In accordance with this invention, the USB flaps 24 and 26 are extendable during normal STOL operation to form a downwardly and rearwardly extending, continuous, upper surface for turning the exhaust stream supplied by engines 12 and 14 and are arranged to provide open spanwise slots during operation of the aircraft when one engine has failed i.e., fails to provide the selected level of thrust.

A pair of conventional slotted flaps 28 and 30 are mounted along the rear of the wings 16 and 18 at a position outboard of each USB flap 24 and 26. More explicitly, each flap 28, herein referred to as a center flap, extends spanwise from a position adjacent a USB flapl (24 or 26) to an inboard edge of an outboard flap 30, with each outboard flap 30 extending spanwise along the wing to the inboard edge of an aileron 32. Preferably, the center flaps 28 and outboard flaps 30 are of the well-known double slotted configuration designed to provide mechanical lift when the flaps 28 and 30 are extended and ambient air passes through the spanwise slots located therein.

The ailerons 32 can be of conventional design and are generally utilized in conjunction with spoilers 34 to effect lateral control of the aircraft 10. In the STOL aircraft 10 of FIGURE 1, five spoiler panels 34 are located in the upper surfaces 20 and 22 of the wings 16 and 18, at a position forward of the center flaps 28 and outboard flaps 30.

Spoilers 34 are of conventional design although various other arrangements of spoilers can be utilized by aircraft employing this invention. In addition to being utilized as a lateral control surface during normal flight and as speed brakes, in one embodiment of the STOL aircraft 10, spoilers 34 can be utilized to provide direct lift control. In particular, in STOL landing approaches in which the center flaps 28 and outboard flaps 30 are fully extended, each spoiler 34 is deployed to project upwardly from the upper surfaces 20 and 22 of the wings 16 and 18 to prcvide additional lift.

The operation of the USB flaps 24 and 26, the conventional slotted flaps 28 and 30, and the spoilers 34 by the control system of this invention during a STOL landing approach is illustrated in FIGURES 2a and 2b. In FIGURE 2a, which depicts the configuration of the aircraft 10 during a normal STOL approach in which both engines 12 and 14 are operating, the flaps 28 and 30 are fully deployed in a manner generally employed in conventional aircraft. USB flaps 24 and 26 are partially extended, e,g., to an angle within the range of 35 to 60 , and the USB flaps are modulated or varied in angle of extension by the aircraft commander or by an automatic flight control system. Such modulation of the USB flaps, used in combination with changes in the thrust or throttle settings of the engines 12 and 14, controls the vectored thrust supplied by the USB flaps 24 and 26 so as to control both aircraft velocity and lift. In one embodiment of the depicted STOL aircraft, both the flap modulation and throttle settings are established by a flight control system which enables the aircraft commander to select a desired approach velocity. The flight control system then operates the engine throttles and USB flaps so as to maintain the selected approach velocity as the aircraft commander controls the aircraft 10 along the desired descent path or glide slope.

FIGURE 2b depicts the deployment of the USB flaps 24 and 26, the center and outboard flaps 28 and 30, and the spoilers 34 when the engine 14 malfunctions prior to or during the execution of a STOL landing procedure.

In FIGURE 2b, the USB flap 26 is extended to a position that corresponds to the position of the adjacent center flap 28 and slots, generally denoted as 36, are opened to effectively convert the USB flap 26 to a conventional double slotted flap arrangement. The opening of slots 36 is accomplished by operating an actuator mechanism which can be a conventional hydraulic, pneumatic or electrical mechanism. As ambient air passes through the slots 36 lift is produced to partially replace the loss of powered lift that is normally supplied when the exhaust stream from engine 14 is discharged over the partially extended USB flap 26.

Simply extending the USB flap behind the malfunctional engine does not provide the most advantageous aerodynamic configuration under engine malfunction conditions. Although such a procedure may allow an essentially normal STOL landing under conditions in which the aircraft is positioned on the proper glide path when the engine fails and under conditions wherein it is not necessary to undertake a go-around maneuver to attempt another landing, it can be easily recognized that maximum flight path performance should be provided to enable the aircraft commander to safely maneuver the aircraft under all flight conditions. In this respect, the control system of this invention automatically retracts the spoilers 34 if they are being deployed as direct lift control devices and automatically partially retracts the flaps 28 and 30 of the powered wing, (i.e., the wing 16 in FIGURE 2b).

Automatically retracting the spoilers 34 causes a glide path flattening to counteract the steepening in glide path angle which would normally be caused by the engine failure and reduced drag on the aircraft which would cause loss in air speed and a further increase in glide path angle. Partially retracting the flaps 28 and 30 of the wing that includes the operative engine (wing 16 of FIGURE 2b), reduces the lift of powered wing to balance a rolling moment which would otherwise be caused by the malfunctional engine.

As shall be described relative to FIGURE 6. it has been found that retraction of flaps 28 and 30 of the powered wing is best accomplished as a function of aircraft velocity with no flap trim being utilized should the aircraft undertake a go-around procedure.

With the abovedescribed activation of the USB flap 26, spoilers 34 add flaps 28 and 30, any remaining roll moment is compensated by manual, lateral control means. For example, in FIGURE 2b, the aileron 32 and the spoilers 34 of wing 16 are partially activated to provide the necessary lateral control action.

It will be recognized by those skilled in the art that, under most conditions, the aileron 32 and spoilers 34 of the powered wing could be utilized to provide the necessary lateral control to offset the rolling moment caused by the malfunctional engine without partial retraction of the flaps 28 and 30. It will be further recognized however that such a pr cedure further increases aircraft drag and thus is not as advantageous as the action taken by the control system of this invention In this respect, it has been determined that, in one particular aircraft in which the invention is employed, the control system of this invention enables the aircraft to attain level flight during a STOL landing if necessary, further permitting a positive climb gradient without an increase in aircraft air speed should a go-around procedure be necessary.

Referring to FIGURES 3a and 3b, the operation of this invention when an engine fails during or prior to a STOL takeoff procedure can be ascertained. As shown in FIGURE 3a, during a normal STOL takeoff operation when both engines 12 and 14 are fully operative, the flaps 28 and 30 of each wing 16 and 18 are deployed in a conventional manner to provide lift. In the particular aircraft which is depicted in FIGURE 3a, USB flaps 24 and 26 are not partially extended to provide additional lift during the takeoff procedure since this particular aircraft provides adequate thrust vectoring (lift) to accomplish a STOL takeoff with the USB flaps fully retracted. Although it should be recognized that other aircraft suitable for the practice of this invention may utilize extended lift augmentation flaps such as USB flaps 24 and 26 during the takeoff maneuver, operation of the present invention would remain essentially as described herein.

As shown in FIGURE 3b, which depicts the STOL takeoff configuration with the engine 14 having failed prior to or during a STOL takeoff procedure, the USB flap 26 which is located behind the malfunctional engine 14 is extended to the same position as the adjacent flaps 28 and 30. As in the STOL landing procedure with a malfunctional engine, slots 36 of the extended USB flap are open to provide lift which partially compensates for lift normally provided by the flow of the exhaust stream of engine 14 across the wing 18.

The previously described operation of the control system of this invention is effected so that the manual control necessary to accomplish the desired maneuver is substantially the same as would be experienced when both engines are fully operational. Yet, operation of the present invention is rapid enough to provide STOL takeoff capability when an engine fails as the aircraft is in the process of executing a STOL takeoff. In this respect, it has been determined that total system operation times on the order of less than 10 seconds are generally satisfactory. For exacple, in the situation in which the invention is embodied in the previously referred to aircraft, an operation time of approximately 7 seconds is utilized with 1 second being alotted to the detection of the engine out condition, and 6 seconds being utilized for properly positioning the control surfaces. In this embodiment, the USB flaps are hydraulically actuated between extension angles of 0 and 70" and can be moved at a rate of 10 per second.

FIGURE 4 graphically depicts the distribution of lift across the wings 16 and 18 during a STOL landing procedure with both engines malfunctional, with one engine maldescribed above to redistribute the lift.

effected by this invention, and with one engine malfunctional and the aircraft configured as described to redistribute the lift.

Referring to the block diagram of FIGURE 5, the control system of this invention includes a flight control unit 40 interconnected with a plurality of signaled input devices enclosed within the dashed outline 42 and a plurality of actuators shown within the dashed outline 44. As will be understood from the following description, the flight control unit 40 is effectively a digital processing unit and can be a programmable digital computer, a wired microprocessor unit, or an arrangement of digital logic circuitry for effecting the operations described herein. For example, in the previously referred to embodiment of the invention, the operations of the flight control unit 40 are performed within three programmable digital computers of the triple redundant, fail/ operation-fail/passive safe flight control system that provides an integrated system to accommodate several other flight control functions, such as the previously mentioned automatic control of the USB flap position and engine throttle settings to orient the thrust vector supplied by the USB flaps, so as to provide the proper forward add lift thrust components.

During a STOL landing maneuver, such a flight control system automatically controls the USB flaps 24 and 26 and engine throttles of the aircraft depicted in FIGURE 1 to maintain the aircraft at a selected appreach velocity while the aircraft is being guided along a desired slide path. In any case, the flight control unit 40 receives signals from the input devices within the dashed outline 42 to detect an engine failure, detect whether the aircraft is executing a STOL takeoff or landing maneuver, and supplied appropriate signals to the actuators within the dashed outline 44 to reconfigure the aircraft as previously described relative to FIGURES 2 and 3.

As shall be described in more detail hereinafter, the failure of an engine is detected on the basis of pressure levels within the engines 12 and 14. In the arrangement of FIGURE 5, pressure sensors 46 and 48, respectively mounted within the engines 12 and 14, supply signals representative of the internal engine pressure which in turn are representative of the thrust being supplied by that particular engine. In the previously referred to embodiment of the invention, the pressure sensors 46 and 48 are conventional triplex pressure sensors mounted in pressure ports between tur bine stages of the engines 12 and 14 with an output signal from each pressure sensor 46 and 48 being supplied to each to the three digital flight control computers. Preferably, as shall be described relative to FIGURE 7, the control system of this invention is arranged to detect engine failure by comparing the difference in engine pressure between the engines 12 and 14 to a predetermined threshold and by comparing the pressure within each engine 12 and 14 with a threshold level that is representative of a minimum thrust level, e.g., engine idle thrust level. Additionally, to prevent spurious changes in engine pressure from causing an incorrect indication of engine failure, the control system is preferably arranged such that an engine failure will not be indicated unless the thrust being supplied by one of the engines 12 or 14 is less than the minimum thrust threshold for a predetermined time interval and the magnitude of the difference between the two pressure signals continuously exceeds the predetermined pressure threshold during this predrermined time interval.

In the embodiment of FIGURE 5, the flight control unit 40 detects whether a STOL landing or takeoff maneuver is being executed on the basis of signals supplied by a flap position selector 49. The flap position selector 49 is the mechanism utilized by the aircraft commander to deploy the USB flaps 24 and 28, the center and outboard flaps 28 and 30, and the spoilers 34 in the proper position to execute a desired maneuver. In particular, as in conventional aircraft, predetermined settings of the aircraft control surfaces are associated with the landing and takeoff maneuvers of a STOL aircraft, which predetermined settings can provide an indication of the procedure being attempted. Specifically, in the previously referred to STOL aircraft in which the invention is embodied the flap position selector 49 includes five positions with one position corresponding to all flaps retracted, two positions corresponding to flap deployment for takeoff manuevers in which the aircraft is carrying a heavy load and a STOL takeoff cannot be undertaken, one position corresponding to deployment of the control surface deployment for STOL takeoff, and two positions corresponding to STOL landing maneuvers. To simplify the identification of these positions, the five positions of the flap position selector 49 are respectively labeled "up" and settings of "10", "20", "30", "45'' and "60", even though the indicated "settings" do not correspond to the angular deployment of either the center and outboard flaps 28 and 30, or the USB flaps 24 and 26. For example, when the flap position selector 49 is set in the STOL takeoff position of "30" or the STOL landing positions of "45" and "60", the center and outboard flaps are fully extended to an angle of approximately 60 , with the USB flaps retracted in the STOL takeoff setting of "30" and extended for thrust vectoring when the flap position selector 49 is set at the STOL landing settings of "45" and "60". As previously described the spoilers 34 are deployed as direct lift control devices only when the flap position selector 49 is set to execute a full STOL landing ("60").

In the arrangement of FIGURE 5, the flight control unit 40 detects that a STOL landing or takeoff maneuver is being executed by comparing a signal representative of the setting of the flap position selector 49 with a first threshold value to determine whether the position selector is set at a position of "30" or greater. To determine whether the STOL procedure is a landing maneuver, the flight control unit 40 compares the setting of the flap position selector 49 with a second threshold value to determine whether the posi tion selector is set at a position of "45" or greater, i.e., "60".

It will be recognized by those skilled in the art that various other arrangements can be utilized to determine whether the aircraft is executing a STOL takeoff or landing pro cedure. For example, the signals supplied by the center and outboard flap position sensors 50 and 52 of FIGURE 5, the USB flap posi tion sensors 58 and 60, and a position sensor associated with the spoilers 34 can be logically processed within the flight control unit 40 to determine if the control surfaces are being de ployed to place the aircraft in a STOL land ing or takeoff configuration.

As previously described, in the STOL air craft 10 of FIGURE 1, spoilers 34 can be utilized as direct lift control devices during a STOL l the STOL aircraft of FIGURE 1, the slots 34 are opened only when the USB flap 24 or 26 is extended to correspond to the position of center and outboard flaps 28 and 30 that are extended to a STOL take-off or landing position, i.e., the flap position signal supplied by the flap position selector 49 exceeds the first predetermined threshold value.

To position the center and outboard flaps 28 and 30 of the aircraft wing which includes the operative engine, the flight control unit 40 supplies a command signal to a flap trim servo actuator 70 for positioning the flaps 28 and 30 of the wing 16 or to a flap trim servo actuator 72 for positioning the flaps 28 and 30 of the wing 18. The flap trim servo actuators 70 and 72 are conventional servo actuators, with the flight control unit 40 supplying a control signal to the flap trim servo 70 or 72 until the appropriate flaps 28 and 30 are retracted to a position which substantially counteracts the roll moment caused by a malfunctional engine.

In the practice of this invention, it has been found satisfactory to retract the flaps 28 and 30 of the powered wing as a predetermined function of aircraft velocity. In particular, it has been found advantageous to retract flaps 28 and 30 of the powered wing (e.g., wing 16 in the situation depicted in FIGURE 2b) by a predetermined amount when the aircraft velocity is within a range which includes typical STOL landing approach speeds and retract the flaps 28 and 30 by a lesser amount when the aircraft velocity is greater than the typical speeds at which the aircraft makes a STOL landing approach. To enable the flight control unit 40 to adjust flap position as a function of aircraft velocity, the flight control unit 40 is interconnected with the aircraft air data system 74.

The flap retraction schedule utilized in the previously described embodimem of this invention is depicted in FIGURE 6 with an denoting the normal (both engines fully opera tive) extension of the flaps 28 and 30. For aircraft velocities less than va, where va is equal to or slightly greater than the maximum normal STOL approach velocity, the flaps 28 and 30 of the powered wing are retracted by an amount equal to an minus am where am is the flap extension angle with maximum flap trim applied. For aircraft velocities exceed ing va but less than a velocity Vb, retraction of the flaps 28 and 30 linearly decreases such that for aircraft velocities exceeding Vb, no flap trim is applied. By way of example only, in one particular embodiment of the invention, the maximum flap trim an- am is established at 30 with the aircraft velocities va and Vb respectively corresponding to 90 and 125 knots.

FIGURES 7, 8, 9 and 10 depict logic dia grams which illustrate the operation of the flight control unit 40 of FIGURE 5 to detect an engine failure and operate the USB flaps 24 and 26, the center and outboard flaps 28 and 30, and the spoilers 34 in the above described manner. Although FIGURES 7 through 10 depict the logic operations in the form of flow charts commonly used in conjunction with programming a digital computer or es fablishing the interconnections within a microprocessor unit, it will be recognized by those skilled in the art that conventional logic circuits such as gate circuits, comparator circuits, and counter circuits can be configured in an equivalent arrangement. In any case, it should be recognized that the logic operations depicted in FIGURES 7 through 10 are sequantially performed at an iteration rate that is compatible with overall system operation.

For example, in the embodiment of the invention in which the flight control unit 40 is embodied within a digital flight control computer that also performs other STOL flight control functions, the iteration rate is 25 iterations per second to provide essentially continuous monitoring of engine operation and operation of the control system of this invention.

Referring to FIGURE 7, the flight control unit 40 first determines whether the aircraft is engaged in a STOL takeoff or landing maneuver by determining whether the "flag" or logic variable CSTOL is equal to a logical 1. The flap CSTOL can alternatively be set equal to 1 whenever the signal supplied by the flap position selector 49 exceeds the first threshold value, or by a switch with which the STOL flight control system is engaged.

If the flag CSTOL is Ifot equal to 1, i.e., CSTOL equals a logical 0, the flight control unit 40 sets the flags LEFAIL and REFAIL, which respectively indicate failure of the left engine (engine 12 of FIGURE 1) and the right engine (engine 14 - of FIGURE 1), and the program variables LCOUNT and RCOUNT equal to 0. As shall be described in more detail hereinafter, the program variables LCOUNT and RCOUNT are utilized to establish a time interval during which an indication must be present that the engine 12 or 14 is malfunctional before the flight control unit 40 generates a signal declaring the engine to be malfunctional, i.e., sets the LEFAIL or REFAIL flap equal to 1.

If the aircraft is executing a STOL takeoff or landing maneuver (CSTOL= 1), the flight control unit 40 determines the difference between the pressure signals supplied by the engine pressure sensors 46 and 48 of FIGURE 5. In the diagram of FIGURE 7, the engine pressure differential is identified by the variable MDP = PL - PR where PL is the pressure supplied by the engine pressure sensor 46 (left engine) and PR is the pressure signal supplied by the engine pressure sensor 48 (right engine). Next, the REFAIL flag is tested to determine if a failure of the engine 14 was declared during the previous iteration period.

If the right engine (engine 14) has previously been declared malfunctional, the flag LEFAIL and the LCOUNT variable are set to 0 and the logic sequence advances to a point A of FIGURE 7, which is effectively the starting point for detecting the failure of the right engine 14.

If the right engine was not declared malfunctional during the previous iteration, the pressure supplied by the engine pressure sensor 46 is compared to a pressure threshold P, where Pt is an engine pressure corresponding to a predetermined minimum engine thrust that must be developed by a functioning engine, e.g., minimum thrust produced by the engine under engine idle condition. If the engine pressure PL iS greater than Pt, the engine 12 is operating during the present iteration period. In this case, the flight control unit 40 detects whether the left engine was declared malfunctional during the previous iteration by testing the LEFAIL flag. If the LEFAIL flag is not equal to 1, (i.e., LEFAIL equals 0), the engine 12 was operative during the previous iteration, the LCOUNT variable is set equal to 0 and the flight control unit 40 sequences to point A for the start of the test sequence for determining failure of the right engine 14.

If it is determined that the left engine 12 is operative during the present iteration but was malfunctional during the previous iteration, i.e., PL not less than pt and the flag LEFAIL is equal to 1, the variable LCOUNT is reduced by 1 and tested to determine whether or not the modified variable LCOUNT is less than or equal to 0. If LCOUNT is presently less than or equal to 0, the flag LEFAIL is set equal to 0 and the logic sequence advances to point A for beginning the test sequence to determine whether the right engine 14 is operative. If the present value of LCOUNT is not less than or equal to 0, the flag LEFAIL remains equal to 1 and the sequence advances to point A of FIGURE 7.

If it is determined that PL is less than Pt, i.e., that the pressure within engine 12 is less than the predetermined threshold of pressure during the present iteration, the flight control unit 40 determines if the magnitude of the pressure differential, MDP, is greater than a predetermined pressure differential ap, where ap is selected to represent a thrust differential between the engines 12 and 14 that occurs when one of the engines is malfunctional. If the magnitude of the engine pressure differential is not greater than tp, the flight control unit 40 advances to the point A of the sequence depicted in FIGURE 7. If the magnitude of the engine pressure differential is greater than Ap, the engine 12 is malfunctional during present iteration and the flight control unit 40 determines whether or not the engine 12 was malfunctional during the previous iteration by determining whether LEFAIL is equal to 1. If it had been determined during the previous iteration that the left engine 12 was malfunctional, i.e., LEFAIL 1, the variable LCOUNT is set equal to a maximum count (24 in the diagram of FIGURE 7) and the flight control unit 40 advances to point A of the sequence. If a failure of the left engine 12 was not declared during the previous iteration, LEFAIL is not equal to 1, the variable LCOUNT is increased by 1 and LCOUNT is then compared to the maximum count (24 in the depicted embodiment), where the maximum count is established to define a suitable time interval. If the present value of the variable LCOUNT is not greater than or equal to the maximum count of 24 the flight control unit 40 advances to point A of the sequence. If however the value of LCOUNT during the present iteration is equal to or greater than 24, the left engine 12 is declared to be malfunctional by setting the flag LEFAIL equal to 1 and advancing the flight control unit 40 to point A of the sequence. With the flight control unit 40 at point A of the sequence of FIGURE 7, the determination of the failure state of left engine 12 is complete for the present iteration.

Examining FIGURE 7, it will be noted that after reaching point A of the sequence, the flight control unit 40 performs identical operations involving the quantities PR, RCOUNT, and REFAIL to determine the operational state of the right engine 14.

In view of the above-described operation of the flight control unit 40, it can be recognized that two conditions must occur for a predetermined period before a failure of either engine is declared. Specifically, the engine thrust must be below a predetermined threshold (as determined by the engine pressure PR or PL being below the threshold Pt), and the magnitude of the difference in thrust between the two engines must be greater than a predetermined thrust differential (as determined by comparing the variable MDP with the threshold value Ap. Further, both conditions must occur for a predetermined time interval as established by accumulating variables LCOUNT and RCOUNT to a maximum value (24 in the diagram of FIGURE 7) before failure of the left or right engine is declared. With respect to the accumulation of the variables LCOUNT and RCOUNT, the sequence depicted in FIGURE 7 operates in the same manner as a conventional, digital up/down counter circuit. Thus it can be seen that, in an embodiment in which the flight control unit 40 is implemented by conventional logic circuits, a conventional up/down counter can be combined with appropriate comparator circuits, gate circuits and timing apparatus, i.e., a clock circuit, to effect a conventional logic circuit implementation of flight control unit 40.

Regardless of whether the flight control unit 40 is realized in digital computer form or as a conventional logic circuit, utilizing both a minimum thrust level and the difference in thrust between the two engines as the failure criteria and declaring an engine malfunctional only if an engine does not exceed both portions of the failure criteria is advantageous.

Specifically, if either the minimum thrust level or thrust differential is utilized as a sole criteria, engine failure could be declared under conditions in which both engines are operative.

For example, if thrust differential is the sole criteria, an engine failure could be declared if the throttles of one engine are retarded dur ing ground operations or even during flight.

On the other - hand, if a minimum thrust differential is utilized as a sole failure criteria, a failure condition in which one engine produces a relatively large amount of thrust while the second engine, due to a failure con dition, produces a substantially lesser amount of thrust would not be detected. Requiring both conditions to exist for a number of itera tions to define a time interval on the order of, for example, one second is advantageous in that an engine failure is not declared in res ponse to spurious pressure changes within one of the engines. Such spurious pressure changes occur, for example, when the air craft is taxiing on a runway and an exhaust stream from another aircraft reaches the engine inlet to cause disruption in the airflow through the engine.

FIGURE 8 depicts an appropriate logic sequence for the flight control unit 40 to effect the previously described downrigging of the spoilers 34 during an STOL landing approach wherein one engine is inoperative and the spoilers 34 are being utilized as direct lift control devices. In the arrangement of FIGURE 8, the failure of one of the engines 12 and 14 is detected by determining whether the flag LEFAIL or the flag REFAIL is equal to 1. If neither flag is equal to 1, a failure flag EFAIL, which denotes the failure of one of the two engines, is set equal to 0.

If either of the engines has failed, the flag EFAIL is set equal to 1 and the flap setting signal supplied by the flap position selector 49 in FIGURE 5 is compared with a threshold value to determine if the center and outboard flaps 28 and 30 are deployed in a position in which the spoilers 34 are uprigged. If the spoilers are uprigged, the flight control unit 40 immediately downrigs the spoilers.

If the center and outboard flaps 28 and 30 are not at a position in which the spoilers 34 are uprigged, the sequence is complete.

FIGURE 9 depicts a suitable logic sequence for positioning the USB flap 24 or 26 behind a malfunctional engine 12 or 14 and opening the appropriate USB flap slots 36 during either a STOL takeoff or landing maneuver.

The sequence is initiated after determining whether either a STOL takeoff or landing is being attempted, i.e., the variable CSTOL =1 as described relative to the logic sequence of FIGURE 7. In the arrangement of FIGURE 9, the engine failure flags LEFAIL and RE FAIL are detected to determine whether the left engine 12 or right engine 14 is malfunctional. If both engines are operative, i.e., neither LEFAIL and REFAIL are equal to 1, no further action is performed by the flight control unit 40. If, however, either LEFAIL or REFAIL is equal to 1, the flight control unit 40 positions the appropriate USB flap 24 or 26 and opens the USB flap slots 36.

For example, if the left engine 12 is malfunctional (flag LEFAIL= 1) the signal supplied by the left engine USB flap positions sensor 58 of FIGURE 5 is compared to the signal supplied by the center and outboard flap position sensor 50 of FIGURE 5. In the diagram of FIGURE 9, the position of the center and outboard flaps 28 and 30 is denoted by AF-If, during the present iteration, the USB flap 24 is not positioned to correspond to the position of the adjacent center and outboard flaps 28 and 30, the flight control unit 401 activates the left USB flap 24 by supplying a signal to the USB flap servo actuator 62 of FIGURE 5. Regardless of whether the USB flap 24 is presently positioned to correspond to the position of the adjacent center flap 28 and outboard flap 30 or whether the USB flap 24 is being so positioned by actuation of the servo actuator 62, the flight control unit 40 next determines whether the present position of the USB flap 24 exceeds a threshold value tl where the threshold a, is established to determine the deployment angle of the USB flap at which the slots 36 will be opened. For ex ample, with respect to the previously described aircraft wherein the flap position selector 49 is set at a position of "45" or "60" for STOL landing maneuvers, the threshold 6t, can be the previously second threshold value which is utilized to determine if a STOL landing is being executed (e.g., bt2=42 ). If the USB flap 24 is set to an angle greater than the threshold but2, the flight control unit 40 supplies a signal to activate the slot actuator 66 of FIGURE 5 to open the slots 36 of the USB flap 24. As can be seen in FIGURE 9, if the right engine is malfunctional and the aircraft is executing a STOL landing or takeoff maneuver, the flight control unit 40 operates the USB flap 26 and the slots 36 therein in the same manner.

In view of the above-described sequence, it can be recognized that the position of a USB flap 24 or 26 located behind a malfunctional engine 12 or 14 will be repositioned with each iteration of the flight control unit 40.

Thus, the USB flaps 24 and 26 are effectively controlled to follow any changes in the position of the flaps 28 and 30. For example, should the aircraft attempt to execute a goaround maneuver during a STOL landing approach with an inoperative engine, the USB flap behind the malfunctional engine will operate in conjunction with the center and outboard flaps 28 and 30.

FIGURE 10 depicts a suitable logic sequence for flight control unit 40 to effect the retraction of the center and outboard flaps 28 and 30 of that wing including the operative engine during a STOL landing maneuver in which one engine has failed. In particular, the arrangement of FIGURE 10 retracts appropriate flaps 28 and 30 in accordance with the schedule illustrated in FIGURE 6. In FIGURE 10, the signal, sF supplied by the flap position selector 49 of FIGURE 5 is compared to the previously described second threshold St2 to determine if the procedure being executed is a STOL landing maneuver.

If tit iS less than the threshold value 6tl, indicating that a STOL landing maneuver is not being attempted, no flap trim is necessary and the flap trim variable aeot is set equal to 00. If tiF iS greater than 6t,, indicating that a STOL landing maneuver is being attempted, the aircraft velocity signal, supplied by the air data system 74 in FIGURE 5, is compared with the velocity va previously defined relative to FIGURE 6. If the present aircraft velocity V0 is less than or equal to Va, the flap trim variable aeot is set for maximum trim of the appropriate flaps 28 and 30, i.e., im- If the aircraft velocity Vc exceeds the velocity Va, and exceeds the velocity vb of FIGURE 6, the flap trim variable tieot is set equal to 0. If however, the aircraft velocity V0 is within the velocity range define by the velocities va and vb, the flap trim variable aeot is set equal to (Vb-VO) (8 m)/ (Vb - Va). Once the proper flap trim variable aeot has been determined, the flight control unit 40 determines whether the malfunctional engine is engine 12 or engine 14 by determining which of the engine failure flaps LEFAIL or REFAIL is equal to 1 and supplies an appropriate signal to the flap trim servo actuator 70 or the flap trim servo actuator 72 of FIGURE 5. That is, if the failure flag LEFAIL is equal to 1, the left engine 12 is malfunctioning and the flap trim signal is applied to the flap trim servo actuator 70 whereas if the failure flag REFAIL is equal to 1, the right engine 14 is malfunctioning and the flap trim signal is applied to the flap trim servo actuator 72.

It will be recognized by those skilled in the art that the embodiment of the invention described herein is exemplary in nature and that many variations can be practiced without departing from the scope of the invention. For example, as previously mentioned the flight control unit 40 can be realized as a programmable digital computer, a micro-processor unit, or as an arrangement of conventional logic circuits. Further, although the invention is described in the context of a STOL aircraft having two engines and conventional slotted flaps including a center and outboard flap on each wing of the aircraft, other multiengine aircraft having various known conventional flap arrangements used in conjunction with lift augmentation flaps (such as the disclosed USB flaps) can advantageously employ the invention.

WHAT WE CLAIM IS: 1. A method for automatically reconfiguring the control surfaces of a STOL aircraft having at least two engines when one of said engines is malfunctional and said aircraft executes a STOL takeoff or landing maneuver, said engines being mounted on left and right wings of said aircraft with each of said engines supplying an exhaust stream to an associated extendable upper surface blown flap, each of said upper surface blown flaps extending spanwise along the trailing edge of the aircraft wing at a position aft of said associated engine, each of said upper surface blown flaps including slot means for opening at least one spanwise slot in said upper surface blown flap in response to a predetermined signal, said STOL aircraft also having at least two conventional slotted flaps, at least one of said conventional slotted flaps extending spanwise along the trailing edge of the aircraft wing at a position outboard of an upper surface blown flap, said method comprising the steps of: detecting that one of said engines is malfunctional; automatically extending the upper surface blown flap located aft of said malfunctional engine to an angle of extension which corresponds to the angle of extension of said conventional slotted flap located outboard of said upper surface blown flap being extended; supplying said predetermined control signal to open said spanwise slot in said upper surface blown flap located aft of said malfunctional engine, automatically retracting each of said conventional slotted flaps located on the wing of said aircraft not including said malfunctional engine, said conventional slotted flap being retracted to a position that at least partially counteracts the roll moment caused by said malfunctional engine.

2. The method of Claim 1 wherein said step of detecting that one of said engines is malfunctional includes the steps of: sensing a pressure internal to each of said engines that is representative of the thrust being supplied by each of said engines; determining the difference between the pressure associated with each particular en

**WARNING** end of DESC field may overlap start of CLMS **.

Claims

**WARNING** start of CLMS field may overlap end of DESC **. Thus, the USB flaps 24 and 26 are effectively controlled to follow any changes in the position of the flaps 28 and 30. For example, should the aircraft attempt to execute a goaround maneuver during a STOL landing approach with an inoperative engine, the USB flap behind the malfunctional engine will operate in conjunction with the center and outboard flaps 28 and 30. FIGURE 10 depicts a suitable logic sequence for flight control unit 40 to effect the retraction of the center and outboard flaps 28 and 30 of that wing including the operative engine during a STOL landing maneuver in which one engine has failed. In particular, the arrangement of FIGURE 10 retracts appropriate flaps 28 and 30 in accordance with the schedule illustrated in FIGURE 6. In FIGURE 10, the signal, sF supplied by the flap position selector 49 of FIGURE 5 is compared to the previously described second threshold St2 to determine if the procedure being executed is a STOL landing maneuver. If tit iS less than the threshold value 6tl, indicating that a STOL landing maneuver is not being attempted, no flap trim is necessary and the flap trim variable aeot is set equal to 00. If tiF iS greater than 6t,, indicating that a STOL landing maneuver is being attempted, the aircraft velocity signal, supplied by the air data system 74 in FIGURE 5, is compared with the velocity va previously defined relative to FIGURE 6. If the present aircraft velocity V0 is less than or equal to Va, the flap trim variable aeot is set for maximum trim of the appropriate flaps 28 and 30, i.e., im- If the aircraft velocity Vc exceeds the velocity Va, and exceeds the velocity vb of FIGURE 6, the flap trim variable tieot is set equal to 0. If however, the aircraft velocity V0 is within the velocity range define by the velocities va and vb, the flap trim variable aeot is set equal to (Vb-VO) (8 m)/ (Vb - Va). Once the proper flap trim variable aeot has been determined, the flight control unit 40 determines whether the malfunctional engine is engine 12 or engine 14 by determining which of the engine failure flaps LEFAIL or REFAIL is equal to 1 and supplies an appropriate signal to the flap trim servo actuator 70 or the flap trim servo actuator 72 of FIGURE 5. That is, if the failure flag LEFAIL is equal to 1, the left engine 12 is malfunctioning and the flap trim signal is applied to the flap trim servo actuator 70 whereas if the failure flag REFAIL is equal to 1, the right engine 14 is malfunctioning and the flap trim signal is applied to the flap trim servo actuator 72. It will be recognized by those skilled in the art that the embodiment of the invention described herein is exemplary in nature and that many variations can be practiced without departing from the scope of the invention. For example, as previously mentioned the flight control unit 40 can be realized as a programmable digital computer, a micro-processor unit, or as an arrangement of conventional logic circuits. Further, although the invention is described in the context of a STOL aircraft having two engines and conventional slotted flaps including a center and outboard flap on each wing of the aircraft, other multiengine aircraft having various known conventional flap arrangements used in conjunction with lift augmentation flaps (such as the disclosed USB flaps) can advantageously employ the invention. WHAT WE CLAIM IS:

1. A method for automatically reconfiguring the control surfaces of a STOL aircraft having at least two engines when one of said engines is malfunctional and said aircraft executes a STOL takeoff or landing maneuver, said engines being mounted on left and right wings of said aircraft with each of said engines supplying an exhaust stream to an associated extendable upper surface blown flap, each of said upper surface blown flaps extending spanwise along the trailing edge of the aircraft wing at a position aft of said associated engine, each of said upper surface blown flaps including slot means for opening at least one spanwise slot in said upper surface blown flap in response to a predetermined signal, said STOL aircraft also having at least two conventional slotted flaps, at least one of said conventional slotted flaps extending spanwise along the trailing edge of the aircraft wing at a position outboard of an upper surface blown flap, said method comprising the steps of: detecting that one of said engines is malfunctional; automatically extending the upper surface blown flap located aft of said malfunctional engine to an angle of extension which corresponds to the angle of extension of said conventional slotted flap located outboard of said upper surface blown flap being extended; supplying said predetermined control signal to open said spanwise slot in said upper surface blown flap located aft of said malfunctional engine, automatically retracting each of said conventional slotted flaps located on the wing of said aircraft not including said malfunctional engine, said conventional slotted flap being retracted to a position that at least partially counteracts the roll moment caused by said malfunctional engine.

gine and the pressure associated with at least one other engine of said engines; comparing the magnitude of said pressure difference with a first predetermined threshold value; comparing the pressure associated with each of said engines with a second predetermined threshold value; supplying a signal indicating a particular engine is malfunctional when said pressure associated with said particular engine is less than said second predetermined threshold value and the magnitude of the difference between the pressure associated with a particular engine and the pressure associated with at least one other engine exceeds said first predetermined threshold value.

3. The method of Claim 1 or 2 further comprising the steps of: determining whether said STOL aircraft is executing a takeoff or landing maneuver; automatically retracting each of said conventional slotted flaps located on the wing of said aircraft not including said malfunctional engine only when said aircraft is executing a landing maneuver.

4. The method of Claim 1, 2 or 3 wherein said STOL aircraft further includes spoilers located on the upper surface of said left and right wing wherein said spoilers are deployable from a position flush with said wing upper surface to an upwardly extending position for direct lift control during STOL landing maneuvers, said method further comprising the step of returning said spoilers to said flush position when said aircraft is executing a landing procedure and one engine is malfunctional.

5. The method of Claim 1, 2, 3 or 4 wherein said step of extending each of said upper surface blown flaps located aft of said malfunctional engine comprises the steps of: determining the difference between the angle of extension of said upper surface blown flap located aft of said malfunctional engine and the angle of extension of said conventional flaps located outboard of said upper surface blown flap being extended; supplying a signal to move said upper surface blown flap aft of said malfunctional engine to an angle of extension more closely corresponding to the angle extension of said conventional flap located outboard of said upper surface blown flap being extended whenever said difference exceeds a predetermined amount.

6. The method of Claim 3, 4 or 5 wherein said step of automatically retracting said conventional slotted flaps located in the wing of said aircraft not including said malfunctional engine includes the steps of: detecting the forward velocity of STOL aircraft; automatically retracting said slotted flaps by a predetermined amount when said forward velocity is less than a predetermined velocity; maintaining said slotted flaps at a fully extended position when said forward velocity is greater than a second predetermined velocity; and automatically retracting said slotted flap by an amount linearly related to the difference between said aircraft forward velocity and said second predetermined velocity when said aircraft forward velocity is greater than said first predetermined velocity and less than said second predetermined velocity.

7. An automatic control system for a STOL aircraft including first and second oppositely disposed wings projecting outwardly from a fuselage wherein at least one gas turbine engine positioned on each of said first and second wings for supplying an exhaust stream to a lift augmentation flap extendable downwardly and rearwardly from the trailing edge of the wing upon which said engine is mounted, said lift augmentation flap including at least one spanwise slot and slot closure means for selectively opening or closing said spanwise slot, and wherein each of said first and second wings of said STOL aircraft further includes at least one slotted flap extendable downwardly and rearwardly from the trailing edge of said aircraft wings at a position in which said slotted flap is not supplied with an engine exhaust stream from said engine supplying said exhaust stream to said lift augmentation flap, said automatic control system activating said lift augmentation flaps and said slotted flaps when one engine is malfunctional, and comprising: engine failure detection means for detecting the malfunctional state of one of said gas turbine engines; and flight control means including; means responsive to said engine failure detection means for extending that lift augmentation flap normally supplied with an exhaust stream by said malfunctional engine to a position wherein the rearward and downward extension of said lift augmentation flap corresponds to the rearward and downward extension of the slotted flap; located on that wing of said aircraft including said malfunctional engine; means for activating said slot closure means to open said spanwise slot of said lift augmentation flap normally supplied with an exhaust stream by said malfunctional engine; and means for partially retracting the slotted flap of the wing not including said malfunctional engine to reduce roll moment associated drag caused by said malfunctional engine.

8. The control system of Claim 7 wherein said engine failure detection means includes a pressure sensor mounted within each of said engines for supplying an exhaust stream to a lift augmentation flap, each of said pressure sensors supplying an electrical signal proportional to the thrust being supplied by the engine in which said pressure sensor is moun- ted, and means for comparing each electrical signal supplied by a pressure sensor with at least one other electrical signal supplied by a pressure sensor of another of said engines.

9. The control system of Claim 8 wherein said engine failure detection means further comprises means for comparing the electrical signals supplied by each of said pressure sensor with a predetermined threshold value.

10. The control system of Claim 9 wherein said means for comparing each of said electrical signals supplied by said pressure sensors with at least one electrical signal supplied by a pressure sensor of one other of said engines and said means for comparing said electrical signals supplied by each of said pressure sensors with said predetermined threshold value is included within a digital flight control unit for processing digital data, said digital flight control unit periodically comparing said pressure sensor electrical signals with one another and periodically comparing each of said pressure sensor electrical signals with said predetermined threshold value.

11. The control system of Claim 7, 8, 9 or 10 further comprising detection means for determining whether said STOL aircraft is executing a STOL landing maneuver and for determining whether said STOL aircraft is executing a STOL takeoff maneuver and means responsive to said maneuver detection means for activating said means for partially retracting the slotted flap of the wing not including said malfunctional engine only when said STOL aircraft is executing a landing maneuver with an inoperative engine.

12. The control system of any of the preceding Claims 7 to 11 wherein said STOL aircraft further includes spoilers mounted on the upper surface of said first and second wings, said spoilers being deployable from a position substantially flush with said upper surface of said first and second wings to an upwardly extending position for direct lift control during a STOL landing maneuver, said control system further comprising means for returning said spoilers to said substantially flush position when said spoilers are deployed for direct lift control and said engine failure means detects a malfunctional engine.

13. The control system of any of the preceding Claims 7 to 12 further including means for supplying a first command signal when said STOL aircraft undertakes a short take off maneuver and for supplying a second command signal when STOL aircraft undertakes a short landing maneuver.

14. The control system of Claim 13 wherein said comparator means supplies an engine failure signal indicative of which one of said gas turbine engines is inoperative in that said inoperative engine is providing a thrust level less than said predetermined thrust level.

15. The control system of Claim 14 wherevS in said means responsive to said engine failure is further responsive to signals representative of the rearward and downward extension of said flaps and said lift augmentation flaps for supplying a signal to that particular lift aug mentation flap normally supplied an exhaust stream by said malfunctional engine, said res ponsive means supplying a signal to deploy said particular lift augmentation flap to an extended position substantially corresponding to the extended position of each flap located on the same wing of said STOL aircraft as said particular lift augmentation flap; and said means for activating including means for supplying a first predetermined control signal to open said spanwise slot of said particular lift augmentation flap normally supplied an exhaust stream by said malfunctional engine when said lift augmentation flap is extended downwardly and rearwardly by a predeter mined amount; and said means for partially retracting further including flap retraction means responsive to said second command signal and said engine failure sginal for supply ing a flap retraction signal to partially retract each of said flaps other than said lift augmen tation flaps that are located on said wing not including said particular lift augmentation flap that is normally supplied an exhaust stream by said malfunctional engine, said flaps being retracted to at least partially alleviate roll moment caused by said malfunctional engine failing to provide said predetermined thrust level.

16. The control system of Claim 15, wherein said flap retraction means is responsive to a signal representative of the forward velocity of said STOL aircraft, said flap retraction means including means for supplying said flap retraction signal to maintain said flaps in a first predetermined extended position when said forward velocity of said STOL aircraft exceeds a first predetermined velocity, said flap retraction means including means for supplying said flap retraction signal to place each of said flaps at a second predetermined extended position when said forward velocity of said STOL aircraft is less than a second predetermined velocity, said flap retraction means further including means for supplying said flap retraction signal to place each of said flaps at an extension position between said first and second predetermined extended positions that is linearly related to said for ward velocity of said STOL aircraft when said forward velocity is greater than said second predetermined velocity and less than said first predetermined velocity.

17. The control system of Claim 15 or 16, wherein said responsive means includes means for comparing a signal representative of the rearward and downward extension of said particular lift augmentation flap that is normally supplied an exhaust stream by said inoperative engine with a signal representative of the rearward and downward extension of each of said flaps located on the same wing of said STOL aircraft as said particular lift augmentation flap, said comparison means including means for supplying a difference signal representative of the difference between the rearward and downward extensions of said particular lift augmentation flap and said flap, said responsive means further including means for supplying a signal to move said lift aug mentation flap toward the same extended position as, said flap when said difference signal exceeds a predetermined threshold value.

18. The control system of Claim 14, 15, 16 or 17, wherein said comparison means further includes means for supplying said engine failure signal only when the thrust representative signal associated with said particular engine is less than said predetermined threshold signal and the thrust level of said particular engine is less than the thrust level supplied by the others of said engine for a predetermined time period.

19. A method for automatically reconfiguring the control surfaces of a STOL aircraft substantially as described herein.

20. An automatic control system for a STOL aircraft substantially as described herein with reference to the accompanying drawings.