1. Field of the Invention
The present invention relates to a method and apparatus for the protection against aircraft parts striking the ground during near-ground maneuvers such as take-offs, landings and go-arounds.
2. Description of the Prior Art
When maneuvering near the ground, great damage to an aircraft may occur if one of its wing tips strikes the ground, or if its nose wheel or tail touches down prior to main gear touch down. Avoidance of these possibilities is therefore of great importance. Present day aircraft use autopilots to help guide the aircraft not only between destinations but also during ground approaches. The autopilot usually provides, among other things, command signals indicative of roll angle, roll rate, pitch attitude, pitch attitude rate and altitude which are used to position the control surfaces (elevators, ailerons etc.) in order to maintain the aircraft in a desired flight path. The autopilot signals may also be useful in determining if wing tips, nose wheels or tails are getting too close to the ground. More particularly, when the aircraft is low enough to strike the ground and the mechanical design of the aircraft is known, then the exact roll and pitch angles are known where an unwanted ground strike will occur. Since aileron control signals and elevator control signals are used to control the roll and pitch of the aircraft, these signals can be used along with altitude signals to prevent unwanted ground strikes.
For purposes of simplicity, the present invention will be described in connection with the prevention of "tail strike", but it will be understood that the methods and apparatus of the present invention may also be used to avoid nose wheel strike and wing tip strike.
The autopilot pitch elevator command, used to control the aircraft pitch is generated using a two stage process the first stage of which, the outer loop, generates the pitch command or pitch rate command signal from the autopilot as a function of path error or speed error and the second of which, the inner loop, generates the elevator command signals as a function of pitch error or pitch rate error to control the angle of the elevators and cause the aircraft to fly along the desired path or at a desired speed. Outer loops are generally the "command outputs" of the autopilot indicative of the error between actual and desired flight path or airspeed. These outputs, usually, are either rate commands or attitude commands. Inner loops are generally of two types, the first, a "rate inner loop" responds to angular rate command signals and will be described in connection with FIG. 2 and the second, an "attitude inner loop" responds to pitch attitude or pitch angle commands and will be described in connection with FIG. 3. In either event, the inner loop generates signals to position the elevators at an angle necessary to bring the aircraft to the desired pitch rate or the desired pitch angle. Under most conditions, these signals will operate the elevators of the aircraft to follow the desired path and will thus avoid any undesired ground strike. However, maneuvering an aircraft near ground during atmospheric disturbances could cause the aircraft to move off course and produce a situation where the tail could prematurely touch ground as shown in FIG. 1. In FIG. 1, an aircraft 10 desires to fly along an approach path 12 but its actual flight path 14 is seen to have dropped below the desired approach path 12 so that the autopilot calls for a nose up command to the elevators to bring it back on course. If the aircraft is too close to the ground, the nose up command can cause the aircraft to over-rotate with the result that the tail could strike the ground 16. To avoid this, some aircraft employ a "tail strike" protection circuit which determines from the aircraft geometry and altitude when the aircraft might encounter a tail strike and produces a signal which is subtracted from the elevator command signals from the inner loop to reduce the nose up command. One such system, the 4059001-902, produced by Honeywell has been utilized on various aircraft and may generally be seen in FIG. 2. In FIG. 2, the autopilot operates in the outer loop system 20 to produce, in this case, a pitch attitude rate command signal, .theta..sub.com, on a line 21 which is fed to the rate inner loop system, shown in dashed line box 22, which converts the signal into a form the elevator can use, i.e. a signal indicative of the number of degrees of elevator per degree of pitch rate command. More particularly, a first summing circuit 23, in rate inner loop box 22, receives the rate command signal on line 21 and also receives the aircraft's actual rate of change of pitch signal, .theta., on a line 24 and subtracts them to produce an error signal on a line 25. The error signal on line 25 is presented to an integrating circuit 26 to remove any steady state errors and the corrected signal from integrator 26 is presented to a second summing circuit 27 which also receives the error signal on line 25. The resulting corrected error signal from summing circuit 27 is presented on a line 28 to a gain box 29 which modifies the signal to correspond to elevator command signal in terms of desired elevator position. The value K of gain box 29 is determined by the elevator effectiveness of the aircraft and is usually a function of dynamic pressure or airspeed. The system is designed such that at low speed, K has a higher value whereas at high speed, K has a lower value. A Honeywell tail strike protection circuit 30 which receives inputs on lines 31, 32, 33 and 34 indicative of pitch angle, pitch rate, altitude and tail strike limit respectively is programmed to compute a correction signal when a tail strike is possible. Pitch angle and pitch rate are usually determined by the inertial reference unit, altitude is usually from a radio altimeter and the tail strike limit is a function of the geometry of the aircraft. The tail strike protection circuitry produces a corrective elevator command signal on a line 35 whenever there is a danger of tail strike. The elevator command signal from gain circuit 29 on a line 36 and the tail strike protection signal on line 35 are presented to a third summing circuit 37 which operates to subtract the tail strike protection signal on line 35 from the elevator command signal on line 36 to produce a total elevator command signal on a line 38 which is presented to the elevator controls shown by box 39. Since the autopilot command signal is reduced by the tail strike protection signal, the elevator command will be less of a nose up signal or may even be a nose down signal to correct for the over rotation caused by being below the desired flight path. This provides the desired protection in nearly all cases but in some situations, such as where the aircraft is far off of its desired flight path, the outer loop command may be so large that the elevator command signal overpowers the tail strike protection signal and damage could still result.
Another Honeywell system, the 4068300-901, has also been used to solve the tail strike problem in a slightly different way as seen in FIG. 3. This system has a pitch attitude inner loop (rather than a rate inner loop used in FIG. 2) and so a signal indicative of pitch angle, .theta..sub.com, is produced by the outer loop circuitry 50 on a line 52 to a pitch command limiter 54. Limiter 54 controls the magnitude of the pitch angle signal so that the outer loop's pitch attitude command cannot exceed the tail strike limit determined by the geometry of the aircraft. Normally this is a relatively small value in the neighborhood of about 10 degrees. This limited signal is then presented on a line 56 to a pitch rate command limiter 58 so that the change in pitch angle command cannot be too fast. More particularly, if the outer loop pitch angle command signal were allowed to change abruptly, causing a large and sudden elevator control change, it could cause discomfort to the passengers. Accordingly, any rapid change in pitch angle command is limited or smoothed so that it takes place less rapidly. The limited pitch angle command signal is presented on a line 60 to the attitude inner loop circuit shown as dashed line box 61. It will be seen that the attitude inner loop circuitry for a pitch angle control is slightly different than the rate inner loop circuitry for the pitch rate control described in connection with FIG. 2. More particularly, the signal on line 60 is presented to a first summing circuit 62 which compares the commanded pitch angle with the actual pitch attitude, .theta., on a line 64 and any difference is a pitch angle error signal presented on a line 65 to a gain box 66 which operates to modify the pitch angle error signal so as to produce a pitch rate command signal on a line 67 to a second summing circuit 68. The value K of gain box 66 is usually determined by the designer for purposes of controlling how fast the pitch angle error should be removed. A larger K value can remove the error faster but it has a greater effect on the overall stability of the aircraft/autopilot system. The remaining part of the inner loop is the same as the inner loop circuit of FIG. 2. More particularly, the pitch rate command signal is compared with the actual pitch angle rate signal, .theta., on a line 70 to produce a resultant pitch rate error signal on a line 72. The signal on line 72 is presented to an integrator 74 which operates to remove any steady state errors in the signal on line 72 and presents the corrected signal on a line 75 to a third summing circuit 76 which also receives the signal on line 72. The corrected rate error signal is presented on a line 78 to a gain box 80 which converts the rate error signal on line 78 to an elevator command signal as described in connection with FIG. 2.
This attitude inner loop system is very reliable since the tail strike angle cannot be exceeded but because the input to the inner loop is an attitude command signal rather than a rate command signal, the system cannot respond as rapidly as a rate inner loop system so that the path tracking performance is not as good. Also there may be situations, for example where a last minute change of plans occurs (e.g. a "go around" is suddenly called for), where a large command results and the system cannot rotate the aircraft fast enough to avoid ground impact.