Abstract: A helicopter flight control system (21) includes a model following control system architecture which operates in a velocity command mode at low ground speeds. The control system processes information from a variety of helicopter sensors (31) in order to provide a command signal to the main rotor (11) of the helicopter which results in a ground speed which is proportional to the input provided via a sidearm controller (29).

Abstract: A ground-plane-referenced, acceleration signal 2 and Doppler velocity signal 56 are utilized by a complementary filter 6 to provide a complementary velocity signal 8 which is transformed 60 to an inertial coordinate referenced velocity signal 62 utilized in combination with a GPS position signal 56 by a complementary filter 68 to provide a complementary, position error signal 130 that is transformed to a ground-plane-referenced, position error signal 168 and summed with a wind speed signal 196 calculated from the difference between air speed 200 and the aforementioned complementary velocity signal 8 to provide a pitch command signal 202. A similar system is utilized to provide a roll command signal. The system is gated by signals 100, 176 which are determinative of the aircraft meeting predetermined flight conditions 250-262.

Abstract: A pitch guidance system for an aircraft utilizes inertially derived pitch information to provide the pilot with information defining the optimum pitch angle for maximum climb during a wind shear condition. The system utilizes a pitch reference modulator that receives a stall warning discrete from a stall warning system to reduce the commanded pitch angle 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 the angle of attack vane signal, and the commanded pitch angle is biased as a function of altitude and vertical speed to optimize the pitch angle for different altitudes and descent rates.

Abstract: A path capture forcing function generator utilizes integrations from an initial altitude rate error signal and an initial altitude error signal to provide an altitude rate signal and an altitude signal which are compared to the altitude rate error and altitude error signals to provide path rate error and path error signals to the autopilot. In the process of providing the altitude rate and altitude error signals an altitude acceleration signal is established which is used as a feed forward predictor for expected normal acceleration of the aircraft.

Abstract: An improved aerodynamic vane is used for sensing the angle of attack of an aircraft during yawed flight. Left and right sensors are rotatably mounted to opposite sides of the fuselage of the aircraft and rotate in the air flow to a position indicative of the angle of attack of the aircraft. Each sensor consists of an arm rotatably attached at one end to the aircraft with a vane at the other end of the arm, which vane is canted at an angle with respect to the axis of rotation of the arm. The vanes on the sensors on the opposite sides of the aircraft are both canted down or up as to cause the sensors to take similar angular positions during yawed flight.

Abstract: A pitch control system for an aircraft flying at a Mach number greater than 0.7 and including a horizontal plane that is adjustable in deflection and air brakes having a nose-down effect. According to the invention, the system comprises:first means providing the aerodynamic incidence (.alpha.) of said aircraft at each instant;second means for generating a nose-down first command (.DELTA.iHC) for said adjustable horizontal plane when said aerodynamic incidence is greater than said first threshold (.alpha.0), the amplitude of said first command being such that if said adjustable horizontal plane were to take up the corresponding position instantaneously, it would compensate the increase in the nose-up effect on said aircraft due to the way in which said aerodynamic pitching couple varies above said first threshold;third means for generating a signal (.DELTA.

Abstract: An adaptive gain and phase controller for an autopilot for a flight vehicle, such as a hypersonic glide vehicle, that includes applying a reference excitation signal to the control system of a flight vehicle, measuring the response of the vehicle to that excitation signal, namely the gain and phase losses through airframe, and making adjustments to the gain and phase inputs to the autopilot based on those measurements. A high gain narrow bandpass filter is incorporated so that the test signal can be extracted from the airframe of the flight vehicle.

Abstract: Disclosed is a method of and an arrangement for generating a ground effects compensated aircraft body angle of attack signal. The disclosed method includes converting an aircraft pitch rate corrected alpha vane signal to a free air body angle of attack signal that is calibrated to the type of aircraft that employs the invention for the two conventional landing flap settings. The free air body angle of attack signal then is processed to obtain an alpha vane correction signal that represents ground effects compensation that would be required at zero altitude (aircraft touchdown). The zero altitude alpha vane correction signal is then multipled by an altitude scaling signal and converted to a ground effects compensation signal that is applicable to the then current altitude of the aircraft. The ground effects compensation signal then is summed with the free air corrected body angle of attach signal to provide the ground effects compensated body angle of attack signal of the invention.

Abstract: A canard based air speed, angle of attack and sideslip measurement system consisting of canard mounted pressure transducers. The canard includes a leading edge, a side surface, a bottom surface and a top surface. The apparatus includes pressure sensors including a plurality of flush orifice pressure sensors mounted on the canard further including a first configuration of pressure sensors flush mounted on the side surface, a second configuration of pressure sensors flush mounted on the bottom surface, a third configuration of pressure sensors mounted on the top surface, and a fourth configuration of pressure sensors mounted on the leading edge and wherein each pressure sensors provides an output signal and the combinations of all of the output signals are proportional to the angles of attack of the aircraft. A shaft encoder provides displacement angles for the canard.

Abstract: A variable bandwidth factor KALT is applied in a total energy control system to obtain a reduction in throttle activity while maintaining system stability. The system has a total energy control loop and an energy distribution control loop. In the former, a net thrust command signal T.sub.c is generated to reduce the total energy error to zero. In the latter, an elevator position command signal .delta.e.sub.c is generated to reduce the energy rate distribution error, i.e. correct the distribution of energy between kinetic energy (speed) and potential energy (altitude). The error signal input into each loop has a flight path component and a speed component. The factor KALT is applied to both components of the total energy error to reduce the bandwidth of the total energy control loop with increasing altitude and thereby reduce throttle activity.

Abstract: An instrument (10) is disclosed for an aircraft having a windshear warning system to vary the threshold alarm setting as a function of a thermodynamic property of the air and as a function of local geographic and seasonal conditions. Specifically, a function generator 16 and other circuitry is disclosed. The output from the function generator 16 and other circuitry is then added to a fixed threshold acceleraton signal to produce a threshold alarm signal which has a value that is a function of the temperature of the air, in a basic embodiment.

Abstract: This invention provides a flare control modification for pilot-in-the-loop aircraft maneuver command electronic flight control systems. During flare, pilot pitch control inputs are interpreted as incremental pitch attitude commands. The flare control modification allows conventional piloting technique (i.e., pitch controller pull-and-hold) to be used during flare, while retaining the benefits of maneuver command system masking effects of gusts, winds, wind shear, and variations in airplane weight, balance, and aerodynamic configuration.

Abstract: An aircraft automatic or semiautomtic vertical path control system which coordinates operation of pitch and engine thrust control systems to transfer speed control from one system to the other depending on a requirement to climb, descend, or maintain altitude as determined by the polarity and magnitude of the difference between a selectable desired altitude and current actual altitude.

Abstract: A flight path control apparatus for controlling an aircraft to execute a capture maneuver of a predetermined heading or track and provide a roll correction signal for controlling the path of said aircraft along an exponential transition curve. The corrective heading or track signal .DELTA..psi. is generated in accordance with ##EQU1## where .DELTA..psi. is the magnitude of the corrective heading or track command signal,g is a constant corresponding to the acceleration of gravity,V.sub.t is the true airspeed of the aircraft,.tau. is a predetermined constant of aircraft roll rate,.phi..sub.i is the roll angle of the aircraft at the initiation of the capture maneuver,.phi..sub.i is the time rate of change of the roll angle of the aircraft at initiation of the capture maneuver, and.phi..sub.M is a predetermined maximum time rate of change of roll rate of the aircraft during the capture maneuver.

Abstract: A display system for use in an aircraft control wheel steering system provides the pilot with a single, quickened flight path angle display to overcome poor handling qualities due to intrinsic flight path angle response lags, while avoiding multiple information display symbology. The control law for the flight path angle control system is designed such that the aircraft's actual flight path angle response lags the pilot's commanded flight path angle by a constant time lag .tau., independent of flight conditions. The synthesized display signal is produced as a predetermined function of the aircraft's actual flight path angle, the time lag .tau. and command inputs from the pilot's column.

Abstract: There is provided a controller for aiding a pilot to control an airplane's track over the ground. While the airplane is in the air, a lateral control subsystem is provided for automatically holding the aircraft aligned with a desired track angle. Control is initiated when the pilot places the airplane in a wings level condition and centers his roll axis controller (e.g., wheel or stick). An in-air track rate control subsystem is also provided. While the airplane is on the ground, a directional control subsystem is provided which enables the pilot to command track rate or heading rate with rudder pedal inputs. The track/heading rate commands are integrated to generate track/heading angle commands. Feedback control is used to position the rudder and nosewheel steering as required to make the airplane's actual track/heading follow the track/heading commands. Provisions are made to initialize the in-air lateral and on-ground directional control subsystems to provide smooth transitions during takeoff and landing.

Abstract: A flight path angle symbol is positioned on an aircraft attitude display indicator as a function of either a measured flight path angle or a flight path angle commanded by the pilot. When the aircraft is operating within a selected airspeed, angle of attack, pitch attitude or load factor range, or the difference between measured flight path angle and commanded flight path angle is within a selected range, the symbol is positioned on the display as a function of commanded flight path angle. However when the aircraft is operating outside any of these ranges, the symbols is positioned on the display as a function of the measured flight path angle. To avoid a rapid jump in symbol position when transitioning between measured and commanded flight path angle displays, a phase lag filter is used to attenuate and lag the symbol positioning signal.

Abstract: This invention relates to a speed reference system for piloting an aircraft, delivering information on piloting in pitch, particularly during the phases of take-off and of go-round, wherein this system comprises a first voter receiving a first difference between a displayed speed and the real aerodynamic speed, a second difference between the real pitch attitude and a reference attitude and a third difference between the real vertical speed and a vertical speed limit; a second voter receiving the output signal of said first voter and said third difference and delivering the smaller of these two signals; and a third voter receiving the output signal of said second voter and a fourth difference between an aerodynamic speed limit and said real aerodynamic speed, and delivering the greater of these two signals.

Abstract: An integrated, remote sensing and reactive detection system is provided for detecting and confirming the presence of hazardous microbursts, macrobursts, and windshears in the general flight path of an aircraft. An infrared remote sensing system is used to seek out, detect, and provide advanced alerts of thermal gradients ahead of an aircraft which correlate with windshear conditions. The measurement of atmospheric temperature is accomplished by a scanning, multi-spectral radiometer that sweeps an approximate 60 degrees path in front of the aircraft at about a 5 hertz rate. The radiometer employs two rows of detectors that are slightly offset resulting in two simultaneous measurements of temperature that are about 7 degrees apart in elevation angle.

Abstract: This invention provides a flare control modification for pilot-in-the-loop aircraft maneuver command electronic flight control systems. During flare pilot pitch control inputs are interpreted as incremental flight path angle commands above a reference flight path angle (nominally a -3.degree. glideslope). The flare control modification allows conventional piloting technique (i.e., pitch controller pull-and-hold) to be used during flare, while retaining the benefits of the maneuver command system masking effects of gusts, winds, windshear, and variations in airplane weight, balance, and aerodynamic configuration.

Abstract: A wind shear warning system monitors the flight path of an aircraft and wind shear in the vicinity of the aircraft and generates an advisory or cautionary message for the pilot based both on the magnitude of the wind shear and the flight path of the aircraft. When the aircraft is flying a normal flight path, a wind shear warning is generated only when relatively high negative wind shears are present. If the flight path of the aircraft is not normal, for example, if the aircraft is below the glide slope or descending too rapidly, a cautionary alert is given at lower values of wind shear. The alert or warning is selected to provide the pilot with information defining the nature of the hazard or potential hazard being encountered.

Abstract: An apparatus for measuring the temperature lapse rate of the environment surrounding an aircraft in flight is described. The measured temperature lapse rate is compared with the dry adiabatic temperature lapse rate. If the measured temperature lapse rate is less than the dry adiabatic temperature lapse rate, a cautionary message is issued to the flight crew alerting them of the probability of microburst windshears. In addition, if the cautionary message has been issued and a rapid reversal of the measured temperature lapse rate occurs, a warning message is issued to the flight crew advising them of the probable penetration of a microburst windshear.

Abstract: A system for elaborating, on board an aircraft, an alarm signal in the event of a mini-tornado, wherebyan alarm threshold (.alpha.+.alpha.w) for mini-tornado is defined which is a function of the present incidence (.alpha.) of the aircraft, incremented or decremented by incidence equivalents (.alpha.w) corresponding to present and past winds and to present and past wind variations; andthe mini-tornado alarm is triggered off when the compensated present incidence (.alpha.+.alpha.w) exceeds a critical threshold, taking the aircraft configuration into account.

Abstract: The invention comprises an aircraft ground collision avoidance system which utilizes a flight control computer for monitoring aircraft flight states and calculating preferred aircraft flyup altitudes at which automatic aircraft flyups should be initiated. The flyup altitude is determined by calculating aircraft altitude to be lost during a recovery maneuver from a current aircraft flight state. The invention also comprises an aircraft autopilot for monitoring aircraft flight states and initiating an automatic aircraft flyup when so directed by the flight control computer. The monitored aircraft states include aircraft airspeed, angle of attack, bank angle and velocity. Sensor lag, aircraft load factor and engine throttle position are also considered in determining when to conduct an automatic aircraft flyup.

Abstract: An indicator meter which may comprise a galvonometer is mounted in a casing having an indexing pointer or "bug" mounted for movement thereon relative to the pointer of an aircraft angle of attack indicator. The settable indexing pointer and the bezel of the casing are integrated with a variable resistor. A reference voltage is fed to the variable resistor such that a voltage is generated which is in accordance with the positioning of the indexing pointer. Thus, a voltage is generated in accordance with the pointer setting which represents the value of a desired angle of attack. The pointer can be set to represent a desired angle of attack with the meter indicating the measured angle of attack. Voltages representing the measured and desired value of angle of attack are compared with a difference or "error" signal is generated.

Abstract: An instrument (10) is disclosed for an aircraft having a windshear warning system to vary the threshold alarm setting as a function of a thermodynamic property of the air and as a function of local geographic and seasonal conditions. Specifically, a function generator 16 and other circuitry is disclosed. The output from the function generator 16 and other circuitry is then added to a fixed threshold acceleration signal to produce a threshold alarm signal which has a value that is a function of the temperature of the air, in a basic embodiment.

Abstract: An aircraft guidance system for flying an optimal inertial flight path in a windshear encounter. Zero inertial flight path angle command is utilized as the basis for guidance commands for either horizontal or vertical windshears; the flight path angle command is constant regardless of the actual magnitude of the shear. The effect of the guidance law described is to provide a zero inertial flight path angle to prolong excitation of the aircraft's phugoid oscillatory mode. By utilizing inertial flight path angle derived from an inertial reference system, errors introduced by vertical gyroscope systems are eliminated and thus a very accurate flight path angle guidance law is produced. The described guidance control law maximizes the aircraft's time in the air during a windshear encounter. Protection against aircraft stall is also provided by using a computed variable maximum allowable angle of attack.

Abstract: A wind shear detection system compares air-speed with a composite signal derived from signals representative of longitudinal acceleration, normal acceleration, angle of attack and flight path angle to generate a shear signal. An enhanced version of the system is also compensated for roll angle, radio altitude and flap position. In a modified system, the accelerations are calculated along the velocity vector of the aircraft rather than along the horizontal axis to compensate for inaccuracies that could occur under extreme flight conditions such as high bank angle turns and dynamic maneuvers.

Abstract: An improved wind shear warning system for aircraft. A wind shear signal representing wind shear as modified by the downdraft drift angle of the aircraft is used to provide a warning signal when the wind shear signal exceeds (negatively) a variable threshold. The threshold for the warning signal is modified in response to the pilot controlled excess of the airspeed of the aircraft over a reference airspeed during the approach to a landing so as to avoid unnecessary warnings. The threshold is further altered when the wind shear signal during approach indicates the high likelihood of a reversal of wind shear conditions.

Abstract: The present invention constitutes a system for generating an elevator command signal for directing a pilot in guiding his aircraft along a path effective for recovering from hazardous windshear conditions. The system includes a descending mode guidance subsystem (10), an ascending mode guidance subsystem (12) and a switching mechanism (14) for shifting between the two subsystems. The descending mode subsystem includes an acceleration generator (20), a flight path command generator (22), a flight path error generator (24), a descending mode pitch error generator (26), an airspeed control device (28) and a descending mode signal controller (30). These components are connected serially together and operate to form an elevator command signal corresponding to the acceleration required to halt the descent of the aircraft by a fixed altitude level.

Abstract: A controller for apportioning high amplitude and low amplitude control signals between two aircraft control surfaces, such as for example, between an elevator and an associated movable horizontal stabilizer. The signals are separated along a first path and a second path. Signals along the first path are fed to a low pass filter where the low frequency signals are separated and an elevator position error signal is generated. Those signals exceeding a predetermined elevator position error are fed in the form of stabilizer rate signals to a stabilizer position module where a stabilizer positioning signal is generated and then fed to a stabilizer servo system to reposition the horizontal stabilizer. The stabilizer positioning signal is also fed to a multiplier where it is multiplied by a predetermined pitch effectiveness ratio representing the amount of elevator deflection required to achieve the same pitch moment as one degree of stabilizer deflection.

Abstract: A method for the automatic control of fuel consumption in an airplane, comprises the determination and indication in real time of the specific range of the airplane as a ratio of the ground speed and the fuel flow rate. A maximum specific range is searched and maintained by producing, in the intervals between two subsequent searches of the maximum, control actions for maintaining an angle of attack, which corresponds to the last maximum of the specific range. When the conditions under which the maximum has been found to change, a new maximum is searched. An apparatus for performing the method, comprises transducers of ground speed and fuel flow rate, and a computing microprocessor connected via an output and an interface to a digital indicator as well as to a device for controlling the thrust of the airplane engines.

Abstract: Method and apparatus for commanding an optimal flight path for an aircraft encountering a windshear condition. An airspeed rate signal equal in magnitude and opposite in sense to the windshear is applied to derive a variable rate of change of airspeed command for application to a flight director indicator. Limits are placed on the derived command such that the minimum allowable speed command is stick shaker speed and the maximum allowable speed command is the nominal speed in the absence of the windshear. The resultant command signal represents a true airspeed rate that yields an optimal flight path for the aircraft to exit the windshear condition.

Abstract: The presently-disclosed invention utilizes sensed pitched rate and sensed vertical acceleration to prevent aberrant pitch rate command outputs as a function of lead compensated sensed acceleration, controlling the vertical acceleration by sensing when a maximum is approached. When the lead compensated acceleration exceeds a first predetermined acceleration factor, the presently-sensed pitch rate replaces the pre-programmed maximum pitch rate factor as the temporary maximum rate, to temporarily prevent additional vertical acceleration due to increasing pitch rate. This apparatus requires no airspeed input and does not interfere with flight performance characteristics during normal flight operations, allowing higher torque servos to be used, and thus improves overall performance.

Abstract: A method for developing desired air data for use in flight control systems is disclosed including determining angle of attack and sideslip angle relative to inertial space and also determining said angles by solving appropriate aerodynamic equations, whereby two independent estimates of the angles are provided. These estimates are combined for providing the desired data.

Abstract: An aircraft control system that simultaneously controls vertical speed (or flight path angle) and airspeed or Mach number by supplying the vertical speed (or flight path angle) command to the autothrottle system and the airspeed or Mach command to the autopilot pitch channel. A thrust reference computer provides an approximation of the thrust required to maintain the reference vertical speed (or flight path angle) and a vertical speed error signal in closed loop fashion provides a vernier adjustment to the approximation to provide the autothrottle channel command.

Abstract: A system particularly adapted for guiding aircraft through wind shear conditions includes: a thrust command control (10) for maintaining airspeed and groundspeed within a defined range; a detection system (12) for detecting and annunciating the presence of moderate and severe wind shear; and a pitch command system (14) which provides pitch command signals for optimum climb out from a wind shear condition.

Abstract: The invention provides for rapid response and reliable detection of vertical windshear during the takeoff and approach phases of flight. Standard instrumentation available on many aircraft is used to calculate the down-draft acceleration and its effect on the aircraft's performance capability. The invention utilizes inputs from vertical acceleration, true airspeed, pitch angle, and angle of attack sensors. The rate of change of vertical wind is derived from the sensed signals and used to determine the margin between the aircraft's present performance capability and the predicted performance capability at stick shaker speed.

Abstract: A wind shear detection and alerting system (30) compares inertially derived accelerations with air speed rate to provide a wind shear warning signal. Inertially derived accelerations are used instead of purely inertial accelerations because purely inertial systems generally require a vertical gyro. By utilizing angle of attack (.alpha.) and flight path angle (.gamma.) instead of pitch angle (.theta.) in the calculation, no vertical gyro signal is required and the system will be responsive to vertical as well as horizontal shear conditions. An enhanced version of the system is also compensated for roll angle (.phi.), with the roll angle being derived from the rate of change of heading (.PSI.) to avoid the need for a vertical gyro. The alerting system is capable of providing visual and aural warnings for a variety of wind shear conditions, such as, head shear, tail shear, head shear followed by tail shear and wind shear trend.

Abstract: Wind tunnel data for a three control surface aircraft is developed for lift, pitching moment, and drag coefficient characteristics. This data is then input into a Lagrange optimization program to determine a unique combination of canard, flap, and strake flap position that trimmed the pitching moment coefficient to zero and provided the minimum drag coefficient as a function of lift coefficient and/or angle of attack, Mach number, and altitude. This program is exercised over the entire Mach number, altitude, and angle of attack range of the aircraft. The output from the Lagrange optimization program are then tabulated and loaded into the memory of a digital flight control computer of an aircraft. As the aircraft flies, the angle of attack sensor, air data sensor and altimeter determine the angle of attack, Mach number and altitude of the aircraft. By means of the computer, the position of the control surfaces are changed to the predetermined settings of the look-up table for minimum drag.

Abstract: A cruise speed control apparatus for aircraft wherein small differences between the target Mach command and the actual Mach number of the aircraft are used to displace the aircraft from its reference altitude to either increase or decrease actual Mach number in the short term. Long term differences are corrected through a special isolation filter to separate the engine control loop and pitch control loop dynamics and an integration technique to adjust the automatic throttles about a thrust target for the commanded Mach.Large differences between the target Mach and the actual Mach or large changes in the commanded Mach are compensated for by a non-linear gain that effectively increases or decreases the target thrust to cause the aircraft to accelerate or decelerate to the commanded Mach.

Abstract: An apparatus for providing asymmetrical torque-switching for an aircraft autopilot. The invention provides an increased amount of available pitch servo torque at high altitude. The direction of pitch of the aircraft is sensed. An increased amount of torque is provided for the pitch-up direction thereby providing more positive control of the aircraft.

Abstract: The rapid and precise control of an aircraft flight path is obtained by a continuous blending of direct lift and conventional longitudinal control system commands by a unit that consists of a combination of a linear gradient and breakout element, a blending integrator, a nonlinear element, a Direct Lift Limit element, an interconnect gain element, and three summing junctions. Direct lift is blended with conventional control as the pilot manual command is conditioned by the linear gradient and breakout element to provide the command signal to the first summing junction. The difference between this command and the output of the blending integrator is applied to the recentering nonlinear element. The direct lift command is obtained by passing the output of the recentering nonlinear element through the direct lift limit element. The second summing junction is used to reform the total system command which is applied to the third summing junction.

Abstract: A system which provides an increased minimum aircraft pitch attitude during take-off or go-around operations in the event that an unsafe wind shear condition is encountered. A control signal from a wind shear warning system, indicative of a dangerous wind shear condition, is fed to a switching device which controls the velocity command signal to the flight director pitch control. This switching device, in response to the wind shear signal, changes the velocity command to a lower predetermined value, such that the attitude of the aircraft is automatically increased to handle the wind shear encounter.

Abstract: The control of the longitudinal or pitching stability of aircraft is imprd by providing a corrected control characteristic for the correction adjustment drive which is interposed between the pilot operated control stick and the main control drive, for example for the blade angle adjustment in a helicopter or for the elevator adjustment in fixed wing aircraft. A control linkage (2) is connected on one end to a main control drive (4) for the elevator or rotor blade adjustment and at its other end to the control stick (1). The correction adjustment drive (3) is interposed in the control linkage (2). A displacement pickup (5) measures the control displacement of the control linkage (2) and a flight speed pickup (6) measures the flight speed.

Abstract: An aircraft is provided with wind velocity measuring equipment which measures the wind velocity both (1) close to the aircraft and (2) at a considerable distance such as one-half mile or a mile away from the aircraft; and a large difference in these two measured velocities indicates dangerous windshear conditions. Either a warning signal is given to the pilot, or automatic control apparatus is actuated, to avoid crash conditions. A single pulsed carbon dioxide (CO.sub.2) laser velocimeter may be mounted in the aircraft for measuring both the close-in and the distant wind speeds. The laser optic focus is shifted between the near and far measurements, and the return signal is time gated to increase signal-to-noise ratio.

Abstract: A display for aiding the pilot of an aircraft in anomalous wind environments. Wind velocity components are measured by an instrument 11, processed by a computer 12 and a vector generator 17 and then displayed as a vector 34 on the display shown in FIG. 3. A signal F indicative of flap position selects a flight validated lift curve which is displayed as a lift curve channel 42. During normal wind conditions the cursor 35 marking the termination of vector 34 is in the channel of curve 42. However, during wind anomalies cursor 35 moves outside the channel. It then becomes the task of the pilot to control the aircraft such that cursor 35 will move back into the channel.

Abstract: An improved flight director go-around mode adjusts the collective stick position in a closed loop to increase lift to a desired value. If the engagement airspeed is below the maximum rate of climb airspeed, the airspeed is increased towards the maximum rate of climb airspeed at a moderate rate. When the maximum rate of climb airspeed is attained, or if the engagement airspeed is at least the maximum rate of climb airspeed, and the rate of climb is satisfactory, the airspeed is maintained. If the engagement airspeed is greater than the maximum rate of climb airspeed and the desired rate of climb is not achieved within ten seconds, the airspeed is decreased towards the maximum rate of climb airspeed. Heading or course is automatically maintained when go-around is engaged.

Abstract: An attitude indicator provides a display which combines the motion of an aircraft symbol with that of a background/horizon symbol to achieve the best features of conventional "inside-out" and "outside-in" types of attitude displays. The present invention is featured in the provision of a cueing control circuitry which, in response to inputted pitch and bank attitude indicative signals, effects the positioning of at least one of the aircraft symbol and the horizon indicating symbol of an attitude indicator as a function of at least one of the magnitude and rate of change characteristics of pitch and bank attitude signals inputted thereto.

Abstract: An improved wind shear warning system for aircraft. A wind shear signal representing wind shear as modified by the downdraft drift angle of the aircraft is used to provide a warning signal when the wind shear signal exceeds (negatively) a variable threshold. The threshold for the warning signal is modified in response to the pilot controlled excess of the airspeed of the aircraft over a reference airspeed during the approach to a landing so as to avoid unnecessary warnings. The threshold is further altered when the wind shear signal during approach indicates the high likelihood of a reversal of wind shear conditions.