Abstract: Methods and a system for a vehicle control system using a reference time profile including an upper control bound and a lower control bound are provided. The system includes an input device configured to receive a required time of arrival at a waypoint and a processor communicatively coupled to said input device wherein the processor is programmed to generate a reference time profile using a first speed profile up to an intermediate control point and a second speed profile between the intermediate control point and an RTA waypoint. The system also includes an output device communicatively coupled to the processor wherein the output device is configured to transmit a speed control signal based on the reference time profile to a vehicle speed control system.

Abstract: A flight control system for an aircraft receives a selected value of a first parameter, which is either the airspeed or inertial velocity of the aircraft. A primary feedback loop generates a primary error signal that is proportional to the difference between the selected value and a measured value of the first parameter. A secondary feedback loop generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and a measured value of a second flight parameter, which is the other of the airspeed and inertial velocity. The primary and secondary error signals are summed to produce a velocity error signal, and the velocity error signal and an integrated value of the primary error signal are summed to produce an actuator command signal. The actuator command signal is then used for operating aircraft devices to control the first parameter to minimize the primary error signal.

Abstract: Systems and methods for controlling activation of a look-ahead function of a terrain alert and warning system. The system receives rate of climb and aircraft speed information, compares the rate of climb and aircraft speed information to a predefined threshold, and deactivates a look-ahead function based on the comparison.

Abstract: In a gas-turbine engine control system having a first control channel inputting the outputs of sensors to calculate and output a first command value indicative of a quantity of fuel to be supplied to the engine and a second control channel inputting the outputs of the sensors to calculate and output a second command value similarly indicative of the quantity of fuel, the second control channel calculates and outputs the second command value using the first command value so long as the instruction to switch the outputs is not generated, while calculates and outputs the second command value, without using the first command value, when the instruction to switch the outputs is generated. With this, immediately after switching to the second command value, the second command value is made substantially equal to and not greatly different from the preceding first command value.

Abstract: A system and method are disclosed for controlling the speed of an aircraft having a preprogrammed speed profile when transitioning from a manually set target speed to the preprogrammed speed profile. In operation, an input is received indicating that a user desires to transition from a manually set target speed to the preprogrammed speed profile. A determination is then made as to whether the manually set target speed satisfies one or more selected conditions for qualifying as a constraint speed of the preprogrammed speed profile. If the manually set target speed satisfies the one or more selected conditions, the preprogrammed speed profile is updated to include the manually set target speed as a constraint speed and the speed of the aircraft is controlled using the updated speed profile.

Abstract: Systems and methods for providing supplemental drag to an aircraft are disclosed. In one embodiment, a method includes detecting changes in at least one throttle resolver angle (TRA). Deflections are determined for one or more flight control surfaces based on the changes in TRA, and accordingly, the one or more flight control surfaces are deflected automatically to generate supplemental drag. The one or more flight control surfaces include at least one at least one of an aileron, a spoiler, and an elevator. Additionally, in one instance, the deflections of the one or more flight control surfaces is implemented as a rated limited time lag function of the changes in TRA.

Abstract: A flight control system for an aircraft receives a selected value of a first parameter, which is either the airspeed or inertial velocity of the aircraft. A primary feedback loop generates a primary error signal that is proportional to the difference between the selected value and a measured value of the first parameter. A secondary feedback loop generates a secondary error signal that is proportional to the difference between the selected value of the first parameter and a measured value of a second flight parameter, which is the other of the airspeed and inertial velocity. The primary and secondary error signals are summed to produce a velocity error signal, and the velocity error signal and an integrated value of the primary error signal are summed to produce an actuator command signal. The actuator command signal is then used for operating aircraft devices to control the first parameter to minimize the primary error signal.

Abstract: This invention relates to the concept of managing the rate of change of energy in a helicopter or other aeronautical vehicle. The invention uses energy management calculations to determine the maximum longitudinal and lateral inputs that can be made while still enabling the vehicle to maintain a desired vertical state. The results of the calculations can be cued to the pilot either tactilely, aurally, or visually, or used for internal software limiting.

Abstract: A positioning control apparatus including feedback loops according to a plurality of control modes which control positioning of an object to be controlled is provided, in which the positioning control apparatus includes a part (121, 122, 123, 124) for reflecting a control process performed by a control mode before being switched in a control process performed by a control mode after being switched when a control mode is switched to another control mode. For example, an operation parameter on the control mode before being switched is dynamically reflected in the control mode after being switched.

Abstract: A method and system is utilized to produce an output corresponding to a safety level, particularly in relation to an activity on a moving body. The method involving producing an output corresponding to the ability to perform an operation within a safe limit on a moving vessel. The method comprising the steps of acquiring real time data from instrumentation on the vessel indicative of first and second elements of vessel motion relevant to the safety of the operation. Processing the data relating to each element of motion. Scaling the data relating to each element to a common scale to provide first and second values relating to the respective elements of vessel motion. Determining which value is of greatest significance and providing a output indicative of the greatest value.

Abstract: An engine thrust management system comprising an engine control device, an aircraft flight manual, a flight management device and a cockpit instrument device. The engine control device is operable to calculate a percent maximum available thrust parameter and a percent indicated thrust parameter. The aircraft flight manual is operable to calculate a required thrust parameter. The flight management device is operable to calculate a percent thrust setting target parameter and a percent commanded thrust parameter. The percent commanded thrust is the amount of thrust requested by an aircraft operator. The percent commanded thrust is varied by the operator according to the value of the percent thrust setting target parameter and the value of the percent indicated thrust parameter in order to produce optimal thrust.

Abstract: An aircraft system for reducing the airspeed of an aircraft as it passes through a preselected altitude includes an automatic throttle system including a computer, a device for inputting a preselected altitude and a preselected airspeed into the computer. An altimeter provides the current altitude of the aircraft and an aircraft instrument provides the current air speed and the vertical speed of the aircraft. The computer in response to the current altitude and vertical speed of the aircraft generates a signal to retard the throttles when the current altitude, the current air speed of the aircraft is equal to 2 times the vertical speed plus a preselected altitude.

Abstract: A system for automatically controlling lift augmentation devices of an aircraft, during a phase of take-off by the aircraft, includes a controllable actuator that shifts the lift-augmentation devices, and a control unit for generating control demands to control the actuator to bring the lift-augmentation devices into a defined position. The control unit further includes a detector that detects actual take-off by the air craft and, if appropriate, signals such detection to the control unit. At start of the take-off phase, the lift-augmentation device are brought into a first position, in which they are deployed. The control unit generates, at least when the detector signals the actual take-off, a control demand making it possible to bring the lift-augmentation devices into a second position, in which the lift-augmentation devices are retracted by comparison with the first position.

Abstract: An apparatus and method for inserting a waypoint into a preexisting flight plan which includes selecting a waypoint on a graphical display of a portion of the flight plan and automatically generating a proposed changed flight plan based upon inserting the waypoint into the nearest leg of the flight plan.

Abstract: The instant invention provides a system and method for controlling the flight of an aircraft to meet an RTA. The system comprises a speed profile generator that communicates with a trajectory generator to produce a speed profile signal that enables the aircraft to reach a waypoint substantially at a predetermined time. In the system of the instant invention, the speed profile generator receives a nominal speed command signal, a time error signal and a sensitivity signal. Based on these inputs, the speed profile generator produces a speed profile signal. The trajectory generator receives the speed profile signal and a required time of arrival signal. Based on these signals, the trajectory generator produces a time error signal and a sensitivity signal. This sensitivity signal represents the sensitivity of the time error signal to changes in the speed profile signal.

Abstract: An airborne power control system for automatically controlling the power of an aircraft during landing is disclssed. The system includes a computer and a minimum airspeed program as a function of altitude. A radio altimeter or the like senses the instantaneous altitude of the aircraft while a pitot tube or the like measures indicated airspeed. A computer and program compare programmed airspeed with actual airspeed for a given altitude. And, a servomotor is provided for decreasing engine thrust where the actual airspeed exceeds the programmed airspeed at any given altitude. An inhibitor inhibits the decrease in engine thrust if the airspeed drops below the programmed airspeed.

Abstract: A method, an apparatus and a computer program product are provided for accurately determining the vertical speed of an aircraft in a manner independent of signals provided by an air data computer, an inertial reference system and an inertial navigation system. Initially, a first vertical velocity of the aircraft is determined based upon a pressure altitude value associated with the aircraft. A second vertical velocity of the aircraft is also obtained from a GPS receiver carried by the aircraft. The first and second vertical velocities are then combined to determine the vertical speed of the aircraft. In this regard, the first and second vertical velocities are combined by complimentarily filtering the first and second vertical velocities. More particularly, the first vertical velocity is typically low pass filtered to remove high frequency noise that is attributable to the relatively low resolution of the first vertical velocity value.

Abstract: In a method of flight control in which a thrust command is computed based on the total aircraft energy error relative to flight path and speed control commands, and an elevator command is computed based on the energy distribution error relative to the same flight path and speed control commands, an improvement is provided including an elevator control command in response to a column control input by the pilot. In the short term, the computing establishes a change in flight path angle beyond the sustainable flight path angle at the trim speed for the prevailing thrust condition. In the long term, the computing establishes a change in speed relative to a set reference speed, the speed change being proportional to the column control input. In the long term, the computing establishes a flight path angle equal to a sustainable value for the prevailing thrust condition and the altered speed condition.

Abstract: An apparatus and method for detecting vertical gusts of wind on board an aircraft in cruising flight is disclosed. The method includes the steps of: calculating an absolute value (.vertline..alpha.-.theta..vertline.) of a difference between a pair of first differentials with respect to time (.alpha. and .theta.) of a current incidence .alpha. and a current pitch attitude .theta. of the aircraft, comparing the absolute value to an upper threshold (Ss), comparing a current Mach number (M) of the aircraft to a Mach number threshold (Mo), and generating an electrical signal that represents presence of a vertical gust of wind when the absolute value is above the upper threshold, when the current Mach number is above the Mach number threshold, and when aerodynamic flaps and slats of the aircraft are in a clean configuration.

Abstract: Disclosed is a pitch-axis stability and command augmentation system (19) in which a pilot column input (.delta..sub.C) is provided to a pitch command processor (26), the output of which is a C*U feedforward command (C*U.sub.FFC) that is supplied as an additive input to a combining unit (20). The combining unit (20) receives a second additive input of an augmented feedback command signal (AFB.sub.COM). The resulting output is filtered and generates an elevator command signal (.delta..sub.e,FILT). The command processor (26) additionally supplies a corrected column position signal (.delta..sub.C,COR) to a pitch command C*U processor (26) that converts the corrected column position signal (.delta..sub.C,COR) into a C*U pitch command (C*U.sub.PilotCmd), representative of the pilot's requested elevator pitch which is generated by movement of the control column. A computed C*U processor (32) forms a computed C*U signal (C*U.sub.Computed) that is based on the current state of the aircraft.

Abstract: An automatic device for maintaining an aircraft at an authorized speed compatible with the technical capabilities of the aircraft includes:a system for comparing the effective speed of the aircraft with reference speeds defined in accordance with the technical capabilities of the aircraft,a system for calculating, on the basis of such comparison, modified control values to align the effective speed with these reference speeds, anda system for imposing one of these calculated values on the aircraft, if this is necessary to maintain the aircraft at an authorized speed.

Abstract: An underspeed protection system for an aircraft under autopilot control selects a target speed based upon the greater of a minimum maneuver speed and a stick shaker speed. The system then compares a monitored speed to the target speed to produce an error signal. The system also monitors vertical speed to determine if tie aircraft begins to descend. If the error signal due to the underspeed condition causes the aircraft to descend, the system provides a hold zero vertical speed signal in place of the error signal such that the aircraft seeks to maintain its altitude. The hold zero vertical speed signal overrides the underspeed error signal such that the aircraft does not pitch forward to seek an increased speed.

Abstract: An aircraft overspeed protection system produces proportional and integral commands for input to an autopilot controlling the aircraft. The proportional and integral commands are produced by comparing actual monitored speed of the aircraft with a target speed from a target speed selector. The target speed selector selects as the target speed a trigger speed above a nominal maximum operating airspeed of the aircraft until the trigger speed is reached by the aircraft. When the trigger speed is reached by the aircraft, the overspeed control goes into Overspeed Protect Command active mode and the target speed selector selects a new target speed below the nominal maximum operating speed of the aircraft. The overspeed protection system remains in Overspeed Protect Command active mode until the pilot takes a positive action by selecting a new autopilot mode or by disengaging the autopilot.

Abstract: An apparatus for compensating for spatial phase angle phase errors in the flapping angle oscillations of a rotor blade is disclosed. The apparatus receives a measured aircraft roll and a measured aircraft pitch. Gyroscopic moments resulting from these measured rolls and pitches are determined, and compensating gyroscopic moments are then generated and summed with commanded pitch and roll accelerations. Flapping angle phase lags of the rotor blade swashplate are added to the summed signals to cancel any flapping angle phase leads, caused by changes in aircraft or rotor blade speed caused by installation configurations of the rotor blade swashplate.

Abstract: A VTOL/STOL free wing aircraft includes a free wing having wings on opposite sides of a fuselage connected to one another respectively for free rotation about a spanwise access. Improved control upon landing of the aircraft is achieved by utilizing a variable pitch propulsion system, enabling the pitch of the propeller to be varied corresponding to the speed of the aircraft and angle of approach upon descent.

Abstract: An underspeed protection system for an aircraft under autopilot control selects a target speed based upon the greater of a minimum maneuver speed and a stick shaker speed. The system then compares a monitored speed to the target speed to produce an error signal. The system also monitors vertical speed to determine if the aircraft begins to descend. If the error signal due to the underspeed condition causes the aircraft to descend, the system provides a hold zero vertical speed signal in place of the error signal such that the aircraft seeks to maintain its altitude. The hold zero vertical speed signal overrides the underspeed error signal such that the aircraft does not pitch forward to seek an increased speed.

Abstract: A synchrophaser for a multi-engine, propeller-driven aircraft including a filter that automatically compensates for misalignment of blade position sensor tabs through derivation and application of a correlation coefficient. The correlation coefficient is calculated in a self-adjustment derivation, utilizing a dynamic tolerance band, providing increased accuracy in propeller speed determination. The correlation coefficient brings about an apparent uniform distribution of blade position sensors, as seen by the synchrophaser, thus eliminating a primary source of cyclic propeller speed disturbances. The correlative filer accomplishes its task without the phase shifting normally encountered in other digital filtering technique.

Abstract: A flight control system that provides energy management for an aircraft. The flight control system includes an energy control system which generates a first error signal that is a function of an actual energy characteristic and a threshold energy characteristic. The energy characteristic may be the bleed rate of the aircraft. The error signal is converted into a control surface command signal which moves the control surface(s) of the aircraft, accordingly. The control system limits the movement of the control surface(s) so that the actual energy characteristic does not exceed the threshold energy characteristic. The energy control system can be coupled to a nominal control system which generates a second error signal that is a function of a pilot input command and a feedback signal which is indicative of the present state of the aircraft. The second error signal can also be converted to a control surface command signal which moves the control surface(s) of the aircraft.

Abstract: The present invention relates to a method and a device for executing phase compensation in a control system of a motor-driven vehicle, in which the time derivative for an output signal (y) from a circuit (1) contained in the control system is limited relative to the time derivative for an input signal (u) to the circuit (1) if the time derivative for the input signal (u) exceeds a predetermined value. A difference signal (.epsilon.) is formed between the output signal (y) and a signal (x) present in the circuit (1), the time derivative of which is not limited in relation to the time derivative for the input signal (u). The difference signal (.epsilon.) is filtered via an asymptotically stable filter (3) to form a signal (.delta.), which is fed back and added to the input signal (u) of the circuit (1) to form the signal (x). A conventional rate limiter may be used to form the output signal (y) from the signal (x).

Abstract: An aircraft automatic throttle system is responsive to a throttle command signal to drive the aircraft throttle through a clutch and detent mechanism which allows for pilot override. Full throttle is achieved only if the throttle is driven against the stop position at the engine with a controlled, limited force. The instant invention is an improved servoamplifier system and method including a position sensor for sensing the position of the throttle, a summing circuit for summing the throttle command signal with the position sensor signal to produce an error signal and an amplifier for amplifying the error signal and producing an output signal to drive the servo driver. Unique signal processing is provided which monitors the amplifier output signal and detects a predetermined condition thereof representative of the throttle being driven against the forward stop at the engine.

Abstract: A Flight Management System for aircraft whereby time-constrained flight can be achieved while maintaining predetermined input parameters selected for minimizing cost of flight, wherein arbitrary points in the flight plan can be designated as time-constraint points, and wherein flight segments can be arbitrarily selected for exclusion from any speed variation. A speed generator derives an initial speed schedule for each segment of a flight plan from inputs of a predetermined cost index, the flight plan, and aircraft performance parameters. The speed schedule is modified by wind data, constant speed segments, and predetermined speed limits. When applied to a profile generator an estimated time of arrival is computed, as well as predicted distance and velocity values for each segment of flight. The predicted values are used to compute a total time of arrival.

Abstract: Process for flying an aircraft in the "elevator speed maintenance mode" during altitude changes, in which for acquiring and/or maintaining a nominal speed it is supplied to the flight computer at the same time as the instantaneous speed of the aircraft, said computer producing the elevator control instruction, characterized in that a nominal speed is also simultaneously transmitted in continuous manner to the automatic thrust control member of the engines and in that the process is performed during two successive sequences, namely:a first sequence during which the elevator only receives a nose up instruction (on climbing) or a dive instruction (on descending) and the engines move to full thrust in the first case or to idling speed in the second anda second sequence, following the first, during which the real nominal speed is supplied to the elevator instruction computer and said same nominal speed increased (on climbing) or decreased (on descending) by a margin is supplied to the automatic thrust control memb

Abstract: A system for integrated pitch and thrust control of an aircraft.

Abstract: A backup flight control system for controlling the flightpath of a multi-engine airplane using the main drive engines. The backup flight control system comprises an input device for generating a control command indicative of a desired flightpath, a feedback sensor for generating a feedback signal indicative of at least one of pitch rate, pitch attitude, roll rate and roll attitude, and a control device for changing the output power of at least one of the main drive engines on each side of the airplane in response to the control command and the feedback signal.

Abstract: A helicopter engine speed reference (66) is increased (117,118) in response to heavy rotor loading (102) in excess of a threshold at the current descent rate (FIG. 4 ). The reference speed is faded up (117,118) at a rapid rate to 107% of rated speed (119). After a fixed time interval (129), reduce rotor loading (122), reduced pitch angle below a threshold magnitude (121) and reduced roll angle below a threshold magnitude (120) will cause the reference speed to be faded down slowly (146,147) to rated speed (148). Automatic increase in engine reference speed is overridden to 107% of rated speed (203) when a battle switch is activated (201) and weapons are ready (202). Increases in engine reference speed are prohibited (207) when the helicopter is operating in quiet mode (206) or the helicopter is resting on its wheels (208).

Abstract: A controller for automatically controlling aircraft thrust during a noise abatement climb operates so as to smoothly reduce aircraft thrust during a reduction in aircraft climb angle at the noise abatement altitude so as to maintain a selected noise abatement climb gradient dependent upon the pilot maintaining a recommended climb airspeed. The selected noise abatement climb gradient is compared with a measured climb gradient to generate an error term. The measured climb gradient is adjusted by a acceleration correction term in order to make the resulting measured climb gradient term independent of airspeed changes. The climb gradient error term is added to the selected climb gradient to generate a commanded climb gradient which is converted to an equivalent thrust (noise abatement thrust). The resulting noise abatement thrust is compared with measured engine thrust to generate an error term for controlling engine thrust so that the aircraft follows the noise abatement climb gradient.

Abstract: A transient free synchronizer stores an input trim signal value on command, and provides an output signal value which transitions in a smooth, continuous manner, when the synchronizer is commanded to the store the input signal. Synchronizers are typically used to capture desired aircraft attitude, position, or velocity trim points for use in aircraft autopilot and automatic flight control systems.

Abstract: A helicopter engine fuel control anticipates sudden changes in engine power demand during yaw inputs to thereby minimize engine and main rotor speed droop and overspeed during yaw maneuvers. The rate (121,123) of yaw control (107) position change generates (110) a yaw component (104) of a helicopter fuel control (52) fuel command signal (70). The magnitude of the yaw component is also dependant upon the rate of yaw control position change (703). The fuel command signal yaw component (104) is overridden (113,115) when rotor decay rate (209,217) has been arrested during a sharp left hover turn (216); when the yaw component is removing fuel (239) during rotor droop (238); and when the yaw component is adding fuel (228) during rotor overspeed (227).

Abstract: A helicopter engine fuel control anticipates changes in main rotor torque in response to lateral cyclic pitch commands, to thereby minimize engine and main rotor speed droop and overspeed during left and right roll maneuvers. A fuel compensation signal (100,101) is summed with a helicopter fuel control (52) fuel command signal (67) in response both to a lateral cyclic pitch command signal (LCP) (107) from a pilot operated cyclic pitch control exceeding a left or right threshold magnitude (201,210) and a total lateral cyclic pitch command signal (TCP) (108) from a lateral cyclic pitch control system exceeding a left or right threshold magnitude (202,207,215,220). The magnitude of the fuel compensation signal is dependant upon the direction of TCP and LCP, e.g., left or right, and helicopter roll acceleration (115). Alternatively, the magnitude and duration of the fuel compensation signal is dependant upon the rate of change in commanded lateral cyclic pitch (107,400,407).

Abstract: A required time of arrival (RTA) control system (200, 220, 240) is disclosed for use with an aircraft on-board computer system (100). The control system (200, 220, 240) is adapted to provide data for controlling aircraft flight in a manner so as to meet time of arrival constraints at selected waypoints. The control system (200, 220, 240) includes means (212) for determining a time error between an estimated time of arrival and a designated RTA. A cost index predictor (216) is utilized to determine an estimated cost index parameter for meeting time of arrival constraints, while maintaining relative minimum fuel consumption. A flight profile predictor (202) is adapted to generate requisite target air speed data based on the estimated cost index parameter. In one embodiment, the system (240) includes determination of a time window based on maximum and minimum permissible cost index and speed limiting parameters.

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: A flight control system of a turboprop airplane includes electronic controlled engines, which are governed by a manual operating device for setting the engine power in order to obtain a certain airspeed, a device for selecting a desired airspeed, and an engine control system for computing and controlling the required engine torque and speed as a function of ambient and engine conditions, the selected engine speed and the setting of said operating device. For automatically controlling the engine speed during the final approach to an airfield, the system includes an electronic approach speed control unit of which the adjustment signal influences the engine control device keeping the speed of the airplane during approach at a selected value whereby said manual operating device has a fixed setpoint. This electronic speed control unit may be carried out as an add-on device for retrofitting on a flight control system.

Abstract: Apparatus and an associated method are described for an aircraft for providing an auto-changeover procedure from a calibrated airspeed control parameter to a Mach number control parameter when the aircraft is ascending; and, from a Mach number control parameter to a calibrated airspeed parameter when the aircraft is descending. The auto-changeover procedure is responsive to application of preselected calibrated airspeed and Mach number parameters. The data processing unit of the aircraft provides a continually updated prediction of the target parameter that would result should the aircraft execute the auto-changeover under the currently existing conditions. When the predicted value and the preselected value are equal, the changeover procedure is invoked, reducing the transients resulting from overshoot of the target parameter value. Provision is made for the circumstance wherein the changeover is determined by an altitude rather than a target parameter.

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: The reference speed for a helicopter engine is incremented or decremented (237, 239) in dependence upon whether a specific range (miles per unit of fuel) has increased or decreased (236a) in a current period of time compared to the next preceding period of time, separated therefrom by at least 20 seconds (243, 244) to thereby set engine speed for minimizing use of fuel over distance traveled. Fuel is sampled only during steady flight (250-255).

Abstract: A helicopter engine speed reference (66) is increased (113, 103-105) in response to heavy rotor loading (108). The reference speed is faded up (113, 104) at a rather rapid rate to 107% of rated speed (114). After a fixed time interval (118), reduced rotor loading (119) will cause the reference speed to be faded down slowly (120, 103-105) to rated speed (121). If torque exceeds 111% of rated torque (117), the reference speed is similarly faded down (120, 103-105).

Abstract: An integrated engine control-autothrottle control system for controlling changes in the thrust of a jet aircraft engine during transients (i.e., changes in speed) in a manner that prevents throttle lead and overshoot (i.e., exceeds the response capability of the engine) is disclosed. The system uses the engine control to continuously monitor current fuel flow (WF), burner pressure (PB), high rotor speed (N2) and engine thrust (EPR), manipulates the signals and produces limiting fuel flow transient error (WFPBERR) and rotor speed transient error (N2DOTERR) signals for both acceleration and deceleration. The lower parameter transient error signal is selected as the controlling limit signal during acceleration and the higher parameter transient error signal is selected as the controlling limit signal during deceleration. The controlling limit signals are used to produce limiting throttle lever angle rate of change signals--+TRADOT for acceleration and -TRADOT for deceleration.

Abstract: An aircraft guidance system for optimizing the flight path of an aircraft in the presence of a windshear maximizes the time the aircraft remains in the air and the distance traveled regardless of the magnitude of the windshear, in the presence of horizontal or vertical windshear components, while effectively minimizing excitation of the aircraft's phugoid mode. A flight path angle is commanded sufficient to clear any obstacle that may be found in the airport vicinity. For longitudinal or horizontal shears, a slightly positive constant flight path angle which is a function of the magnitude of the vertical wind is added to the slightly positive flight path angle command to produce a modified command that compensates for the decrease in flight path angle relative to the ground caused by the vertical wind. The system inhibits exceeding stick shaker angle of attack by reducing the command signal until the actual angle of attack is equal to or less than the stick shaker angle of attack.

Abstract: There is proposed a method of controlling an engine aerodyne in the climb phase, wherein a velocity variation law as a function of altitude is imposed. Moreover, there is set a law of variation of the engine speed corresponding generally to progerssive decrease in such speed as altitude increases. Optimization of exploitation costs in the climb phase can be reached mainly by taking into account the engine maintenance costs by means of a wear model.

Abstract: A warning system for rotary wing aircraft compares the accumulated altitude loss after take-off of the aircraft with its altitude above ground, and generates a warning if the altitude loss is excessive for the altitude above ground at which the aircraft is flying. The position of the landing gear, the speed of the aircraft and its altitude enable the system only during the take-off and missed approach phases of operation in order to minimize nuisance warnings during other phases. The relationship between radio altitude and altitude loss required to generate a warning is optimized for rotary wing aircaft such as helicopters.