Abstract: A method of controlling a path of a rotorcraft. The method comprises a step of generating a first path setpoint and a step of automatically generating a first autopilot command from the first path setpoint. The method comprises a step of generating a first pilot setpoint during which a movement of a control member is transformed into a first pilot setpoint, the first pilot setpoint and the first flight parameter being homogenous in that they are expressed in the same measurement unit. The method includes a step of generating a first human pilot command from the first pilot setpoint followed by a step of generating a path command by combining the first autopilot command and the human pilot command.

Abstract: User interface devices using magnetic sensing to provide output signals associated with motion and/or deformation of an actuator element of the interface devices are described. The output signals may be provided to an electronic computing system to provide commands, controls, and/or other data or information. In one embodiment, a user interface device may include a plurality of permanent magnets and a plurality of multi-axis magnetic sensors to generate motion and/or deformation signals to be provided to a processing element to generate the output signals.

Abstract: A servo-controlled system is disclosed for providing simulated feel equivalent to that of traditional mechanical hand controllers. Processed position and force sensor signals are used in a feedback loop that controls the motor mechanically connected to the stick. The feedback loop comprises a low-level motor feedback loop, and high-level force feel loop. The latter comprises static and dynamic performance components. The system allows variable and additional force cues to be specified externally to the system and felt by the operator, and/or external signal to backdrive die stick to follow a specified motion. The control framework permits the electronic coupling of motion and applied forces by pilots in a dual arrangement while retaining the above-mentioned features. It simulates cross-coupling mechanical compliance due to force fight between pilots detents and asymmetric force feel gradients. Parameters associated with loops and performance components can be specified independently.

Abstract: The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch attitude.

Abstract: A method and an apparatus for controlling aircraft rudder movement are disclosed. The system including a yaw damping control portion integrated with a directional compensation rudder control portion, such that the system may simultaneously provide yaw damping control and directional compensation.

Abstract: Process and device for controlling at least one aerodynamic elevator surface of an aircraft during a takeoff. The device (1) comprises a control member (3), calculation means (5) and motor means (7) for controlling the aerodynamic elevator surface (2), as a function of an attitude angle control law received from the calculation means (5). If the duration T between the moment at which the pilot actuates the control member (3) so as to control the aerodynamic elevator surface (2) in such a way as to increase the pitch attitude angle of the aircraft and the moment at which the aircraft should reach the minimum takeoff speed is less than or equal to a nominal duration, the calculation means (5) determine and transmit a minimum control law. Otherwise, they determine and transmit a modified control law which is such that at a duration T after the actuation of the control member (3), the pitch attitude angle is substantially equal to a nominal value.

Abstract: An electrical fly-by-wire system for operating an aircraft rudder includes a low-pass filter, arranged between a rudder bar and an actuator of a rudder. The low-pass filter receives a control command from the rudder bar corresponding to the degree of travel the rudder bar has experienced from a neutral position. Based on the amplitude of the control command, the filter generates an operating command for the actuator. Additionally, the filter operates such that the higher the fraction of the rudder bar's travel away from the neutral position, with respect to its maximum value of travel, the higher the filter's time constant is set.

Abstract: In a method for current regulation of permanently excited synchronous motors for guided missiles with an electromechanical actuating drive for the rudder, the commanded current is always guided orthogonally to the magnet-wheel flow. Further: a) from measured current of two voltage phases, a measured magnitude of a current-space pointer is determined and an error value for the current is added to a desired value of the amount via a correction regulator; and b) from measured current of two voltage phases as well as from the angle command of the current-space pointer, a measurement for the angle error is determined and added to the desired angle value via a correction regulator. The method according to the invention can be implemented with little hardware expenditure and in a compact design size, even for high rotational speeds of >10,000 min−1. Current-regulating methods used heretofore, which involve frequency response compensation in polar coordinates, cannot be used with synchronous motors, i.e.

Abstract: The actuation system of the present invention is based upon sensing relatively small delta errors over relative short strokes as opposed to absolute travel error.

Abstract: A hand controller system which provides the proper feel of the hand controller during aircraft operation. In both non-redundant and redundant embodiments, torque, and position measurements are made from the hand controller movements and processed to provide a feedback to a control motor in mechanical connection with the hand controller. The system includes self-monitoring of the motor, as well as the signals provided by the position, and torque sensors. Connections are provided from the hand controller system so that cross-coupling may be provided between the pilot's and copilot's hand controllers.

Abstract: The control system includes a control member connected to a control centering device which resists operational control member movement and returns it to a neutral position upon release thereof. Force and position sensors connected to the control member determine operator force exerted thereon as well as the position thereof. Output from the sensors is transmitted to a computer which also receives data relating to position of aircraft controlled surfaces and aerodynamic forces acting on the aircraft as well as flight data, etc. The computer combines and compares that data with other data relating to frictional forces produced by movement of the mechanical control system and data relating to desired control member resistive forces under predetermined operational conditions. The computer thereby determines the magnitude of the control member resistive forces required to produce the desired tactile feel and provides an output based thereon to an actuator which applies these forces to the control member.

Abstract: A servo controlled system is provided for electrically coupling the motions and applied forces of pilot and co-pilot control sticks. Servo-coupling between the two control sticks is accomplished by detecting the position of the control sticks of respective controllers and/or detecting torque exerted on the control sticks. The position and torques signals are processed and employed in a feedback loop for each control stick to generate a feedback signal to motors to which the control sticks are mechanically coupled. The motors act on the control stick to simulate the feel of traditional mechanical controllers, and to cause the controllers to track each other's movements.

Abstract: The apparatus controls the positioning of a device in response to a commanded input signal, and comprises an element for moving the device in response to an error signal. A sensor component determines the actual position of the device, the sensor component outputting a primary signal and a derivative signal corresponding to the actual position and velocity of the device, respectively. A complementary filter system filters the primary signal and the derivative signal to generate a derived position signal, the derived position signal being combined with the commanded input signal thereby generating the error signal. The filtering results in effectively increasing the resolution of the sensor component thereby reducing the limit cycle caused by quantization of the sensor component.

Abstract: A full authority engine-out control augmentation subsystem for use in a flight control system to assist a pilot and/or lateral-directional control subsystem in compensating for the yawing moment produced by the differential engine thrust that occurs when the power produced by an engine located on one side of a multi-engine aircraft is lost (or reduced) with respect to the other engines is disclosed. The engine-out control augmentation subsystem produces a compensating rudder command (27) based on straight-gain (21), lag filter (15) and washout filters (23 and (25) modifications of a differential engine thrust signal (11). The engine-out control augmentation subsystem will function with both conventional mechanical and electronic (fly-by-wire) flight control systems.

Abstract: A device for compensating for actuator rate saturation is disclosed. The ice is an improvement to a servomechanism of the type wherein a command signal and a position feedback signal from a member being positioned are summed to produce an error signal for driving an actuator to position the member in accordance therewith. The improvement is a means for inverting the error signal when the ratio of the slope of the plot of the position feedback signal versus time to the slope of the plot of the command signal versus time is less than one and the polarity of the error signal is opposite to the polarity of the slope of the command signal, and the absolute value of the error signal is not increasing. Application of the invention to the flight controls of an aircraft is disclosed.

Abstract: An electric system for controlling a depth aerodynamic surface (1) of an aircraft. The system includes an attitude sensor (3) and an incidence sensor (6). A calculation means (7) in communication with the sensor controls a motor (4) so as the depth aerodynamic surface of the aircraft reaches an incidence according to a programmed depth control law. The system utilizes an incidence protection device (35) able to control the incidence of the aircraft within a range of values limited upwards by a limit value, beyond which the aircraft could become unstable.

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 flight control system utilizing multi-controlled surfaces to provide the necessary control power and feedback logic to stabilize the aircraft with minimal noise amplification. In addition, a unique performance optimization loop is integrated into the stability logic to position the multi-control surfaces for optimum maneuver performance.

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: Strake, flap and canard control surfaces of an aircraft are driven by an error signal which is comprised of a pilot stick command signal and feedback components including measured vertical acceleration and pitch rate of the aircraft and a third feedback component derived from a complementary filter. The filter has a high-pass filter section which operates upon canard position data and a low-pass filter section which operates upon pitch rate. A summation of the signals passing both filtering sections is summed with the vertical acceleration and pitch rate feedback signals to form a combined feedback signal.

Abstract: An aircraft automatic trim system is operable during autopilot control to transfer trim from the elevators to the stabilizers. The thresholds for trim transfer are based on the ratio of elevator demand to available elevator control. Trim commands are produced responsive to this ratio exceeding an amplitude and time dependent limit.

Abstract: An authority limiter includes an input shaft clutched to an output shaft. The output shaft is geared to a motor. Opto-sensors provide control signals to a current source supplying power to the motor. Different amounts of opposing torque are applied by the motor dependent upon the positioning of a disc, mechanically attached to the output shaft, effecting the optical path in the opto-sensors. In another embodiment, the opto-sensor control is replaced by an electro-mechanical mechanism which senses various thresholds corresponding to angular positions of the coupled input and output shafts.

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: Deceleration control apparatus for an aircraft having an automatic altitude capture and hold system and an airspeed hold system, both systems controlled by controlling pitch attitude wherein during a descent from a higher altitude under airspeed-on-pitch control with throttle set at idle thrust, and at some existing negative altitude rate, a slower speed is commanded, a synthetic altitude based on the existing descent rate is computed and the attitude capture and hold system operation is switched into control in place of the airspeed on pitch control to cause the aircraft to flare toward the synthetic altitude. The resultant loss of airspeed, i.e., craft deceleration, during the altitude capture flare is monitored and when the commanded airspeed reduction is achieved, the airspeed-on-pitch control is resumed to thereby hold the commanded lower airspeed.

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: An improved flight control system for initiating and thereafter controlling the automatic capture of a preselected altitude is described. At any point in the flight and in any autopilot pitch submode, the remaining time before an altitude capture is to commence is computed and used as a limiting signal such that the aircraft is precisely aligned with the commanded capture flight path when the capture maneuver commences thereby obviating mode to mode transients. The aircraft's altitude rate is continuously monitored during the capture maneuver and the flight path command is automatically altered between asymptotic and circular paths to assure that the acceleration normal to the flight path will not exceed a predetermined value while at the same time allowing a timely altitude capture.

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: Control force errors are automatically eliminated in aircraft, especially licopters. Such errors may occur because of disturbing forces that act on a control stick (1) for the elevator and/or rotor blade adjustment. The following features are combined. A main control drive (3) for the adjusting of the elevators and/or rotor blades is connected through a control linkage (2) with the control stick (1). A spring box (5) with a pretensioned spring (17) is connected through a control rod (14) to the control stick (2) and to one end of the spring box (5). A trimming rod (15) of a trimming motor (4) is operatively connected to the other end of the spring box (5). To compensate the disturbing forces caused by the inertial mass forces acting upon the control stick (1) as much as possible there are further provided two displacement pickups (6) and (7) or a single displacement difference forming pickup (16). A first summing circuit (8) receives the output signals from the two pickups (6, 7).

Abstract: A two axis joystick controller (14) produces at least one output signal in relation to pivotal displacement of a member (32) with respect to an intersection of the two axes. The member (32) is pivotally movable on a support (42) with respect to the two axes. The support (42) has a centrally disposed aperture (46). A light source (34) is mounted on the pivotally movable member (32) above the aperture (46) to direct light (47) through the aperture. A light sensor (48) is mounted below the aperture (46) in the support (42) at the intersection of the two axes to receive the light (47) from the light source (34) directed through the aperture (46). The light sensor (48) produces at least one output signal related to a location on the sensor (48) at which the light (47) from the light source strikes the sensor (48). The two axis controller (14) is small enough in size to be mounted on another, hand-operated X-Y controller (10) for thumb operation by a hand positioned to operate the other X-Y controller (10).

Abstract: Aircraft attitude along the pitch axis is controlled by inputting a velocity command to a pitch control subsystem that includes a velocity side channel and a gain side channel of a pitch servo amplifier. An output from each channel of the pitch servo amplifier is applied to a pitch servo motor for positioning a primary control surface. The velocity of the pitch servo motor is detected by a velocity detector that generates a velocity feedback signal and control signals for each channel of the pitch servo amplifier. Also responsive to the velocity signal of the detector is a pre-flight test circuit for pre-flight checkout of the pitch control subsystem. In each line from the output of the channels of the pitch servo amplifier is a current detector that is part of a pitch trim control subsystem that responds to the current through the pitch servo motor to generate a trim velocity command signal. This velocity command signal is input to a velocity side channel and a gain side channel of a trim servo amplifier.

Abstract: A maneuver force gradient system causes a helicopter, that otherwise tends to pitch up in a banked turn, to pitch nose-down. A command signal (25) is provided as a function of the roll angle to operate the longitudinal trim actuator (27) which automatically moves the cyclic control (50) resiliently (28) forward to push the nose down. The pilot must consequently pull back on the cyclic control (50) to achieve a desired pitch attitude, thereby establishing a longitudinal positive maneuver-force gradient. The system is operable only when the pilot initiates a roll (33, 38, 50, 51) and the roll angle equals or exceeds 30.degree. (30, 33, 36, 37). Both analog (FIG. 1) and digital (FIG. 2) embodiments are disclosed, and the invention may be practiced in association with an AFCS (101) having the longitudinal trim actuator (27) and resilient linkage (28).

Abstract: Control apparatus of the kind disclosed in our U.K. patent application No. 2173887 wherein the `mover` or rotor of a linear or rotary electrical stepper motor is coupled to a manual operating member so that, not only can the motor drive the operating member to give an automatic mode of the apparatus, but can exert a controlled resistance to manual movement of the operating member thereby giving a degree of `feel` to the operator and/or one or more electrically generated discreet detent positions within the range of movement. Herein the apparatus incorporates an improvement whereby the movable operating member is pivotably mounted so that the manual movement is arcuate--flexible cables connected to the drive motor mover and/or control switches incorporated in the operating member can pass from the operating member to a fixed part of the apparatus at or near the pivotable mounting where the movement to be taken up is less.

Abstract: An altitude preselect apparatus for use with or incorporation into an automatic flight control system combines the incremental altitude signal from a pneumatic encoding altimeter with a vertical speed computation to generate a smooth, continuous instantaneous altitude signal. A vertical speed reference signal is generated to control the vertical flight path of the aircraft when the instantaneous altitude reaches a prescribed proximity to a desired flying altitude which is manually selected by the pilot. The apparatus initiates a capture sequence to hold the aircraft at the desired altitude once it is reached.

Abstract: A controller is described for limiting the rate and acceleration of a device positioned by at least one actuator which responds to a sequence of position commands separated by time intervals incorporating a subtractor for taking the difference of two consecutive position commands to provide a difference signal in vector form, means for determining the magnitude of the difference signal, a comparator for comparing the magnitude of the difference signal with a predetermined value, a switch for coupling the desired position command to the actuator at times the magnitude of the difference signal is less than or equal to the predetermined value, two dividers, two multipliers, an adder and normalizer for providing a new position command between the past command and the desired position command at times the magnitude of the difference signal exceeds the predetermined value.

Abstract: A method for augmenting a noise-adaptive predictive proportional navigation terminal guidance scheme. In a second order configuration with noise adaptive varying gain parameters calculated as a function of time-to-go, this system implements an optimal control law to augment proportional navigation with a missile acceleration feedback term and a navigation ratio to account for first order autopilot time lag. This configuration results in nearly optimal miss distance behavior, with sensitivity to unmodeled errors being substantially less than that in modern guidance systems. The system is particularly useful for high altitude, minimally maneuvering targets.

Abstract: An automatic aircraft pitch axis flight stabilization system which does not require the use of a gyroscope is described. Stability signals are derived from an altitude signal responsive to atmospheric pressure. Means responsive to rate of change of altitude produces a vertical rate signal and means responsive to rate of change of vertical rate produces a vertical acceleration signal. The vertical speed signal is combined with a command signal which causes the aircraft to fly at constant altitude, at constant vertical speed, or to descend along a glide slope, and the vertical acceleration signal to produce a control signal which operates a means, such as a servomechanism for positioning the vertical control surfaces of the aircraft.

Abstract: In a servoactuator system for an aircraft, a strapdown series servoactuator having an external linkage assembly is coupled between the pilot's control stick and a control surface member. The external linkage assembly includes a differential link and servo link having specially tapered stop surfaces adapted to be varied in dimensions to accommodate a wide variety of installations, which stop surfaces cooperate with fixed stops on the servoactuator housing whereby to provide a variety of pilot authority limits and series servo authority limits depending upon particular installation requirements.

Abstract: An azimuth data stabilization system for a shipborne mechanically scanned arch-track system. A unitary package is provided mounted on the roll and pitch stabilized search-track system platform. The package includes a gyro stabilized turntable with the axis of rotation thereof aligned parallel to the axis of rotation of the search-track system. A gyro-torquer loop is provided to inertially stabilize the turntable and a synchro or resolver chain is provided to align the turntable with an azimuth reference. Corrected ships heading signals, derived from the angular position of the turntable, are either coupled directly to appropriate data processing equipment or are combined differentially as compensating signals with the azimuth data signals from the search-track system to provide a true azimuth data output.

Abstract: In an automatic flight control system of the type having automatic stabilizer trim control, the long term, constant altitude cruise trim threshold limits which alter the maximum and minimum pull-in/drop-out deflection of the elevator relative to the stabilizer minimize the steady-state downward deflection of the elevator, thereby reducing unnecessary aerodynamic drag and realizing significant savings in fuel costs.

Abstract: Strapdown, multifunction actuator apparatus comprising one or two integral units adapted for installation in an aircraft, particularly a helicopter, and coupled between the pilot's control stick linkage and the aircraft control surface (or surface servo boost) linkage to perform the functions of series actuation, trim actuation, artificial feel, control position sensing, and control authority limits. The apparatus includes a series actuator installed in the vehicle and connected to the vehicle control linkages in the conventional parallel actuator manner; that is, it is secured directly or strapped down to the airframe and therefore greatly simplifies control system installation and reduces problems such as those associated with control rod vibration resonances normally encountered with conventional integral-with-linkage series actuator installations.

Abstract: A helicopter elevator (12, 13) is driven by an electrical actuator (20) to position the elevator (12, 13) to an optimum angle of incidence for given flight conditions. The actuator (20) is included within a servo loop which responds to a command signal derived from an electronics module (69). Transducers which measure longitudinal cyclic position (40), collective pitch position (41) and air speed (42) generate signals which are summed in a summation unit (43). The output of the summation unit (43) is passed through a signal multiplier (45) which has the gain thereof set inversely proportional to air speed. A trigger unit (56) activates the signal multiplier (45) at an air speed above a preset threshold. An integrator (54) integrates the sum of the outputs of the longitudinal cyclic position transducer (40) and air speed transducer (42) to compensate for different center of gravity aircraft loadings.

Abstract: A pilot controlled stability control system that employs direct lift control (spoiler control) with elevator control to control the flight path angle of an aircraft. A computer on the aircraft generates an elevator control signal and a spoiler control signal, using a pilot-controlled pitch control signal and pitch rate, vertical velocity, roll angle, groundspeed, engine pressure ratio and vertical acceleration signals which are generated on the aircraft. The direct lift control by the aircraft spoilers improves the response of the aircraft flight path angle and provides short term flight path stabilization against environmental disturbances.

Abstract: A single actuator unit is installed in an aircraft, for example, a helicopter, and is coupled to the control linkage from the pilot's stick to perform the functions of series actuation, trim actuation, artificial feel, control position sensing, control limits and, if needed, force boost. Although the device performs as a series actuator, it is installed in and connected to the vehicle surface control linkages in the conventional parallel actuator manner and therefore greatly simplifies control system installation and reduces problems associated with control rod vibration resonances normally encountered with series actuator installations.

Abstract: The outputs of a pair of linear accelerometers are utilized to provide measures of craft vertical acceleration and body axis pitch angular acceleration. The pitch angular acceleration signal is integrated to provide a measure of body axis pitch rate. The pitch rate signal is in turn integrated to provide a measure of body axis pitch attitude. An earth referenced pitch attitude sensor provides a measure of earth referenced pitch attitude and through a rate taker a measure of earth referenced pitch rate. The earth referenced sensor is utilized to calibrate out errors and effectively align the linear accelerometers so that in the event of a loss of the earth referenced sensor, the calibrated accelerometers have adequate accuracy to compute pitch rate and pitch attitude within the long period bandwidth needed to provide the stability augmentation function for a relaxed static stability aircraft.

Abstract: A system for generating an optimum vertical speed command and controlling the aircraft to this command by allowing the aircraft to assume an optium trimmed condition for the thrust energy available which system utilizes the aircraft autopilot system.

Abstract: The altitude acquisition and control system cooperates in an aircraft flight controller for permitting a previously selected altitude to be quickly captured without excessive acceleration of the craft normal to the flight path, even though the human pilot may have neglected to arm the automatic capture mode of the apparatus until after the time that the automatic capture mode is normally commanded.

Abstract: Aircraft altitude is maintained through control surfaces responsive to an altitude hold control voltage varying with the output of a pressure transducer. The pressure transducer provides an output varying in a linear relationship with absolute atmospheric pressure and this output is applied to the input of a nonlinear circuit that generates a voltage varying linearly with altitude. A digital-to-analog converter generates an altitude reference voltage in response to an altitude hold engage signal and this voltage is combined in a summing amplifier with the voltage varying linearly with altitude. Also summed with the altitude reference voltage and the voltage varying linearly with altitude in the summing amplifier is a standard reference voltage. The total summation that is made in the summing amplifier produces the altitude hold control voltage as an output thereof.

Abstract: The linear signal output range of a rotary transformer type device, such as a sine/cosine synchro, is considerably extended beyond its normal 45.degree. linear range by electronic circuit means. Specifically, the electronic circuit means utilizes the sine signal, which is generally linear from 0.degree. to 45.degree., as a portion of the output signal and utilizes a resultant signal computed from the cosine signal, which is linear from 45.degree. to approximately 135.degree. for the remaining portion of the output signal.

Abstract: A multiple function generator is described in the environment of an airspeed indicator for controlling the maximum allowable airspeed needle in accordance with a plurality of selectable functions of altitude and comprises a cam having a plurality of surface segments shaped in accordance with the functions positioning a cam follower coupled to the maximum allowable airspeed needle. A null-seeking electro-mechanical feedback servo mechanism having an altitude signal applied to its input synchro drives the cam, thereby positioning the needle in accordance with the function of altitude represented by the cam surface segment in contact with the follower. Means are included for altering the null index of the servo thereby selectively engaging the desired cam functionsurface segment with the cam follower in accordance with the function desired. A selection knob is coupled with the indexing means for selecting the function.

Abstract: A throttle governor/collective pitch control apparatus for radio controlled model helicopters for proportionally controlling either the model helicopter rotor speed or the rotor collective pitch, including sensing and timing means determining the rotor speed and any changes thereof, comparison means for comparing a subsequently sensed rotor speed with a first sensed rotor speed and developing an error signal, and control means responsive to the error signal to provide a control signal proportional to any changes in the rotor speed. In the throttle governor mode any variations in rotor speed result in a proportional control signal to vary the model helicopter throttle so as to maintain constant rotor speed. In the auto-collective mode changes in rotor speed result in a proportional control signal coupled to the collective pitch servo for proportionally varying the model helicopter collective pitch.