Abstract: The present invention relates to a device for reducing the vibration in the cabin (2) of a rotary-wing aircraft, especially a helicopter (He).According to the invention, said device (1) comprises at least one sensor (C) arranged on the structure of the aircraft (He) and capable of measuring the values of at least one parameter that represents the vibration in said cabin (2); at least one controllable resonator (R), mounted directly in said cabin (2) and capable therein of generating a force which opposes said vibration; and a control unit (UC) capable of controlling said resonator (R) on the basis of the values measured by said sensor (C), so as to reduce and minimize said vibration.

Abstract: The stress test rig consists of a rigidly mounted part which contains the devices uniformly used for all types of test gears and an assembly truck (37) which holds the test gear (16) and the required adapter devices. A test rig gear unit (2) forms a continuous stress circuit with a superimposed gear unit (9), an adaptation gear (23) and a second test rig gear (19) wherein the stress torque is set by a control unit (15) through the superimposed gear unit (9). The regulation variable for the stress torque is determined up to more than 80% by a control network (31). The test rig gear unit (2) is a bevel gear wheel and the superimposed gear unit (9) has a bevel gear step (13) so that the output shaft (14) of the superimposed gear unit (9) is pivoted and has a height adapted to an input shaft of an adaptation gear unit (23) or of a test gear unit (16).

Abstract: The invention is directed to a helicopter health and usage monitoring system utilizing a neural network for estimating gross weight and center of gravity location from measured flight condition parameter inputs; and includes means for measuring a plurality of variable flight condition parameters during flight of the helicopter; memory means for successively receiving and storing parameter input signals as well as estimates of gross weight and center of gravity location; and processing means responsive to the signals received from the measurement means for generating the gross weight and center of gravity location estimates.

Abstract: A flight control indicator for twin-engine aircraft, in particular a rotary-wing aircraft comprising two turboengines for driving at least one rotor, intended to provide the power margin available on the engines of the aircraft as a function of the flight conditions. Display means (6) presents a single dial (8) furnished with two needles (9.sub.1, 9.sub.2), each of them corresponding to a respective engine, and positioned in such a way that, when the engines are operating normally, the two needles are merged.

Abstract: A horizontal speed indicator for a rotary-wing aircraft includes first and second sensors for determining, respectively, the heading and the horizontal speed of the aircraft. The indicator processes the signals delivered by the first and second sensors, and displays the processed signals on a screen, which includes a compass rose which can rotate when the aircraft alters its heading, and a scale which represents the horizontal speed of the aircraft. The scale includes cross wires and concentric circles, and is capable of moving with respect to the compass rose.

Abstract: Gravitation in three dimensions is measured using an accelerometer attached to a carrier, a satellite navigation receiver and a computer. The satellite navigation receiver determines a position and attitude, and changes in the position and attitude over time. The computer calculates complete kinematics in inertial space from the timely changes of the determined position and attitude and calculates the gravitation by subtracting the kinematic acceleration from the one observed by the accelerometer. In a preferred embodiment the accelerometer is connected tightly to the carrier. At most, damping elements may be mounted between the accelerometer and the carrier. The accelerometer can be a single accelerometer or a triplet of non-parallel accelerometers. The satellite navigation receiver may be one instrument equipped with three or more antennae or a set of three or more single instruments connected to each other.

Abstract: A flight indicator for an aircraft, intended to give the power margin available on at least one engine of the aircraft as a function of the flight conditions. The indicator includes sensors (2, 3, 4) capable of delivering data relating to various monitored parameters of the engine, device (5) for processing the data coming from the sensors (2, 3, 4), and display device (6) showing, on a display screen (7) the processed data relating to that parameter among the monitored engine parameters, whose current value is closest to the limiting value for such parameter, each limiting value being established taking the ambient pressure (P0) and ambient temperature (T0) into account, in real time.

Abstract: The invention is directed to a system, utilizing a neural network, for estimating helicopter airspeed in the low airspeed flight range of below about 50 knots using only fixed system parameters as inputs to the neural network.

Abstract: The present invention relates to an aircraft flight indicator intended to supply a data item regarding the power margin available on at least one engine of the aircraft as a function of the flight conditions. According to the invention, the indicator comprises: a sensor (2) capable of delivering data relating to the speed (Ng) of the gas generator of the engine; means (3) for calculating, using the data relating to the speed (Ng) of the gas generator, a data item .DELTA.Ng representing the difference between the actual value of Ng and the reference value on take-off and for processing said data so that it can be displayed; and display means (4) showing the data relating to the values of the speed Ng of the gas generator and/or of .DELTA.Ng on a display screen (5).

Abstract: An airspeed and direction indicating system for an aircraft of the rotary wing type includes a vertical accelerometer for monitoring the up and down movement of an outer portion or tip of the rotary wing. The vertical accelerometer produces a generally sinusoidal signal in response to the up and down movement of the wing tip as it rotates about an axis. The amplitude of the signal corresponds to the airspeed and the phase angle indicates direction. Suitable readouts display the information for consideration by a pilot.

Abstract: The present invention provides ground proximity warning system and method for aircraft, particularly rotary wing aircraft such as helicopters. The warning system generally includes a barometric altitude rate detector, such an altimeter and a rate circuit, disposed to detect the rate of change of the barometric altitude of the aircraft. A controller is coupled to the detector for receiving the barometric altitude rate signals and adjusting the signals based on the ground effect from the main rotors when the aircraft is near the ground. Typically, the controller decreases the barometric altitude rate signals to compensate for the apparent increase in barometric altitude rate caused by ground effects. This reduces the number of nuisance warnings that may otherwise occur during takeoffs, landings or other low altitude flight conditions.

Abstract: A method and apparatus for determining the airspeed of a rotary wing aircraft having a sustaining rotor with a plurality of radially extending rotor blades for rotation about a substantially vertical axis comprises the steps of measuring an airflow sideslip angle established during rotation of each rotor blade by the vector of a radial flow component due to the relative wind vector and the rotational velocity perpendicular to a blade feathering axis, producing a signal representative of the sideslip angle, comparing the signal in a processor containing information on a known relationship between the sideslip angle and airspeed, providing an output signal representative of the airspeed and wind direction to a display which may display both the airspeed and wind direction.

Abstract: The invention is directed to a method, utilizing a neural network, for estimating helicopter airspeed in the low airspeed flight range of below about 50 knots using only fixed system parameters as inputs to the neural network.

Abstract: The invention relates to novel basic architecture for helicopter autopilots, transforming the conventional objective of maintaining fuselage trim to an objective of maintaining airspeed, and without making use of coupler techniques. On each of the pitch and roll control systems, a known fraction of the output signal (24S) from the flight control is injected by being summed with the standard fuselage trim input (1S). The signal (4S) formed in this way possesses the novel static and dynamic properties required for the control loop to make the helicopter statically stable in speed throughout its flight envelope. The novel architecture makes it possible to replace the conventional fuselage trim signal (1S) with an accelerometer signal. The signal (4S) is displayed in particular on a standard "artificial horizon" type instrument (33), thereby making it possible simultaneously to monitor rapid variations of fuselage trim and to control airspeeds.

Abstract: The rig comprises a vertical electric motor in a shell-type framework carrying a force balance itself supporting either a main rotor in dynamic tests or a complete propulsion assembly with main rotor in endurance tests, or a height extender supporting, on the one hand, an upper shaft line being coupled to a lower shaft line supported in the framework and driven by the motor and, on the other hand, an inverse-thrust rotor mast, driving a hub jig linked to blades mounted inverted so as to exercise a thrust downwards. The mast is linked to the shaft line by a torque meter and is mounted rotating in a bearing supported by a second balance fixed to the height extender. The mast and the upper shaft line are removable with the height extender in order to make the rig multi-purpose and useable for performance, endurance and dynamic tests of helicopter rotors.

Abstract: Static pressure sensing and free airstream temperature sensing is obviated in a method for determining aircraft velocity relative to an airmass by utilizing a static pressure and temperature component present in pressure variations sensed by a total pressure sensor located on a rotating arm mounted on the fuselage of an aircraft. A harmonic analysis of a quasi-sinusoidal total pressure variation includes a determination of steady state and both first and second harmonic components. These harmonic components are used to determine the static pressure, the free airstream temperature and the aircraft airspeed without separately sensing these parameters.

Abstract: A bearing assembly monitoring system provides continuous, real time monitoring of the operating condition of a powertrain support bearing assembly and real time alarm signals indicative of degradation and/or impending catastrophic failure thereof. One embodiment of the monitoring system includes a vibration sensing device, first and second primary temperature sensing devices, an ambient temperature sensing device, vibration and temperature signal processing subcircuits, and an alarm subsystem. The vibration sensing device monitors complex vibration signals emanating from the bearing assembly and provides signals corresponding thereto. The vibration signal processing circuit filters such complex vibration signals to provide a predetermined bandwidth thereof that are converted to stabilized DC voltage signals which are compared to reference voltages corresponding to higher than normal and excessive vibration levels. Exceedance of such reference voltages generates a corresponding vibration alarm signal.

Abstract: An apparatus for the detection of the icing-up of an aircraft rotor. The apparatus has a sensor (10) for the tractive force and a sensor (11) for torque; for measuring respectively the axial force and the torque exerted on the shaft of the rotor. A device (12) calculates the ratio between these two quantities. A determination is also made, as a function of the flying conditions, of what should be the nominal value of this ratio in the absence of icing-up. Finally a device (12) compares the calculated ratio with the nominal ratio.

Abstract: A pattern of light is projected from multiple sources located on an aircraft to form two clusters. The pattern of each cluster changes as the aircraft flies above and below a predetermined nominal altitude. The initial patterns are two horizontal, spaced apart lines. Each is capable of changing to a delta formation as either the altitude or the terrain varies. The direction of the delta cues the pilot as to the direction of corrective action.

Abstract: The invention is a helicopter blade crack detection system for blades that are hollow and pressurized or evacuated to a low pressure A rotating assembly having a rotating indicator assembly photo-optically communicates to an airframe detection assembly providing high reliability, ease of maintenance, and EMI secured.

Abstract: A power management symbology (PMS) display system for helicopters that provides a continually updated, combined analog/digital symbolic display of selected status information vis-a-vis the helicopter powerplant. The PMS display system includes a multi-dimensional, continuously updated, combined analog/digital display symbol that is optimized for discriminability and compatibility with respect to the pilot's visual system.

Abstract: A method and device are disclosed for determining the airspeed of a helicoper, in which, for a plurality of flight configurations of a calibration flight, magnitudes for estimating the airspeed and the collective P.sub.O and cyclic pitches P.sub.X and P.sub.Y of the lift rotor, the pitch P.sub.Z of the antitorque rotor, the angles of bank .theta..sub.X and .theta..sub.Y and the acceleration .gamma..sub.Z are measured. Correlation matrices between the speed estimated with respect to the air and variables related to the pitches, angles and acceleration measured during the calibration flight are then formed. Formulae are derived, for calculating the estimated speed, by a stepwise regression algorithm, which are used in normal flight for calculating, in response to the pitches, angles and acceleration measured during normal flight, the speed with respect to the air, if required after correction of the values of the multiplicative coefficients of these calculation formulae.

Abstract: A device and method of back-up fly by wire control, especially useful in fly by wire systems when a failure mode is encountered. The problem of back-up system controllability with the type of compliant sidestick which has minimal displacement and returns to zero upon release of the control force is solved by reducing the authority of the sidestick to a fraction of full authority actuator position capability, to provide good control sensitivity. Full authority actuator position control is provided by automatically activating a trim integrator when the sidestick reaches its displacement stops.

Abstract: A warning system for rotary wing aircraft that monitors the altitude above ground of the aircraft and issues a warning when there is insufficient terrain clearance. The warning criteria are optimized for rotary wing aircraft and are altered as a function of landing gear position and airspeed. Two distinct warnings are given. One indicates insufficient terrain clearance and the other warns the pilot if he attempts to land with the landing gear up. Logic circuitry is provided to modify the criteria necessary to generate a warning, and to inhibit certain warnings, as required by the mode of operation of the aircraft.

Abstract: An autonomous kinematic measurement system for helicopters is disclosed. A helicopter comprises a lifting rotor, an anti-torque rotor, first means to control the cyclical pitch and the collective pitch of the rotor and second means to control the pitch of the anti-torque rotor. The system comprises means to measure the helicopter's ground speed, means to measure the longitudinal cyclical pitch of the lifting rotor blades, means to measure the lateral cyclical pitch of the lifting rotor blades, means to measure the collective pitch given to the lifting rotor blades, and means to measure the pitch of the anti-torque rotor. An inertial unit, provided with ancillary computing means, uses a direction signal and a helicopter attitude signal corresponding to the longitudinal acceleration signal and lateral acceleration signal of the helicopter's ground speed signal to establish the helicopter's longitudinal and lateral airspeed values, following a determination of the coefficients .alpha.,.beta.

Abstract: Visually observable safe limits for the proximity of an aircraft, such as a helicopter, to obstructions are provided by pairs of lights which provide collimated beams intersecting at limit points located radially outwardly of the arc of the helicopter rotor blades so that light scattered from the intersecting beams at the limit points provides the helicopter pilot with visual references of the rotor arc. Each light of each pair has a concave reflector and a lamp at the focal point of the reflector to provide the collimated beams and transparent plastic shields over the open ends of the reflectors have concavities in central portions to provide diverging beams that illuminate the obstructions.

Abstract: Aircraft speed and direction are determined by sampling the sinusoid-like pressure variations at the end of a rotor and performing a Fourier analysis on the pressure samples. Fourier quadrature component signals are used to determine aircraft direction and are also used to determine, along with a steady state component, the aircraft's total velocity. A discrete Fourier analysis may be performed on the pressure samples using a plurality of subrevolution pressure samples averaged over each subrevolution interval.

Abstract: A measuring device (1, 12, 13, 14) is equipped with a blade angle indicator (1) indicating the blade angle adjustment of the rotor blades of a helicopter, a pressure altimeter (13) and an outside temperature probe (14). The measuring device (1, 12, 13, 14) is linked to a control device (4-6) which has a central processing unit (4) and a program memory (5). The memory (5) stores data processing instructions for determining density height from pressure height and outside temperature as well as data processing instructions corresponding to the helicopter operating instructions for determining the permissible maximum value of the blade angle from the density height. The control device (4-6) periodically determines the density height on the basis of the prevailing pressure height and outside temperature and the corresponding permissible maximum value of the blade angle.

Abstract: A visual alarm suitable for use with night vision goggles comprises a red incandescent lamp (105); a blue, green or blue-green electroluminscent panel (130); and control means to sequentially activate the above-mentioned light sources. Brief activation of the red lamp causes momentary partial blooming of the night vision goggles to warn of an emergency condition. Subsequent steady energization of the electroluminescent panel provides a continuous observable alarm which does not interfere with operation of the night vision goggles.

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.

Abstract: A helicopter is provided with a lift rotor (Ros), a tail rotor (Rac), first control means for the longitudinal (Pcx) and lateral (Pcy) cyclic pitch of the lift rotor and second control means for the pitch (Pac) of the tail rotor. The parameters such as the longitudinal cyclic pitch (Pax), the lateral cyclic pitch (Pcy), the longitudinal acceleration (Vx) and the acceleration of the helicopter (Vy) as well as the collective pitch (Pcol) of the lift rotor (Ros) and the pitch of the tail rotor (Rac) are measured. The longitudinal (Vax) and lateral (Vay) speed of the helicopter with respect to the air are determined from a biunivocal function of the measured values, particularly in linear combination.

Abstract: A helicopter weight and torque advisory system for connection to cargo hook load sensors, fuel flow, air temperature and altitude indicating systems of a helicopter. The advisory system provides the flight crew of the helicopter information as to available engine torque, gross weight, gross weight center of gravity condition, and weight supported by each cargo hook as an external sling load is raised and supported by the helicopter. Also information as to required engine torque for maximum helicopter range and endurance performance is provided.

Abstract: A system for warning the pilot of a rotary wing aircraft such as a helicopter of a flight condition that, if left uncorrected, would result in the tail of the aircraft striking the ground monitors the altitude above ground, descent rate, pitch and pitch rate of the aircraft and provides a warning if the relationship between the aforementioned parameters is such that a tail strike is imminent.

Abstract: A method for determining the horizontal airspeed of a helicopter in the high speed (v.sub.e >20 m/s) range. Pitch angle and the longitudinal control signals are recorded by sensors such as an inertial navigation system and a potentiometer connected to the pilot's control lever. Equivalent speed signals are calculated from the measured pitch angle and longitudinal control signals by means of calibration equations. These signals are then filtered and weighted. The horizontal airspeed is then derived from the sum of the weighted signals divided by the sum of weighting factors k.sub..theta. and k.sub..delta.x determined in accordance with the sensed pitch angle and longitudinal control signals. The horizontal velocity signal may be filtered, if necessary.

Abstract: When helicopters are accomplishing certain special maneuvers such as powered descent, or the like involving low "g" conditions, the helicopter blades may teeter excessively with respect to the mast, and under severe conditions, the mast may shear. To sense such low "g" conditions, which may be dangerous, an accelerometer is mounted on the helicopter frame close to the center of gravity of the helicopter and is coupled to data processing circuitry which classifies the low-g conditions in bands or levels of severity, recording the duration of exceedances in which the low-g conditions are below predetermined levels, and issues warning signals to the helicopter pilot through the helicopter intercom system and a suitable warning light.

Abstract: The disclosure relates to a helicopter (5) having a flight path computer to provide signals for pilot or automatic flying in accordance with a predetermined flight path, a forward looking infra-red camera (6) for scanning ahead and to either side of the flight for an object such as a person in the sea, a screen display for the camera and a control to be operated when the helicopter over-flies the person to cause the flight path computer to provide control signals to take the helicopter through a circuit down-wind of the person and then back into-wind at pre-selected heights and speed towards the location of the person and then to hover over the person.

Abstract: An improved system and technique is disclosed which enables an aircraft control system to bring an aircraft to hover with no direct measurement of ground speed. The system includes a device for numerically calculating a value for ground speed based on a differentiation of aircraft position as measured from conventional ground reference navigational aids. The computed value of ground speed is compared with values representing heading and true air speed to generate a computed value of wind. The computed value of wind is then resolved into vector components of headwind and crosswind and the headwind component and air speed are summed to provide a signal used to control aircraft speed until zero ground speed is achieved.

Abstract: The power required for a helicopter to hover is generated (14, 82) as the ratio of current operating power in forward flight (12, 77) determined (10, 73) from data relating operating power in forward flight to power required for hover for the aircraft. The power required to hover is compared (18, 83) with the maximum power available developed (16, FIG. 2; FIG. 3) by an engine model algorithm utilizing actual engine parameters. The comparison of maximum power to power required for hover is utilized to provide an indication (22) to the pilot. The viability of the indication is indicated by a "ready" indication (26).

Abstract: A method is disclosed for determining the horizontal airspeed of helicopters in low speed ranges. At least three control functions, .theta..sub.o,i (collective pitch), .theta..sub.s,i (longitudinal control) and .theta..sub.c,i (lateral control) are measured. The magnitude of the horizontal airspeed, V.sub.H,i, is determined from its known relationship to .theta..sub.o,i. This magnitude is included in non-linear calibration equations .theta..sub.s (V.sub.H, .psi.) and .theta..sub.c (V.sub.H, .psi.) and control profiles .theta..sub.s (V.sub.H,i, .psi.) and .theta..sub.c (V.sub.H,i, .psi.) are derived therefrom. The measured values .theta..sub.s,i and .theta..sub.c,i are applied to the respective control profiles and corresponding phase pairs .psi..sub.s,1, .psi..sub.s,2 and .psi..sub.c,1, and .psi..sub.c,2 located. The pairs are compared and the incident flow angle .psi..sub.i is determined as the value common to the two pairs.

Abstract: The invention relates to the technical field of pilotage aids for helicopters. In a speed measuring device having two measuring devices for longitudinal speed and lateral speed of the helicopter, each measuring device comprises a single sensor (9a, 9b) for sensing the position of the cyclic rotor pitch control along the axis concerned, and for sensing the acceleration component along the same axis. Each measuring device further includes a computer to provide the air speed components of the helicopter. Applications include flight testing, and the construction of firing computers and flight computers.

Abstract: A hub moment sensor is disclosed for an aircraft having a horizontal rotor, a rotor fixed mast for supporting the rotor above the aircraft body, and a support ring formed on the aircraft body for attachment to the rotor fixed mast. The sensor comprises an annular flange formed on the end of the rotor fixed mast remote from the rotor, and coupling means for securing the flange to the support ring, the coupling means including stress measuring means, whereby the moments acting on the aircraft may be sensed.

Abstract: A warning system for providing an aural warning to the pilot of a rotary wing aircraft monitors the altitude above ground and barometric descent rate of the aircraft and generates the warning if the descent rate of the aircraft is excessive for the altitude at which the aircraft is flying. The system provides two distinct warnings, one being an early warning advising the pilot of a hazardous condition, and the other being a warning alerting the pilot of imminent danger. The system also includes tactical and non-tactical modes of operation, wherein the warning criteria are adjusted to meet tactical and non-tactical flight conditions.

Abstract: A warning device, for a helicopter with a tail rotor and a mechanical protection device (15) therefor, for giving warning if the helicopter tail approaches too closely the ground or an obstruction. The warning device comprises a height-finder (11) with a transmitting/receiving antenna (26) mounted at the helicopter tail to produce a height-finding beam backwards and obliquely downwards. Height signals are processed by computing means (12) to determine the sink rate of the tail and in dependence upon this sink rate and the instantaneous height, a warning signal is initiated if this corresponds to a predetermined relative relationship. The higher the sink rate the greater is the height at which the warning signal is initiated and vice versa.In the preferred arrangement the beam is directed substantially along or parallel to the axis of the mechanical protection device (15).

Abstract: An overload warning system for fighter type aircraft combines analytical techniques of computing structural loads with flight conditions and aircraft control inputs to calculate the allowable load factor/control input on a real time basis. The flight conditions and aircraft control inputs are monitored by a microprocessor, which continuously monitors actual load factor, compares the answers with a computed allowable value and generates an output signal whenever a prescribed value is exceeded. This signal is available for transmission to aural or visual warning devices and for permanent recording.

Abstract: In an excessive descent-rate warning systems particularly useful for helicopters, a descent-rate signal is produced corresponding to the rate of change of altitude, and is used for producing a warning signal whenever the descent-rate exceeds the maximum one for the respective altitude according to a predetermined relationship which relationship includes an offset value defining the maximum descent rate at "0" altitude. An increase in the forward-speed of the aircraft is used to reduce the offset in the above predetermined relationship, and thereby to decrease the maximum descent rate at "0" altitude producing the warning signal. A described system may be operated according to a "Take-Off Mode" effective upon take-off to cause a warning signal to be produced upon any descent of the aircraft until it reaches a preselected altitude, at which time the system is automatically actuated to operate according to the above-described Flight Mode.

Abstract: The invention resides in apparatus for directly measuring density altitude in an aircraft and includes a housing in communication with ambient conditions of air temperature, pressure and humidity outside the aircraft. A rotor having dynamically balanced blade elements is positioned rotatably within the housing. An electric motor is connected so as to drive the rotor within the housing against the resistance therein of the air under the ambient conditions of temperature, pressure and humidity. An operational characteristic of the motor, such as the current which it draws, is continuously sampled, amplified and changed in shape and applied as the input to a display device which is calibrated to display the input in terms of density altitude.

Abstract: The factors (including collective pitch, weight, altitude density, rotor speed and airspeed) which determine the extent to which the rotor of a helicopter may be loaded, are fed into a look-up table, and the permissible acceleration normal to the rotor is compared against actual acceleration normal to the rotor, an alarm signal resulting whenever the actual exceeds the permissive; the alarm signal may be utilized to shake the collective stick so as to simulate pre-overload vibration as a warning to the pilot.

Abstract: A system for detecting ice on exterior surfaces of aircraft by transmitting a relatively low power microwave electromagnetic signal into a dielectric layer functioning as a surface waveguide, and monitoring the signals transmitted into and reflected from the waveguide. The waveguide includes a termination element which is mismatched with the waveguide impedance, resulting in partial or total reflection of the microwave energy from the remote end of the waveguide. As ice builds up on the surface waveguide, the impedance or reflection characteristics of the composite waveguide comprising the ice layer and the permanent surface waveguide give a reliable indication of the presence and location of the ice. The reflection characteristics are conventionally monitored utilizing a dual directional coupler and a reflectometer.

Abstract: An omnidirectional cockpit display has perpendicular bars independently moving across a plate for providing indications of speed in lateral, forward and rearward directions. A zero point is centered between lateral edges of the plate and is displaced vertically from the center of the plate to provide greater speed indications in a forward direction than in a rearward direction. A safe flight envelope is indicated on the plate, extending around the zero point in lateral and rearward directions and extending toward the edges of the plate in the forward flight direction. Indicator scales are mounted along edges of the plate, and pointers move along the scales to indicate vector sum of component speeds, and density altitude. Other scales and pointers indicate ground speed in orthogonal directions, while the crossbars indicate air speed in those directions. Other peripheral scales and pointers indicate directional control position, collective thrust control position, rate of climb and rate of closure.

Abstract: Power required curves for ratios of helicopter weight to air density ratio form a background over which move perpendicular bars representing air speed and engine torque. The intersection of the relatively movable bars indicates which power required curve the helicopter is operating on. An area to the right and above that curve represents power margin available for maneuvering. A knob positions a mask over a portion of the background to indicate the actual maximum power limitation of the helicopter. A maximum power setting is varied according to ambient pressure and temperature. A minimum speed warning is produced by comparing air speed and engine torque. A rapid descent automatically raises the minimum speed warning.