Abstract: An aircraft includes a display, a support system, and an avionics system. The support system includes at least one operable component, where at least one of the support system and the at least one operable component is configured to operate with an operating value falling within a predefined normal operating range. The avionics system is programmed to: calculate a rate of change of a value of at least one of the support system and the at least one operable component; determine whether the value will exit a predetermined normal operating range within a predetermined time based on the rate of change of the value; and indicate on the display that the value is expected to exit the predetermined normal operating range. The avionics system can leverage this function to declutter displays, and then present additional information to the crew when a value is trending out of range.

Abstract: A technique for controlling a rotating stall in a compressor of a gas turbine engine. In the technique a flow injection is introduced into an axial air flow path of the compressor via a flow-injection opening located at a pressure side of a guide vane in the compressor and directed towards a leading edge of a compressor rotor blade located adjacently downstream of the guide vane. The flow injection is introduced when the rotating stall is detected and/or when the compressor is being operated at a speed lower than full load speed. The flow injection reduces an angle of incidence of compressor air on the leading edge of the downstream rotor blade and hence the rotor sees a more favorable velocity. The favorable velocity results into an increase in the operating range of the rotor and hence of the compressor by mitigating and/or reducing the rotating stalls.

Abstract: A surge avoidance control method that includes a surge detection step of determining whether the operation point of the compressor is positioned in a surge operation region at each predetermined timing; a surge avoidance opening-degree calculation step calculating a surge avoidance opening degree of the operation device necessary for moving the operation point of the compressor determined to be positioned in the surge operation region in the surge detection step out of the surge operation region; a correction opening-degree calculation step calculating a correction opening degree of the operation device on the basis of a moving velocity of the operation point of the compressor which is determined to be positioned in the surge operation region in the surge detection step; and an opening-degree command value calculation step calculating an opening-degree command value of the operation device on the basis of the surge avoidance opening degree and the correction opening degree.

Abstract: A cyber-life electronic commerce and system control device is disclosed having a three dimensional or other image of an area stored on a server. An operating exchange interfaces with the image to enable an individual to manipulate virtual control elements within the operating exchange. A visible light embedded communication system is used as the backbone to communicate commands from the operating exchange to a building operating system control item in order to alter the setting or status of a building operating system. The operating exchange may also be used by a consumer to virtually explore a cyber-retail location and to engage in commercial activities.

Abstract: Cockpit display systems and methods are provided for performing Glide Slope (G/S) validation processes during Instrument Landing System (ILS) approaches. In one embodiment, the cockpit display system utilizes validated G/S signals to selectively correct the viewpoint of a Synthetic Vision System (SVS) scene generated on a Synthetic Vision Primary Flight Display (SV-FPD). In such an embodiment, the cockpit display system may include an ILS receiver, a cockpit display device on which the SV-PFD is generated, and a controller operably coupled to the cockpit display device and to the ILS receiver. During an ILS approach, the controller selectively performs a G/S validation algorithm to determine the validity of the G/S signals received during the ILS approach. If determining that the G/S signals are valid, the controller then repeatedly updates the SVS viewpoint during the ILS approach based, at least in part, on the validated G/S signals.

Abstract: Methods are provided for identifying degradation in components of a compressor recirculation valve (CRV). One method may include inferring degradation of the CRV based on adaptation of a compressor surge line outside an expected range.

Abstract: An aircraft safety system includes a Stall/Spin Warning & Recovery Computer, an angle of attack sensor, a side slip sensor and an angle of attack display. The Stall/Spin Warning and Recovery Computer detects whether the aircraft is in a safe condition or an unsafe condition. The angle of attack display includes a first configuration that displays information indicative of safe aircraft condition when the orientation detector detects a safe condition and a second configuration that displays information indicative of unsafe aircraft condition when the computer detects an unsafe condition. The display may transition from the first configuration to the second configuration when the orientation detector detects an unsafe condition and then to a third configuration, which may include visual control-input instructions that enable recovery to a safe condition.

Abstract: To provide a warning system that can issue a stall warning taking a flight environment into account. The warning system according to the present invention is a warning system for an aircraft, for issuing a warning in the case where there is a possibility of the aircraft stalling, and includes a selecting section for selecting one of two or more calculation criteria based on an icing state of the aircraft, a calculating section for calculating a stall angle based on the selected calculation criterion, and a warning section for comparing the calculated stall angle with a current angle of attack of the aircraft and issuing a stall warning if the current angle of attack exceeds the stall angle.

Abstract: A method and system for indicating a potential stall condition for an aircraft during flight. An alert lift coefficient is identified for the aircraft. The alert lift coefficient is adjusted in response to a number of changes in a current state of the aircraft. A set of thresholds is identified for use in generating an alert indicating that the aircraft has reached the potential stall condition using the alert lift coefficient.

Abstract: An electronic multi-function probe (MFP) is provided for positioning adjacent an aircraft skin. The electronic MFP includes a plurality of pressure sensing ports and an electronics housing. Positioned within the electronics housing are air data processing circuitry and stall protection system processing circuitry. The air data processing circuitry is configured to generate air data parameters, including airspeed and altitude, as a function of pressures at the plurality of pressure sensing ports. The stall protection system processing circuitry of the electronic MFP is configured to detect stall conditions as a function of angle of attack and at least one air data parameter from the air data processing circuitry, and to generate stall protection system outputs indicative of the detected stall conditions.

Abstract: A method and system for indicating a potential stall condition for an aircraft during flight. An alert lift coefficient is identified for the aircraft. The alert lift coefficient is adjusted in response to a number of changes in a current state of the aircraft. A set of thresholds is identified for use in generating an alert indicating that the aircraft has reached the potential stall condition using the alert lift coefficient.

Abstract: There is described a method for generating a parameter for airfoil performance monitoring, the method comprising: receiving at least a first signal indicative of a first airflow parameter measured at one of an aft position and a side-facing position of a sensor mast mounted to an airfoil and at least a second signal indicative of a second airflow parameter measured at a fore position of the sensor mast; generating a turbulence intensity ratio for each one of the first airflow parameter and the second airflow parameter; and combining the turbulence intensity ratio of the first airflow parameter and the turbulence intensity ratio of the second airflow parameter to generate a combined turbulence intensity ratio for airfoil performance monitoring.

Abstract: A device for confirming the engine thrust of an aircraft is disclosed. The device, via an operating logic unit (14) of each engine processing different information, including a warning indicator (15) connected to the unit (14) delivers or not thrust confirmation information at the level of the cockpit.

Abstract: A method and device control the thrust of a multi-engine aircraft. At least one engine command is determined by a command determining unit, with the determined command commanding each engine that has not failed to deliver a thrust substantially equal to a reduced thrust value. A command application unit is configured to apply the determined command to each aircraft engine that has not failed. The reduced thrust value is determined, by a weight determining device, according to the current weight of the aircraft and according to a reduced thrust value FOEI, which is calculated by a thrust calculation unit.

Abstract: There is described a method for generating a parameter for airfoil performance monitoring, the method comprising: receiving at least a first signal indicative of a first airflow parameter measured at one of an aft position and a side-facing position of a sensor mast mounted to an airfoil and at least a second signal indicative of a second airflow parameter measured at a fore position of the sensor mast; generating a turbulence intensity ratio for each one of the first airflow parameter and the second airflow parameter; and combining the turbulence intensity ratio of the first airflow parameter and the turbulence intensity ratio of the second airflow parameter to generate a combined turbulence intensity ratio for airfoil performance monitoring.

Abstract: A reference value is arbitrarily selected from a range of possible aircraft rotation speeds. A position of a trimmable horizontal stabilizer is angled in accordance with a centering of the reference value. A deviation between the reference value and an accelerating speed value of the aircraft is determined. Elevators or the horizontal stabilizer are controlled, prior to rotation, in accordance with the determined deviation.

Abstract: A system and method of at least reducing the likelihood of a rotor lock in an aircraft gas turbine engine is provided. A determination is made that each propulsion gas turbine engine on an aircraft has experienced a flameout. The rotational speed of each propulsion gas turbine engine is compared to a predetermined threshold rotational speed and, if the rotational speed of each propulsion gas turbine engine is below the predetermined threshold rotational speed, a warning is supplied to a flight crew.

Abstract: A stall protection system for use with an aircraft may have at least one sensor configured to provide an indication of a pressure differential associated with airflow across an airfoil, and a controller in communication with the at least one sensor. The controller may be configured to compare the pressure differential to a pressure differential threshold associated with a stall condition of the airfoil, and to selectively generate a recovery response signal based on the comparison.

Abstract: Methods and systems for a position indicating display system for an aircraft are provided. The system includes a map display unit configured to display a map representative of an area being traversed by the aircraft, and an overlay comprising an own ship depiction, said overlay displayed on the map for a period of time in response to an input from at least one of a user and an aircraft sensor.

Abstract: An airspace awareness and warning system (“AAWS”) provides input to an airspace alert (“AA”) processor from at least one real-time aircraft system or sensor, a navigation system, and an airspace database containing three-dimensional delineations of defined airspace; the processor determines an airspace clearance surface and an aircraft airspace alert surface, and if one surface penetrates the other, the processor generates an alert signal and provides an alert signal to a crew alerting system. The two surfaces are determined by the processor by executing an algorithm(s) embedded in software containing the disclosed embodiments and methods. At least one criterion used to define an aircraft airspace alert surface is programmed to include real-time and/or static input factor data provided by at least one system or sensor input from an aircraft. Such input factor could be used to define an airspace clearance surface.

Abstract: A method and system for displaying aircraft engine information. In one embodiment, the method includes directing the display of an icon that represents an overall operational state of an aircraft engine, with the overall operational state including one of a started state and an unstarted state. The method can further include directing a change in a displayed characteristic of the icon when the overall operational state of the aircraft engine changes from the one of the started state and the unstarted state to the other of the started state and the unstarted state. In a further embodiment, the method can further include repeatedly directing changes in the display of the icon as the aircraft changes from the unstarted state to the started state.

Abstract: A method and system compares combinations of vehicle or aircraft state variables against known combinations of potentially dangerous states. Alarms and error messages are selectively generated based on such comparisons. In one embodiment, pairs of aircraft state variables are selected and compared to known undesirable pairs of state combinations that indicate an error or a condition that a crew should monitor closely. The combinations and error messages are provided by the database. The comparisons are conducted on a periodic basis monitoring real time states of the parameters collected from various sensors and commands. Experts considering a matrix that provides an exhaustive pairwise comparison of potentially important state variables initially identify undesirable state combinations. Error messages and identification of potential alarms are generated based on both knowledge of actual accidents, and on use of expert knowledge to predict potentially dangerous states.

Abstract: A method for displaying flight information includes displaying a first characteristic sign illustrating a speed vector of an aircraft. The method also includes determining a first longitudinal margin of maneuver of the aircraft as a load factor, and displaying a second characteristic sign such that a position of the second characteristic sign relative to the first characteristic sign illustrates the first longitudinal margin of maneuver.

Abstract: The present invention is directed to novel primary flight displays, moving map displays and engine parameter displays for aircraft cockpits. These displays process data gathered from various aircraft sensors and, using an various electronic databases, generate intuitive symbols that aid the pilot. The present invention also uses various novel algorithms that use data gathered from inexpensive sensors to derive parameters normally requiring far more expensive sensors.

Abstract: A display system for an aircraft includes a central unit configured to determine a speed vector of the aircraft and to determine a longitudinal margin of maneuver of the aircraft. The longitudinal margin of maneuver is expressed as a load factor and is related to one of a pitch-up maneuver and a pitch-down maneuver. The display system also includes a display unit connected to the central unit. The display unit includes a display screen configured to display a first characteristic sign illustrating the speed vector and a second characteristic sign illustrating the longitudinal margin of maneuver.

Abstract: A ground proximity warning system including a pitch angle alerting device, the pitch angle alerting device including warning logic having an input that is structured to receive a signal that is representative of the indicated pitch angle of the aircraft, the warning logic being structured to determine a difference between a pitch angle signal received on the input and a predetermined threshold, the warning logic being structured also to output a signal representative of the difference between the two signals; and an advisory indication generator coupled to accept the output of the warning logic, the advisory indication generator being structured to generate an advisory indication signal as a function of the difference between the two signals.

Abstract: The device (1) for aiding piloting comprises means (2) for determining a speed vector of the aircraft, means of depiction (4) for presenting on a display screen (6) a first characteristic sign illustrating the speed vector of the aircraft, and auxiliary means (7) for determining at least one longitudinal margin of maneuver of the aircraft, which is expressed as a load factor and which relates to one of the two maneuvers of pitch-up and of pitch-down of the aircraft. The means of depiction (4) moreover present, on the display screen (6), at least one second characteristic sign which is associated with said first characteristic sign and which illustrates the longitudinal margin of maneuver.

Abstract: An aerial sampler system includes an enclosure, a transfer system, and a measurement system. The enclosure is connected to an external surface of an aerial vehicle and receives atmospheric flow from the external surface of the aerial vehicle. The enclosure also directs at least some of the atmospheric flow into an aperture in the external surface. The transfer system transfers some of the atmospheric flow from the aperture to a measurement system. The measurement system is internal to the external surface of the aerial vehicle and measures atmospheric trace gases in the atmospheric flow.

Abstract: A stall indicator for use on an aircraft comprises: a structure for mounting a housing at a leading edge of a wing of the aircraft; a paddle disposed to air flow at the leading edge of the wing and movable in response to the air flow; and a proximity sensor disposed at the housing for detecting movement of the paddle at a sensing area. Also disclosed is a method of indicating when an aircraft is approaching a stall condition comprising the steps of: permitting a paddle, disposed at a leading edge of a wing of the aircraft, movement with respect to a sensing area in response to air flow at the wing; detecting movement of the paddle by proximity sensing with electrical energy in the vicinity of the sensing area; and effecting an electrical signal representative of the detected movement of the paddle, which signal being indicative of an approaching stall condition at the wing.

Abstract: A method is disclosed for monitoring and controlling a compressor including (a) monitoring at least one compressor parameter; (b) storing data indicative of the monitored parameter in a database system; (c) processing the collected data using a stall precursor detection algorithm to determine a stall precursor; (d) comparing the stall precursor with at least one of a corresponding average compressor value, and a corresponding unit specific value to determine the level of compressor operability, and (e) if the stall precursor varies from at least one of the average compressor value and the unit specific value, performing corrective actions to vary the level of compressor operability to prevent a surge condition.

Abstract: The present invention provides a method for determining information for an air vehicle including the steps of providing a first coefficient of pressure data point, providing a situation involving the air vehicle, deriving pressures during the situation and deriving a second coefficient of pressure data point from the pressures derived during the situation, and defining a line passing through the first and second coefficient of pressure data points, whereby a selected coefficient of pressure data point on the line corresponds to the information. The invention encompasses apparatus for accomplishing the method. In some embodiments, the information relates to angles from zero lift and/or zero yaw.

Abstract: The present invention is a system for monitoring, storing, and reporting the accumulation of fatigue occurrences experienced by an aircraft. The system according to the present invention uses existing aircraft instrumentation to monitor fatigue occurrences. Data is collected on the magnitude and cycles of any turbulence encounters; magnitude and number of gloading accrued by aircraft maneuvering; the number of pressurization cycles that the aircraft has experienced; and the number of takeoffs/landing cycles that have occurred. These data are stored in electronic memory for later review by the flight crew or maintenance personnel. Additionally, the number of accumulated wing flap cycles and landing extensions/retractions may also be monitored to facilitate the maintenance operations of these two critical flight controls.

Abstract: An autopilot controller, for controlling a vehicle traveling in a fluid medium, designed for controlling the vehicle in any one of pre stall, stall and post stall regions. The autopilot controller provides an approach for one-to-one mapping between a target force or moment command which is received and control surface deflections for effecting the command. In one embodiment, the autopilot controller determines linear functions representative of the vehicle dynamic response for the current operating conditions in order to determine a direction of change of fin deflection for effeeting the command.

Abstract: A flight warning system comprises an outside temperature probe for sensing outside temperature, a gyroscope for sensing aircraft bank angle, a static transducer for sensing static pressure, and a dynamic pressure sensor for receiving total pressure from a pitot tube. A computer is responsive to the gyroscope, static transducer, outside temperature probe and dynamic pressure sensor calculates whether the aircraft is entering a stall condition by comparing the aircraft's bank angle with the stall bank angle. The computer can also calculate whether sufficient altitude exists for the aircraft to safely return to a field after a power failure. An output device can also be provided for indicating the existence of said sufficient altitude, and a stall indicator can be provided for indicating whether the aircraft is entering a stall condition. A method for determining the existence of a stall condition with the aircraft is also disclosed.

Abstract: A method and system for sensing surface contamination on an aircraft having a control surface. A control element including the control surface is connected to the aircraft by a hinge. The method and system sense a control surface hinge moment about a line on the hinge of the aircraft. A control surface steady hinge moment coefficient is calculated from control surface hinge moment data representing the hinge moment over a period of time. An unsteady hinge moment is calculated which is dependent on the control surface steady hinge moment coefficient. By using a value which is relatively constant during uncontaminated surface conditions, this value can be compared against calculated values to check for variations. If the calculated unsteady value varies from the known uncontaminated values, a warning may be given, alerting an operator of unsafe conditions. This warning is given before a critical value is reached, allowing the operator a period of time to react to the warning.

Abstract: A method and apparatus for predicting the onset of stall in an axial flow compressor and stabilizing stall so as to achieve the maximum pressure rise to the combustor. Stall in an axial flow compressor is predicted by obtaining pressure or axial velocity measurements at a plurality of circumferential angles around a compressor stage, fourier transforming the measurements with respect to the circumferential angles to obtain a time history of the zeroth, first and higher modes, and determining amplitude gain at twice the rotor frequency of the transfer function from the product of the zeroth and first modes to the first mode, with an increase in amplitude gain depicting a right skewed compressor going into stall and a decrease in amplitude gain depicting a left skewed compressor going into stall.

Abstract: The present invention provides a method of detecting an engine stall based on the intensity of ultraviolet light being sensed in the augmentor. The method is used to confirm a stall when the gas generator burner pressure sensor is operating properly, and as a back-up detection method when the burner pressure sensor fails, or when the burner pressure line to the control system ruptures and fails.

Abstract: A two-stage alarm for a rotary wing aircraft, such as a helicopter which operates on a collective control arm. A moveable mass or weight is operatively connected to the collective control arm. The moveable weight may be of a reciprocating type or a rotor driven mass which is mounted eccentrically for rotation about an axs which is spaced from the center of gravity of the weight. In the latter case, a D.C. motor spins the weight rapidly about the axis at a first speed to produce a shaking action to the collective control arm as a warning of an approaching danger. The weight is moved more rapidly to produce a much more noticeable action at a second frequency, or at an increased amplitude as a warning of a more imminent or serious condition which requires immediate action.

Abstract: The present invention provides a device for indicating an operating point of an aircraft comprising a sensor to measure air speed and a sensor to measure vertical acceleration. The device also comprises a processor in communicative connection with each of the airspeed sensor and the vertical acceleration sensor. The processor is also in communicative connection with a display. The display presents an operating envelope diagram of the aircraft. The display also displays the current operating point of the aircraft relative to the operating envelope diagram. The present invention also provides a method for indicating an operating point of an aircraft in which the operating envelope diagram and the measured operating point of the aircraft are displayed upon the same display.

Abstract: In an aircraft under autopilot control, a stall protection system overrides established autopilot parameters in response to a monitored angle of attack for the aircraft exceeding a trigger angle of attack. The trigger angle of attack is established with reference to a stick shaker angle of attack for the aircraft. When the aircraft exceeds the trigger angle of attack, the stall protection system produces an error signal causing the autopilot to seek an angle of attack 1 degree below the stick shaker angle of attack. Initially, the stall protection system provides a boost command to the error signal to accelerate the response of the aircraft to the error signal. The boost command is produced by a driving signal generator that selects one of four driving signals in response to monitored flap angle and impact pressure.

Abstract: An aircraft stall warning system includes an angle of attack sensor which provides a signal to an angle of attack processor which provides an angle of attack signal to a display. An air pressure sensor provides an air turbulence intensity ratio signal to an aerodynamic performance processor which provides an aerodynamic performance signal to the display. The display therefore simultaneously displays both theoretical stall margin and actual stall margin to thereby provide a pilot with wing contamination information.

Abstract: Method, apparatus, and sensor mast for airfoil aerodynamic performance monitoring are based on pressure measurement. Pressure transducers mounted within the sensor mast measure real-time pressure, to produce an output signal. The signal is separated into steady state and overlaid ripple components. The steady state component corresponds to the mean dynamic pressure of the airflow, while overlaid ripple component corresponds to the superimposed turbulent flow component. The overlaid ripple component is transformed to an amplitude equivalent. A dimensionless ratio of the two components forms a pressure turbulence intensity factor, which may be used consistently to predict and warn of imminent stall conditions. Pressure measurement was found to be an improvement over airspeed measurement. Pressure sensors proved more robust and less fragile than airspeed sensors.

Abstract: An audible indicator is arranged in combination with an electronic stall indicator, wherein the audible device includes a conical entrance and exit tube coaxially aligned with a central cylindrical tube, with a whistle member positioned within the tube, with the whistle member defined by an elongate body oriented parallel to the axis of the exit, entrance, and central tube structure to provide audible indication in a range of ten to fifteen knots in advance of stall indication by the electronic stall indicator device.

Abstract: A pitch limit symbol is generated at an aircraft attitude indicator which provides a pilot with display of a stall warning angle of attack margin which is referenced to an airplane symbol on the primary attitude indicator. The position of the pitch limit symbol relative to the aircraft symbol is a function of the difference in angle of attack between the current angle of attack at the aircraft, and a predetermined angle of attack at which the onset of stall warning occurs. Additional information is provided on the attitude indicator by a flight path angle symbol. The pitch limit symbol together with the flight path angle symbol provides the pilot with information to achieve a positive rate of climb during a pitch up maneuver while adjusting the pitch attitude of the aircraft to avoid a stall situation. The position of the pitch limit symbol may be compensated by various means to decrease the distance between the aircraft symbol and the pitch limit symbol.This is a continuation of application Ser. No.

Abstract: A stall warning "stick shaker" reference level signal is produced as a function of both flap position and total aircraft thrust. More specifically, a correction term .DELTA..alpha. is combined with a standard stick shaker reference level signal which is scheduled as a function of aircraft flap position. The .DELTA..alpha. signal is scheduled as a function of the total aircraft gross thrust coefficient. The resultant reference level signal produces advance stall warnings that are essentially invariant with engine power level.

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

Abstract: To detect a surge in a turbocompressor, the temperature at the compressor intake, or the difference between temperatures measured at different distances from the compressor intake, is subjected to a differentiation or to an operation equivalent to differentiation, thereby determining the temperature change rate. A surge is indicated when the temperature change rate is above a specified value and possibly other plausibility criteria are met.

Abstract: The airspeed over an external surface of the aircraft is sensed at a predetermined location spaced from the surface, signals are produced representative of average airspeed and of airspeed fluctuations caused by air turbulence, and the signals are also used to produce a signal representative of the ratio of the airspeed fluctuations to the average airspeed.

Abstract: In a pressure measuring system which includes a pressure transducer for generating an analogue signal as a function of pressure, circuitry for generating a non-linear error signal as a function of the difference between the input signal and a feedback signal, a feedback circuit which generates the feedback signal as a function of said error signal to null the error signal and generates a digital output signal as a function of the pressure, a pressure surge detecting circuit having an amplitude detector connected to said pressure measuring system to receive the error signal and generating an output signal when the error signal is equal to and greater than a predetermined magnitude; a timer circuit connected to receive said output signal and generate an alarm signal when the output signal continues for a predetermined period of time, a digital computing circuit connected to receive the digital output signal for computing the rate of change of pressure and the change in pressure, a memory circuit storing predete

Abstract: Pressure tapping for rigid sails, the tappings (10, 11, 12, 15, 16) being located to either provide a stall warning or to enable tuning of the sailset to the desired operating angles. A particular flow-state sensitive stall warning device is also described.