Abstract: An avionics display is disclosed. A synthetic rendering of terrain is created using stored terrain data. A horizon line is overlaid upon the synthetic rendering of terrain. The horizon line represents zero degrees pitch attitude in a direction of travel of an aircraft in which the avionics display is installed. The horizon line is configured to move in response to attitude-changing movement of the aircraft. An attitude cue is overlaid upon the synthetic rendering of terrain and depicted on first and second sides of the horizon line such that the synthetic rendering of terrain remains substantially visible. The attitude cue has a first appearance on the first side of the horizon line and a second appearance on the second side of the horizon line. The first appearance is visually contrasted from the second appearance.

Abstract: Systems and methods for determining center of gravity for an aircraft. An example system includes one or more load measurement devices that generate one of nose gear or main gear weight information and a processing device that determines center of gravity of the aircraft based on previously received gross weight information and the generated nose or main gear weight information. The number of gear sets with load measurement devices is one less that the total number of gear sets having distinct longitudinal positions along a fuselage of the aircraft. The processing device further determines center of gravity based on temperature and/or pitch attitude information. The system includes a user interface that allows a user to enter the gross weight information that might be included in a flight manifest or a load and trim sheet and/or a communication component that receives the gross weight information from a ground-based system.

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: In distance maps used in order to facilitate the navigation of craft such as an aircraft, the distances estimated take into account obstacles to be circumvented but not the maneuverability of the craft. The adaptation route required by the craft in order to take the right direction is not taken into account so that certain distance estimates for accessible points located in the neighborhood of the craft are unrealistic. In order to make a distance map more realistic, it is proposed that an obstacle of concave shape associated with the craft be added behind the position of the craft, forcing the estimations of distances to be circumvented an area that is inaccessible to the craft for reasons of maneuverability.

Abstract: A method and device for detecting, on the ground, the obstruction of a pressure tap of a static pressure sensor of an aircraft. The device includes a pressure sensor that measures the pressure inside a probe and a heating system for heating the probe. When the heating system is activated, the sensor carries out a first measurement of a parameter dependent on the pressure inside the probe. After a predetermined duration from the activation of the heating system, the sensor carries out a second measurement of the parameter. A central processing unit calculates the difference between the first and second measurements and compares this difference to a predetermined value. A warning device emits a warning signal indicating the detection of an obstruction if the difference is greater than the predetermined value.

Abstract: The present invention relates to a standby instrument fitted to the instrument panel of an aircraft detecting overloads. The standby instrument includes a processor to calculate and display flight information based on data provided by associated sensors integrated into a standby system. It additionally includes a module for detecting overload of the aircraft on the basis of the data provided to display the flight information. The invention applies to the detection of overloads during a heavy landing. More generally, it applies to the detection and indication of overloads of an aircraft in various types of environment.

Abstract: The device relates to the signaling, to the pilot of a craft, of the risks that it incurs due to obstacles placed in its zone of deployment. This device determines, on the zone of deployment of the craft, the contours of several types of risk region as well as the contours of two lateral regions of tight deployment tied to the craft and produces alarms as a function of the types of zones at risk found in the lateral regions of tight deployment.

Abstract: A display system and method for an aircraft simultaneously displays the terrain under the flight plan, or under the current track of the aircraft, as both a two-dimensional lateral situation view image and as a perspective vertical situation view image. This provides improved tactical flight planning, and that does not erroneously show the flight plan passing through the displayed terrain.

Abstract: The present invention is directed to a computerized system that includes a graphical user interface (GUI) having a display and a data input device. The computerized system performs a method including the step of providing a plurality of aircraft parameters. At least one vertical g-force value is calculated in real time based on the plurality of aircraft parameters. A lateral g-force value is calculated in real time based on the plurality of aircraft parameters. The at least one vertical g-force value is displayed relative to at least one predetermined vertical reference value. The lateral g-force value is displayed relative to at least one predetermined lateral reference value.

Abstract: A method is disclosed for checking model planes entering into a dive, as well as an anti-dive method and apparatus. The method of checking planes includes installation of a first optical sensor in the upper part of the plane, with installation angle upwardly intersecting with the front of the plane at ? degrees, that checks light intensity and outputs a first checked optical signal; installation of a second optical sensor in the bottom part of the plane, with its installation angle downwardly intersecting with the back of the plane at ? degrees, which checks light intensity and outputs a second checked optical signal. A comparison is made between the first and second checked optical signals, and when the difference in light intensity is smaller than a threshold, a warning signal is sent out indicating that the plane has entered into a dive.

Abstract: A radio frequency identification (RFID) system having the capacity to detect conditions of alignment, wherein the system may be used with hand-held, fixed-in-place, stationary, and permanently mounted apparatus. The system includes an RF interrogator configured for use with dental x-ray, medical imaging, film, and digital radiography apparatus, and may include a multiplicity of RF transponders or interrogators. An RF interrogator, an RF transponder, and an x-ray sensitive imaging device, and its holder are configured to be critically aligned to a dental x-ray machine head apparatus, rendering repeat imaging unnecessary. The x-ray emitter may be further configured to automatically obtain a desired x-ray image or configured so that the device cannot activate and provide a radiograph until alignment with the transponder and associated x-ray sensitive imaging device has occurred.

Abstract: A flight control system which detects a failure of a flight control surface and performs at least one action in response to the detected failure.

Abstract: The disclosed embodiments relate to a method of determining a maximum stabilization height hstab of an airplane in a final phase of said airplane in order to inform the pilot as early as possible of conditions that would not enable him to make a landing in stabilized approach conditions. The disclosed embodiments also relate to a device which includes means (2) of acquiring parameters necessary for the method, calculation means (3) which determine a stabilization height hstab from the information received from means (2) and means (4) of presenting the information to warn the pilot, the stabilization height being established according to the method.

Abstract: A method of dynamically displaying lateral deviation in formats on a display is provided. The method comprises assigning one or more positional boundaries for displaying the lateral deviation information, wherein the one or more positional boundaries define areas on a display. The method also comprises calculating display positions of a lateral deviation symbology and determining if each calculated display position is allowed based on where the lateral deviation display would be located, if displayed at each calculated position, in relation to the one or more positional boundaries. The method further comprises, displaying the lateral deviation symbology at the calculated positions when the calculated positions are allowed, and displaying a lateral deviation indicator when the calculated positions are not allowed.

Abstract: Methods and systems are provided for displaying an unusual attitude of an aircraft in a head-up display system. The system comprises a projection unit configured to project the unusual attitude of the aircraft and a processor coupled to the LCD projection unit. The processor is configured to control the LCD projection unit during the unusual attitude of the aircraft to produce an attitude indicator of the aircraft and a horizon symbol. The horizon symbol divides the attitude indicator into a first region having a first display characteristic and a second region having a second display characteristic different from the first display characteristic. The first display characteristic is substantially uniform throughout the first region, and the first region indicates a ground with respect to the aircraft.

Abstract: A method of automatically guiding an aircraft for flight at low altitude determines an LLF (low level flight) trajectory section for following terrain to be overflown by the aircraft. The LLF trajectory section takes account of predictions of the aircraft's mass and speed and climb and descent performance. The LLF trajectory section includes an entry point and an exit point. The LLF trajectory section is integrated between a first trajectory part corresponding to a first phase of flight and a second trajectory part corresponding to a second phase of flight, by providing respectively first and second transition phases. The aircraft is automatically guided so that it successively captures, follows, and leaves the LLF trajectory section.

Abstract: A method of aiding the piloting of an aircraft during a maneuver, bringing about an increase in the attitude of the aircraft, is carried out automatically and repeatedly. According to the method, an attitude deviation is determined by subtracting a first value of attitude from a determined current attitude of the aircraft. A horizon line is presented on a display screen of a head-up display device, superimposed with the environment seen in front of the aircraft, as is an indication of the attitude deviation. The first value of attitude corresponds to the minimum value of (1) an optimal value of attitude at the end of the maneuver and (2) a value of attitude of protection against a rear collision.

Abstract: A method for providing on-demand audio information to a pilot of a first aircraft during aerial combat engagement with a hostile aircraft. The method includes receiving a pilot-initiated request for information, identifying the pilot-initiated request as either a first request-type or a second request-type. If the pilot-initiated request is identified as the first request-type, retrieving output information relating to current flight parameters of the first aircraft; if the pilot-initiated request is identified as the second request-type, retrieving output information relating to current flight parameters of the hostile aircraft; and generating a verbal message audible to the pilot containing the output information.

Abstract: An aircraft terrain avoidance method and device may employ a collision warning section that transmits a caution signal when the aircraft risks colliding with the terrain at the end of a first predetermined period of time. A warning signal is transmitted when the aircraft risks colliding with the terrain at the end of a second predetermined period of time, which is shorter than the first predetermined period of time. A piloting system causes the aircraft to automatically climb with a first gradient corresponding to a predetermined value upon the transmission of the caution signal. An automatic pilot is engaged automatically, if currently disengaged, to force the aircraft to climb with a second gradient, corresponding to the maximum gradient possible in the current flight conditions of the aircraft, upon the transmission of the warning signal.

Abstract: A flight instrument displays azimuth information in combination with horizon information and other optional information such as altitude and/or airspeed. Azimuth information is displayed using an elliptical pattern placed below a horizon line displaying horizon information. The horizon information may be based on gyro enhanced vertical speed and/or vertical axis information. The elliptical pattern may be representative of a horizontal circular heading indicator viewed from above and behind the circular indicator. In one embodiment, the elliptical azimuth pattern moves in response to movement of the horizon line to maintain position below the horizon line. Airspeed and/or altitude information may be displayed at the sides of the screen area using a circular arc display pattern.

Abstract: A system and method for assisting a pilot to recover the aircraft from a variety of dangerous situations involving one or more of the following: altitude, airspeed, attitude, roll rate, and a partial equipment failure in the control panel. In accordance with the invention, these factors are sensed automatically, and voice instructions are provided to the pilot to direct him or her to get the aircraft out of the dangerous situation sensed by this system.

Abstract: A vertical deviation indicator and predictor includes a vertical deviation scale having a plurality of vertically spaced markers, one of the markers indicating the vehicle's present vertical position. Current vertical flight path segment symbols are selectively superimposable over the vertical deviation scale in accordance with the vehicle's current vertical flight path segment. Next vertical flight path segment symbols are selectively superimposable over the vertical deviation scale in accordance with the vehicle's next vertical flight path segment, the position thereof being determined by backward extrapolation of the next vertical flight path segment. Thus, the type and position of the current and next vertical flight path segment symbols on the vertical deviation scale provide situational awareness of the present vertical flight path deviation and an indication of an efficient and timely manner for intercepting the desired flight path.

Abstract: Methods and systems for displaying the source of aircraft control instructions are disclosed. A method in accordance with one embodiment to the invention includes receiving a first target to which an aircraft will be automatically directed upon authorization from an operator, and receiving a second target to which the aircraft will automatically be directed upon meeting a threshold condition. The method can further include displaying the first and second targets at least proximate to each other, with a target to which the aircraft is currently being controlled being displayed in a different manner than a target to which the aircraft is not currently being controlled.

Abstract: A process and device for detecting on an aircraft an overshoot of design loads at the level of a structural part of the aircraft. The detection device includes a speed measuring device for measuring an effective speed of the aircraft. A first comparator compares the measured effective speed with a maximum speed relating to the current flight configuration of the aircraft. A vertical load measuring device measures a vertical load factor of the aircraft, and a second comparator compares the measured vertical load factor with a limit value of the vertical load factor. An inspection determining device determines whether a structural inspection needs to be performed at the level of the structural part, based of the results of the first and second comparisons.

Abstract: A sensor for measuring yaw rate or roll rate of an automotive vehicle comprises a freely rotating inertial disk and an angular rate sensor responsive to the rotation of the inertial disk relative to a housing. In one embodiment the inertial disk presents an alternating magnetic field at its circumference. The rate and direction of rotation of the inertial disk relative to its housing is determined by three magnetic field sensors such as linear Hall Effect sensors responsive to the field presented by the inertial disk. In another embodiment electronic cameras measure movement of fiducial marks on the inertial disk. Air surrounds the inertial disk and air viscosity gradually brings rotation to a stop. For yaw rate measurement the disk axis is oriented vertically and the inertial disk is supported in the radial direction by low friction bearings such as ball bearings or magnetic bearings and in the axial direction by substantially frictionless bearings such as magnetic bearings.

Abstract: A display control device, having a processor structured for receiving terrain elevation information and samples of current position and altitude above ground data; and one or more algorithms resident on the processor for generating, as a function of the current position and altitude above ground data, one or more display control signals for causing a display device to display, in a monochromatic scale graduated as a function of terrain elevation relative to mean sea level, a strategic portion of the terrain elevation information extending between a minimum elevation of the strategic portion of the terrain elevation information and a pre-selected strategic altitude threshold less than the altitude above ground data.

Abstract: A flight control system includes a primary flight display having a hover display which facilitates operations in DVE conditions. The hover display includes a radar altitude display, heading drift symbology and lateral drift symbology which assists pilots in the detection of drift, which is common in DVE situations and to provide the pilot with a salient cue pertinent to the approach to hover and landing.

Abstract: The flight control indicator may include a set of information sources, a computation unit connected to said set of information sources, and a head-up display device. The computation unit determines, on the basis of information emanating from the set of information sources, at least one maximum total slope of the aircraft. The maximum total slope corresponds to the maximum thrust and is obtained on the basis of the measurement of the current performance of the aircraft, extrapolated to the condition of maximum slope and in anticipation of an engine fault. The head-up display device presents, on a display screen, at least one indicator which indicates the maximum total slope and which is presented superimposed on the environment existing in front of the aircraft.

Abstract: There is provided a method for displaying aircraft positions, which is capable of effectively avoiding the occurrence of near miss or collision. The method displays the terrain of an air space under air traffic control apparently in three dimensions on a display screen, displays aircraft marks at positions on the display screen so as to correspond to the three-dimensional positions of the respective aircraft, and displays a warning mark when a distance between two aircraft is shorter than a threshold value, wherein the warning mark comprises a triangle having three apexes, which comprise the aircraft marks, and a position that is apart from the aircraft mark of one of the aircraft along a vertical line extending perpendicular to the one aircraft toward a ground surface of the terrain by a distance corresponding to an altitude difference between both aircrafts on the display screen.

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: In an image compensation apparatus for controlling an attitude of an image pickup section capable of rotating around mutually orthogonal three axes, the apparatus comprises a sight line direction information detection section for detecting information about a sight line direction of a pilot, a driving section which outputs a driving signal for rotating and driving the image pickup section with reference to the information about the sight line direction and rotates and drives the image pickup section based on this driving signal, an image pickup attitude information detection section for detecting information about an attitude of the image pickup section, an aircraft attitude information detection section for detecting information about an attitude of an aircraft, and a correction section for correcting the driving signal with reference to the information about the attitude of the image pickup section and the information about the attitude of the aircraft.

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: Method and indicator for displaying information showing the airspeed tolerance margins for an aircraft. The indicator comprises a central processing unit which selects, from a plurality of longitudinal scales, the longitudinal scale representative of the current aerodynamic configuration of the aircraft and a display means which presents on a display screen the selected longitudinal scale, which is mobile in the longitudinal direction, and whose position on the display screen depends on the current angle of attack of the aircraft, shown by a characteristic marker in a fixed position on the display screen across the longitudinal scale, the latter scrolling up and down relative to the characteristic marker as a function of the current angle of attack of the aircraft.

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 flight situation presentation system displays a large majority of the information required for piloting an aircraft in an enhanced manner to reduce potential for misinterpretation and instances of piloting error. Approaches include visual integration of flight data indicators into various functional display areas and further visual integration of the various functional display areas. Visual integration includes creating a synergism amongst the flight data indicators being presented so that the indicators tend not to distract from each other, but rather help to reinforce each other in presenting fully representative flight situation information. Techniques for visual integration include placement of flight data indicators in certain positions on a display relative to one another based upon functional relationships between the operational parameters being monitored.

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

Abstract: The invention provides an attitude sensing apparatus for determining the attitude of a mobile unit that can reliably estimate an alignment angle between a GPS antenna coordinate system and an IMU coordinate system with good accuracy regardless of the magnitude of the alignment angle. Based on observation of the difference between a GPS angular velocity and an IMU angular velocity, an alignment angle estimating section estimates an alignment angle and sensor errors. An alignment angle adder and a sensor error adder cumulatively add and update the estimated alignment angle and sensor errors, respectively. The estimated alignment angle is fed back to an inertia data converter while the estimated sensor errors are fed back to an inertia data correcting section. The apparatus repeatedly performs estimation until the estimated alignment angle gradually approaches a true alignment angle by successively feeding back estimated values to a flow of alignment angle estimation process.

Abstract: In one embodiment, a method for controlling an aircraft comprises providing an attitude error as a first input into a neural controller and an attitude rate as a second input into the neural controller. The attitude error is calculated from a commanded attitude and a current measured attitude, and the attitude rate is derived from the current measured attitude. The method also comprises processing the first input and the second input to generate a commanded servo actuator rate as an output of the neural controller. The method further comprises generating a commanded actuator position from the commanded servo actuator rate and a current servo position, and inputting the commanded actuator position to a servo motor configured to drive an attitude actuator to the commanded actuator position. The neural controller is developed from a neural network, wherein the neural network is designed without using conventional control laws, and the neural network is trained to eliminate the attitude error.

Abstract: Attitude acquisition methods and systems are provided which reduce the time generally required to acquire spacecraft attitude estimates and enhance the probability of realizing such estimates. The methods and systems receive, over a time span &Dgr;t, successive frames of star-sensor signals that correspond to successive stellar fields-of-view, estimate spacecraft rotation &Dgr;r throughout at least a portion of the time span &Dgr;t, and, in response to the spacecraft rotation &Dgr;r, process the star-sensor signals into a processed set of star-sensor signals that denote star positions across an expanded field-of-view that exceeds any of the successive fields-of-view. The expanded field-of-view facilitates identification of the stars that generated the processed set of star-sensor signals to thereby acquire an initial attitude estimate.

Abstract: The present invention includes a display for viewing information, where a number of lines on the display has a variable line resolution scale. Each of the number of lines corresponds to a value of information, such as angle degrees relative to the horizon line of the sky in the context of avionics, for example. In addition, the present invention may include a value indicator positioned within the number of lines on the display such that the value indicator designates a value of information. Each distance between each of the number of lines may vary, so that the line resolution of the number of lines near the value indicator, for example, is greater than the line resolution further removed from the value indicator. By varying the line resolution using a distortion function, values of information on the display may be more precisely read.

Abstract: A system of pitch attitude symbols, which can be generated by a head-up display unit of an air vehicle, is provided. The system of pitch attitude symbols can be displayed by the head-up display as a virtual image in the direction of flight, behind the cockpit screen in a focal plane in front of the background. An unequivocal system of pitch attitude symbols is provided for the pilot, for a display area of an HUD, which display area is arranged below the natural horizon during neutral flight attitude of an air vehicle. Within the focal plane, the system of pitch attitude symbols comprises at least one pitch attitude representation area, arranged in perspective, which pitch attitude representation area can always be arranged parallel to the surface of the earth. The rear edge of the pitch attitude representation area is tiltably fixed in a rear edge axis and, depending on the pitch attitude of the air vehicle, the pitch attitude representation area can be tilted on the focal plane.

Abstract: A system, method, and computer program product for alerting an aircraft flight crew of an inconsistency during an instrument landing system (ILS) approach to a selected runway. The system includes a component configured to receive glideslope and localizer signals, and a component configured to determine aircraft approach angle to runway end. Also, a component configured to alert the aircraft flight crew of an approach inconsistency, if the determined aircraft approach angle to runway end is less than a threshold angular value and one or more parameters based on aircraft position and the received glideslope and localizer signals are met. The component configured to determine aircraft approach angle to runway end determines the aircraft approach angle to runway end based on aircraft geometric altitude, a vertical figure of merit, a horizontal figure of merit, runway elevation, and aircraft distance to runway threshold.

Abstract: A method and computer program product are provided for controlling the actuators of an aerodynamic vehicle to affect a desired change in the time rate of change of the system state vector. The method initially determines the differences between anticipated changes in the states of the aerodynamic vehicle based upon the current condition of each actuator, and desired state changes. The differences between the anticipated and desired state changes may be weighted based upon a predetermined criteria, such as the importance of the respective states and/or the weight to be attributed to outliers. The differences between the anticipated and desired state changes are converted to the corresponding rates of change of the actuators. These changes in the actuators may be limited to within predefined bounds. Control signals are then issued to the actuators to affect the desired change in the time rate of change of the system state vector.

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: In an iterative method of determining aircraft flight data parameters using first and second multi-function probes, an assumed value of a first aircraft parameter is defined to be equal to an initial value. Using the assumed value of the first aircraft parameter together with the respective local angles of attack determined at first and second multi-function probes, first and second estimates of a second aircraft parameter are calculated and compared. If the first and second estimates of the second aircraft parameter are within tolerance of each other, then the first aircraft parameter is approximately equal to the assumed value, and the second aircraft parameter is determined from the first and second estimates. If the first and second estimates of the second aircraft parameter are not within tolerance of each other, then an iterative process is continued to correctly determine the first and second parameters.

Abstract: An apparatus, method and computer program product useful for supplying the pilot of an aircraft with aircraft attitude information in the event of full or partial failure of the gyroscopic system normally used to supply such data. The pitch angle, roll angle and heading angle estimates provided by the apparatus, method and computer program product can be displayed to the pilot thereby alleviating the requirement that the pilot mentally integrate such data from the remaining aircraft instruments.

Abstract: In an adaptive vehicle speed control system, a method and system for controlling the speed of the vehicle while the vehicle is traversing a curved path. The method and system include determining a yaw acceleration of the vehicle, and determining a maximum allowed speed of the vehicle on the curved path based on the yaw acceleration. The method and system also include limiting the speed of the vehicle on the curved path to a value no greater than the maximum allowed vehicle speed.

Abstract: A system for determining an altitude of an aircraft using barometric pressure measurements is disclosed comprising: a transmitter disposed at a ground station for transmitting a signal to the aircraft, which signal including first data representative of barometric pressure at the ground station; a receiver disposed at the aircraft for receiving the transmitted signal; sensing apparatus disposed at the aircraft for measuring barometric pressure at the altitude of the aircraft and generating second data representative thereof; and processing circuitry for determining the altitude of the aircraft based on the first and second data and data representative of the elevation of the ground station. The ground station includes sensing apparatus for measuring barometric pressure at the ground station and generating pressure data representative thereof which is superimposed onto a carrier signal of the transmitter for transmission to the aircraft.