85 results about "Inertial frame of reference" patented technology

Deriving inertial-aided deviations for autoland systems during GPS signal interruptions involving an inertial reference system and a GPS landing system can include converting position coordinates of an aircraft from the inertial reference system to runway, lateral, and vertical coordinates. After this conversion, the method calibrates runway distance and lateral distance based on the converted position coordinates from the inertial reference system using runway distance and lateral distance from the GPS landing system with a third-order calibration filter when the aircraft is below a first height above terrain and calibrates vertical distance based on the converted position coordinates from the inertial reference system using vertical distance from the GPS landing system with a second-order calibration filter when the aircraft is below the first height above terrain. The method uses the calibrated runway, lateral, and vertical distances for deviation computations when GPS signals are interrupted below a second height above terrain.

The invention pertains to the monitoring of the integrity of position and speed information arising from a hybridization between an inertial reference system and a satellite-based positioning receiver. The invention relates more precisely to a navigation apparatus known in the art by the name INS/GNSS system (for “Inertial Navigation System” and “Global Navigation Satellite System”) hybridized in closed loop.

It is presumed and commonly accepted by those skilled in the art of satellite navigation and Kalman filter design that the filter must be provided with the tracker position and velocity a priori in order to determine target position and velocity. Indeed, it is generally asserted that without a priori knowledge (known or measured values) of the tracker position and velocity, line of sight measurements between satellites do not contain adequate information to infer target states. Passive and autonomous navigation of space vehicles without a priori values for the position and velocity of either the target or tracker vehicle is achieved by reconfiguring the extended Kalman filter, or more generally any predictor/correction class filter, to include states for both the target and tracker vehicles. The target and tracker vehicles must both follow trajectories in an inertial frame of reference through the gravitational field of a gravitational body having a known gravitational model. The reconfigured filter simultaneously estimates the position and velocity of both tracking and target space-based vehicles from line-of-sight measurements.

Method and device for detecting an erroneous speed generated by an air data inertial reference system. The device (1) comprises means (2) for detecting an erroneous speed provided by an air data and inertial data system (3) and means for generating a substitution speed.

The invention discloses an attitude parameter calibration method and an attitude parameter calibration device of an airborne inertial navigation system. A high-dynamic star sensor is used as a data acquisition method and fixed stars are used as measuring objects; the plurality of fixed stars on a celestial sphere are detected, and start point positioning, star map identification and attitude calculation are carried out to provide attitudes of the star sensor relative to an inertial reference system; then, airplane attitude standard data under an inertial navigation coordinate system is obtained according to a conversion relation between a star sensor coordinate system and the inertial navigation coordinate system; finally, the standard data is used as reference and is compared with the attitude data of the inertial navigation system, so that the dynamic online calibration of the attitude precision of the inertial navigation system is realized under an airborne environment.

A navigation technique for a vehicle employs an inertial reference system to derive a first position indication and a first velocity value. A first receiver processes signals of a global positioning system from which a second position indication and a second velocity value are derived. A second receiver processes signals from a plurality of distance measuring equipment stations at fixed positions on the earth and determines the distance between the vehicle and each of those stations. A third position indication is derived from those distances. A Kalman filter function is applied to the first, second and third position indications and to the first and second velocity values to compensate for uncertainty in the first position indication and in the first velocity value and thereby produce a vehicle position estimate and a vehicle velocity estimate.

In a first aspect, the invention uses the thrust and axial aerodynamic forces of and on the missile system as the primary mechanism for altering the missile system's flight path as opposed to the missile system's aerodynamic normal force. In accordance with this first aspect, a first embodiment is a method for changing the direction of travel of a vehicle moving through a fluid medium to a desired direction of travel measured in the vehicle's inertial frame of reference, the vehicle comprising a body including a nose, a moment generating control device, and an axial propulsive device. The method comprises actuating the moment generating control device to align the nose substantially with the desired direction of travel, and applying the axial propulsive device as the primary mechanism to redirect the vehicle's velocity to the desired direction of travel. In a second aspect, the invention provides a missile control system implementing the first aspect of the invention. In a third aspect, the invention simulates in software a missile control system implementing the first aspect of the invention.

Apparatus is disclosed for monitoring strain and/or load applied to the body of a vertebral mammal. The apparatus includes means (10, 11) for measuring position of the body relative to an inertial frame of reference and for providing first data indicative of the position, wherein the means for measuring position includes at least one inertial sensor such as an accelerometer. The apparatus includes means (12, 15-19) for measuring muscle activity and for providing second data indicative of the muscle activity. The apparatus also includes means for storing the data and means for processing the data to provide a measure of strain and/or load to the body. The strain and/or load is typically applied to the back of the mammal. A method of monitoring strain and/or load applied to the body of the vertebral mammal is also disclosed.

Disclosed is a hybrid inertial measurement system including a cold atom interferometric inertial sensor having a laser source generating a sequence of laser pulses towards a cold atom burst and a conventional inertial sensor attached to the inertial reference frame of the interferometric inertial sensor. The hybrid system includes a signal processing system suitable for receiving an inertial measurement signal from the conventional inertial sensor and for generating in real time a non-linear frequency modulation signal, the feedback loop electronic system being configured to modulate in real time the central optical frequency of the laser according to the modulation signal, such that the cold atom interferometric inertial sensor generates a first hybrid inertial measurement signal by atomic interferometry corrected for the relative movements of the inertial reference frame.

The invention belongs to the technical field of spacecraft flight, and particularly relates to a modeling method of a space three-body flexible tethered satellite formation system under a non-inertialreference coordinate system. The modeling method comprises the following steps that firstly, a satellite formation system model and a non-inertial reference system fixed to the system centroid are established; secondly, a flexible tether connecting a satellite is discrete into a plurality of tether units, and the satellite and tether unit particles are numbered; and finally, dynamic equations ofdiscrete element particles under the non-inertial reference system are established, and then the dynamic equations of the satellite formation system are derived. According to the modeling method of the space three-body flexible tethered satellite formation system under the non-inertial reference coordinate system, modeling is carried out under the non-inertial reference system, redundant steps such as selecting generalized coordinates and coordinate transformation are not required, and simultaneously flexibility of the space tether is considered, so that the method is faster and more accuratein describing various space dynamic behaviors of a formation system and revealing spin stability and instability states of the system. In addition, the modeling method of the space three-body flexibletethered satellite formation system under the non-inertial reference coordinate system can further be extended to modeling of a four-body or multi-body space tethered formation system and has great practical value.

The invention proposes a method for tracking the navigation of a mobile carrier, in which an extended Kalman filter estimates, over successive iterations, a navigation state of the carrier, one iteration of the filter comprising steps of: propagating a previous navigation state of the carrier into a propagated state according to a kinematic model and/or measurements acquired by at least one inertial sensor, updating the propagated state according to measurements acquired by at least one navigation sensor, in which, for at least one navigation variable of the carrier contained in the propagated state, the updating comprises sub-steps of: calculating a linear correction term from an innovation representative of a difference between the measurements acquired by the navigation sensor and the propagated state, calculating an exponential of the linear correction in terms of a Lie group, calculating a first correction term expressed in a reference frame that is fixed relative to the carrier, the first correction term depending on said exponential, calculating a second correction term expressed in an inertial reference frame in which the carrier is mobile, by changing the reference frame applied to the first correction term, and adding the second correction term to the value of the variable contained in the propagated state, the state containing the result of this addition being used as the output state of the iteration.

An input device for controlling an indicia shown by a presentation system is described and includes the following: a device housing; a wireless communication component configured to establish a wireless communication channel between the input device and the presentation system; an inertial sensor configured to monitor rotation of the input device relative to an inertial reference frame defined by orthogonal first, second and third axes; and a processor configured to generate a signal for transmission by the wireless communication component to the presentation system to adjust an attribute of the indicia when rotation of the input device about the first axis exceeds a threshold value that varies based on an amount of rotation detected relative to the second and third axes.

A method of presenting attitude and heading information of a vehicle on a display to a viewer inside the given vehicle, comprising the steps of: using a 2D display, using data provided by an inertial reference system, using a computer and graphical software for treating the data and defining graphical elements, depicting the graphical elements in 3D on the 2D display, and presenting the attitude and heading information on the 2D display by associating them to at least one graphical element.

A method for determining orbit of a spacecraft by using relative position increment belongs to the technical field of determination of spacecraft orbit, and aims at solving the problem that a conventional orbit determination method based on measured epoch difference needs to employ other orbit determination method for obtaining initial position information. The method for determining orbit of the spacecraft by using relative position increment comprises: firstly, determining the motional orbit plane of the spacecraft in a three-dimensional space; then, converting the three-dimensional relative position vector into a spacecraft relative position coordinate in a two-dimensional plane, fitting the spacecraft motion orbit shape in the two-dimensional plane by employing a least square method, and determining eigenvalue of the fit matrix by employing lagrangian multiplier; and finally, determining the spacecraft orbit type according to the eigenvalue of the fit matrix, and obtaining the absolute position information of the spacecraft in an inertia reference system according to the position of a central celestial body, so as to determine the initial orbit of the spacecraft. The method is applicable to orbit determination of spacecraft.

The invention relates to a calibration method for an inertial reference system, which is characterized in that on the premise of ensuring that the installation precision of the inertial reference component meets the requirement, a mutual relation is established between a standing ground target point and the symmetry axis of the aircraft, the course angle of a target datum line is measured, the distance between the aircraft and the datum line of the aircraft is converted, and the position parameters of the aircraft are input into calibration software to complete the calibration of the inertialreference system. The method guarantees the calibration precision of the inertial reference system, is convenient and practical, is low in cost by using the existing production conditions, and saves the fund.

A method of calibrating an inertial unit is provided. During a first static stage, in which the inertial unit is in a first orientation, measurements are taken by means of the accelerometers and the inertial rotation sensors. During a dynamic stage, the orientation of the inertial unit is changed, at least in part in azimuth, from the first orientation towards a second orientation, while taking measurements by means of the inertial rotation sensors. During a second static stage, in which the inertial unit is in the second position, measurements are taken by means of the accelerometers and of the inertial rotation sensors. For each static stage, a direction, an amplitude, and a mean speed of rotation for apparent gravity in an inertial frame of reference is estimated, variation is calculated in orientation between the static stages, and the accelerometer biases is deduced therefrom.

The invention provides an inertial navigation method and device based on real-time gravity compensation. The inertial navigation method comprises the following steps: receiving a navigation request, wherein the navigation request includes a motion carrier identifier; measuring to acquire accelerated speed of a motion carrier corresponding to the motion carrier identifier in an inertial reference system according to the navigation request; establishing a Kalman state equation and a measurement equation of gravity disturbance, and acquiring an optimal estimated value of the gravity disturbance; compensating the accelerated speed according to the optimal estimated value of the gravity disturbance, and carrying out corresponding navigation treatment on the compensated accelerated speed. According to the inertial navigation method based on the real-time gravity compensation, the navigation precision of an inertial navigation system is improved.

A long-term navigation method using an inertial unit associated with a system of axes X, Y, Z and mounted on a vehicle traveling relative to the Earth in order to measure its movements relative to an inertial frame of reference having axes Xi, Yi, and Zi. The method includes the steps of acting in permanent manner to measure an orientation of the system of axes X, Y, Z in the inertial frame of reference and applying a predetermined sequence of turnovers to the inertial unit in the inertial frame of reference about first and second axes that are substantially perpendicular to each other and in such a manner that at the end of the sequence the inertial unit is in a final orientation identical to its initial orientation relative to the inertial frame of reference, with the turnovers canceling within the sequence.