Showing papers on "Inertial reference unit published in 1991"

Gravity aided inertial navigation system

Abstract: An autonomous covert Inertial Navigation System uniquely suited for underwater applications wherein Schuler and siderial errors are bounded without external navigation aids or active instrumentation of ground speed is achieved by integrating a conventional Inertial Navigation System with a gravity gradiometer capable of measuring gravity field components independently of platform accelerations.

Procedure for measuring angles and trajectories by means of gyros and inertial systems

Abstract: The calibration of time-depending measurement uncertainties of gyros and inertial systems by the differences of angular and/or velocity and/or positional measurements which are carried out and repetitiously executed with a view to undetermined reference directions and/or fixed locations varying with time. The differentiation allows to forego precise reference data. Measurement uncertainties can be corrected through calibration subsequently to the measurement process. For the measurement of spatial angles a three-axis gyro measurement package comprising a computer and the possibility of feeding or automaticly reading-in of reference data characterizing the reference directions. The reference data can be a number, a marker of the measurement point or in the case of positional measurements a terrestrial aiming point. When measuring angular characteristics, i.e. the interdependence of angles and external forces or moments, the latter are considered as references. For the measurement of contours the distance has to be fed in or read in, or computed by accelerometers within an integrated inertial system, respectively. By such a system and through the application of the invention the measurement accuracy of movements and gravity anomalies can be increased while refraining from external references otherwise necessary under the present state of the art.

Design and Flight Test of a Differential GPS/Inertial Navigation System for Approach/Landing Guidance

Abstract: NASA-Langley has conducted a flight test program evaluating a differential GPS/inertial navigation system's (DGPS/INS) utility as an approach/landing aid. The DGPS/INS airborne and ground components are based on off-the-shelf transport aircraft avionics, namely a global positioning/inertial reference unit (GPIRU) and two GPS sensor units (GPSSUs). Systematic GPS errors are measured by the ground GPSSU and transmitted to the aircraft GPIRU, allowing the errors to be eliminated or greatly reduced in the airborne equipment. Over 120 landings were flown; 36 of these were fully automatic DGPS/INS landings.

System for aligning the inertial unit of a carried vehicle on that of a carrier vehicle

Abstract: The system may be used, for example, to initialize the inertial unit of a missile from the inertial unit of a vehicle carrying the missile. According to the invention, a servo-control circuit is provided for slaving data in the missile inertial unit on data from the carrier inertial unit, together with a Kalman filter connected in parallel with said servo-control circuit and designed to identify harmonization errors between the two inertial units.

GPS Inertial Attitude Estimation via Carrier Accumulated-Phase Measurements

Abstract: This paper describes the design and simulation of an integrated GPS/INS that accomplishes space vehicle navigation and attitude estimation. The simulation contains models of an inertial measurement unit (IMU), a GPS receiver processing signals from an array of antennas placed on the spacecraft structure, and an integrated navigation filter. Interferometric equations are employed to process GPS carrier accumulated-phase measurements from precise antenna locations in the user vehicle. Vehicle orientation in inertial space is then obtained by relating baseline-difference phase measurements to orientations of antenna baselines relative to satellite line-of-sight vectors. The method of GPS-aided attitude estimation described herein is distinguished from other methods in that it does not employ search algorithms to resolve the carrier phase cycle ambiguity. Instead, the ambiguity is modeled as a random process and is included as an error state of a statistical filter. As a result, the attitude estimation time history evolves directly for any dynamic environment.

Inertial navigation system with a compensation filter for its alignment during flight

Abstract: In a navigation system, the speed error DELTA VN resulting from a position misalignment DELTA psi is conducted via a compensation filter X the transfer characteristic of which is selected in such a manner that the resultant speed error is now only dependent on time.

System for the transfer of alignment between the inertial system of a carried vehicle and that of the carrier vehicle

Abstract: System permitting, for example, initialisation of the inertial unit (CNm) of a missile from the inertial unit (CNp) of the carrier of this missile. According to the invention, a servo control device (A) is provided for the data from the inertial unit of the carrier (CNp) in the inertial unit (CNm) of the missile and a Kalman filter (K) is provided, mounted in parallel on the said servo control device (A) and intended to identify the harmonisation errors between the two inertial units.

Integrated System for Automatic Landing Using Differential GPS and Inertial Measurement Unit

Abstract: GPS-receivers have accuracy problems in high precision flight gui­dance applications. In dynamic flight maneuvers they show not only operational problems due to satellite masking but also a reduction in accuracy in accelerated flight and turn flight.

An Observability Analysis of a GPS-Aided Geodetic Inertial Strapdown Measurement Unit

Abstract: A widely used tool to combine data of different measurement devices in real time is the Kalman-Filter-Technique. To guarantee a uniformly asymptotically globally stable filter, the system model upon which the Kalman filter is based has to be both observable and controllable. Thus a first step to design a filter is to check the observability and controllability of the system model. In this paper we restrict ourselves to the observability problem, but because of the Kalman duality theorem the results can easily be extended to the controllability problem. The structure of the system model is not only important in real time filtering, but also for the post-mission adjustment, if, e.g., the state vector is introduced as an unknown quantity in the adjustment.