Showing papers on "Inertial navigation system published in 1988"

Navigation System Integrity Monitoring Using Redundant Measurements

Abstract: The advantages of a navigation system that can monitor its own integrity are obvious. Integrity monitoring requires that the navigation system detect faulty measurement sources before they corrupt the outputs. This paper describes a parity approach to measurement error detection when redundant measurements are available. The general form of the detector operating characteristic (DOC) is developed. This equation relates the probability of missed detection to the probability of false alarm, the measurement observation matrix, and the ratio of the detectable bias shift to the standard deviation of the measurement noise. Two applications are presented: skewed axis strapdown inertial navigation systems, where DOCs are used to compare the integrity monitoring capabilities of various redundant sensor strapdown system configurations; and GPS navigation sets, where DOCs are used to discuss GPS integrity monitoring for meeting non-precision approach requirements. A fault identification algorithm is also presented.

Control theoretic approach to inertial navigation systems

Abstract: In this work, the analysis of inertial navigation systems (INS) is approached from a control theory point of view. Linear error models are presented and discussed and their eigenvalues are computed in several special cases. It is shown that the exact expressions derived for the eigenvalues differ slightly from the commonly used expressions. The observability of INS during initial alignment and calibration at rest is analyzed. A transformation that is based on physical insight is introduced that enables us to determine the unobservable subspace and states rather easily by inspection of the new dynamics matrix. Finally, the relationship between system observability and quality of estima- tion is presented. ever, as will be shown in the sequel, the inclusion of the vertical channel does alter the eigenvalues slightly. In the examination of system observability, we use a straightforward transforma- tion into observable and unobservable subsystems that, in turn, expose the states that hamper the estimation of INS errors during the initial alignment and calibration phase of operation. This approach was adopted successfully in the past by Kor- tum3 who considered the problem of INS platform alignment in which the measurements were the horizontal accelerometer outputs, whereas in the present case the measurements are the INS horizontal velocity components. In addition, the compari- son of this approach to the classical one that is presented in the present paper, as well as the discussion of uniqueness and the relationship between observability and quality of estimation, provide additional insight into the observability issue. It is hoped that the examination of INS as a unified system from a control theory point of view will shed more light on the system and contribute additional insight into the analysis of INS. In the next section, we describe the INS linear error model that will be the investigated plant. In Sec. Ill, we investigate the eigenvalues of INS in various phases of operation, and in Sec. IV, the issues of controllability and observability of the system are discussed. The relation between system observability and the ability to estimate its states during initial alignment is discussed in Sec. V. Finally, in Sec. VI, the conclusions are presented.

Federated filter for fault-tolerant integrated navigation systems

Abstract: An efficient, federated Kalman filtering method is presented, based on rigorous information-sharing principles. The method applies to decentralized navigation systems in which one or more sensor-dedicated local filters feed a larger master filter. The local filters operate in parallel, processing unique data from their local sensors, and common data from a shared inertial navigation system. The master filter combines local filter outputs at a selectable reduced rate, and yields estimates that are globally optimal or subset-optimal. The method provides major improvements in throughput (speed) and fault tolerance, and is well suited to real-time implementation. Practical federated filter examples are presented, and discussed in terms of structure, accuracy, fault tolerance, throughput, data compression, and other real-time issues. >

Low altitude warning system for aircraft

Abstract: To provide a predictive emergency warning to the pilot of flight and terrain conditions which will result in a collision with the ground unless the pilot takes immediate action, without issuing excessive nuisance warnings, the system uses input parameters from other aircraft systems such as the radar altimeter. Inertial Navigation System, and Central Air Data Computer which are processed in an on-board computer to determine when a warning is required. A software program which is part of the warning system provides a logic link between the on-board aircraft parameters and the on-board voice command ("pull-up, pull-up"). The software program is readily adaptable to all aircraft applications with varying amounts of modification depending on specific mission requirements for which this protection is required. A feature is the use of a continuously computing predictive warning algorithm (based on classical flight dynamics equations) in combination with unique inhibit logic equations. Another feature is the introduction of "extended coverage" logic which permits the altitude dependent, time limited use of an alternate (other than radar) altitude reference signal when the radar altimeter is beyond limits.

High-Accuracy Kinematic Positioning by GPS-INS

Abstract: Recent results achieved with relative GPS positioning techniques indicate that accuracies at the meter level are possible in land vehicle mode if the cycle slip problem can be minimized. One of the possible solutions to the problem is the integration of GPS and inertial data. Results from inertial surveys show that, with regular, accurate coordinate or range updates, an INS will give velocity estimates at the cm/s level [1]. By integrating differential GPS measurements with an INS, the effect of cycle slips over short intervals may be eliminated from the positioning results. A Kalman filter-smoother to handle this problem has been developed. It integrates differential range and phase measurements with data from an inertial navigation system. The optimal backward smoother improves the filter estimates for periods of poor geometry and multiple cycle slips. The package has been tested with GPS and INS data from a baseline survey near Calgary, Canada. Results of the test show that sub-meter kinematic positioning accuracies and cm/s velocity accuracies are achievable with an integrated GPS-INS.

Strapdown inertial measurement units for motion compensation for synthetic aperture radars

Abstract: It is shown that synthetic-aperture radar (SAR) motion can be compensated by using an antenna-mounted strapdown inertial measurement unit (IMU) as the motion sensing system, but sensor and system errors affect SAR image quality. A strapdown IMU consists of three accelerator channels and three gyro channels. Strapdown IMU errors include gyro-scale and accelerometer-scale factor and bias errors, velocity error, platform tilt, and errors induced by limited inertial sensor bandwidth. The effects of these errors on the SAR image quality are presented in terms of the SAR impulse response. IMU errors that cause low-frequency phase errors (less than one cycle per array time) are categorized in terms of quadratic and cubic phase errors. IMU errors that cause high-frequency phase errors (greater than one cycle per array time) are categorized in terms of the integrated sidelobe ratio and peak sidelobe ratio. A motion compensation system conceptualization is described wherein a strapdown IMU is attached to an antenna and transfer-aligns to the aircraft's master navigator. >

Integration of GPS receivers into existing inertial navigation systems

Abstract: Many inertial navigation systems of both platform and ring laser strapdown types are currently in service. This paper discusses the possibility and desirability of incorporating a small GPS receiver in these systems. Advances in technology such as microprocessors, gate arrays, and surface mount devices allow the existing INS electronics to be replaced in a reduced volume. The remaining space in many cases is sufficient to permit the insertion of a small GPS receiver. Locating the GPS receiver in an inertial navigation system (INS) solves many of the usual system integration problems. Tight coupling between the GPS and INS can be achieved since data latency is minimized and well controlled. In such a configuration, rate aiding of the GPS is easily achieved. This approach also leads to greater flexibility and enhanced overall performance since all GPS and INS data are simultaneously available. While not providing the ultimate in redundancy, the integrated INS/GPS approach does offer greater simplicity with enhanced performance. This discussion primarily focuses on military systems. Nevertheless, the proposed techniques also are applicable to commercial units.

Terrain-Aided Navigation Test Results in the AFTI/F-16 Aircraft

Abstract: Sandia National Laboratories has developed a low-level attack aircraft terrain-aided navigation algorithm for the Air Force Wright Aeronautical Laboratories, Avionics Laboratory. This algorithm is based on the Sandia Inertial Terrain-Aided Navigation (SITAN) approach to terrain-aided navigation that was initially formulated at Sandia in the mid-1970s. Extended Kalman filter theory is used to provide essentially continuous terrain-aided navigation through estimation of inertial navigation system errors from radar altimeter ground clearance measurements and on-board digital terrain elevation data (DTED). The SITAN algorithm is integrated into the Advanced Fighter Technology Integration (AFTI)/ F-16 aircraft, and called AFTI/SITAN to distinguish it from other SITAN designs. AFTI/SITAN performance, as determined by Air Force tracking radar during real-time testing in the AFTI/F-16 aircraft at Edwards Air Force Base from September 1986 through April 1987, is presented. Required performance is less than 100 m median horizontal radial error over 200 nmi trajectories that are flown over gently rolling terrain with available DTED. The flight data support the conclusions that AFTI/SITAN reliably determines aircraft position within an initial 0.5 nmi, CEP horizontal position error, and tracks aircraft horizontal position with an accuracy of 75 m median radial error using DTED supplied by the Defense Mapping Agency. Although accurate estimation of absolute altitude is not required for proper SITAN operation, altitude is estimated with less than 17 m root-mean-square error.

Computerized flight inspection system

Abstract: A computerized flight inspection system is disclosed. The system of the present invention may be utilized to generate an accurate reference location with respect to an airport runway for an aircraft having an inertial navigation system. A selected geometric pattern having a highly unambiguous autocorrelation function is placed on at least one end of the runway. A video line scanning camera mounted to the aircraft is then utilized to scan the geometric pattern in a line generally perpendicular to the line of flight. The output of the video line scanning camera is then correlated with a stored reference pattern to generate a signal indicative of the detection of the geometric pattern on the runway. A laser altimeter is mounted to the aircraft and utilized to generate an accurate signal indicative of the aircraft altitude with respect to the runway pattern. The outputs of the correlation circuit and the laser altimeter are then utilized to correct data from the inertial navigation system.

Sensor array dynamic position and orientation determination system

Abstract: A system, which uses inertial measurement units, is shown for determining the position and orientation of a towed array of sensors used for target detection. The system uses an onboard master inertial navigation system and a relative position determination mechanism to generate a first estimated position for each inertial measurement unit within the array. Each inertial measurement unit measures force and angular change information used by an onboard computer to create a second estimated position by known methods for each inertial sensor. An error signal represented by the difference between the two estimated positions for each inertial unit is processed over time by a Kalman filter to reduce the error in the heading and attitude determined for each inertial unit to establish an accurate location for each inertial unit and, thus, the towed array of such units.

Applications of Dual Aircraft Formation Flights

Abstract: We discuss procedures for analyzing dual aircraft formation flights using time-lapse photographs of one aircraft from the other, combined with inertial navigation system position measurements, to estimate the displacement vector between the two aircraft. We show that accuracies of a few percent of the separation distance can be readily achieved, and we develop a technique for aligning the datasets from the two aircraft to correct for variations in the longitudinal component of the displacement vector. We then derive an expression for the variance of the difference between measurements of the same variable on each aircraft as a function of averaging time and separation distance. An example of data from a series of formation flights over eastern Colorado is used to demonstrate the techniques for estimating the displacement vector, aligning the datasets, and calculating. lateral coherences and phase angles.

Recent Developments in the "TERPROM" Integrated Navigation System

Abstract: The British Aerospace "TERPROM" system is briefly introduced, and the operation of the "continuous mode" Kalman filter is explained. The advantages of TERPROM as a navigation system are summarised, and its performance over flat ground is assessed. A number of ways of enhancing the TERPROM technique are in traduced, such as the addi tiona! use of Doppler, Scene Matching or GPS data, and the resultant improvement over flat ground is demonstrated. Alternative forms of TERPROM - which do not rely on the use of an Inertial Navigation System are described, and it is shown that performance with these reversionary modes is equally reliable. Finally a new navigation system known as BRAINS The British Aerospace Integrated Navigation System is introduced. This new system has the advantages that: it does not rely upon any one sensor to perform the basic navigation; sensors can be added or removed at will, with a minimum of interference; and tests for integrity and sensor reliability can conveniently be carried out.

Navigation: ships to space

Abstract: An examination is made of old and new navigation technology that considers most factors affecting this technology, including cost, accuracy, autonomy, time-delay, global coverage, and the human interface. The author reviews cartography and the navigation of land, ship, air, and space vehicles, concluding with a forecast of navigation in the twenty-first century. An extensive annotated bibliography is included. >

Application of fixed point smoothing to the calibration, alignment and navigation data of inertial navigation systems

Abstract: The authors present an application of the fixed-point smoothing technique to calculate the relative azimuth of inertial measurement units (IMU), by postprocessing of calibration and alignment data. It is shown that the smoother performs best when the filter and smoother error covariances for gyro fixed-drift-rate parameters are increased. A test scenario was created with the calibration and alignment data from an IMU simulator. The data were postprocessed and the smoother estimated all but 0.1 arc seconds of the filter azimuth error. This is well within the expected accuracy for relative azimuth computation. >

Using low cost magnetic sensors on magnetically hostile land vehicles

Abstract: The author discusses methods used to compensate for a land vehicle's magnetic signature and gives results obtained with various types of vehicles in several world locations. Advantages and disadvantages of pendulous mounting are considered and vehicle magnetic changes caused by driving in areas of extremely strong magnetic fields, such as those generated by DC-powered trains, are discussed. It is suggested that dead-reckoning sensors will be needed on land because of the line-of-sight characteristics of GPS (Global Positioning System) and the difficulty land vehicles have in operating with as many as four satellites constantly in view. Flux gate magnetic sensors will provide a reliable means of measuring heading in many areas where vehicle navigation and tracking needs exist. >

Inertial navigation system synthesis approach and gravity-induced error sensitivity

Abstract: A design criterion for improving the performance of the speed-damped inertial navigation system is presented. The single-axis speed-damped system is approached by optimizing the response of the system to a step-function disturbing signal. Butterworth, integral of time-multiplied absolute-value of error (ITAE), and solution-time standard forms are assumed to be the figures of merit for optimizing the system performance. The steady-state RMS (root-mean-square) gravity-induced navigation errors that are excited in the speed-damped system are determined for two gravity uncertainty models. The proposed figures of merit are compared. These comparisons reveal the sensitivity of predicted navigation errors to uncertainties in the gravity statistics, and simplify the choice of a suitable figure of merit for use in the design and error analysis of inertial navigation systems. >

Angle of Attack and Sideslip Estimation Using an Inertial Reference Platform

Abstract: : The purpose of this research is to develope and flight test concepts for the estimation of angle of attack (alpha) and sideslip (beta) using an inertial reference platform. This development was further broken down into real- time, inflight estimation of alpha and post-flight estimation of alpha and beta. Following theoretical development, the concepts were tested with NASA F-15A flight data and examined real-time during a nasa Highly Integrated Digital Engine Control (HIDEC) flight test using the F-15A aircraft. Angle of attack is a critical parameter in the maneuverable, high performance aircraft of today. Yet many errors are present in the current methods of obtaining this angle. An accurate method of alpha and beta estimation could eliminate the need for such probes, and allow these quantities to be used for a broad range of applications. An inflight estimator was developed for computational speed and accuracy using inertial navigation system linear accelerations and angular rates. A second system based on linear recursive modeling was developed for post-flight estimation of alpha and beta. The data and programs specified in this research are applicable only to those aircraft mentioned, but the methods of estimation are universal.

Fault tolerant inertial navigation system

Abstract: This paper describes the sensor redundancy management concept employed in a fault tolerant inertial reference system built by Honeywell's Commercial Flight Systems Minneapolis Operation to be flight tested in 1988. vided. No attempt was made to use state of the art packaging. Existing circuits were used wherever feasible, and low voltage power is provided using external commercial power supplies. The hexad IRU is shown in figurel. The concept uses a sensor hexad and employs separate fault detection and isolation (FDI) algorithms for navigation computations and outputs to the flight control system. The flight control FDI employs large thresholds and short time constants consistent with flight control requirements. In order to achieve the low detection levels required to maintain navigation accuracy, the navigation FDI employs heavy filtering, low thresholds, Kalman filtering to correct the parity residuals for sensor misalignment and scale factor errors, and dynamic parity gains to minimize the effects of maneuvers. Two redundant nav solutions using sub-optimal sensor sets are also included to reset the primary solution after failures. The FDI algorithm implementation is based on the generalized likelihood ratio test (GLRT) that has been the subject of several papers in recent years. Enhancements to the failure isolation logic virtually eliminate the probabilities of false detection and wrong isolation inherent with the basic GLRT concept, thus making the GLRT scheme practical in the real world of high reliability requirements.

The Cost and Benefits of Reliability in Military Equipment

Abstract: : This report investigates the costs of achieving greater reliability in military equipment, the benefits of improved reliability in reduced support costs and increased availability, and strategies for attaining reliability goals. Three kinds of evidence were examined: reliability improvement programs, new product developments, and statistical analyses of reliability costs and outcomes in new programs. Literature reviews of both military and commercial experience, and interviews with reliability experts and managers in government and industry, provided additional information to supplement detailed case studies. Performed were detailed case studies of seven systems: F-18 aircraft, CH-47D helicopter modernization, F100 turbine engine, Phalanx Mk15 Close-in Weapon System, LAMPS MKIII helicopter antisubmarine warfare system, Minuteman I inertial navigation system, and the Carousel inertial guidance system, Minuteman I inertial navigation system, and the Carousel inertial guidance system. Information produced in other studies was reanalyzed in the context of the present research; these included F-16 aircraft reliability improvements, spacecraft reliability costs, Duane models of reliability growth, and a study of 19 Navy systems. These cases covered several different technologies and were drawn from all three military services.

Use of an external cascade Kalman filter to improve the performance of a Global Positioning System (GPS) inertial navigator

Abstract: The accuracy of in-air inertial navigation system (INS) alignment has been analyzed for several methods of GPS/INS integration. It is found that an external filter which operates on the outputs of the GPS receiver achieves a faster alignment than the GPS receiver alone. Steady-state accuracy is not significantly improved by the use of the external filter. The data rate of this external filter should be one update per second, which is the same as the GPS receiver's update rate and is the fastest that the external filter could be updated. In-air alignment of a second INS can be performed as a transfer alignment. In this procedure, the INS, which is integrated with the GPS, transfers its alignment to the second INS using an external filter. The performance of this external filter is analyzed and it is found that for a medium-accuracy INS, and also for a low-accuracy attitude heading reference system (AHRS), the fastest update rate provides the best alignment response. >

Angle of attack estimation using an inertial reference platform

Abstract: This paper presents the mathematical development and flight test results of an angle of attack estimation system based on inertial navigation system inputs. The estimator uses these inputs to determine the coefficient of lift required at any instant inflight. Angle of attack is then modeled through a regression analysis based on coefficient of lift, altitude and Mach. Overall correlation of the estimator as tested was generally within 0.5 degrees through 17 degrees angle of attack on an F-15A aircraft. A robustness analysis indicates that the system can be used adequately in maneuvering flight.

Correction for deflections of the vertical at the runup site

Abstract: The effects of unmodeled deflections of the vertical on aircraft inertial navigation system performance are considered. It is argued that, with few exceptions, the work on the area to date has ignored an important aspect, the correlated effects of deflections of the vertical at the runup site interacting with the systems's self-alignment algorithm. The author reviews statistical analysis and flight test results, examines the effects of self-alignment on navigation errors, and describes a methodology that applies appropriate corrections during self-alignment and static navigation testing, and removes those corrections during flight. An implementation that substantially reduced the resulting errors for navigation time of one or two Schuler cycles was successfully demonstrated during flight test of the Mini-GEANS (Gimbaled Electrostatic Gyro Aircraft Navigation System). >

Development and field testing of a RLG strapdown inertial survey system

Abstract: The authors describe the successful modification of a ring laser gyro (RLG) strapdown inertial reference unit into a strapdown inertial survey system (SISS) for geodetic positioning. The system, a Litton LTN-90-100, has been modified into a land-vehicle mode inertial survey system by developing specialized software and error control techniques. The software package integrates the raw body rate and acceleration data from the LTN-90-100 into velocity and coordinates of the system. The error states of the SISS are estimated by a Kalman filter-smoother using regular zero velocity measurements and occasional control coordinates as updates. The system was tested on an L-shaped baseline near Calgary and results show that a positioning accuracy of better than 1 m (1 sigma ) is achievable. >

An integrated gps/irs design approach

Abstract: Development and testing of a prototype GPS receiver housed in a commercial inertial reference system has recently been completed. A flight test program in 1986 demonstrated dynamic performance in maneuvering flight, operation in low signal/noise environments, and operation at zero degree satellite elevation angles with 20–30 m typical accuracies using the Block I GPS satellites. The effects of altitude and clock aiding were also demonstrated. Many of the recognized limitations of an inertial system (including unbounded position error and pilot initialization as well as stand-alone GPS limitations such as poor satellite geometry, dynamic noise and satellite masking) are overcome with this integrated design approach. This paper describes the advantages of GPS/IRS integration as well as a design approach that utilizes them. The paper also summarizes some test results of the prototype GPS receiver.

Functional Performance Specification for an Inertial Navigation System

Abstract: : This document defines the functional and performance requirements for the inertial navigation system simulator that interfaces with the external computer system (ECS) simulator. Both the INS simulator and the ECS simulator are being developed in Ada by the Real-Time Embedded Systems Testbed Project. The INS simulator is similar to a real-world InS, but has reduced functionality. This document provides specifications for the major functions on the INS simulator. Keywords: functional performance specification, Inertial navigation system, External computer system.

Development of navigation guidance and control technology for Indian launch vehicles

Abstract: The orbital injection accuracy of any payload depends on the calibre of the inertial guidance system used on board the launch vehicle. This paper outlines the mission definition and the rationale for the selection of a proper guidance system to meet final mission objectives. The functions and the architecture of the navigation, guidance and control are discussed. The developmental aspects of the sophisticated inertial sensors, inertial systems, associated complex electronics, on-board computers, control actuators and systems are reported. The complexity of the on-board control and guidance software and the test and evaluation procedures used for their validation are included. The general scheme of the inertial guidance systems and the critical role played by them in the realization of Indian satellite launch vehicles SLV-3, ASLV and PSLV are presented in brief.

Inertial Drift Correction for Aircraft-Derived Wind Fields

Abstract: During the last three decades, aircraft have assumed an increasingly important role as platforms for atmospheric measurement. Because of the nature of the wind measurement process, however, low-frequency winds remain susceptible to significant errors, which have their origin in the inertial navigation system (INS) of the aircraft. Computations are presented showing that the Electra, a state-of-the-art aircraft for meteorological measurements, experiences low-frequency wind measurement errors typically as large as 3 m s−1 in each component of the horizontal wind. Errors of this magnitude can produce can produce line-integral divergence errors of 10−4 s−1 in typical flight geometries. Use of the LORAN-C navigational information now received on the aircraft offers an order-of-magnitude reduction of errors from the INS.

Low cost high accuracy integrated GPS-inertial navigator for reconnaissance missions

Abstract: The authors affirm that the ability to deliver a reconnaisance payload over 300 miles from a launch point to better than 35 m can be achieved by using a tightly integrated GPS (global positioning system)-inertial system. This system provides navigation, guidance, airframe and payload management functions for an unmanned air vehicle (UAV). High accuracy is achieved by the integration of a 22-state Kalman filter for two low-cost sensors: a single-channel GPS receiver and a multisensor inertial measurement unit. The authors discusses various aspects of the UAV navigator functions and the hardware that will be used. The simulations run for a typical mission scenario, exclusive of satellite errors, demonstrated results well within specified accuracy. >