Showing papers on "Inertial navigation system published in 1973"

Tracking system utilizing Kalman filter concepts

Abstract: A monopulse radar tracking system is disclosed employing four tracking channels, one for each of range, velocity, azimuth and elevation. The two angle channels are referenced directly to antenna coordinates. Each channel is mechanized by a stored program in a digital computer, and the mechanization employs Kalman filtering with gain factors continually optimized for measured signal-to-noise ratios. Angular rate commands for the antennas are obtained by passing pointing error estimates from the angle channels through a compensator for antenna motion and adding line-of-sight rate estimates from the angle filter channels. Cross-coupling between channels is provided, and each channel is aided by outputs from an inertial navigation platform. Preaveraging of discriminants received between computational cycles is provided.

Effects of Geodetic Uncertainties on a Damped Inertial Navigation System

Abstract: Gravity uncertainties are an inexorable source of error in all inertial navigation systems and are particularly important in high-quality inertial navigation systems. In this paper the steady-state rms errors that are excited in a damped inertial navigation system are analytically determined for four gravity uncertainty models and two vehicle maneuver models. The statistical approach used in this paper is compared with an alternate scheme (?direct simulation?) that does not require statistical models for gravity uncertainties.

Star Scanner Attitude Determination for the OSO-7 Spacecraft

Abstract: Calculation of the inertial orientation of the OSO-7 spacecraft from the times at which known stars or planets transit planes fixed in the spacecraft. Both the reference planes and the timing information are provided by the star scanner instrument aboard the spacecraft, while the star identification and the statistical estimation of a set of parameters describing the spacecraft attitude are accomplished in a ground station computer facility. A recursive least-squares determination is made of a vector of first-order differential corrections to the attitude state vector. Preliminary analysis indicates the system accuracy to be 3 arc min in each attitude Euler angle.

Cross track strapdown inertial quidance system

Abstract: A simplified strapdown inertial guidance system is provided for the cross track or inertial ''''line-of-sight'''' planar guidance of an unmanned tactical vehicle, or the like; wherein one of the attitude angles can be of large magnitude. The system includes three orthogonally-mounted accelerometers and two gyroscopes. The large attitude angle is accommodated by torquing the appropriate gyroscope in a capture loop. The system implements the body attitude angles to inertial attitude angles in two steps. First a transformation is made from body to gyroscope coordinates, and then a transformation is made from gyroscope to inertial coordinates.

Behaviour of the spinning gyro rotor

Abstract: An elementary understanding of the physical laws governing the spinning rotor 1S necessary to grasp the essence and function of the various gyroscopic instruments found in to-day's aerospace and maritime vehicle control systems. This report gives a brief exhibition of the very basic phenomena of gyro behaviour in a way that may especially appeal to the control engineer with a limited educational background in the field of mechanical engineering.

Updating Inertial Navigation Systems with VOR/DME Information

Abstract: Updating an inertial navigation system (INS) with VOR/DME information (from one or two stations) by means of a maximum-likelihood filter is shown to result in substantial improvements in navigational accuracy over that obtained by the use of a single VOR/DME (current practice). When continuously updating, the use of a high- quality INS (0.01 °/hr gyro drift) instead of a low-quality INS (1.0°/hr gyro drift) does not substantially improve position accuracy. In-flight alignment (or realignment) of an INS to an accuracy comparable to that of ground alignment can be accomplished by using two DME's. Several reduced-order suboptimal filters were found to perform nearly optimally. HE primary navigation aid for civil aircraft flying in the airspace of most of the developed countries of the world is the VOR/DME system. The VOR (Very high-frequency Omni- Range) and the DME (Distance Measuring Equipment) enable onboard determination of an aircraft's bearing relative to north at the fixed ground station and slant range from the station, respectively. Current use of the VOR/DME system involves primarily radial navigation, i.e., aircraft fly directly to or from the ground stations. However, some beginnings have been made in using the VOR/ DME system for area navigation, i.e., use of the system without being restricted to fly directly to or from the ground stations. The number of commercial airliners equipped with inertial navigation systems (INS's) is steadily increasing. The systems now onboard aircraft utilize a gyro-stabilized platform on which the accelerometers are mounted. The platform is aligned before takeoff to the desired orientation. Due to alignment errors and in-flight random disturbances such as gyro drift, scale-factor errors, and accelerometer bias errors, errors in the desired orientation of the platform increase with time. This results in increasing position and velocity errors. Thus, improving navi- gational accuracy by in-flight realignment of an INS is an interesting possibility. Position errors are generally greater for area than for radial navigation. This comes about because the position error resulting from a VOR angular error increases with distance from the station; and an aircraft is farther, on the average, from the VOR stations for area than for radial navigation. Hence, improved navigational accuracy is required to obtain an accuracy for area navigation comparable to that of present-day radial navigation. The availability of an onboard computer to do the triangulation computations required for area navigation suggests Presented as Paper 72-846 at the AIAA Guidance and Control

Redundancy management of inertial systems.

Abstract: The paper reviews developments in failure detection and isolation techniques applicable to gimballed and strapdown systems. It examines basic redundancy management goals of improved reliability, performance and logistic costs, and explores mechanizations available for both input and output data handling. The meaning of redundant system reliability in terms of available coverage, system MTBF, and mission time is presented and the practical hardware performance limitations of failure detection and isolation techniques are explored. Simulation results are presented illustrating implementation coverages attainable considering IMU performance models and mission detection threshold requirements. The implications of a complete GN&C redundancy management method on inertial techniques are also explored.

Tilt-Table Alignment for Inertial-Platform Maintenance without a Surveyed Site

Abstract: A method for aligning the rotary tilt table for an inertial-platform maintenance facility without surveying the site is analyzed and evaluated. The method utilizes multiple measurements of the tilt-table azimuth alignment error with different inertial platforms to determine a best estimate of the alignment error. Error analysis indicates that useful facility performance can be obtained with only a small number of measurements. The tilt-tablealignment accuracy can be improved as more measurements of the alignment error are made during normal facility operation.

Altitude Damping of Space-Stable Inertial Navigation Systems

Abstract: In order to stabilize the altitude calculation in an inertial navigation system, an altimeter is commonly used. In a conventional local-level mechanization, this is generally accomplished by correcting the vertical channel integrators with the difference between the inertial system and altimeter indication of vertical position. However, in a space-stable system the procedure is not as clear since a vertical channel is not physically present. Three altitude damping mechanizations for a space-stable inertial navigation system are proposed. The equivalent local-level mechanizations are then found by comparing error propagation equations in a common coordinate frame.

Stellar scanning-chopping system

Abstract: A star scanning device for use in inertial navigation systems and the like in which a scanning disk containing a V slit and having a position known at all times is used to determine the X and Y position of a star with respect to a fixed platform. Also shown is a chopping disk used with the scanning disk to improve performance and the system to demodulate and process the signals obtained in order to provide the coordinate outputs for further use in the inertial system.

Direct measurement of earth{40 s vertical deflection using ship{40 s inertial navigation system

Abstract: A method for obtaining at a point the value of earth''s vertical deflection, which is the angular difference between the local or astronomic vertical at a point and the normal to the reference ellipsoid of the earth at that point. The position coordinates of the point is measured by using a ship''s inertial navigation system (SINS) which is compared with corresponding the position coordinates of the same point measured by using a geodetic reference apparatus, such as a long-range electronic navigation system (LORAN) or high precision short-range navigation (HIRAN). The difference between the two values of the position coordinates of the point; also called dPT total position error; is due to vertical deflection; i.e., the difference between local vertical and geodetic vertical or normal to the reference ellipsoid of the earth at that point; error due to gyro drift of the SINS; and reference velocity error. Contributions due to error because of gyro drift and to error in reference velocity are subtracted from dPT, the value of the difference between the two values of position coordinates of the point, using SINS RESET technique and a standard reference velocity technique respectively so as to obtain dPV, position error due to earth''s vertical deflection. The value of dPV is then mainly due to earth''s vertical deflection at that point, from which the value of vertical deflection at that point is obtained using SINS INVERSE FILTER technique.

Application of Extended Kalman Filtering to a Dynamic Laboratory Calibration of an Inertial Navigation System

Abstract: : The report describes a data reduction technique that obtains estimates of inertial sensor error model coefficients from a dynamic laboratory calibration of a typical Inertial Navigation System. The error model coefficients are those associated with gyros, accelerometers, and their misalignment errors that have been found by test and analysis to be the predominant sources of error affecting system accuracy. All the error terms considered are categorized as either fixed (independent of applied acceleration) , first-order (proportional to the first power of acceleration), or higher-order terms, which are proportional to the square or cube of acceleration. In the case of the higher-order terms, the error model coefficients of inertial grade sensors are from one to four orders of magnitude smaller than the fixed and first-order terms. (Modified author abstract)

Kalman Filter Design Considerations for Space-Stable Inertial Navigation Systems

Abstract: Terrestrial inertial navigation is typically performed using an instrumented platform stabilized in a ?local-level? configuration for convenient generation of geographic navigation information. The local-level geographic reference must be maintained by torquing the system gyros, a requirement which may be incompatible with high-precision inertial sensors currently under development. Gyro torquing in a gimballed navigation system can be avoided by employing a ?space-stable? mechanization of the platform where an inertial, rather than geographic, reference is used for navigation calculations. The software design problems associated with this concept, especially those related to the application of Kalman filtering, are the principle focus of this paper. Although the space-stable configuration has been used extensively for spacecraft navigation and guidance, it has not been widely applied to terrestrial navigation, either for air or marine applications. The chief problem in this application is to perform navigation in local-level coordinates, using system outputs generated in an inertial reference frame. It can be demonstrated that, although the navigation system error dynamics are identical for the local-level and space-stable configurations, the dynamics of the sensor errors which drive these systems are quite different. These differences in sensor error propagation characteristics impose new requirements for the design of procedures to accomplish system calibration, alignment, and reset. This paper outlines a Kalman filtering approach which is applicable to all of the above procedures, and presents numerical results to demonstrate its effectiveness.

New mechanization equations for aided inertial navigation systems

Abstract: Inertial navigation equations are developed which use area navigation (RNAV) waypoints and runway references as coodinate centers. The formulation is designed for aided inertial navigation systems and gives a high numerical accuracy through all phases of flight. A new formulation of the error equations for inertial navigation systems is also presented. This new formulation reduces numerical calculations in the use of Kalman filters for aided inertial navigation systems.

Flight results from a study of aided inertial navigation applied to landing operations

Abstract: An evaluation is presented of the approach and landing performance of a Kalman filter aided inertial navigation system using flight data obtained from a series of approaches and landings of the CV-340 aircraft at an instrumented test area. A description of the flight test is given, in which data recorded included: (1) accelerometer signals from the platform of an INS; (2) three ranges from the Ames-Cubic Precision Ranging System; and (3) radar and barometric altimeter signals. The method of system evaluation employed was postflight processing of the recorded data using a Kalman filter which was designed for use on the XDS920 computer onboard the CV-340 aircraft. Results shown include comparisons between the trajectories as estimated by the Kalman filter aided system and as determined from cinetheodolite data. Data start initialization of the Kalman filter, operation at a practical data rate, postflight modeling of sensor errors and operation under the adverse condition of bad data are illustrated.

Strapped Down Inertial Navigation Systems

Abstract: The heart of an inertial navigation system is the Inertial Measuring Unit (I.M.U.) as this carries out the fundamental tasks of measuring the vehicle acceleration and providing a spatial reference. The I.M.U. thus determines the system performance and accuracy; it also accounts for about two thirds of the initial cost of the system and is a major factor in the cost of ownership.Up to now the majority of IN systems have used a stable platform I.M.U. where the gyros and accelerometers are mounted on a gimbal suspended platform which is servo controlled from the gyros. This has greatly eased the task of developing suitable gyros of the required accuracy, as the gimbal stabilization system isolates the gyros from the angular motion of the vehicle and greatly simplifies the subsequent computation in terms of axis transformations; it provides a direct read out of the euler angles—heading, pitch and roll. However such a system is inevitably mechanically complex and its ultimate reliability is constrained by such components as slip rings, servo motors, synchros, resolvers, encoders, gimbal bearings &c.

Flight test results from the CV990 simulated space shuttle during unpowered automatic approaches and landings

Abstract: Unpowered automatic approaches and landings with a CV990 aircraft were conducted to study navigation, guidance, and control problems associated with terminal area approach and landing for the space shuttle. The flight tests were designed to study from 11,300 m to touchdown the performance of a navigation and guidance concept which utilized blended radio/inertial navigation using VOR, DME, and ILS as the ground navigation aids. In excess of fifty automatic approaches and landings were conducted. Preliminary results indicate that this concept may provide sufficient accuracy to accomplish automatic landing of the shuttle orbiter without air-breathing engines on a conventional size runway.

A low-cost inertial smoothing system for landing approach guidance

Abstract: Accurate position and velocity information with low noise content for instrument approaches and landings is required for both control and display applications. In a current VTOL automatic instrument approach and landing research program, radar-derived landing guidance position reference signals, which are noisy, have been mixed with acceleration information derived from low-cost onboard sensors to provide high-quality position and velocity information. An in-flight comparison of signal quality and accuracy has shown good agreement between the low-cost inertial smoothing system and an aided inertial navigation system. Furthermore, the low-cost inertial smoothing system has been proven to be satisfactory in control and display system applications for both automatic and pilot-in-the-loop instrument approaches and landings.

Accuracy of ESG monitor/sins inertial navigation system

Abstract: Discussed and analyzed is a marine navigation system in which a high performance system, the Ships Inertial Navigation System (SINS), is monitored by an even more accurate system now under development, the Electrostatically Supported Gyro Navigator (ESGN). An advantage of this monitor configuration is a potentially substantial improvement in navigational accuracy, while imposing a minimal change to the existing system. General equations for performing a covariance error analysis of the SINS, ESGN, and ESG Monitor/SINS systems are described. Results are presented which draw relative accuracy comparisons between these systems.

Differential velocity meter

Abstract: An inertial guidance system for measuring the difference between programmed and actual velocity change in a missile including a sensor for converting a velocity change to a shaft angle located between the sensor case and a pendulous gyro mounted on an internal turntable and means for converting the analog angle indication to digital read out data.

The Use of the 3.4 kHz Omega Difference Frequency for Aircraft Navigation.

Abstract: : The Naval Research Laboratory (NRL) conducted an investigation to determine the effect of multi-mode propagation on the usefulness of the Omega 3.4 KHz difference frequency for aircraft navigation. This investigation was conducted under the sponsorship of the Federal Aviation Administration (FAA) and with the cooperation of the National Oceanic and Atmospheric Administration (NOAA). A NOAA aircraft that was equipped with Omega receivers was used for this investigation. Flights were made between Grand Forks AFB, Grand Forks, North Dakota and Piarco airport in Trinidad. Position fixes, for comparison with those obtained from the Omega system, were obtained from the standard VOR/VORTAC stations when the aircraft was over the continental United States and doppler radar and inertial systems were used on the portion of the flights that were over water. The flights made during nighttime conditions when the effects of multi-mode propagation were the most pronounced. The data for this specific path and under nighttime conditions show that the 3.4 kHz difference frequency can be used for navigation using only simple correction factors if an accuracy of + or - 5 miles is acceptable. More data are required to determine the accuracy that can be obtained in other areas and during other propagation conditions. (Author)

Inertial reference unit

Abstract: The inertial reference unit is a high performance gyro attitude reference system for use on the OAO spacecraft. The IRU is a three axis system, which provides both rate and attitude information for spacecraft control. The purpose of the IRU is to reduce the dependency on gimballed startrackers and, in turn, simplify the OAO ground operations by eliminating the need for the continual programming of gimballed startracker assignments in accordance with computed occultation schedules. During normal operations, it is used to control the pitch and yaw axes during experiment occultations and during spacecraft reorientations. The roll axis is continuously under control of the IRU except during brief periods for attitude update. To provide for these capabilities the IRU must be able to perform two basic functions. One is to maintain an inertially fixed reference for spacecraft control and the second is to accurately reorient the reference upon command.

KalmanFilter Design Considerations for

Abstract: Terrestrial inertial navigation istypically performedusingan instrumented platform stabilized ina "local-level" configuration for convenient generation ofgeographic navigation information. The local-level geographic reference mustbemaintained bytorquing the systemgyros,a requirement whichmay be incompatible with high-precision inertial sensorscurrently underdevelopment. Gyro torquing in a gimballed navigation systemcan be avoidedby employing a "space-stable" mechanization oftheplatform wherean inertial, rather thangeographic, reference isusedfornavigation calculations. The softwaredesignproblems associated withthis concept,especially thoserelated to theapplication of Kalman filtering, aretheprinciple focusofthis paper. Althoughthe space-stable configuration has been used extensively forspacecraft navigation andguidance, ithasnotbeen widelyapplied to terrestrial navigation, eitherforairor marine applications. The chiefprobleminthisapplication istoperform navigation inlocal-level coordinates, using systemoutputsgenerated inan inertial reference frame.Itcan bedemonstrated that, although thenavigation system errordynamics areidentical forthelocal-level andspace-stable configurations, thedynamics ofthesensorerrors whichdrivethesesystemsarequitedifferent. Thesedifferences in sensor error propagation characteristics imposenew requirements forthedesignof procedures to accomplish systemcalibration, alignment, and reset.Thispaper outlines a Kalmanfiltering approach whichisapplicable to alloftheaboveprocedures, and presents numerical results todemonstrate itseffectiveness.

Strapdown system performance optimization test evaluations (SPOT), volume 1

Abstract: A three axis inertial system was packaged in an Apollo gimbal fixture for fine grain evaluation of strapdown system performance in dynamic environments. These evaluations have provided information to assess the effectiveness of real-time compensation techniques and to study system performance tradeoffs to factors such as quantization and iteration rate. The strapdown performance and tradeoff studies conducted include: (1) Compensation models and techniques for the inertial instrument first-order error terms were developed and compensation effectivity was demonstrated in four basic environments; single and multi-axis slew, and single and multi-axis oscillatory. (2) The theoretical coning bandwidth for the first-order quaternion algorithm expansion was verified. (3) Gyro loop quantization was identified to affect proportionally the system attitude uncertainty. (4) Land navigation evaluations identified the requirement for accurate initialization alignment in order to pursue fine grain navigation evaluations.

Analysis of aided inertial navigation systems performance on international routes

Abstract: The performance of hybrid navigation systems for commercial transoceanic flights is evaluated by means of a digital computer simulation. Error models are developed for aided-inertial navigation systems with external measurements from Doppler radar, Omega and satellite-ranging. Key features of the simulation program AIRNAV (Aided-Inertial Reference NAVigation) are described. Covariance matrix error analysis is used to obtain the navigation error histories, with recursive navigation techniques to incorporate the measurements. A 34th-order error state vector is defined, which requires the numerical integration of up to 585 first-order differential equations to propagate the covariance matrix. Example computer results are presented for a typical North Atlantic flight.