Showing papers on "Inertial measurement unit published in 1983"

Guided missile subsystem

Abstract: A guided missile subsystem including a Kalmanized radar track loop driven by acceleration signals of the missile generated by an inertial measuring unit (IMU), and a missile control loop driven by estimates of the relative kinematics of the missile and target computed by the radar track loop is disclosed. The IMU driven Kalmanized radar track loop accommodates the use of a high performance radar, like a synthetic aperture radar, for example, which operates to measure radar data at a low rate on the order of 1 Hz, to generate estimates of relative target and missile kinematics to drive the control loop at rates compatible with high performance missile kinematics. The Kalmanized track loop effects an exchange of IMU errors for "dynamic lag" errors of conventional track loops which cannot be modeled very well, and can change very rapidly. In contrast, the IMU errors can be modeled well, and in addition change very slowly which is what permits the Kalmanization function to work well in the track loop at reduced rates. Because of the dynamic exactness of the track loop, very good estimates of the relative kinematics of the missile may be supplied to the control loop to effect more accurate computations of maneuver commands which drive the controls of the missile. Moreover, the Kalmanized track loop does not let large amounts of angle glint noise into the control loop prior to missile impact. An effective bandwidth decrease as glint noise increases is provided without incurring a dynamic lag error penalty.

Inertial navigation

Abstract: Inertial Navigation Systems have found universal application both militarily and commercially. They are self-contained, nonradiating, nonjammable, and sufficiently accurate to meet the requirements of users in a most satisfactory manner. An overview of inertial navigation is provided, followed by several sections detailing a specific, but different mechanization approach. A Ring Laser Gyro (RLG) based navigation system design is reviewed with special emphasis directed at requirements for navigation accuracy and alignment time. Along with discussions of the RLG unit, an introduction to a novel accelerometer approach, the Vibration Beam Accelerometer (VBA), is provided. A gimballed, self-contained High Accuracy Inertial Navigation System, denoted HAINS, represents one approach toward achieving navigation capability of 0.2 nmi / h and an rms velocity of 1.5 ft / s per axis while retaining the form and fit and affordability of standard inertial tactical flight navigators. The Stellar-Inertial Navigation section illustrates the bounding of position and verticality errors thus achieving exceptional accuracies. Two gyroscopic approaches, presently in development are finally discussed. The Fiber Optic Gyroscope (FOG) and Magnetic Resonance Gyroscopes (MRG's) are of interest for navigation because of their potential for low cost and excellent reliability.

A Microprocessor-Based Data-Acquisition System for Stall/Spin Research

Abstract: Accurate data are necessary for the identification of aerodynamic parameters at high angle of attack and in spinning flight. Modern sensors and microelectronic devices are employed in a stall/spin data acquisition system that will be tested in a Schweizer 2-32 sailplane. This instrumentation package will use a quaternion-based ?strapdown IMU? algorithm to derive vehicle attitude from angular rate data, and it will use fault-detection logic to identify inflight sensor failures. Details of the system and flight test program are presented, together with results of preflight digital simulation analysis.

Data acquisition system and methodology for high angle of attack parameter estimation

Abstract: A digital data acquisition system has been flight tested in a Schweizer 2-32 sailplane and the algorithms, designed for high angle of attack parameter identification, are described in this paper. The system is portable and self-contained. Inertial sensors are mounted in a strapdown unit. Two Kiel-probes are used for airspeed measurement. "Off-the- shelf" microelectronics components are configured to operate as a multiprocessing system. The on-board computing power is used to monitor the sensor and support system and to improve flight test effectiveness. The Estimation-Before-Modeling technique has been selected for parameter identification. A time history of a flight test maneuver that will be used in the estimation of the aerodynamic parameters is also presented.

Shuttle Orbiter stellar-inertial reference system

Abstract: The Space Shuttle stellar-inertial reference system is a velocity and attitude data source during flight operations. An overview of the reference system is presented as well as specifics discussing design concepts, functional operation, and performance capabilities. Techniques for star sighting and inertial measurement unit alignment and calibration are described, and alignment accuracy, star tracker capability, and gyro and accelerometer accuracy are discussed, with emphasis on flight test results. Test programs have confirmed that the system meets performance requirements such as being accurate to within 0.26 degree at the 400,000 foot altitude entry interface in order to execute an accurate touchdown, as well as demonstrating reusability, payload capability, and operational flexibility. Growth possibilities, such as the implementation of rendezvous target tracking, are discussed.

Application of small-diameter inertial grade gyroscopes significantly reduces borehole position uncertainty

Abstract: Initial tests with a new directional survey tool show a significant enhancement in attainable accuracy over conventional instrumentation. Two prototype systems, developed over the last two years by Gyrodata, Inc., have recently been tested in a well in West Texas. Although many more tests are required, preliminary results indicate that the original design objective for borehole position uncertainty less than 1.7 feet per 1,000 feet of hole--has been met. The Gyrodata Wellbore Surveyor employs an inertial grade rate gyro adapted from the aerospace industry. In combination with its other sensors and electronics, the device can sense the orientation of the earth's spin vector at each independent survey station. As a result, the major systematic errors associated with conventional gyros--geographical reference and unaccountable drift--are eliminated. Other sources of inaccuracy are minimized by the system's measuring techniques and operational procedures, and additional benefits should arise from faster survey speed and increased reliability. A true north reference device can also employ a small outside diameter since it requires only one gyro and one accelerometer, rather than the two or three of each needed in an inertial navigation system.

A Wellbore Inertial Navigation System

Abstract: A prototype wireline tool which includes a downhole inertial platform and a surface computer to spatially map a well is described. The hardware consists of a single-gimballed inertial platform with accelerometers and gyros to obtain three-axis motion information. The gyroscope and accelerometer outputs are transmitted to a computer at the surface which calculates probe attitude relative to north, east, and vertical. Double integration of the accelerometer data provides the position information. A conventional 7-conductor wireline is used for the system data transmission. System accuracy is enhanced by advances made in the computer software which processes the data received from the tool. The software uses statistical sampling estimation to obtain optimal estimates of the system errors. Measurement errors are determined by periodically stopping the tool during the logging procedure and observing the indicated velocity measurements. This procedure, known as Kalman filtering, results in increased accuracy of the data. Present mapping systems have an X-Y-Z location accuracy of 100 to 200 feet for a typical well depth of 10,000 feet. Test results show that the new system is accurate to about 1 foot per 1000 feet of well depth. Unlike conventional systems, the inertial navigator does not require any sort of projectionmore » of the cable length (which may not be accurately known). Also, this system provides continuous data throughout the wellbore and logging speeds on the order of 10 ft/sec appear possible. The hardware and software associated with this mapping system are described and the recent field test results are reported.« less

Unified analysis methods for a fault tolerant redundant strapdown inertial measurement unit

Abstract: The use of a redundant strapdown inertial measurement unit (RSDIMU) has been considered as a source of sensor information for future aircraft. It is pointed out that failure detection and isolation (FDI) should be accomplished at the sensor level to ensure the accurate transfer of information to elements of the integrated avionics system. The present investigation is concerned with the impact of the FDI algorithm on system reliability and vice versa. A description is given of a technique to identify and assess critical areas of reliability requirements for the RSDIMU. Attention is given to the development of an FDI algorithm to meet these requirements.

Minuteman inertial guidance assessment - The next best thing to flight tests

Abstract: Techniques have been developed and implemented to enable the continuous assessment of Minuteman inertial guidance performance and in-flight reliability in a situation of relatively few flight tests planned for the remaining service lifetime of the weapon system. The Minuteman force consists of 1000 deployed missiles with guidance systems in full-time operation. Inertial Measurement Unit Performance Data are collected continuously for processing by the Minuteman Data Analysis System. Output data products provide observability of critical parameters measured during the calibration mode and during the normal strategic alert operating mode. Algorithms facilitate the prompt identification of performance anomalies as well as the ensemble statistical characterization of performance parameters. A simulated flight test is used to augment the in-flight reliability data base and to detect possible incipient degradation under the stresses representative of flight. The simulated flight environment is modeled from actual flight shock and vibration data. Systems are selected from the force for this non-destructive test on an unbiased sampling basis.

Development and test of an integrated sensory system for advanced aircraft

Abstract: This paper presents the results of the second phase of an integrated sensory subsystem (ISS) development effort. The ISS is a combination of redundant inertial sensors, air data probes, transducers, other flight control related sensors, interfaces, and the associated data handling system (DHS). The sensor data derived within the ISS meets the requirements (performance, redundancy, survivability, etc.) for digital fly-by-wire flight control systems. The key task of this development phase was the synthesis and development of a DHS for dispersed arrays of flight control inertial sensors subjected to dissimilar motions due to body bending of the aircraft structure. The system design is described in terms of the hardware and data handling system synthesis, followed by a discussion of the methods and results utilized to verify the system design.

Reliability analysis and fault-tolerant system development for a redundant strapdown inertial measurement unit

Abstract: A methodology is developed and applied for quantitatively analyzing the reliability of a dual, fail-operational redundant strapdown inertial measurement unit (RSDIMU). A Markov evaluation model is defined in terms of the operational states of the RSDIMU to predict system reliability. A 27 state model is defined based upon a candidate redundancy management system which can detect and isolate a spectrum of failure magnitudes. The results of parametric studies are presented which show the effect on reliability of the gyro failure rate, both the gyro and accelerometer failure rates together, false alarms, probability of failure detection, probability of failure isolation, and probability of damage effects and mission time. A technique is developed and evaluated for generating dynamic thresholds for detecting and isolating failures of the dual, separated IMU. Special emphasis is given to the detection of multiple, nonconcurrent failures. Digital simulation time histories are presented which show the thresholds obtained and their effectiveness in detecting and isolating sensor failures.

Calibration of the aerodynamic coefficient identification package measurements from the shuttle entry flights using inertial measurement unit data

Abstract: The Aerodynamic Coefficient Identification Package (ACIP) is an instrument consisting of body mounted linear accelerometers, rate gyros, and angular accelerometers for measuring the Space Shuttle vehicular dynamics. The high rate recorded data are utilized for postflight aerodynamic coefficient extraction studies. Although consistent with pre-mission accuracies specified by the manufacturer, the ACIP data were found to contain detectable levels of systematic error, primarily bias, as well as scale factor, static misalignment, and temperature dependent errors. This paper summarizes the technique whereby the systematic ACIP error sources were detected, identified, and calibrated with the use of recorded dynamic data from the low rate, highly accurate Inertial Measurement Units.