Showing papers on "Inertial reference unit published in 2002"

The Microwave Anisotropy Probe (MAP) Mission

Abstract: The Microwave Anisotropy Probe mission is designed to produce a map of the cosmic microwave background radiation over the entire celestial sphere by executing a fast spin and a slow precession of its spin axis about the Sun line to obtain a highly interconnected set of measurements. The spacecraft attitude is sensed and controlled using an inertial reference unit, two star trackers, a digital sun sensor, twelve coarse sun sensors, three reaction wheel assemblies, and a propulsion system. This paper presents an overview of the design of the attitude control system to carry out this mission and presents some early flight experience.

A wearable attitude-measurement system using a fiberoptic gyroscope

Abstract: An attitude-measurement system (TISS-5-40) has been developed to achieve a wearable sensor for individuals. This equipment is one of the inertial sensor systems having three fiberoptic gyroscopes and three accelerometers. Heading stability of 1 deg./hr. (1 σ) and attitude accuracy of ±0.5 deg. have been demonstrated. At present, some of the attitude-measurement Systems have been applied in the field of mixed-reality technology, and the users confirm and report its effectiveness (Hara, Anabuki, Satoh, Yamamoto, & Tamura, 2000).

Geophysical inertial navigation system

Abstract: An inertial navigation system employs geophysical (earth's magnetic and gravity) field sensors to measure body rotations of an aircraft or other vehicle. In effect, these replace expensive laser ring gyros used in conventional inertial navigation systems. The vector components of the geomagnetic field may be incorporated directly into the state vector describing the vehicle's attitude, acceleration, velocity, and position. A predictive filter (e.g. Kalman) operates on measured and predicted geophysical field components and on-board accelerometers. Predicted geophysical field components are obtained from maps and/or models of the earth's magnetic and gravity fields.

Method to calculate sideslip angle and correct static pressure for sideslip effects using inertial information

Abstract: A method of the invention allows the determination of aircraft sideslip using an air data probe and an inertial reference unit. A lateral sideslip component β L of the aircraft is calculated as a function of inertial flight information. An angular sideslip component β A of the aircraft is also calculated as a function of inertial flight information. The lateral sideslip component β L and the angular sideslip component β A are combined to obtain a total sideslip angle β TOTAL for the aircraft. The total sideslip angle β TOTAL can be used to compensate static pressure, angle of attack and other aircraft parameters for sideslip effects.

Miniature attitude sensing suite

Abstract: The present invention provides methods of and apparatus for determining the inertial attitude of an aerospace vehicle. In one embodiment, the invention provides a rotational astronomical object-sighting concept to determine the inertial attitude of an axis of the aerospace vehicle without the star identification or dragback. In another embodiment, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor.

Noise estimation for star tracker calibration and enhanced precision attitude determination

Abstract: This paper presents the design, development, and validation of a nonlinear least square estimation scheme applied to star tracker noise extraction and identification. The paper is the by-product of a Post-Launch Test (PLT) tool development effort conducted by two independent teams, Swales/NASA and Boeing. The main objective is to have a set of tools ready to provide on-orbit support to the GOES N-Q Program. GOES N-Q employs a stellar inertial attitude determination (SIAD) system that achieves high precision attitude estimation by processing attitude and rate data provided by multiple star trackers (ST) and an inertial reference unit (IRU), respectively. The key component of SIAD is the ST. The ST's star position vector is corrupted by three major noise sources: temporal noise (TN), high spatial frequency noise (HSF), and low spatial frequency (LSF) noise. The last two noise sources are not while and correlated. As a result, the performance of the SIAD filter is no longer optimal, causing the reconstructed attitude knowledge to potentially satisfy requirements with a narrow margin. This tight margin is critical and may affect the GOES N-Q mission, particularly the Image Navigation and Registration (INR) system performance. The PLT toolset is expected to provide the capability to mitigate this potential problem during PLT time.

Fault tolerant attitude control system for zero momentum spacecraft

Abstract: A fault tolerant attitude control system for a zero momentum spacecraft. A zero momentum attitude control system is operable to control spacecraft attitude utilizing data received from an earth sensor, a sun sensor, and the inertial measurement unit. A gyroless attitude control system is operable to control spacecraft attitude without receiving data from the inertial measurement unit. A redundancy management system is operable to monitor an inertial measurement unit to detect faults and to reconfigure the inertial measurement unit if a fault is detected and operable to determine when an inertial measurement unit fault is resolved. A controller is operable to automatically switch the spacecraft from the zero momentum attitude control to the gyroless attitude control when a fault in the inertial measurement unit is detected and is operable to automatically switch the spacecraft from the gyroless attitude control to the zero momentum control upon resolution of the fault.

A merging of system technologies: all-accelerometer inertial navigation and gravity gradiometry

Abstract: Significant research has been conducted in recent decades on developing a practical all-accelerometer inertial navigation system and on developing a practical moving-base gravity gradiometer. The former strives to measure kinematic motion in the presence of unwanted variations in gravity; the latter endeavors to measure variations in gravity in the presence of unwanted motion. Thus, goes the adage, one person's signal is another person's noise. In this paper, a mathematical relationship is derived that links the two concepts together. It is demonstrated that a complete solution to real-time navigation and gravity gradient determination can be performed simultaneously and unambiguously using all-accelerometer inertial measurements only. Although the separation of gravity from kinematic motion appears to violate Einstein's principle of equivalence, this is not the case. Nonetheless, accelerometer technology is not yet of sufficient maturity to allow a practical implementation of this concept, but advances certainly are being made that could result in a practical implementation perhaps within a decade.

Testing of the attitude and heading reference system

Abstract: Rzeszow University of Technology has undertaken the task of designing and providing of equipment to a flying laboratory. This paper presents basic design principles of the Inertial Reference Unit (IRU) which employs measuring signals from the Fiber Optic Gyros (FOG), accelerometers and electronic compass module. A microcomputer follows the algorithm of complementary filtration for of calculating the Euler angles for the aircraft attitude (pitch, roll and heading), angular rates, and linear accelerations. The correction systems that minimize error of the steady‐state measuring have been employed. The results of computer simulations, lab tests and selected flight tests have also been presented. The Inertial Reference Unit μIRU‐1 was tested in flight on board of the general aviation aircraft PZL‐110 “Koliber”. It has been confirmed that metrological properties of the system are appropriate for the purposes of teaching process. Currently, a modified version of the unit is being prepared. The new IRU is planned as a main reference unit for integrated flight control system of general aviation aircraft.

Integrated inertial vms navigation with inertial odometer correction

Abstract: A system for determining accurate inertial navigation data of a vehicle (1) while moving must compensate for the orientation differences between an upper body and a lower body of the vehicle (1). These orientation differences are used to increase the accuracy of the aided inertial navigation solution. An odometer on the lower body is used to reduce errors in the inertial date arising from the movement between the upper and low bodies and errors in the odometer associated with vehicle pitch are reduced by sensing pitch changes with the inertial navigation data.

Proposed IEEE Coriolis Vibratory Gyro standard and other inertial sensor standards

Abstract: A new Coriolis Vibratory Gyro (CVG) Specification Format Guide and Test Procedure is being developed by the IEEE/AESS Gyro and Accelerometer Panel for consideration by the IEEE Standards Association as IEEE Std. 1431. It provides guides for specifying CVG performance and gives test procedures for verifying compliance with requirements. It also describes the various types of CVG, ranging from vibrating shell macro-sized devices to low-cost silicon and quartz chip micro-sized devices. Model equations are given for the various CVG operating and readout modes: whole-angle, open-loop, force-rebalance, and ratiometric. Also discussed in this paper are the IEEE standards that have been published for single- and two-degree-of-freedom spinning wheel gyros, laser gyros, interferometric fiber optic gyros, angular accelerometers, linear accelerometers, accelerometer centrifuge testing, and inertial sensor terminology, and an inertial sensor test equipment and analysis document and an inertial systems terminology document that are under development.

Optimum technology for non-gyro micro inertial measuring unit

Abstract: An inertial navigational system with high anti g capability, low power consumption, small volume and low cost is needed on certain specific occasions. By using linear accelerometers unit at the different positions in space, a non-gyro micro inertial measuring unit is used to calculate the angular velocity and sense the axial acceleration of the carrier. A feasible inertial navigational system with medium accuracy can be designed in this way. The nine accelerometers were improved based on researched by foreign scholars and a novel calculation method was proposed to control the iterative error. The redundant information of sensors was used directly to calculate the absolute value of measured angular velocity and correct the solutions of differential equations thereby diminishing the accumulative errors for the solution of differential equations and reducing the shift of an inertial measurement unit.

Inertial measurement system

Abstract: An apparatus for accurately measuring inertial measurements for use in an inertial measurement system includes a high range sensor, a low range sensor, and an apparatus for combining data from the high range sensor and the low range sensor to provide optimal accurate inertial measurements.

Satellite TV Antenna Attitude Stabilization System Based on Micromachined Inertial Sensors

Abstract: This paper describes a satellite TV antenna attitude stabilization system based on micromachined inertial sensors. The composition, principle and specification are introduced, and the key parts of the system are discussed, including micro-machined inertial attitude measurement unit, servo control of the stepping motor and RF automatic tracking. The prototype experiment shows that the system has reached the expected specification and meets the requirement of its application on boats. Since the micromachined devices are used, the system has the advantages of small size, low cost, and high reliability.

Analysis of compound altitude channel of A340 air data and inertial reference system

Abstract: Inertial reference system is a navigation system with high precision and independent characteristic and with the capability of working on any climate and the entire world.It is also the system,which provides the most output of navigation information.It is usually called as a center information resource.Air Data System(ADS)is integrated in Inertial Reference System(IRS)as an integrated navigation system.ADIRS makes altitude channel become perfect.Most of Boeing and Airbus aircraft are equipped with ADIRS.So it is the trend of the development of avionics that improving the function and precision of ADIRS.

Observations on one-dimensional counterflow diffusion problem

Abstract: The proposed paper will describe development of a time-dependent, one-dimensional counterflow diffusion model meant to help study the issue of helium exposure of the IRU (Inertial Reference Unit) used in the Earth Observing-1 (EO-1) spacecraft. The IRU features sensitive quartz crystal hemispherical resonating gyros (HRG's). Although the IRU enclosure was purged with high-purity liquid boil-off nitrogen, the HRG's were still quite susceptible to contamination generated by exposure of the enclosure to levels of helium above atmospheric background levels. This helium would be preferentially collected by the HRG's, changing their mass, and hence the driving voltage required for operations. The paper will discuss a comparison of theoretical results with test data for an IRU enclosure, and how contaminant gases can enter vent holes despite the presence of a purge. These observations are then used to describe a possible improvement for purge effectiveness.

Strapdown inertial navigation system on laser gyros in the attitude definition mode

Abstract: An errors of strapdown inertial navigation system (SINS) working in the attitude definition mode and intended for use in a structure of a magnetic measuring complex are considered. An algorithm and block diagram of the inertial- magnetic system on the basis of Kalman filtering method and the geomagnetic field components definition method are developed. Features of the measuring operational mode and algorithms of SINS working in a structure of a magnetic measuring complex are cited in this paper. A results of research and mathematical modeling both algorithms and errors of SINS are shown.