Showing papers on "Inertial reference unit published in 1998"

A low-cost GPS/inertial attitude heading reference system (AHRS) for general aviation applications

Abstract: An inexpensive Attitude Heading Reference System (AHRS) for general aviation applications is developed by fusing low cost ($20-$1000) automotive grade inertial sensors with GPS. The inertial sensor suit consists of three orthogonally mounted solid state rate gyros. GPS is used for attitude determination in a triple antenna ultra short baseline configuration. A complementary filter is used to combine the information from the inertial sensors with the attitude information derived from GPS. The inertial sensors provide attitude information at a sufficiently high bandwidth to drive an inexpensive glass-cockpit type display for pilot-in-the-loop control. The low bandwidth GPS attitude is used to calibrate the rate gyro biases on-line. Data collected during laboratory testing is used to construct error models for the inertial sensors. Analysis based on these models shows that the system can coast through momentary GPS outages lasting 2 minutes with attitude errors less than 6 degrees. Actual performance observed during ground and flight tests with GPS off was found to be substantially better than that predicted by manufacturer supplied specification sheets. Based on this, it is concluded that off-line calibration combined with GPS based in-flight calibration can dramatically improve the performance of inexpensive automotive grade inertial sensors. Data collected from flight tests indicate that some of the automotive grade inertial sensors (180 deg/hr) can perform near the low end of tactical grade (10 deg/hr) sensors for short periods of time after being calibrated on-line by GPS.

The role of dead reckoning and inertial sensors in future general aviation navigation

Abstract: Possible configurations for a general aviation autonomous navigation system are studied. Doubtless, an advanced GPS receiver is a "must have" system component. GPS has had some outages due to unintentional interference or even intentional jamming and aircraft should be able to navigate through such an event. Natural candidates for GPS backup are inertial sensors, magnetic compass, and air-speed sensors. All these sensors can be calibrated during GPS availability. Moreover, for dead reckoning systems, wind velocity can be estimated as well. This paper presents an original statistical model for wind variations that matches actual data very well. Using this model and a parametric family of inertial measurement sensors, horizontal position errors during a GPS outage are compared for a variety of configurations: a dead reckoning system, stand alone inertial sensors and inertial sensors integrated with the dead reckoning system.

F-16 flight tests of a rapid transfer alignment procedure

Abstract: This paper presents the results of an effort directed at developing and flight-testing an innovative rapid transfer alignment algorithm for inertially-guided air-launched munitions. The algorithm, referred to as RAP (Rapid Alignment Prototype), employs a 17-state Kalman filter designed to accurately align a weapon-grade Inertial Measurement Unit (IMU) relative to an aircraft-grade Inertial Navigation System (INS) within five seconds. The alignment procedure requires the pilot to execute only a brief wing-rock maneuver. No time-consuming heading changes or lengthy s-turns are required. The RAP Kalman filter achieves the rapid convergence time by recursively processing both velocity-match and attitude-match measurements at a 12.5 Hz rate to estimate and correct IMU velocity, attitude, and inertial sensor errors. Following laboratory and van testing at Eglin AFB, a series of F-16 flight tests were conducted. Flight test results demonstrated that the RAP filter achieved sub-milliradian alignment accuracy in less than 10 seconds. As further confirmation of alignment accuracy, IMU position error statistics were computed over a 100-sec post-alignment captive-carry trajectory. Test results indicated that the mean radial position error after 100-sec of unaided navigation was roughly 70 ft with an associated CEP of 61 ft. RAP's unprecedented alignment accuracy and reduced launch timeline provide a rapid-response capability for time-critical targets such as mobile launchers and troop emplacements.

Constrained navigation algorithms for strapdown inertial navigation systems with reduced set of sensors

Abstract: This paper develops a family of algorithms for low-cost strapdown inertial navigation system for land vehicles. Constraints on the motion of land vehicles are defined. They include constraints on vehicle's orientation relative to the Earth surface, and relationship between vehicle's attitude and its velocity direction. Navigation equations are derived that assume validity of these constraints on the vehicle's motion. Compared to standard strapdown inertial navigation, these algorithms reduce navigation errors in the presence of relatively high instrument noise, and at the same reduce number of required inertial sensors. Navigation errors are analyzed and used in an error model for the Kalman filter. Derived algorithms are applied to processing of experimental data.

Strapdown inertial navigation system. Part 1 – Navigation equations

Abstract: Almost all aircraft are equipped with the Inertial Navigation Systems. The autonomous Inertial Systems are capable of calculating the navigational parameters of aircraft: position, velocity, and attitude, without external sources of information. In the paper the equations and algorithms of strapdown navigation system are derived and analysed. The research is focused on the calculation methods of the aircraft attitude.

GPS analytic redundancy for gyroscope failure detection

Abstract: An inertial measurement system for determining the attitude and rate of change of attitude of a vehicle such as an aircraft is disclosed. Redundancy of the inertial measurement system may be accomplished by utilizing a global positioning system having three receivers and a set of antennas strategically disposed on the vehicle. The global positioning system having receivers may be utilized for accomplished failure detection and isolation of the inertial measurement system rather than providing a redundant inertial measurement system. Failure detection of the IMUs may be accomplished with only dual antenna/receiver configuration.

Vital Advanced Inertial Network

Abstract: Many aircraft sensors and weapons require accurate, local, time-varying inertial information relative to a principal reference, usually the main inertial navigation system. This data may be required during severe manoeuvres and despite other structural deflections due to aircraft aging, changing stores configurations, and fuel burnoff. Past solutions to this problem have included co-location, structural stiffening, or transfer alignment using high cost navigation grade inertial sensors. The Vital Advanced Inertial Network (VAIN) program sponsored by the Reference Systems Branch of the Wright Laboratory Avionics Directorate, is developing and testing an advanced inertial network to support precision targeting and weapon delivery under the influence of structural bending. This paper addresses some of the VAIN work and describes the sensor requirements for and performance of a transfer alignment based algorithm for compensating for aircraft flexure using low cost inertial sensors.

Surveying system with an inertial measuring device

Abstract: A surveying system has a sighting device and a position measuring unit. The position measuring unit includes an inertial measuring device.

System for accurately determining missile vertical velocity and altitude

Abstract: A system (30) for determining a missile kinematic property. The inventive system includes a radar system (36) for obtaining radar range measurements (34). An inertial reference unit (40) tracks changes in missile acceleration without the use of the radar range measurements and provides a signal (38, 42) in response thereto. An altitude Kalman filter (32) and a subtractor circuit (48) are used to combine the signal (38, 42) and the radar range measurements (34) to provide an accurate estimate (50) of the property. In a specific embodiment, the kinematic property includes missile vertical velocity and missile altitude. A radar system (36) supplies radar range information (34) to the Kalman filter (32). A loop (52) from the Kalman filter (32) to the radar system (36) facilitates estimating the radar range measurements (34) when invalid radar range measurements are provided by the radar system (36). The Kalman filter (34) is a two state filter, one state corresponding to missile altitude, and the other state corresponding to missile vertical velocity bias error in the inertial reference unit's (40) velocity measurements (42). In the illustrative embodiment, the Kalman filter (32) combines radar pseudo-measurement estimates of missile altitude with estimates (38) of missile altitude obtained from the inertial reference unit (40) and provides an output signal (46) representative of an accurate missile altitude estimate in response thereto. The Kalman filter (32) includes a velocity bias output (44) which provides an estimate of the error in a velocity measurement (42) obtained from inertial reference unit (40) to the subtractor circuit (48). The subtractor circuit (48) subtracts the estimate (44) of the bias error from the velocity measurement (42) and provides the missile velocity estimate (50) in response thereto.

Development of a FOG-based GPS/INS

Abstract: Honeywell and its teammates, Allied Signal and Trimble, have developed a design for a compact (100 in/sup 3/) integrated Global Positioning System/Inertial Navigation System (GPS/INS) under the GPS Guidance Package (GGP) Program funded by DARPA. This effort included developing new inertial sensor technology and implementing inertial aiding of the GPS receiver's tracking loops. The inertial sensors consist of fiber-optic gyroscopes and solid-state silicon accelerometers suitable for a 1-nmi/hr-class navigation system. Inertial aiding of the GPS receiver's carrier tracking loop was used to enhance tracking under severe dynamics without compromising anti-jam performance. This paper reviews the system architecture, provides a brief description of the key subassemblies and describes the test results from a prototype unit. These prototype units demonstrated the capabilities of the overall system, including integrated GPS/INS navigation and enhanced performance of the GPS receiver as a result of carrier loop aiding.

Development of a low-cost integrated GPS/IMU system

Abstract: This paper describes the development of a low-cost and small size integrated Global Positioning System (GPS)/inertial measurement unit (IMU). The developed strapdown Attitude and Heading Reference System (AHRS) is capable of providing attitude and heading accuracy at /spl plusmn/0.2 and /spl plusmn/0.4 degrees, respectively. The prototype of low-cost integrated GPS/IMU can give positioning accuracy of 10 metres.

Precision Spacecraft Attitude Estimators Using an Optical Payload Pointing System

Abstract: The advancement of satellite program technology mandates that issues affecting future spacecraft system development cost be resolved, while continuing to meet smaller size, lower power, and more stringent mission requirements. In the specie c area of the spacecraft attitude determination system, an optical payload pointing system can servea dual purpose: to support the payload mission and to providepreciseattitude information for the spacecraft functions. As a result, star trackers orotherprecisereferencesystems can beeliminated to reduce the development cost. Additional benee ts include performance robustness to spacecraft motion or disturbance and potential cost saving due to weight reduction, e.g., launch cost. Attitude estimators using an optical payload pointing system and strap-down gyros are derived in detail for two distinct formulations: spacecraft body and inertial formulations. Alternative Kalman e lter equation derivations are shown using the skew symmetric matrix properties of the attitude kinematics equation. It is also shown that both formulations produce identical solutions. However, the body formulation is conceptually easier to understand, and the inertial formulation requires less processing time. To show thespacecraft attitudeestimatorperformanceusing an optical payload pointing system, an exampleisshown for an agile low-Earth-orbit satellite.

The expected performance of Gravity Probe B electrically suspended gyroscopes as differential accelerometers

Abstract: Four cryogenic gyroscopes on the Gravity Probe B satellite will be used to measure the precession of the local inertial reference frame with respect to a distant inertial reference frame. One of these four gyroscopes will serve as the drag-free sensor for the satellite. The other three gyroscopes, which are separated from each other by 8.25 cm, will be electrostatically supported by a digital control system. Although the gyroscopes and the electrostatic suspension system are designed to measure a precession as small as 0.1 mas/yr, any pair of these gyroscopes may also be used as a differential accelerometer. This paper analyzes the expected performance of these gyroscopes as differential accelerometers for accelerations in the frequency band from 2×10−3 to 2×10−2 Hz. The three contributions to the specific force on any one of the gyroscopes are the residual acceleration of the spacecraft, the specific forces acting between the gyroscope and the satellite, and the noise in the capacitance bridge which sens...

On the INS and GPS Data Processing for Airborne Gravity Applications

Abstract: Airborne gravimetric and altimetric measurements depend heavily on the determination of the instantaneous position and attitude of the aircraft. Here, a methodology to process raw inertial data together with GPS data collected from at least two antennas/receivers in the aircraft is presented. This methodology is primarily aimed at the attitude determination, but has potential to be applied in precise position determination as well. The main idea behind the methodology is to process GPS carrier phase measurements in differential mode between a pair of aircraft antennas to obtain precise heading angles. These, together with doppler velocity readings and approximate position estimates, serves to maintain stability of the attitude angles obtained from filtering raw inertial measurements. On the other hand, the availability of precise attitude estimates from the inertial sensors allows keeping phase integer ambiguities fixed through cycle slips and satellite configuration changes. As a side effect, it is also possible to evaluate the quality of each phase measurement in order to exclude readings corrupted by noise or multipath effect. This is important for absolute position determination where the aircraft GPS measurements are processed against a reference station, namely when long baselines are used. The integration of displacement estimates computed from inertial data is potentially helpful to solve and keep integer phase ambiguities between the aircraft and the reference station. This methodology is being employed to process data from the Azores campaign of the AGMASCO project.

Testing Inertial Sensors

Abstract: In Chapter 2 we described inertial systems error modeling and the generation of specifications for the sensors in an inertial navigator, specifications which could (and should) follow the Institution of Electrical and Electronic Engineers (IEEE) standard formats for inertial sensors. In this chapter we will describe the testing of accelerometers and gyros for conformance to these specifications; we will describe the tests that we would perform for acceptance test procedures (ATPs) on a pendulous accelerometer and on three types of gyro, the single-degree-of-freedom (SDFG), the dynamically tuned gyro (DTG), and the ring laser gyro (RLG). This description is general and takes account of neither specific design issues nor of particular customer needs that would expand the ATP in real life.

MAPS goes underground

Abstract: The Honeywell Ore Recovery/Tunneling Aid (HORTA) is a precision inertial positioning and pointing system for use on mining equipment. The HORTA is based on the Honeywell Modular Azimuth Positioning System (MAPS) that is used in military applications. The HORTA is a completely autonomous self-contained Dynamic Reference Unit (DRU) inertial navigator mechanized using strapdown inertial algorithms, three ring laser gyroscopes (RLGs) for angular motion sensing, three Q-Flex accelerometers for translation measurements and special software for mining applications. The HORTA is used to determine the position of a continuous miner (CM) from remote controlled equipment. This allows the mining industry the ability to keep the workers at safe distances during the mining process and to have an accurate, reliable navigation system providing heading and location of the ore extraction equipment.