Ferranti Laser Gyro Flies

In modern aerospace technology, strapdown inertial navigation systems (SINS) are widely used, using fiber-optic (FOG) or ring laser (CLG) gyroscopes. During the operation of such systems, the sensitivity axes are rotated relative to the basic coordinate system. The resulting angles between the axes of the base coordinate system and the axes of sensitivity of the navigation system (non-orthogonality) are one of the factors leading to an increase in the measurement errors of the device, which affects the measurement accuracy. During operation, the system is affected by vibrations of various nature, the impact of which can contribute to the appearance of non-orthogonality. The purpose of this work is to determine the maximum permissible vibration amplitudes affecting the SINS body according to the permissible values ​​of the deviation of the FOG sensitivity axes for two variants of the SINS layout. An approach to determining the permissible amplitudes of an external harmonic impact on the unit of a strapdown inertial navigation system based on fiber-optic or ring laser gyroscopes is considered. A design scheme, mathematical and finite element models for calculating natural frequencies and forced oscillations of a strapdown inertial navigation system unit have been developed. In various frequency ranges, numerical calculations have determined the boundary values ​​of the amplitudes of the external harmonic impact on the base of specific configurations of the SINS assembly. It has been established that dangerous states take place in the region of the 1st natural frequency of the system, as well as near higher frequencies. Comparison of the results for design options 1 and 2 allows us to conclude that in order to weaken the effect of vibrations on the accuracy of the SINS unit, it is advisable that the lowest natural vibration frequencies for the SINS assembly be as high as possible (more than 1000 Hz).
 Key words: vibration; fiber optic gyroscope; strapdown inertial navigation system; finite element method; natural frequencies and modes of vibration.

The ring laser gyro (RLG) dither axis will bend and exhibit errors due to the specific forces acting on the instrument, which are known as g-sensitive misalignments of the gyros. The g-sensitive misalignments of the RLG triad will cause severe attitude error in vibration or maneuver environments where large-amplitude specific forces and angular rates coexist. However, g-sensitive misalignments are usually ignored when calibrating the strapdown inertial navigation system (SINS). This paper proposes a novel method to calibrate the g-sensitive misalignments of an RLG triad in linear vibration environments. With the SINS is attached to a linear vibration bench through outer rubber dampers, rocking of the SINS can occur when the linear vibration is performed on the SINS. Therefore, linear vibration environments can be created to simulate the harsh environment during aircraft flight. By analyzing the mathematical model of g-sensitive misalignments, the relationship between attitude errors and specific forces as well as angular rates is established, whereby a calibration scheme with approximately optimal observations is designed. Vibration experiments are conducted to calibrate g-sensitive misalignments of the RLG triad. Vibration tests also show that SINS velocity error decreases significantly after g-sensitive misalignments compensation.

An idea of using computer mechanics for inertial navigation systems is given and examples are presented. The equations, algorithms, and properties of the pendulum-type strapdown inertial navigation system are analyzed. As a result, it has been stated that such a system is analogous to analytical systems. A similar comparison for both semi-analytical and strapdown inertial navigation systems, in which the Schuler’s pendulum models are described in a horizontal coordinate system, is carried out. An analogy of their properties is established. By comparing the analytical system, platform axes of which are directed along the axes of the inertial coordinate system (orientation of the Schuler’s pendulum is also described in the inertial coordinate system), and the strapdown inertial navigation system with the same orientation of the platform’s computer model (the comparison is also made for the operation algorithms of such systems), an analogy of such systems has been established. The degree of use of computer mechanics in all types of strapdown inertial navigation systems is much greater than in platform ones. According to the degree of utilization for principles of computer mechanics, types of inertial navigation systems can be arranged in the following order: strapdown pendulum, other strapdown systems, semi-analytical, analytical, and geometric platform inertial navigation systems. Their accuracy depends on the degree of sophistication of the element base, that is, on sensitive elements and on-board computers. We claim that the smaller the volume and mass of the mechanical part of the system the better its weight-and-dimensional characteristics and cost.

A correcting digital filter for a mechanically dithered single-axis ring laser gyro (RLG) is proposed. The filter is intended to determine the angular position of a strapdown inertial navigation system (SINS) in the RLG sensitivity axis using the known value of the RLG angular position with respect to inertial space, taking into account the models of the suspension elastic forces nonlinearity, the hysteresis model of vibration energy dissipation, and the model of the suspension piezoceramic vibratory drive. Parameterized models of nonlinearity, dissipation, and the output characteristic of the suspension piezoceramic drive have been developed. A method for finding nonlinearity parameters of elastic forces using a set of experimental amplitudefrequency characteristics (AFC) is proposed. Numerical simulations of the AFCs of an elliptic eliminator and a correcting filter under sinusoidal rotation of the SINS with the amplitude of 1 arcmin in the frequency band of 80–3600 Hz have been performed. The efficiency of the proposed filter has been verified experimentally under pulse action on the SINS; the test results are discussed. Comparison of the experimental and calculated data shows the adequacy of the proposed models and satisfactory operation of the correcting digital filter.

In order to solve the strapdown inertial navigation system (INS) alignment problem under the marine mooring condition, the rapid strapdown INS fine alignment method is proposed. This method uses the gravity in the inertial frame to deal with the lineal and angular disturbances. Also the forward and backward processes for strapdown INS calculation and filter estimation are designed. Making use of the forward and backward processes to repeatedly process the saved inertial measurement unit (IMU) data sequence could quickly obtain the initial strapdown attitude matrix. Simulation and trial test validate the performance of the proposed rapid fine strapdown INS alignment.

Strapdown inertial navigation systems (SINS) are widely used in modern air transport. The service life of such systems is decades. During this time, the characteristics of such systems may deteriorate due to the degradation of inertial sensors - gyroscopes and accelerometers. Under these conditions, in order to maintain the required accuracy of the SINS during the entire period of operation, it is necessary either to periodically carry out routine maintenance with the onboard system, or to continuously adapt the used sensor measurement model in terms of compensating for errors that arise. The second direction, which we will call calibration, has the advantage over the first one that it is carried out automatically by the system itself and does not require any additional equipment or additional maintenance work. The known method of SINS recalibration is based on the integration of information from the SINS and from the receiver of satellite radio navigation signals (SRNS). Such calibration, firstly, significantly depends on the availability and quality of satellite information, and secondly, it is possible only for integrated navigation systems in which the inertial subsystem is supplemented by a SRNS receiver. In contrast to the known method, this article solves the problem of autonomous SINS calibration, which is carried out after each flight. The solution is based on a linearized inertial navigation error model based on the boundary conditions of the state vector (at the beginning and end of the flight). The external information used consists only of the initial values of the coordinates and the current measurements of the baroaltimeter. As a result of the decision at the end of the flight, the errors of gyroscopes and accelerometers are determined, as well as the error in determining the final values of latitude and longitude in the SINS. Sensor errors can be used in subsequent system activations, which improves the accuracy of its operation as a whole. The analysis of the suitability and effectiveness of the developed methodology was carried out on a typical flight simulation program.

This paper describes the theoretical and practical stages through development to testing of an integrated navigation system, specifically composed of an Inertial Navigation System (INS), and Global Positioning System (GPS). Among various INS/GPS integration strategies, our aim is to construct a mathematical model of INS and GPS systems. A Kalman filter is designed and implemented depending upon these models. Besides these, the performance of the developed system is evaluated with real data recorded by the sensors. A comparison with a reference system and also with a loosely coupled system is done to show the Superiority of the tightly coupled structure. The strap down inertial navigation system (SINS) uses (ENU) frame and the integrated system uses position and velocity as measurements. The system model of the integrated system for Kalman filtering are derived and modeled as 27-states. These states estimation system showed clearly the application of fundamental modeling and filtering techniques. The simulation is built on the integrated system INS/GPS/GNSS and the trajectory generator data for flight simulation to be used as a base line algorithm contributing for design, analysis and develop a guidance and control strategy which compensates for the effects of the uncertainties accompanying the navigation systems produced by INS alignment.

Navigation of Hypersonic vehicles in the radio frequency (RF) blackout area during atmospheric reentry is challenging as the vehicles can only use the inertial navigation system (INS) as autonomous navigation method in this area. In this paper, strapdown inertial navigation system (SINS) based on the Fiber Optic Gyroscope (FOG) is used for navigation in blackout area. However, without external navigation measurement, the errors of SINS caused by the FOG drift and accelerometer bias would cumulate with time and degrade navigation accuracy. To solve this problem, single axis rotation modulation along with the azimuth axis of the body frame is adopted. The Generic Hypersonic Vehicle (GHV) model designed by NASA Langley Research Center is used to build the reentry fight model which can generate navigation information for simulation. Through derivation the error equations of FOG SINS in the North-East-Down (NED) navigation frame, the principle of error compensation by rotation modulation can be well understood. The simulation results show that rotation modulation can effectively decrease the impact of inertial sensor drift and improve the navigation accuracy in blackout area.

Navigation parameters such as vehicle heading, linear speed and latitude and longitude are determined today, as a rule,using satellite navigation systems. However, there is a need for autonomous determination of navigation parameterswhen satellite navigation system signals are unavailable for one reason or another. For example, satellite navigationsystem signals can be jammed by electromagnetic countermeasures.Autonomous systems for determining the vehicle navigation parameters are inertial navigation systems, the basis ofwhich is the double integration of signals from accelerometers, as well as the integration of signals from gyroscopes toform one or another coordinate system, if we are talking about strapdown inertial navigation systems (SINS).Integration of the output signals of accelerometers and gyroscopes, which contain errors, leads to the accumulation ofSINS errors. To increase accuracy, SINS are integrated, for example, with satellite navigation systems. However, suchsystems cease to be completely autonomous.The article uses the method of autonomous determination of vehicle latitude and longitude, which is a new method ininertial navigation. The method is based on the strapdown inertial technology, which, unlike the standard methods usedin SINS, does not use the integration of the output signals of accelerometers and gyroscopes. To determine the angles oflatitude and longitude, the signals of the gyroscopes included in the inertial measurement unite (IMU) are used.As a result of a semi-natural experiment, the efficiency of the method of autonomous determination of latitude andlongitude was verified. Flight data of a small-sized aircraft were used for experimental verification. The data containedinformation about the angles of the heading, pitch, roll of the aircraft, the output signals of three micromechanicalgyroscopes, three micromechanical accelerometers, information about latitude, longitude and altitude measured by thereceiver of the satellite navigation system, as well as information about the horizontal and vertical components of speedand current time. Plots of calculated values of latitude and longitude and values measured using the receiver of thesatellite navigation system are given. From the given plots and tables, it can be seen that from a certain time, thecalculated values coincide with the "reference" ones, which confirms the efficiency of the method of autonomousdetermination of latitude and longitude.

The basic operation principles of the strapdown inertial navigation systems (SINS) without platform and models of the measuring sensors errors, including offsets of zero signals, errors of scale factors and cross-links connections, are considered. The influence of various components of these errors and their linear combinations on the numeration accuracy of coordinates, velocity and orientation angles is investigated by the method of mathematical modeling of the SINS operation at one and a half hour intervals. To assess the linear properties of SINS algorithms, a quantitative measure is introduced — the value of the numeration error deviation, caused by a linear combination of input errors, relative linear combination of errors caused by separate input errors, as a percentage is expressed. Numerical estimates for typical trajectories of a maneuverable aircraft are obtained for this value. The obtained results can serve as the basis for the formation of technical requirements, used at the design and testing of aviation navigation systems and complexes. Keywords strapdown inertial navigation system, potential of the gravity field, gravitational acceleration, apparent acceleration, numeration errors, math modeling

With the rapid development of computer technology and communication technology, pedestrian navigation system(PNS) began to enter the real practical stage. PNS is a method of track reckoning based on strapdown inertial navigation system. However, inertial navigation system is prone to error accumulation with time, resulting in low positioning accuracy and practicability. Low cost inertial measurement unit(IMU) device has low accuracy, and the inertial navigation system has accumulated error, and the system is easy to diverge. Apart from those above, civil GPS accuracy is not stable, and poor GPS signal in a multipath environment may lead to errors, and GPS signals can not be used indoors. It is necessary to design a kind of navigation method to adapt low cost IMU. In this paper, it is presented that a foot inertial navigation system which is based on low cost MEMS inertial sensors is presented. According to the characteristics of pedestrian zero-velocity during walking steady state, an inertial navigation method based on Zero Velocity Update (ZUPT) is adopted. This paper uses air pressure, GPS to assist ZUPT method to correcting navigation parameters through Kalman filter. Based on the implementation scheme of pedestrian navigation system proposed in this paper, the steady state of walking were experimentally verified through the actual test data. The experimental results show that the proposed method can provide efficient, reliable and continuous position and navigation services.

Navigation equipment of a road surface testing vehicle (car) includes a strapdown inertial navigation system (INS), a global navigations satellite system (GNSS) receiver and an odometer. In this work we analyse a low-cost INS for using in the vehicle navigation equipment. An INS/GNSS/odometer integration algorithm was developed and applied to postprocessing data collected from a typical testing survey. The computed integrated navigation solution was compared with a reference solution provided by a NovAtel navigation system. The reference navigation system consisted of a tactic-grade strapdown INS, an odometer and two high precision GNSS receivers operating in the differential mode. Numerical results presented in the paper show that the low-cost INS/GNSS/odometer integration allows to determine the vehicle position in an urban area with an accuracy of 1 m and the roll and pitch angles with an accuracy of 0.1° that satisfy accuracy requirements.

The US Army has a requirement for a highly accurate inertial/TACAN hybrid navigation system. This paper describes the navigation system architecture developed to meet this requirement, the specific hardware selected and the development status. At the beginning of the Integrated Inertial Navigation System (IINS) development program it was anticipated that significant technological advances would take place in its life cycle. In addition, mission requirements were expected to change as the system evolved. Particular attention was therefore given to the IINS system architecture, specifically the partitioning of hardware and software functions as well as the means of data transmittal between the subsystem elements. The paper describes in detail the design trade-offs made as well as hardware selections. The system is configured using several different manufacturers' equipment making the data interface and its configuration control of great importance. The above considerations lead to the conclusion that the appropriate data bus interface standard would be MIL-STD-1553. This interface standard not only facilitates configuration control during system development, but will allow rapid integration with future aircraft systems that will use it as a primary means of data transmittal. The paper describes in detail the signal interface between the subsystem element, onboard mission equipment and flight instruments. The most significant change anticipated in the course of the IINS life cycle is the substitution of Global Positioning System User Equipment for the external update function currently provided by TACAN. Another area where technology insertion may be deemed appropriate is the substitution of a strapdown inertial navigation unit (INU) employing ring laser gyros (RLG) for the currently employed conventional gimballed INU. The INU employed in the IINS is built to a form, fit and function specification, which will allow substitution of a RLG INU when such a unit is qualified and in production. The system hardware and software is structured so that these changes can be accomplished with a minimum of effort. The paper also describes the federated processing structure which allows this flexibility with particular attention to the inertial/TACAN update algorithm and data bus control functions. The paper concludes with a brief summary of the development status of the IINS, currently in full scale engineering development.

Initial alignment is of vital importance to inertial navigation systems (INS). The two main methods of self-alignment, the traditional gyro-compassing alignment method and the Kalman filter method, are discussed. We apply and compare the two methods via simulation and test in use of a strapdown inertial navigation system (SINS) based fiber optic gyroscope (FOG), respectively. The simulation and test show that Kalman filter method is better in convergent time and accuracy of attitude and azimuth, however, gyro-compassing alignment method is better in robustness. This provides a useful reference for designing alignment algorithm for FOG SINS.

Basic features of implementing the modes of operation of a strapdown inertial navigation system (SINS) under polar conditions are considered. The above-mentioned SINS based on fiber-optic gyros has been the subject of our studies. Approaches to the improvement of the accuracy characteristics of such SINS are proposed, which are based on the use of geophysical invariants and on the integrated processing of appropriate observations. The results of field development of the SINS presented in this paper are shown and analyzed.