Showing papers on "Inertial reference unit published in 1992"

Multiaxis redundant fly-by-wire primary flight control system

Abstract: A redundant, multi-channel fly-by-wire control system for use in an aircraft is disclosed. The system includes a left flight control channel (60), a center flight control channel (80) and a right flight control channel (90). Each control channel is capable of flying the aircraft in the event the other two channels fail. Included within each control channel is an actuator controller electronics unit (ACE (62, 82, 92)), which transmits a series of pilot control transducer signals to a set of primary flight computers (64, 84, and 94). The primary flight computers combine the pilot control transducer signals with data obtained from an air data and inertial reference unit (145) to generate a set of flight surface commands. Each ACE selects a set of flight surface commands to control the movement of a set of flight control surfaces on the aircraft.

Fault-tolerant inertial navigation system

Abstract: Disclosed herein is a Fault-Tolerant Inertial Navigation System comprising, in a preferred embodiment, a Redundant Set of at least two Inertial Navigation Systems, from which one may identify and isolate at least one instrument within an Inertial Navigation Unit which shows substantial performance degradation. The two inertial navigation units are fully capable of performing navigational functions. Each of these inertial navigation units has a plurality of navigational instruments, including at least three linear sensors (such as accelerometers) and three angular change sensors (such as gyroscopes or ring laser gyroscopes). No two linear sensors nor any two angular change sensors of either unit are aligned colinearly. Each of the inertial navigation units produces a set of independent navigational solutions at each of their respective outputs. The independent navigational solutions of each of the navigation units are compared and any significantly degraded performance of any one linear sensor or any one angular change sensor is detected.

Self-contained method for correction of an inertial system over a body of water

Abstract: A method and apparatus for performing on-board corrections to the computed navigation variables of an inertial system on an aircraft while flying over a body of water. Onboard instruments, including a barometric altimeter and a radar altimeter, measure the vertical distance of the aircraft above an ellipsoidal model of the earth and above the body of water respectively. An on-board computer calculates the differences between such heights over a plurality of points along the path the aircraft travels over the water as indicated by its inertial navigation system. The differences are compared with a map of the undulation of the geoid encompassing the region to determine the deviation of the navigated course from the true course. Appropriate corrections to the aircraft's inertial system may then be made to reduce error.

Airdata Calibration Techniques for Measuring Atmospheric Wind Profiles

Abstract: The research airdata system of an instrumented F-104 aircraft has been calibrated to measure winds aloft in support of the Space Shuttle wind measurement investigation at the National Aeronautics and Space Administration Ames Research Center Dryden Flight Research Facility. For this investigation, wind measurement accuracies comparable to those obtained from Jimsphere balloons were desired. This required an airdata calibration more accurate than needed for most aircraft research programs. The F-104 aircraft was equipped with a research pilot-static noseboom with integral angle-of-attack and flank angle-of-attack vanes and a ring-laser-gyro inertial reference unit. Tower fly-bys and radar acceleration-decelerations were used to calibrate Mach number and total temperature. Angle of attack and angle of side slip were calibrated with a trajectory reconstruction technique using a multiple-state linear Kalman filter. The F-104 aircrat and instrumentation configuration, flight test maneuvers, data corrections, calibration techniques, and resulting calibrations and data repeatability are presented. Recommendations for future airdata systems on aircraft used to measure winds aloft are also given.

Comparison of wideband inertial line-of-sight stabilization reference mechanizations

Abstract: One subsystem critical to the performance of a precision electro-optical line-of-sight (LOS) pointing system is a wide-band inertial stabilization reference. This paper compares, in terms of relative performance of LOS stabilization in the presence of vehicle jitter, three inertial LOS stabilization reference mechanizations for space-based optical systems. The three mechanizations are: an inertially stabilized platform, the Inertial Pseudo-Star Reference Unit (IPSRU) under development at Draper Laboratory; a device called the Optical Reference Gyro (ORG), also developed at Draper Laboratory; and a strapdown wide-band inertial sensor assembly. Each of the three stabilization reference mechanizations generates a collimated alignment beam that is injected into the entrance aperture of the optical system. In the stabilized platform mechanization, the alignment beam emanates from a platform inertially stabilized from vehicle jitter in two axes, and thus the alignment beam becomes a jitter- stabilized pseudo-star. An alignment loop closed around the pseudo-star image and a steering mirror in the optical path stabilizes the LOS against vehicle jitter. The ORG alignment beam projects from the gyro rotor, which is decoupled from case motion and is effectively inertially stabilized. The ORG spin-speed noise is compensated with phase-lock technology. In the strapdown mechanization, the alignment beam source is hard-mounted to the vehicle. Inertial measurement of the local vehicle motion is fed forward, open loop, to a steering mirror in the optical path to compensate for alignment beam jitter.© (1992) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Method for selfguidance of missile towards a target by means of range measurements

Abstract: The missile (E) comprises a recognition and tracking unit (13) which receives an infrared image in order to extract from it objects having characteristics identical to predetermined targets (CB). When identification is established, the recognition and tracking unit (13) then controls the antenna control unit (14) so that the object identified is permanently in the centre of the image received. The identification unit (2) establishes the attitude (S, G) of the antenna (11) with respect to the missile (E), and the inertial reference unit (3) establishes the attitude of the missile (E) with respect to an inertial reference axis system linked to the centre of gravity of the missile so that a computer (4) calculates the azimuth angle ( alpha G) of the presumed target in the inertial axis system and deduces the distance (D) separating the missile (E) from the presumed target (CB). This distance information (D) allows a more precise authentication of the target. Additional means (5, 6) make it possible to orient the missile (E) with a view to reaching the target (CB) according to predetermined criteria, particularly by skirting round the target towards a favourable impact point on the target.

Inertial pseudo star reference unit

Abstract: Summary form only given. The authors present the development of a precision pointing system, the inertial pseudo-star reference unit (IPSRU). The IPSRU implements a collimated light source mounted on an inertially stabilized platform that is configured using a centrally located two-degree-of-freedom (2-DOF) flexure assembly. The platform inertial sensing is achieved through the combined use of a precision, low-noise 2-DOF dry-tuned rotor gyro (DTG) and two angular displacement sensors (ADS). The composite gyro and ADS implementation optimally combines low-frequency gyro measurements with ADS higher frequency sensing to achieve a composite wide band, extremely low noise, and inertial sensing capability. The collimated light beam in effect becomes a jitter-stabilized pseudo-star. In addition, its pointing direction in inertial space can be changed at a precise rate by commands applied via torquing signals to the gyro. >

High speed missile inertial reference optical support and re-initialising system - derives alignment error estimates from comparison of outputs of video camera sensors of reference points located on launcher with values from guidance system

Abstract: Sensors mounted on the missile body detect reference points with known coordinates. The missile body's orientation is derived from the sensor signals and compared with that produced by the inertial reference system. Alignment errors estimated from the results of this comparison are used to re-initialise the inertial reference system. The missile body can be fired from a weapon launcher on which the reference points are arranged. In this case the sensors view in the opposite direction to the flight direction. For an unpowered guided weapon system, the guided weapon forms the reference points and a sensor views in the flight direction and responds to the reference points on the weapon. ADVANTAGE - Ensures highly accurate alignment of inertial reference system nd hence targeting reliability.

Wideband control of gyro/accelerometer multisensors in a strapdown guidance system

Abstract: This paper describes an inertial measurement unit based on three dry tuned gyroscopes, two of which are unbalanced. This configuration, without the three usual accelerometers, leads to savings in space, weight, and cost that are necessary in guidance systems for agile missiles. Design objectives are wideband estimation of angular rate and linear acceleration applied to the inertial measurement unit together with high stiffness of caging loops. The contribution of this paper is to show how to address the coupled high-frequency dynamics of the six measurement outputs of the system, by applying the so-called linear quadratic Gaussian technique. First the linearized model of the rotor motion of each dry-tuned gyroscope is derived from a mechanical analysis. Then the design of the controller/estimator is described. Performances are illustrated with test data.

Navigational needs for current and future high-G inertial measurement units

Abstract: The author uses the ADKEM (advanced kinetic energy missile) IMU (inertial measurement unit) performance requirements as a platform to discuss various error sources and their contribution to the ADKEM system performance and to indicate possible applications of this technology for the future. Particular attention is given to the ADKEM IMU error budget. >

Fiber optic rotation sensor for long lifetime space missions

Abstract: The present F-O rotation sensors (FORS) are all-solid state devices for measuring rotations and rotation rates in inertial space that may reach the 0.003 deg/hr (1-sigma) accuracies required for NASA's Saturn-orbiting Cassini mission. Attention is presently given to the mission, inertial reference unit, and FORS instrument optoelectronic component requirements envisioned for such spacecraft applications.

Application of Zero-Lock Laser Gyro Technology to High Accuracy Stellar-Inertial Systems

Abstract: The four-mode multioscillator ring laser gyro, designated as the Zero-lock Laser GyroTM (ZLG) by Litton, enables new and innovative approaches to mechanizing high accuracy stellar-inertial navigation systems. Absent any requirement for dithering or rate biasing to circumvent the lock-in phe- nomenon inherent in conventional two-mode ring laser gy- ros (RLG), the ZLG offers many advantages for stellar- inertial systems. The ZLG does not impose any disturbing dynamic motion on the stellar sensor, facilitating high accu- racy star tracking. In addition, the extremely precise scale factor of the ZLG allows comounting of the stellar sensor with the inertial sensors on a common instrument cluster and sequentially pointing the whole stellar-inertial sensor clus- ter at multiple target stars to measure the inertial system errors. This eliminates the need for the highly accurate angle encoders used in currently operational stellar-inertial sys- tems to transform stellar observations to their inertial refer- ence axes. Factory production of medium-accuracy ZLG inertial navi- gation systems (INS) was initiated in early 1991 for several military andcommercial contractual programs. Since stellar updates can correct for gyro drift errors, the same medium- accuracy ZLG is being utilized in Litton developmental pro- grams for high accuracy stellar-inertial systems. This paper describes the significant differences between four- and two-mode RLGs, the tradeoffs in mechanizing RLG/ZLG-based stellar-inertial systems, the specific applications of ZLG technology to stellar-inertial programs, and the resultant benefits.