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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Failure detection and isolation of redundant inertial systems for Space Shuttle

A failure detection and isolation technique is presented for application to 'off-the-shelf' type four-gimbaled inertial measurement units (IMUs) such as the KT-70 or Carousel IV. This study concentrates initially upon actual four-gimbaled IMU performance requirements for current Space Shuttle booster and orbiter mission phases. When this information is obtained for a simplex system, the remaining area of study necessary to attain redundant IMU capabilities is the establishment of mission-dependent performance failure detection thresholds. These thresholds, then, permit a careful evaluation of the capabilities of various 'off-the-shelf' four gimbaled IMUs to satisfy the Shuttle mission performance requirements with various failure detection and isolation methods.

Brown, H. E.↗
NASA NTRS1975

A three-axis flight simulator

A simulator is described, which was designed for testing and evaluating inertial measuring units, and flight platforms. Mechanical and electrical specifications for the outer, middle, and inner axis are presented. Test results are included.

Mason, M. G.↗
NASA NTRS1975

Maximum likelihood failure detection techniques applied to the shuttle orbiter reaction control subsystem

A technique for on-board detection and identification of hard failures and leaks of the shuttle orbiter reaction control subsystem jets, during the orbital flight phase, is presented. The method uses gimbal angle and linear accelerometer measurements from the orbiter inertial measurement unit and requires no additional hardware. Extended Kalman filters with residual traps are employed for state estimation, and generalized likelihood ratio tests for jet failure identification. Rigid body simulation results indicate identification times of less than 2 seconds for hard jet failures and less than 70 seconds for jet leaks.

Deckert, J. C.↗
NASA NTRS1973

Space shuttle post-entry and landing analysis. Volume 1: Candidate system evaluations

The general purpose of this study is to aid in the evaluation and design of multi-sensor navigation schemes proposed for the orbiter. The scope of the effort is limited to the post-entry, energy management, and approach and landing mission phases. One candidate system based on conventional navigation aids is illustrated including two DME (Distance Measuring Equipment) stations and ILS (Instrument Landing System) glide slope and localizer antennas. Some key elements of the system not shown are the onboard IMUs (Inertial Measurement Units), altimeters, and a computer. The latter is programmed to mix together (filter) the IMU data and the externally-derived data. A completely automatic, all-weather landing capability is required. Since no air-breathing engines will be carried on orbital flights, there will be no chance to go around and try again following a missed approach.

Crawford, B. S.↗
NASA NTRS1975

Dispersion analysis and linear error analysis capabilities of the space vehicle dynamics simulation program

Previous error analyses conducted by the Guidance and Dynamics Branch of NASA have used the Guidance Analysis Program (GAP) as the trajectory simulation tool. Plans are made to conduct all future error analyses using the Space Vehicle Dynamics Simulation (SVDS) program. A study was conducted to compare the inertial measurement unit (IMU) error simulations of the two programs. Results of the GAP/SVDS comparison are presented and problem areas encountered while attempting to simulate IMU errors, vehicle performance uncertainties and environmental uncertainties using SVDS are defined. An evaluation of the SVDS linear error analysis capability is also included.

Snow, L. S.↗
NASA NTRS1975

Space Tug laser gyro IMU

A redundant inertial measuring unit (IMU) incorporating six strapdown laser gyros and six accelerometers, arranged so that sensitive axes are normal to the faces of a dodecahedron, provides enhanced reliability with reduced hardware weight. Software monitoring of sensor outputs senses failure of sensors and the system is designed for triple redundancy, with built-in test equipment. Attention is centered on redundancy and fail-safe features, and on the closed-path ring laser gyro arrangement.

Morrison, R.↗
NASA NTRS1976

Redundant Strapdown Laser Gyro Navigation System

For the last several years, NASA has pursued the development of low-cost high-reliability inertial navigation systems that would satisfy a broad spectrum of future space and avionics missions. Two specific programs have culminated in the construction of a Redundant Strapdown Laser Gyro Navigation System. These two programs were for development of a space ultrareliable modular computer (SUMC) and a redundant laser gyro inertial measurement unit (IMU). The SUMC is a digital computer that employs state-of-the-art large-scale integrated circuits configured in a functional modular breakdown. The redundant laser gyro IMU is a six-pack strapdown sensor package in a dodecahedron configuration which uses six laser gyros to provide incremental angular positions and six accelerometers for linear velocity outputs. The sensor arrangement allows automatic accommodation of two failures; a third failure can be tolerated provided it can be determined. The navigation system also includes redundant power supplies, built-in test-equipment (BITE) circuits for failure detection, and software which provides for navigation, redundancy management, and automatic calibration and alignment.

Mcpherson, B. W.↗
NASA NTRS1976

A pseudo root-locus method for the design of a class of two-input-two-output systems

This paper presents the design of digital rebalance loops for a tuned-rotor gyro which is intended to be an attitude sensor in a strapdown inertial measurement unit. A tuned-rotor gyro is a two-degree-of-freedom gyro which has two input axes, with cross-coupling between them. The rebalance loop of the gyro serves for two purposes. First, it restores the position of the rotor to its null position after experiencing an attitude input. Second, the rebalancing signal, which is proportional to the attitude change, is calibrated to give the desired attitude information. Since the gyro has more than one input and one output, a multivariable control technique is needed for designing the rebalance loop. The analytic model for the gyro is described and the available design methods are examined. A new root-locus design technique specially developed for the present problem is outlined. The design of the rebalance loop using the new technique is given. The computer simulation result of the designed system is presented and discussed.

Kao, M. K.↗
NASA NTRS1978

Orbit IMU alinement interpretation of onboard display data

The space shuttle inertial measurement unit (IMU) alinement algorith was examined to determine the most important alinement starpair selection criterion. Three crew displayed parameters were considered: (1) the results of the separation angle difference (SAD) check for each starpair; (2) the separation angle of each starpair; and (3) the age of each star measurement. It was determined that the SAD for each pair cannot be used to predict the IMu alinement accuracy. If the age of each star measurement is less than approximately 30 minutes, time is a relatively unimportant factor and the most important alinement pair selection criterion is the starpair separation angle. Therefore, when there are three available alinement starpairs and all measurements were taken within the last 30 minutes, the pair with the separation angle closest to 90 degrees should be selected for IMU alinement.

Corson, R.↗
NASA NTRS1980

Shuttle orbit IMU alignment. Single-precision computation error

The source of computational error in the inertial measurement unit (IMU) onorbit alignment software was investigated. Simulation runs were made on the IBM 360/70 computer with the IMU orbit alignment software coded in hal/s. The results indicate that for small IMU misalignment angles (less than 600 arc seconds), single precision computations in combination with the arc cosine method of eigen rotation angle extraction introduces an additional misalignment error of up to 230 arc seconds per axis. Use of the arc sine method, however, produced negligible misalignment error. As a result of this study, the arc sine method was recommended for use in the IMU onorbit alignment software.

Mcclain, C. R.↗
NASA NTRS1980

Onorbit IMU alignment error budget

The Star Tracker, Crew Optical Alignment Sight (COAS), and Inertial Measurement Unit (IMU) from a complex navigation system with a multitude of error sources were combined. A complete list of the system errors is presented. The errors were combined in a rational way to yield an estimate of the IMU alignment accuracy for STS-1. The expected standard deviation in the IMU alignment error for STS-1 type alignments was determined to be 72 arc seconds per axis for star tracker alignments and 188 arc seconds per axis for COAS alignments. These estimates are based on current knowledge of the star tracker, COAS, IMU, and navigation base error specifications, and were partially verified by preliminary Monte Carlo analysis.

Corson, R. W.↗
NASA NTRS1980

Orbit IMU alignment: Error analysis

A comprehensive accuracy analysis of orbit inertial measurement unit (IMU) alignments using the shuttle star trackers was completed and the results are presented. Monte Carlo techniques were used in a computer simulation of the IMU alignment hardware and software systems to: (1) determine the expected Space Transportation System 1 Flight (STS-1) manual mode IMU alignment accuracy; (2) investigate the accuracy of alignments in later shuttle flights when the automatic mode of star acquisition may be used; and (3) verify that an analytical model previously used for estimating the alignment error is a valid model. The analysis results do not differ significantly from expectations. The standard deviation in the IMU alignment error for STS-1 alignments was determined to the 68 arc seconds per axis. This corresponds to a 99.7% probability that the magnitude of the total alignment error is less than 258 arc seconds.

Corson, R. W.↗
NASA NTRS1980

Space shuttle navigation analysis. Volume 2: Baseline system navigation

Studies related to the baseline navigation system for the orbiter are presented. The baseline navigation system studies include a covariance analysis of the Inertial Measurement Unit calibration and alignment procedures, postflight IMU error recovery for the approach and landing phases, on-orbit calibration of IMU instrument biases, and a covariance analysis of entry and prelaunch navigation system performance.

Jones, H. L.↗
NASA NTRS1981

GPS aiding of ocean current determination

The navigational accuracy of an oceangoing vessel using conventional GPS p-code data is examined. The GPS signal is transmitted over two carrier frequencies in the L-band at 1575.42 and 1227.6 MHz. Achievable navigational uncertainties of differenced positional estimates are presented as a function of the parameters of the problem, with particular attention given to the effect of sea-state, user equivalent range error, uncompensated antenna motion, varying delay intervals, and reduced data rate examined in the unaided mode. The unmodeled errors resulting from satellite ephemeris uncertainties are shown to be negligible for the GPS-NDS (Navigation Development) satellites. Requirements are met in relatively calm seas, but accuracy degradation by a factor of at least 2 must be anticipated in heavier sea states. The aided mode of operation is examined, and it is shown that requirements can be met by using an inertial measurement unit (IMU) to aid the GPS receiver operation. Since the use of an IMU would mean higher costs, direct Doppler from the GPS satellites is presented as a viable alternative.

Mohan, S. N.↗
NASA NTRS1981

An extended BET format for La RC shuttle experiments: Definition and development

A program for shuttle post-flight data reduction is discussed. An extended Best Estimate Trajectory (BET) file was developed. The extended format results in some subtle changes to the header record. The major change is the addition of twenty-six words to each data record. These words include atmospheric related parameters, body axis rate and acceleration data, computed aerodynamic coefficients, and angular accelerations. These parameters were added to facilitate post-flight aerodynamic coefficient determinations as well as shuttle entry air data sensor analyses. Software (NEWBET) was developed to generate the extended BET file utilizing the previously defined ENTREE BET, a dynamic data file which may be either derived inertial measurement unit data or aerodynamic coefficient instrument package data, and some atmospheric information.

Findlay, J. T.↗
NASA NTRS1982

Reconstruction of the 1st Space Shuttle (STS-1) entry trajectory

A discussion of the generation of the best estimate trajectory (BET) of the first Space Shuttle Orbiter entry flight is presented. The BET defines a time history of the state, attitude, and atmospheric relative parameters throughout the Shuttle entry from an altitude of approximately 183 km to rollout. The inertial parameters were estimated utilizing a weighted least squares batch filter algorithm. Spacecraft angular rate and acceleration data derived from the Inertial Measurement Unit were utilized to predict the state and attitude which was constrained in a weighted least squares process to fit external tracking data consisting of ground based S-band and C-band data. Refined spacecraft altitude and velocity during and post rollout were obtained by processing artificial altimeter and Doppler data. The BET generation process is discussed. Software and data interface discussions are included. The variables and coordinate systems utilized are defined. STS-1 mission peculiar inputs are summarized. A listing of the contents of the actual BET is provided.

Findlay, J. T.↗
NASA NTRS1981

Shuttle /STS-1/ entry trajectory reconstruction

The methods and data used to reconstruct the Shuttle entry trajectory aerodynamics from 180 km to landing are reviewed. Onboard accelerometer and gyro measurements are used with a weighted least squares filter along with ground tracking data to provide the evolving spacecraft state. The software development is described, noting that computations include linear acceleration and angular rate data, aerodynamic coefficient, and inertial measurement units, sampled at 170 Hz. Ground tracking data was gathered on S- and C-bands; Guam data is considered critical as it was taken during the blackout at the beginning of entry. The generation of a best estimate trajectory, combining flight and ground data, yields inplane parameters of altitude, flight path angle, and velocity for comparison with navigational instrumentation. Predicted lift/drag ratios were found to be accurate from Mach 1-25, and up to 23% low below Mach 1.

Compton, H. R.↗
NASA NTRS1982

Shuttle orbiter stellar-inertial reference system

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.

Smith, F. E.↗