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At least 91 records · Page 5

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NASA NTRS1978

Evaluation of display and control concepts for a terminal configured vehicle in final approach in a windshear environment

A revised treatment of nonrandom inputs was incorporated in the model. Response behavior was observed for two display configurations (a pictorial EADI presentation and a flight-director configuration requiring use of a panel-mounted airspeed indicator), two control configurations (attitude and velocity control wheel steering), and two shear environments, each of which contained a head-to-tail shear and a vertical component. In general, performance trends predicted by the model were confirmed experimentally. Experimental and analytical results both indicated superiority to the EADI display with respect to regulation of height and airspeed errors. Velocity steering allowed tighter regulation of height errors, but control parameters had little influence on airspeed regulation. Model analysis indicated that display-related differences could be ascribed to differences in the quality of speed-related information provided by the two displays.

Levison, W. H.↗

NASA NTRS2024

Quality Assessment of the Nominal Attitude Model of TOPEX/Poseidon Using Quaternion Data

The altimetry mission TOPEX/Poseidon was operational between 1992 and 2005 and provided innovative insight into global and regional sea level changes and their variability. The correct modelling of the spacecraft orbit in precise orbit determination requires detailed information on the attitude of the spacecraft body and the solar array. Since the availability of attitude observations in form of quaternion data for this mission is limited, the nominal orientation model has to be used mainly. In this study, we compare the nominal and quaternion-based approaches of the TOPEX/Poseidon attitude realisation and validate the nominal model w.r.t. the observed attitude. We found good agreement between both approaches in the spacecraft-related roll, pitch, yaw, and solar array angles with slight differences of up to 0.31◦ on average. The best accordance is obtained in the yaw angles. A general angular offset of 0.12◦ and −0.19◦ on average is found in the pitch and solar array angle differences, respectively. Parameters estimated in the orbit computation process also reflect the good agreement between both orientation models with slightly better results when using the observed attitude. However, in some intervals, the nominal model lacks detailed information about the correct attitude. These can be adjusted using the information from the quaternion data. Such intervals indicate off-nominal spacecraft attitude events like drifts or offsets. Purely geometrically analysed, mispointing in the roll and pitch components of about 0.25◦ results in a change of the altimeter phase centre position in the radial direction of about 7 mm, which directly impacts the sea-level heights determined from altimetry measurements.

TOPEX/Poseidon↗

NASA NTRS1989

Measurements of pressures on the tail and aft fuselage of an airplane model during rotary motions at spin attitudes

An investigation was conducted in the Langley Spin Tunnel to measure the pressures on the surface of the horizontal and vertical tail and the aft fuselage of an aircraft model. The pressures were measured on a model of a proposed Australian Primary Trainer airplane configuration while the model was rotating at spinning attitudes. The test results indicate that the presence of the horizontal tail adversely modifies the surface pressure on the vertical tail. The presence of the wing also adversely modifies the pressures on the horizontal and vertical tails.

Bowman, James S., Jr.↗

NASA NTRS1995

Vertical Scales of Turbulence at the Mount Wilson Observatory

The vertical scales of turbulence at the Mount Wilson Observatory are inferred from data from the University of California at Berkeley Infrared Spatial Interferometer (ISI), by modeling path length fluctuations observed in the interferometric paths to celestial objects and those in instrumental ground-based paths. The correlations between the stellar and ground-based path length fluctuations and the temporal statistics of those fluctuations are modeled on various timescales to constrain the vertical scales. A Kolmogorov-Taylor turbulence model with a finite outer scale was used to simulate ISI data. The simulation also included the white instrumental noise of the interferometer, aperture-filtering effects, and the data analysis algorithms. The simulations suggest that the path delay fluctuations observed in the 1992-1993 ISI data are largely consistent with being generated by refractivity fluctuations at two characteristic vertical scales: one extending to a height of 45 m above the ground, with a wind speed of about 1 m/ s, and another at a much higher altitude, with a wind speed of about 10 m/ s. The height of the lower layer is of the order of the dimensions of trees and other structures near the interferometer, which suggests that these objects, including elements of the interferometer, may play a role in generating the lower layer of turbulence. The modeling indicates that the high- attitude component contributes primarily to short-period (less than 10 s) fluctuations, while the lower component dominates the long-period (up to a few minutes) fluctuations. The lower component turbulent height, along with outer scales of the order of 10 m, suggest that the baseline dependence of long-term interferometric, atmospheric fluctuations should weaken for baselines greater than a few tens of meters. Simulations further show that there is the potential for improving the seeing or astrometric accuracy by about 30%-50% on average, if the path length fluctuations in the lower component are directly calibrated. Statistical and systematic effects induce an error of about 15 m in the estimate of the lower component turbulent altitude.

Treuhaft, Robert N.↗

NASA NTRS1999

Entry Dispersion Analysis for the Genesis Sample Return Capsule

Genesis will be the first mission to return samples from beyond the Earth-Moon system. The spacecraft will be inserted into a halo orbit about the L1 (Sun- Earth) libration point where it will remain for two years collecting solar wind particles. Upon Earth return, the sample return capsule, which is passively controlled, will descend under parachute to Utah. The present study describes the analysis of the entry, descent, and landing scenario of the returning sample capsule. The robustness of the entry sequence is assessed through a Monte Carlo dispersion analysis where the impact of off-nominal conditions is ascertained. The dispersion results indicate that the capsule attitude excursions near peak heating and drogue chute deployment are within Genesis mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 47.8 km in downrange by 15.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗

NASA NTRS1999

Entry Dispersion Analysis for the Genesis Sample Return Capsule

Genesis will be the first mission to return samples from beyond the Earth-Moon system. The spacecraft will be inserted into a halo orbit about the L1 (Sun- Earth) libration point where it will remain for two years collecting solar wind particles. Upon Earth return, the sample return capsule, which is passively controlled, will descend under parachute to Utah. The present study describes the analysis of the entry, descent, and landing scenario of the returning sample cap- sule. The robustness of the entry sequence is assessed through a Monte Carlo dispersion analysis where the impact of off-nominal conditions is ascertained. The dispersion results indicate that the capsule attitude excursions near peak heating and drogue chute deployment are within Genesis mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 47.8 km in downrange by 15.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun, N.↗

NASA NTRS1949

Supplementary Ditching Investigation of a 1/18-Scale Model of the North American B-45 Airplane

A supplementary investigation of a 1/18-scale dynamically similar model of the North American B-45 airplane was made in calm water to observe the ditching behavior and to determine the effect of further conditions of simulated damage. The behavior of the model was determined from motion-picture records, time-history deceleration records, and visual observations. It was concluded that, during a ditching, the nose-wheel doors, bomb-bay doors, under surface of the aft fuselage and lower part of the bomb-bay aft bulkhead will be torn away. When such damage occurs, the diving found in the previous tests will not occur and the airplane will porpoise during the first third of the landing run and then run smoothly. Because of the extensive damage which occurs, the fuselage aft of the pilots' compartment will fill with water very rapidly and is, therefore, an undesirable ditching station. No change in the previously recommended ditching attitude and flap setting (6 deg attitude, flaps down) is indicated by these supplementary tests.

Fisher, Lloyd J.↗

NASA NTRS2005

Entry Dispersion Analysis for the Genesis Sample Return Capsule

Genesis will be the first mission to return samples from beyond the Earth-Moon system. The spacecraft will be inserted into a halo orbit about the L1 (Sun-Earth) libration point where it will remain for two years collecting solar wind particles. Upon Earth return, the sample return capsule, which is passively controlled, will descend under parachute to Utah. The present study describes the analysis of the entry, descent, and landing scenario of the returning sample capsule. The robustness of the entry sequence is assessed through a Monte Carlo dispersion analysis where the impact of off-nominal conditions is ascertained. The dispersion results indicate that the capsule attitude excursions near peak heating and drogue chute deployment are within Genesis mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 47.8 km in downrange by 15.2 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗

NASA NTRS1993

Handling qualities effects of display latency

Display latency is the time delay between aircraft response and the corresponding response of the cockpit displays. Currently, there is no explicit specification for allowable display lags to ensure acceptable aircraft handling qualities in instrument flight conditions. This paper examines the handling qualities effects of display latency between 70 and 400 milliseconds for precision instrument flight tasks of the V-22 Tiltrotor aircraft. Display delay effects on the pilot control loop are analytically predicted through a second order pilot crossover model of the V-22 lateral axis, and handling qualities trends are evaluated through a series of fixed-base piloted simulation tests. The results show that the effects of display latency for flight path tracking tasks are driven by the stability characteristics of the attitude control loop. The data indicate that the loss of control damping due to latency can be simply predicted from knowledge of the aircraft's stability margins, control system lags, and required control bandwidths. Based on the relationship between attitude control damping and handling qualities ratings, latency design guidelines are presented. In addition, this paper presents a design philosophy, supported by simulation data, for using flight director display augmentation to suppress the effects of display latency for delays up to 300 milliseconds.

King, David W.↗

NASA NTRS1985

Simulation evaluation of display/FLIR concepts for low-altitude, terrain-following helicopter operations

A piloted simulation of three head-down display (HDD) concepts with flight-director guidance superimposed on forward-looking infrared (FLIR) imagery was performed to evaluate the task of low-level, terrain-following (TF), manual helicopter flight. The three display concepts were examined for the purpose of finding ways by which aircraft flight-attitude and command symbols and FLIR imagery could be integrated onto one instrument. In all cases, the FLIR imagery was centered on the flight-path vector of the aircraft. The three displays were then characterized by having: (1) pitch attitude conformal to the FLIR imagery; (2) pitch attitude conformal to the FLIR imagery, but with an increase in the scaling; and (3) pitch attitude nonconformal to the FLIR imagery with the same pitch scaling as in (2). The simulation was conducted on the Vertical Motion Simulator (VMS) at Ames Research Center, using NASA and Air Force test pilots. The pilots indicated that the nonconformal pitch attitude and FLIR display was the preferred way to display information because of the absence of pitch-attitude information on displays (1) and (2) during some portions of the operational flight envelope.

Swenson, H. N.↗

NASA NTRS1975

Attitude and orbit estimation using stars and landmarks

An extended Kalman filter is used to process line-of-sight measurements to stars and known landmarks providing a statistical indication of performance in estimating spacecraft attitude, orbital ephemeris, and the bias drift of a set of three strapdown gyros. The landmark measurements were assumed to have been taken from the imagery of an Earth-observing multispectral scanner. It is shown that filtering of these noisy measurements results in highly accurate estimates of the above parameters. Results are given showing the sensitivity of performance to various system parameters such as star tracker accuracy, errors in the knowledge of landmark position, and number of stars and landmarks processed.

White, R. L.↗

NASA NTRS2000

Quasi-Steady Acceleration Direction Indicator in Three Dimensions

Many materials processing and fluids physics experiments conducted in a microgravity environment require knowledge of the orientation of the low-frequency acceleration vector. This need becomes especially acute for space experiments such as directional solidification of a molten semiconductor, which is extremely sensitive to orientation and may involve tens of hours of operations of a materials furnace. These low-frequency acceleration data have been measured for many Shuttle missions with the Orbital Acceleration Research Experiment. Previous attempts at using fluid chambers for acceleration measurements have met with limited success due to pointing and vehicle attitude complications. An acceleration direction indicator is described, which is comprised of two orthogonal short cylinders of fluid, each with a small bubble. The motion and the position of the bubble within the chamber will indicate the direction of the acceleration experienced at the sensor location. The direction of the acceleration vector may then be calculated from these data. The frequency response of such an instrument may be tailored for particular experiments with the proper selection of fluid and gas parameters, surface type, and geometry. A three-dimensional system for sensing and displaying the low-frequency acceleration direction via an innovative technique described in this paper has advantages in terms of size, mass, and power compared with electronic instrumentation systems.

DeLombard, Richard↗

NASA NTRS1978

Summary of system state-of-health measurements STRATCOM 8: A balloon experiment

Various measurements were made to determine the temperature and attitude of the gondola and the status of primary power and control equipment. Bead thermistors were used to measure temperatures at selected points throughout the gondola. A two-axis magnetometer and a two-axis pendulum were used to measure gondola attitude. Voltage and current measurements indicated the status of the primary power sources and associated power converters.

Herrington, P. B.↗

NASA NTRS1992

Signal processing for airborne doppler radar detection of hazardous wind shear as applied to NASA 1991 radar flight experiment data

Radar data collected during the 1991 NASA flight tests have been selectively analyzed to support research directed at developing both improved as well as new algorithms for detecting hazardous low-altitude windshear. Analysis of aircraft attitude data from several flights indicated that platform stability bandwidths were small compared to the data rate bandwidths which should support an assumption that radar returns can be treated as short time stationary. Various approaches at detection of weather returns in the presence of ground clutter are being investigated. Non-coventional clutter rejection through spectrum mode tracking and classification algorithms is a subject of continuing research. Based upon autoregressive modeling of the radar return time sequence, this approach may offer an alternative to overcome errors in conventional pulse-pair estimates. Adaptive filtering is being evaluated as a means of rejecting clutter with emphasis on low signal-to-clutter ratio situations, particularly in the presence of discrete clutter interference. An analysis of out-of-range clutter returns is included to illustrate effects of ground clutter interference due to range aliasing for aircraft on final approach. Data are presented to indicate how aircraft groundspeed might be corrected from the radar data as well as point to an observed problem of groundspeed estimate bias variation with radar antenna scan angle. A description of how recorded clutter return data are mixed with simulated weather returns is included. This enables the researcher to run controlled experiments to test signal processing algorithms. In the summary research efforts involving improved modelling of radar ground clutter returns and a Bayesian approach at hazard factor estimation are mentioned.

Baxa, Ernest G., Jr.↗

NASA NTRS1998

A Novel Sensor for Attitude Determination Using Global Positioning System Signals

An entirely new sensor approach for attitude determination using Global Positioning System (GPS) signals is developed. The concept involves the use of multiple GPS antenna elements arrayed on a single sensor head to provide maximum GPS space vehicle availability. A number of sensor element configurations are discussed. In addition to the navigation function, the array is used to find which GPS space vehicles are within the field-of-view of each antenna element. Attitude determination is performed by considering the sightline vectors of the found GPS space vehicles together with the fixed boresight vectors of the individual antenna elements. This approach has clear advantages over the standard differential carrier-phase approach. First, errors induced by multipath effects can be significantly reduced or eliminated altogether. Also, integer ambiguity resolution is not required, nor do line biases need to be determined through costly and cumbersome self-surveys. Furthermore, the new sensor does not require individual antennas to be physically separated to form interferometric baselines to determine attitude. Finally, development potential of the new sensor is limited only by antenna and receiver technology development unlike the physical limitations of the current interferometric attitude determination scheme. Simulation results indicate that accuracies of about 1 degree (3 omega) are possible.

Crassidis, John L.↗

NASA NTRS1997

Aerodynamics of a Simple Three-Dimensional High-Lift System

Experimental results are presented for a three-element high-lift system with a half-span flap and both a full- and three-quarter span slat. The simplified wing spanned the wind-tunnel test section with no wing sweep or wing tip. Full- and part-span slat configurations were tested to investigate the three-dimensional effects of a finite-span slat in a landing configuration. With a part-span slat, flow visualization and surface pressures revealed two vortices at the flap tip and one at the slat tip. Compared to the full-span slat case, the part-span slat configuration stalled at a lower angle of attack and yielded reduced lift at all angles of attack. At a representative approach attitude (10 deg), wake surveys indicated comparable induced drag for the full- and part-span slat configurations while the profile drag was more than doubled with a part-span slat. The vorticity contribution of the slat-tip vortex was minimal, suggesting that the vortex dissipates through viscous interaction with the main-element boundary layer. The addition of a lower-surface fence at the slat tip yielded two co-rotating vortices and partially alleviated the profile drag increase of the part-span slat.

Storms, Bruce↗

NASA NTRS1986

Magellan spacecraft - Attitude determination, updates and gyro parameter calibration

A detailed mathematical algorithm for onboard autonomous attitude determination and parameter calibration in the Magellan Venus-orbiting spacecraft has been developed, and its performance has been estimated. This algorithm uses a contamination reduction attitude profile and retractive star scan maneuvers to minimize the interparameter error contamination and to maximize the parameter observability. The details of the underlying theory and strategy are described, and an extensive performance analysis with worst-case parameter errors is reported which indicates that the in-flight attitude determination unit/parameter calibration (ADU/PC) algorithm will be able to meet all the different mission phase requirements with comfortable margins. The problem of nonexact mapping retrace due to uncertain performance of control authority is addressed, and its lack of impact on the system performance is demonstrated.

Huang, W. H.↗

NASA NTRS1982

Application of matrix singular value properties for evaluating gain and phase margins of multiloop systems

A stability margin evaluation method in terms of simultaneous gain and phase changes in all loops of a multiloop system is presented. A universal gain-phase margin evaluation diagram is constructed by generalizing an existing method using matrix singular value properties. Using this diagram and computing the minimum singular value of the system return difference matrix over the operating frequency range, regions of guaranteed stability margins can be obtained. Singular values are computed for a wing flutter suppression and a drone lateral attitude control problem. The numerical results indicate that this method predicts quite conservative stability margins. In the second example if the eigenvalue magnitude is used instead of the singular value, as a measure of nearness to singularity, more realistic stability margins are obtained. However, this relaxed measure generally cannot guarantee global stability.

Mukhopadhyay, V.↗