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An innovative procedure is described for measuring blaze angle that is free of systematic effects associated with an alternative approach. The method uses the effect of facet diffraction on the blaze function when an echelle grating is used off-Littrow. The angle measurement in the Littrow condition is wavelength independent and locates the actual direction of the grating facet normal.
Developments reported in modeling of fluid turbulence as superposition of number of interpenetrating velocity fields. Multivalued-velocity model has practical implications in design of aircraft, turbines, nozzles, pumps, and other systems that involve turbulent flow. In both viscous and inviscid cases, equations made to predict stable regine of oscillations, both finite and independent of initial conditions as in case of real turbulence.
The reattachment of a fully turbulent, two-dimensional shear layer downstream of a backward-facing step has been studied experimentally. The work examines the effect of modifications in shear layer structure upstream of reattachment. A series of experiments were conducted in a low-speed wind tunnel using specialized instrumentation suited to the highly turbulent, reversing flow near reattachment. Accurate characterization of the time-mean features of the reattaching flows was possible. Assuming linear scaling normalized on distance from reattachment, distributions of the normalized pressure coefficients, the time-averaged skin friction coefficient, and the forward flow fraction appear universal for two-dimensional reattachment, independent of initial conditions and steep height, for given duct geometry (area ratio) and for high step height Reynolds numbers. The results suggest universal flow structure in the reattachment zone.
The results of an experimental study of aerodynamic (surface velocity) and heat transfer distributions over the surfaces of two different, highly loaded, low-solidity contemporary turbine vane designs are presented. The aerodynamic configurations of the two vanes were carefully selected to emphasize fundamental differences in the character of the solution surface pressure distributions and the consequent effect on surface heat transfer distributions. The experimental measurements were made in moderate-temperature, three-vane cascades under steady-state conditions. The principal independent parameters (Mach number, Reynolds number, turbulence intensity, and wall-to-gas temperature ratio) were varied over ranges consistent with actual engine operation, and the heat matrix was structured to provide an assessment of the independent influence of each parameter. These measurements are intended to serve as verification data for a parallel analytical code development effort. The results of this parallel effort are briefly reviewed, and the principal conclusions to date are summarized.
Radar remote sensors may provide valuable input to crop classification procedures because of (1) their independence of weather conditions and solar illumination, and (2) their ability to respond to differences in crop type. Manual classification of multidate synthetic aperture radar (SAR) imagery resulted in an overall accuracy of 83 percent for corn, forest, grain, and 'other' cover types. Forests and corn fields were identified with accuracies approaching or exceeding 90 percent. Grain fields and 'other' fields were often confused with each other, resulting in classification accuracies of 51 and 66 percent, respectively. The 83 percent correct classification represents a 10 percent improvement when compared to similar SAR data for the same area collected at alternate time periods in 1978. These results demonstrate that improvements in crop classification accuracy can be achieved with SAR data by synchronizing data collection times with crop growth stages in order to maximize differences in the geometric and dielectric properties of the cover types of interest.
A pointing error model is used in the antenna calibration process. Data from spacecraft or radio star observations are used to determine the parameters in the model. However, the regression variables are not truly independent, displaying a condition known as multicollinearity. Ridge regression, a biased estimation technique, is used to combat the multicollinearity problem. Two data sets pertaining to Voyager 1 spacecraft tracking (days 105 and 106 of 1987) were analyzed using both linear least squares and ridge regression methods. The advantages and limitations of employing the technique are presented. The problem is not yet fully resolved.
Even though the Pioneer Venus plasma wave instrument returns a measure of only one component of the electric field due to waves, for the 100 Hz channel it is possible to calcualte an approximate Poynting flux of the waves by making several assumptions. When this is done the Poynting flux at lowest altitudes in a 30 Hz band centered on 100 Hz is found to be about 10 to the -7th W/sq m, independent of ionospheric conditions as indicated by the strength of the magnetic field. The local time variation of the approximate Poynting flux shows a maximum from about 2000 to 2200 LT in accord with inferences from the higher frequency channels. The observed wave energy flux is consistent with that expected from lightning given rates and flash strengths on Venus that are similar to or greater than that of the earth.
Under some circumstances the absorption of radiation momentum by an absorbing medium opens the possibility of a dynamical instability, sometimes called 'mock gravity'. Here, a simplified abstract model is studied in which the radiation source is assumed to remain spatially uniform, there is no reabsorption or reradiated light, and no forces other than radiative pressure act on the absorbing medium. It is shown that this model displays the unique feature of being not only unstable, but also self-organizing. The structure approaches a statistical dynamical steady state which is almost independent of initial conditions. In this saturated state the absorbers are concentrated in thin walls around empty bubbles; as the instability develops the big bubbles get bigger and the small ones get crushed and disappear. A linear analysis shows that to first order the thin walls are indeed stable structures. It is speculated that this instability may play a role in forming cosmic large-scale structure.
The asymptotic behavior of the curvature of material elements in turbulence is investigated using Lagrangian velocity-gradient time series obtained from direct numerical simulations of isotropic turbulence. Several material-element ensembles of different initial curvatures and shapes are studied. It is found that, at long times, the (first five) moments of the logarithm of characteristic curvature and shape factor asymptote to values that are independent of the initial curvature or shape. This evidence strongly suggests that the asymptotic pdf's of the curvature and shape of material elements are stationary and independent of initial conditions. Irrespective of initial curvature or shape, the asymptotic shape of a material surface is cylindrical with a high probability.
We evaluate the three-point function in Fourier space for an ensemble of three-dimensional 128 exp 3 numerical simulations with initial power spectra characterized by spectral index n = +1, 0, -1, -2, -3, with no high-frequency cutoff and with cutoff k(c) = 16 or k(c) = 4. To remove dependences on scale and on time, we present results as the reduced amplitude Q in the hierarchical model as a function of the dimensionless variable kd(rms), where d(rms) is the mean square displacement of a particle from its initial position. For scale-free initial conditions, there is no evolution in Q. For initial conditions with a cutoff, Q evolves until the scale of the cutoff is in the nonlinear regime; the results afterwards are no different from those with no initial cutoff. The transition from quasi-linear to nonlinear regimes is followed. In the quasi-linear regime, our results agree well with gravitational perturbation theory predictions, including a marked dependence on the shape of the configuration. In the nonlinear regime, the value of Q for scale-invariant initial conditions is remarkably independent of evolution epoch, of scale, and of configuration shape, and depends on spectral index roughly as Q = 3/(3 + n).
Measurement of the jet thrust of a turbojet engine in flight becomes more difficult as the number and complexity of the engine components increase. It is desirable, therefore, that a general correlation of jet thrust be developed which is applicable to a simple direct reading thrustmeter. In view of this need a correlation is presented which is independent of flight conditions and applicable to both non-afterburning and afterburning engines equipped with nonejector type fixed- and variable-area convergent exhaust nozzles. The general equation used in this correlation was derived from the theoretical jet-thrust equation for a choked convergent nozzle. The data used to verify the correlation were obtained over a range of altitudes from 10,000 to 54,000 feet and a range of flight Mach numbers from 0.4 to 1.1. A thrustometer based on this equation was installed on an afterburning turbojet engine equipped with a fixed area convergent exhaust nozzle. The results indicated that a meter based on this correlation would be applicable to a flight installation and that the probable error in thrust measurement would be approximately + or -1.5 percent provided the exhaust-nozzle thrust coefficient is known.
A simple evaporative cooling scheme was incorporated into a dynamic model to estimate orographic precipitation in mountainous regions. The orographic precipitation model is based on the transport of atmospheric moisture and the quantification of preciptable water across a 3D representation of the terrain from the surface up to 250 hPa. Advective wind fields are computed independently and boundary conditions are extracted from radiosonde data. Precipitation rates are obtained through calibration of a spatially distributed precipitation efficiency parameter. The model was applied to the central Sierra Nevada. Results show a gain of the order of 20% in threat-score coefficients designed to measure the forecast ability of the model. Accuracy gains are largest at high elevations and during intense storms associated with warm air masses.
A flight investigation was made to obtain experimental information on the handling qualities of a normal-acceleration type of automatic longitudinal control system. The control system was installed in a subsonic fighter-type airplane. In hands-off (stick-free) flight the normal-acceleration control system attempted to regulate the normal acceleration to a constant value which is dependent on the automatic-control-system trim setting. In maneuvering flight a given pilot's stick deflection produced a proportional change in normal acceleration, the change in acceleration being independent of flight condition. A small side-located controller stick was used by the pilot to introduce signals into the automatic control system. In the flight program emphasis was placed on the acceleration-limiting capabilities of the control system. The handling qualities were investigated in maneuvers such as slow and rapid pull-ups and turns and also in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking. Good acceleration limiting was obtained with the normal-acceleration control system by limiting the magnitude of the input signal that the pilot could introduce into the control system. The same values of control-system gain settings could be used from an acceleration-limiting stand-point at both 10,000 and 30,000 feet for the complete speed range of the airplane. The response characteristics of the airplane-control system combination were also satisfactory at both high and low altitude with these same values of control-system gain setting. In the pilot's opinion, the normal-acceleration control system provided good stability and control characteristics in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking.
The spectral reflectance of oceanic whitecaps in the visible and near infrared was investigated using high-altitude, 20 m resolution Airborne Visible/Infrared Imaging Spectrometer (AVIRIS) measurements off the Southern California coast. The whitecap effect on surface reflectance was expressed as a function of the difference between the reflectance of pixels contaminated by whitecaps and of adjacent pixels free of whitecaps. Whitecap reflectance was found to decrease substantially in the near infrared, by about 40% at 850 nm and 80% at 1,600 nm, in agreement with previous measurements in the coastal zone and the open ocean. The spectral dependence of whitecap reflectance appears to be fairly independent of environmental conditions, making it easy to take into account the resulting and significant effects in ocean color and aerosol remote sensing algorithms.
Single-trial event-related responses collected during the course of an experiment are typically averaged before analysis resulting in a rather crude picture of event-related brain dynamics. It has been quite clear for some time that these responses exhibit trial-to-trial variability: however, the computational techniques necessary to deal with such responses in noisy conditions have not been available. To this end we have developed the multiple-component, event-related potential model (mcERP), which assumes that the each event-related response consists of a sum of multiple evoked components each described by a stereotypical waveshape. These waveshapes are allowed to vary in amplitude and onset latency from trial to trial, which allows us to capture, to first-order, the trial-dependent variations in event-related brain dynamics. We have constructed many sets of synthetic data designed to simulate intracortical recordings from a 15 channel, linear-array multielectrode implanted acutely in V1 of an awake-behaving macaque undergoing visual stimulation with a red light flash. This synthetic data was used to characterize the performance of the mcERP algorithm. First we quantified the degree to which such trial-to-trial variability aids in the identification of multiple components, and we demonstrate that amplitude variability is a more important factor in component separation than latency variability. Second, we quantified the behavior of the algorithm under two distinct signal-to-noise ratio (SNR) conditions: Gaussian noise independently present in each channel, and highly correlated (1/f distributed), far-field noise presented identically in each channel of the array. The mcERP algorithm was found to be robust to noise accurately identifying all component waveshapes and their associated single-trial characteristics down to SNR levels of -20dB for Gaussian noise and -7dB for 1/f far-field noise. Comparisons of the performance of this algorithm with factor analysis (FA) and independent component analysis (ICA) will be described by Knuth et al. (SFN abstracts, 2002). In addition, the advantages of application of mcERP to real data will be described by Shah et al, (these abstracts, 2002: SFN abstracts, 2002).
The National Aeronautics and Space Administration (NASA) at the Dryden Flight Research Center (DFRC) has been conducting flight-test research using an F-15 aircraft (figure 1). This aircraft has been specially modified to interface a neural net (NN) controller as part of a single-string Airborne Research Test System (ARTS) computer with the existing quad-redundant flight control system (FCC) shown in figure 2. The NN commands are passed to FCC channels 2 and 4 and are cross channel data linked (CCDL) to the other computers as shown. Numerous types of fault-detection monitors exist in the FCC when the NN mode is engaged; these monitors would cause an automatic disengagement of the NN in the event of a triggering fault. Unfortunately, these monitors still may not prevent a possible NN hard-over command from coming through to the control laws. Therefore, an additional and unique safety monitor was designed for a single-string source that allows authority at maximum actuator rates but protects the pilot and structural loads against excessive g-limits in the case of a NN hard-over command input. This additional monitor resides in the FCCs and is executed before the control laws are computed. This presentation describes a floating limiter (FL) concept1 that was developed and successfully test-flown for this program (figure 3). The FL computes the rate of change of the NN commands that are input to the FCC from the ARTS. A window is created with upper and lower boundaries, which is constantly floating and trying to stay centered as the NN command rates are changing. The limiter works by only allowing the window to move at a much slower rate than those of the NN commands. Anywhere within the window, however, full rates are allowed. If a rate persists in one direction, it will eventually hit the boundary and be rate-limited to the floating limiter rate. When this happens, a persistent counter begins and after a limit is reached, a NN disengage command is generated. The tunable metrics for the FL are (1) window size, (2) drift rate, and (3) persistence counter. Ultimate range limits are also included in case the NN command should drift slowly to a limit value that would cause the FL to be defeated. The FL has proven to work as intended. Both high-g transients and excessive structural loads are controlled with NN hard-over commands. This presentation discusses the FL design features and presents test cases. Simulation runs are included to illustrate the dramatic improvement made to the control of NN hard-over effects. A mission control room display from a flight playback is presented to illustrate the neural net fault display representation. The FL is very adaptable to various requirements and is independent of flight condition. It should be considered as a cost-effective safety monitor to control single-string inputs in general.
A display system for use in an aircraft control wheel steering system provides the pilot with a single, quickened flight path angle display to overcome poor handling qualities due to intrinsic flight path angle response lags, while avoiding multiple information display symbology. The control law for the flight path angle control system is designed such that the aircraft's actual flight path angle response lags the pilot's commanded flight path angle by a constant time lag .tau., independent of flight conditions. The synthesized display signal is produced as a predetermined function of the aircraft's actual flight path angle, the time lag .tau. and command inputs from the pilot's column.
The National Aeronautics and Space Administration (NASA) at the Dryden Flight Research Center (DFRC) has been conducting flight-test research using an F-15 aircraft (figure 1). This aircraft has been specially modified to interface a neural net (NN) controller as part of a single-string Airborne Research Test System (ARTS) computer with the existing quad-redundant flight control system (FCC) shown in figure 2. The NN commands are passed to FCC channels 2 and 4 and are cross channel data linked (CCDL) to the other computers as shown. Numerous types of fault-detection monitors exist in the FCC when the NN mode is engaged; these monitors would cause an automatic disengagement of the NN in the event of a triggering fault. Unfortunately, these monitors still may not prevent a possible NN hard-over command from coming through to the control laws. Therefore, an additional and unique safety monitor was designed for a single-string source that allows authority at maximum actuator rates but protects the pilot and structural loads against excessive g-limits in the case of a NN hard-over command input. This additional monitor resides in the FCCs and is executed before the control laws are computed. This presentation describes a "floating limiter" (FL) concept that was developed and successfully test-flown for this program (figure 3). The FL computes the rate of change of the NN commands that are input to the FCC from the ARTS. A window is created with upper and lower boundaries, which is constantly "floating" and trying to stay centered as the NN command rates are changing. The limiter works by only allowing the window to move at a much slower rate than those of the NN commands. Anywhere within the window, however, full rates are allowed. If a rate persists in one direction, it will eventually "hit" the boundary and be rate-limited to the floating limiter rate. When this happens, a persistent counter begins and after a limit is reached, a NN disengage command is generated. The tunable metrics for the FL are (1) window size, (2) drift rate, and (3) persistence counter. Ultimate range limits are also included in case the NN command should drift slowly to a limit value that would cause the FL to be defeated. The FL has proven to work as intended. Both high-g transients and excessive structural loads are controlled with NN hard-over commands. This presentation discusses the FL design features and presents test cases. Simulation runs are included to illustrate the dramatic improvement made to the control of NN hard-over effects. A mission control room display from a flight playback is presented to illustrate the neural net fault display representation. The FL is very adaptable to various requirements and is independent of flight condition. It should be considered as a cost-effective safety monitor to control single-string inputs in general.