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Harris, C. D.

Publications and source records attributed to Harris, C. D..

At least 19 records

Investigation of heat transport using directly driven gold spheres

Recently, heat transport was investigated using a directly driven beryllium sphere [Farmer et al., Phys. Plasmas 27, 082701 (2020)]. Models that overly restrict heat transport were rejected. This paper extends work to directly driven gold spheres where radiation loss is more important. Here, gold coated spheres are directly driven at the OMEGA laser facility at intensities of 5×10 14 W/cm 2 . Plasma conditions, laser coupling, and x-ray flux are all measured. Additionally, comparisons to 2D radiation-hydrodynamic simulations are performed. Simulations use three common heat transport models: local transport with flux limiters of f = 0.15 and f = 0.03, and the nonlocal Schurtz–Nicolai–Busquet (SNB) model. It is shown that both the SNB model and f = 0.15 match the measured plasma conditions with the SNB model better capturing the temporal evolution of electron temperature. The f = 0.03 model predicts too low of an electron density and too hot of a temperature. The measured scattered light is roughly 6% of the incident energy, the f = 0.15 and SNB models predict 0.5% uncoupled light, and f = 0.03, 38% uncoupled light. The x-ray fluxes in the f = 0.15 and SNB simulations rise too quickly and are just outside the measurement's error, while the x-ray flux in the f = 0.03 simulation is low by a factor of two-three. For these reasons, the f = 0.03 model is rejected.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

Spectral sensor error analysis for measuring x-ray radiation drive using the DANTE diagnostic toward inertial confinement fusion experiments

DANTE is a diagnostic used to measure the x-radiation drive produced by heating a high-Z cavity (“hohlraum”) with high-powered laser beams. It records the spectrally and temporally resolved radiation flux at x-ray energies between 50 eV and 20 keV. Each sensor configuration on DANTE is composed of filters, mirrors, and x-ray diodes to define 18 different x-ray channels whose output is voltage as a function of time. The absolute flux is then determined from the photometric calibration of the sensor configuration and a spectral reconstructing algorithm. The reconstruction of the spectra vs time from the measured voltages and known response of each channel has presented challenges. Here we demonstrate a novel approach here for quantifying the error on the determined flux based on the channel sensor configuration and most commonly used reconstruction algorithm. In general, we find that the integrated spectral flux from a hohlraum can robustly be reconstructed (within ~14%) using a traditional unfold approach with as few as ten channels due to the underlying assumption of a largely Planckian spectral intensity distribution.

47 OTHER INSTRUMENTATION↗

A genetic algorithm approach for reconstructing spectral content from filtered x-ray diode array spectrometers

Filtered diode array spectrometers are routinely employed to infer the temporal evolution of spectral power from x-ray sources, but uniquely extracting spectral content from a finite set of broad, spectrally overlapping channel spectral sensitivities is decidedly nontrivial in these under-determined systems. Here, we present the use of genetic algorithms to reconstruct a probabilistic spectral intensity distribution and compare to the traditional approach most commonly found in the literature. Unlike many of the previously published models, spectral reconstructions from this approach are neither limited by basis functional forms nor do they require a priori spectral knowledge. While the original intent of such measurements was to diagnose the temporal evolution of spectral power from quasi-blackbody radiation sources—where the exact details of spectral content were not thought to be crucial—we demonstrate that this new technique can greatly enhance the utility of the diagnostic by providing more physical spectra and improved robustness to hardware configuration for even strongly non-Planckian distributions.

47 OTHER INSTRUMENTATION↗

Design and experimental evaluation of a swept supercritical Laminar Flow Control (LFC) airfoil

A large chord swept supercritical laminar flow control (LFC) airfoil was designed, constructed, and tested in the NASA Langley 8-ft Transonic Pressure Tunnel (TPT). The LFC airfoil experiment was established to provide basic information concerning the design and compatibility of high-performance supercritical airfoils with suction boundary layer control achieved through discrete fine slots or porous surface concepts. It was aimed at validating prediction techniques and establishing a technology base for future transport designs and drag reduction. Good agreement was obtained between measured and theoretically designed shockless pressure distributions. Suction laminarization was maintained over an extensive supercritical zone up to high Reynolds numbers before transition gradually moved forward. Full-chord laminar flow was maintained on the upper and lower surfaces at M sub infinity = 0.82 up to R sub c is less than or equal to 12 x 10 to the 6th power. When accounting for both the suction and wake drag, the total drag could be reducted by at least one-half of that for an equivalent turbulent airfoil. Specific objectives for the LFC experiment are given.

Harvey, W. D.↗

Wind tunnel testing of low-drag airfoils

Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.

Harvey, W. Donald↗

Two-dimensional aerodynamic characteristics of the NACA 0012 airfoil in the Langley 8 foot transonic pressure tunnel

Data are presented for lift coefficients from near zero through maximum values at Mach numbers from 0.30 to 0.86 and Reynolds numbers of 3.0 x 10 to the sixth power with transition fixed. A limited amount of data is presented near zero and maximum lift for a Reynolds number of 6.0 x 10 to the sixth power with transition fixed. In addition, transition free data is presented through the Mach number range from 0.30 to 0.86 for near zero lift and a Reynolds number of 3.0 x 10 to the sixth power.

Harris, C. D.↗

Low-speed aerodynamic characteristics of a 14-percent-thick NASA phase 2 supercritical airfoil designed for a lift coefficient of 0.7

The low speed aerodynamic characteristics of a 14 percent thick supercritical airfoil are documented. The wind tunnel test was conducted in the Low Turbulence Pressure Tunnel. The effects of varying chord Reynolds number from 2,000,000 to 18,000,000 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6,000,000 are included.

Harris, C. D.↗

Comparison of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an analysis code

Comparisons of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an advanced analysis code. Comparisons are made at near-design normal-force coefficients for Reynolds numbers from 2 to 11 million. Comments are made concerning various input parameters to the code.

Harris, C. D.↗

Aerodynamic Characteristics of a 14-Percent-Thick NASA Supercritical Airfoil Designed for a Normal-Force Coefficient of 0.7

This report documents the experimental aerodynamic characteristics of a 14 percent thick supercritical airfoil based on an off design sonic pressure plateau criterion. The design normal force coefficient was 0.7. The results are compared with those of the family related 10 percent thick supercritical airfoil 33. Comparisons are also made between experimental and theoretical characteristics and composite drag rise characteristics derived for a full scale Reynolds number of 40 million.

Harris, C. D.↗

Transonic aerodynamic characteristics of the 10-percent-thick NASA supercritical airfoil 31

Refinements in a 10 percent thick supercritical airfoil (airfoil 31) have produced significant improvements in the drag characteristics compared with those for an earlier supercritical airfoil (airfoil 12) designed for the same normal force coefficient of 0.7. Drag creep was practically eliminated at normal force coefficients between about 0.4 and 0.7 and was greatly reduced at other normal force coefficients. Substantial reductions in the drag levels preceding drag divergence were also achieved at all normal force coefficients. The Mach numbers at which drag diverges were delayed for airfoil 31 at normal force coefficients up to about 0.6 (by approximately 0.01 and 0.02 at normal force coefficients of 0.4 and 0.6, respectively) but drag divergence occurred at slightly lower Mach numbers at higher normal force coefficients.

Harris, C. D.↗

Aerodynamic characteristics of 10 percent thick NASA supercritical airfoils with different aft camber

The aerodynamic characteristics of several supercritical airfoils interim to the improved 10-percent thick NASA supercritical airfoil 26a are discussed. The airfoils have related slope and curvature distributions over the rear which result in different aft camber. For identification, the airfoils are designated supercritical airfoils 12, 13, 21, 22, and 24. Data is presented without analysis.

Harris, C. D.↗

Aerodynamic characteristics of the 10-percent-thick NASA supercritical airfoil 33 designed for a normal-force coefficient of 0.7

A 10-percent-thick supercritical airfoil based on an off-design sonic-pressure plateau criterion was developed and experimental aerodynamic characteristics measured. The airfoil had a design normal-force coefficient of 0.7 and was identified as supercritical airfoil 33. Results show the airfoil to have good drag rise characteristics over a wide range of normal-force coefficients with no measurable shock losses up to the Mach numbers at which drag divergence occurred for normal-force coefficients up to 0.7. Comparisons of experimental and theoretical characteristics were made and composite drag rise characteristics were derived for normal-force coefficients of 0.5 and 0.7 and a Reynolds number of 40 million.

Harris, C. D.↗

Wind-tunnel investigation of aerodynamic load distribution on a variable-wing-sweep fighter airplane with a NASA supercritical airfoil

Wind-tunnel tests have been conducted at Mach numbers of 0.85, 0.88, and 0.90 to determine the aerodynamic load distribution for the 39 deg swept-wing configuration of a variable-wing-sweep fighter airplane with a NASA supercritical airfoil. Chordwise pressure distributions were measured at two wing stations. Also measured were the overall longitudinal aerodynamic force and moment characteristics and the buffet characteristics. The analysis indicates that localized regions of shock-induced flow separation may exist on the rearward portions of the supercritical wing at high subsonic speeds, and caution must be exercised in the prediction of buffet onset when using variations in trailing-edge pressure coefficients at isolated locations.

Hallissy, J. B.↗

Effects of wing trailing-edge truncation on aerodynamic characteristics of a NASA supercritical-wing research airplane model

The Langley 8-foot transonic pressure tunnel was used at Mach numbers from 0.80 to 1.00 to determine the effects of wing trailing-edge truncation on the aerodynamic characteristics of a 0.0625-scale model of a NASA TF-8A supercritical-wing research airplane. The effects of trailing-edge truncations of 1, 2, and 3 percent of the local streamwise chord on the longitudinal aerodynamic characteristics and the wing section characteristics are presented.

Bartlett, D. W.↗

Comparison of the experimental aerodynamic characteristics of theoretically and experimentally designed supercritical airfoils

A lifting airfoil theoretically designed for shockless supercritical flow utilizing a complex hodograph method has been evaluated in the Langley 8-foot transonic pressure tunnel at design and off-design conditions. The experimental results are presented and compared with those of an experimentally designed supercritical airfoil which were obtained in the same tunnel.

Harris, C. D.↗

Aerodynamic characteristics of two 10-percent-thick NASA supercritical airfoils with different upper surface curvature distributions

In order to assess the degree to which the characteristic region of low curvature of the supercritical airfoil can be practically extended on the upper surface, the aerodynamic characteristics of two supercritical airfoils with different upper surface curvature distributions were measured at Mach numbers from 0.60 to 0.81. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.

Harris, C. D.↗

Aerodynamic characteristics of an improved 10-percent-thick NASA supercritical airfoil

Refinements in a 10 percent thick supercritical airfoil produced improvements in the overall drag characteristics at normal force coefficients from about 0.30 to 0.65 compared with earlier supercritical airfoils which were developed for a normal force coefficient of 0.7. The drag divergence Mach number of the improved supercritical airfoil (airfoil 26a) varied from approximately 0.82 at a normal force coefficient to of 0.30, to 0.78 at a normal force coefficient of 0.80 with no drag creep evident. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.

Harris, C. D.↗

Wind-tunnel development of underwing leading-edge vortex generators on a NASA supercritical-wing research airplane configuration

A program for the wind tunnel development of underwing leading-edge vortex generators on an NASA supercritical-wing research airplane model was conducted in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.25 to 0.99. The effects on the longitudinal aerodynamic characteristics of vortex-generator wing semispan location, distance from wing leading edge, leading-edge sweep angle, toe-in angle, and span are presented. The effects of the vortex generators on the lateral-directional aerodynamic characteristics of the model for a sideslip angle of 5 deg are included.

Bartlett, D. W.↗