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NASA NTRS1990

Development of microchannel plates in advanced wind-tunnel instrumentation

Microchannel plate (MCP) electron multiplier dynamic range has been increased 3 to 4 orders of magnitude at ambient temperatures, through enhanced input count rate capability and reduced background or 'dark' noise. The previous upper limit of roughly 10(exp 7) - 10(exp 8) cm(exp -2)s(exp -1) at ambient has been extended to levels approach 10(exp 10) cm(exp -2)s(exp -1) under continuous dc operation. The lower limit, previously set by an irreducible background component (approximately 0.6 cm(exp -2)s(exp -1)), has been lowered to the cosmic ray limit of .01 cm(exp -2)s(exp -1). The high end improvement was achieved by conductively cooling a very low resistance MCP by bonding it to a heat sink, while maintaining pulse-counting operation with multianode readouts. The low-end improvement was achieved by removing all radioisotopes from the MCP matrix glass. The detectors will benefit optical and mass spectrometry, flow visualization, plasma diagnostics, magnetometry, and other high signal flux applications. Very low MCP background noise will benefit X-ray and UV astronomy, medical imaging, trace isotope mass spectrometry, and other applications where the signal flux is often extremely low.

Feller, W. Bruce↗

NASA NTRS1993

The Effects of Gravity on Wrinkled Laminar Flames

The effects of gravity are significant to the dynamics of idealized unconfined open premixed flames. Moderate to low turbulence Reynolds number flames, i.e., wrinkled laminar flames, of various unconfined geometries have been used extensively for investigating fundamental processes of turbulent flame propagation and to validate theoretical models. Without the wall constraints, the flames are free to expand and interact with surrounding ambient air. The flow field in which the flame exists is determined by a coupling of burner geometry, flame orientation and the gravity field. These complex interactions raise serious questions regarding the validity of comparing the experimental data of open flames with current theoretical and numerical models that do not include the effects of gravity nor effects of the larger aerodynamic flowfield. Therefore, studies of wrinkled laminar flame in microgravity are needed for a better understanding of the role of gravity on flame characteristics such as the orientation, mean aerodynamics stretch, flame wrinkle size and burning rate. Our approach to characterize and quantify turbulent flame structures under microgravity is to exploit qualitative and quantitative flow visualization techniques coupled with video recording and computer controlled image analysis technologies. The experiments will be carried out in the 2.2 second drop tower at the NASA Lewis Research Center. The longest time scales of typical wrinkled laminar flames in the geometries considered here are in the order of 10 msec. Hence, the duration of the drop is sufficient to obtain the amount of statistical data necessary for characterize turbulent flame structures.

Kostiuk, Larry W.↗

NASA NTRS1992

Surface and flow field measurements in a symmetric crossing shock wave/turbulent boundary-layer interaction

Results of an experimental investigation of a symmetric crossing shock/turbulent boundary layer interaction are presented for a Mach number of 3.44 and deflection angles of 2, 6, 8, and 9 degrees. The interaction strengths vary from weak to strong enough to cause a large region of separated flow. Measured quantities include surface static pressure (both steady and unsteady) and flowfield Pitot pressures. Pitot profiles in the plane of symmetry through the interaction region are shown for various deflection angles. Oil flow visualization and the results of a trace gas streamline tracking technique are also presented.

Davis, D. O.↗

NASA NTRS1993

Neural networks for calibration tomography

Artificial neural networks are suitable for performing pattern-to-pattern calibrations. These calibrations are potentially useful for facilities operations in aeronautics, the control of optical alignment, and the like. Computed tomography is compared with neural net calibration tomography for estimating density from its x-ray transform. X-ray transforms are measured, for example, in diffuse-illumination, holographic interferometry of fluids. Computed tomography and neural net calibration tomography are shown to have comparable performance for a 10 degree viewing cone and 29 interferograms within that cone. The system of tomography discussed is proposed as a relevant test of neural networks and other parallel processors intended for using flow visualization data.

Decker, Arthur↗

NASA NTRS1993

Calibration of a shock wave position sensor using artificial neural networks

This report discusses the calibration of a shock wave position sensor. The position sensor works by using artificial neural networks to map cropped CCD frames of the shadows of the shock wave into the value of the shock wave position. This project was done as a tutorial demonstration of method and feasibility. It used a laboratory shadowgraph, nozzle, and commercial neural network package. The results were quite good, indicating that artificial neural networks can be used efficiently to automate the semi-quantitative applications of flow visualization.

Decker, Arthur J.↗

NASA NTRS1992

Hover performance validation of an ejector configured STOVL fighter aircraft

The hover thrust performance of a full-scale model of an ejector-lift/vectored thrust Short Take-Off and Vertical Landing (STOVL) fighter configuration was measured at the Outdoor Aerodynamic Research Facility of the National Full-Scale Aerodynamics Complex located at NASA Ames Research Center. The static thrust of the lifting ejector and ventral nozzle was calibrated at an out-of-ground-effect height. The test permitted comparison with static thrust calibrations obtained in the 80- by 120-Foot Wind Tunnel. The average static thrust augmentation ratio of the ejector system was 1.55 compared with the design goal of 1.63. The outdoor test showed that the addition of short extensions to the diffuser on the fuselage sidewall decreased suckdown and increased ejector performance. As expected, an increase in the ventral nozzle thrust deflection angle from nearly horizontal to the vertical-lift hover position caused a significant decrease in thrust. Flow visualization using smoke revealed greater secondary airflow into the left ejector and the penetration of hot ventral gases along the ground beneath and between the cool plumes from the ejector system.

Naumowicz, Tim↗

NASA NTRS1992

A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift.

Rao, Dhanvada M.↗

NASA NTRS1991

A high-frequency, secondary instability of crossflow vortices that leads to transition

Three-dimensional boundary-layer transition experiments are currently being conducted on a 45 deg swept wing in the Arizona State University Unsteady Wind Tunnel. Crossflow-dominated transition is produced via a model with contoured end liners to simulate infinite swept-wing flow. Fixed-wavelength, stationary and travelling crossflow vortices are observed. The frequencies of the most amplified travelling waves are in agreement with linear-stability theory; however travelling waves at frequencies an order of magnitude higher than predicted are also observed near transition. Near the transition location, the distorted boundary layer is due to the stationary crossflow vortex and is subject to a Rayleigh-type instability in the stream direction. As a result, a high-frequency secondary instability is detected in the transition region and spatial relations of the process are well documented by coupled use of flow visualization and hot-wire measurements.

Kohama, Yasuaki↗

NASA NTRS1992

Experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor flow field

An experimental and computational investigation of the NASA Low-Speed Centrifugal Compressor (LSCC) flow field has been conducted using laser anemometry and Dawes' 3D viscous code. The experimental configuration consists of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational analysis, and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the rotor as well as surface flow visualization along the impeller blade surfaces provide independent confirmation of the laser measurement technique.

Hathaway, M. D.↗

NASA NTRS1993

Effect of micron-sized roughness on transition in swept-wing flows

Boundary-layer transition-to-turbulence studies are conducted in the Arizona State University Unsteady Wind Tunnel on a 45-degree swept airfoil. The pressure gradient is designed so that the initial stability characteristics are purely crossflow-dominated. Flow visualization and hot-wire measurements show that the development of the crossflow vortices is influenced by roughness near the attachment-line. Comparisons of transition location are made between a painted surface, a machine-polished surface, and a hand-polished surface. Then, isolated 6 micron roughness elements are placed near the attachment line on the airfoil surface under conditions of the final polish (0.25 micron rms). These elements amplify a centered stationary crossflow vortex and its neighbors, resulting in localized early transition. The diameter, height, and location of these roughness elements are varied in a systematic manner. Spanwise hot-wire measurements are taken behind the roughness element to document the enhanced vortices. These scans are made at several different chord locations to examine vortex growth.

Radeztsky, Ronald H., Jr.↗

NASA NTRS1992

High-performance computing for flight vehicles; Proceedings of the Symposium, Washington, Dec. 7-9, 1992

The present conference discusses high-performance computing systems for flight vehicles, large-scale simulations on high-performance flight computers and software, multidisciplinary and design/optimization applications of computers, computational electromagnetics and acoustics, the simulation of aircraft powerplant turbomachinery and reacting flows, and flow calculations on parallel machines. Also discussed are direct flow simulation Monte Carlo methods, structural mechanics sensitivity and fracture calculations on parallel machines, grid-generation and advanced algorithms for CFD, advanced solid-mechanics and structures applications, and advancements in flow visualization technology and neural networks.

Noor, Ahmed K.↗

NASA NTRS1993

Measurements in the near-field of a turbulent wingtip vortex

The roll-up of a wingtip vortex, at Reynold's number based on chord of 4.6 million, was examined with emphasis on suction side and near wake measurements. Surface oil flow and laser smoke flow studies are performed, demonstrating the highly 3D nature of the flow around the wingtip, the formation of the feeding sheet, and the viscous core region of the vortex. Flow visualization and correlation measurements show that meandering of the vortex is small and does not appreciably contribute to the turbulence measurements.

Chow, Jim S.↗

NASA NTRS1992

Flight research on natural laminar flow applications

Natural laminar flow (NLF) is clearly one of the most potentially attractive drag reduction technologies by virtue of its relative simplicity. NLF is achieved passively, that is, by design of surface shapes to produce favorable pressure gradients. However, it is not without its challenges and limitations. This chapter describes the significant challenges to achieving and maintaining NLF and documents certain of the limitations for practical applications. A brief review of the history and of more recent NLF flight experiments is given, followed by a summary of lessons learned which are pertinent to future applications. The chapter also summarizes important progress in test techniques, particularly in flow visualization and hot-film techniques for boundary-layer measurements in flight.

Holmes, Bruce J.↗

NASA NTRS1993

Actuated forebody strake controls for the F-18 high alpha research vehicle

A series of ground-based studies have been conducted to develop actuated forebody strake controls for flight test evaluations using the NASA F-18 High-Alpha Research Vehicle. The actuated forebody strake concept has been designed to provide increased levels of yaw control at high angles of attack where conventional rudders become ineffective. Results are presented from tests conducted with the flight-test strake design, including static and dynamic wind-tunnel tests, transonic wind-tunnel tests, full-scale wind-tunnel tests, pressure surveys, and flow visualization tests. Results from these studies show that a pair of conformal actuated forebody strakes applied to the F-18 HARV can provide a powerful and precise yaw control device at high angles of attack. The preparations for flight testing are described, including the fabrication of flight hardware and the development of aircraft flight control laws. The primary objectives of the flight tests are to provide flight validation of the groundbased studies and to evaluate the use of this type of control to enhance fighter aircraft maneuverability.

Murri, Daniel G.↗

NASA NTRS1993

Flow field measurements in a crossing shockwave turbulent boundary layer interaction at Mach 3

An experimental study has been conducted to examine the flow field of the 3D crossing shock wave/turbulent boundary layer interaction. A symmetric pair of 9-deg fins were used to generate the crossing shocks. The incoming boundary layer was developed on the tunnel sidewall and thus was relatively thick, 0.49 arcsec, and suited for pitot probe surveys. The test conditions were a nominal Mach number of 3 and unit Reynolds number of 1.2 x 10 exp 7/ft. The measurements obtained included surface oil flow visualizations, surface static pressures, and boundary layer pitot pressure profiles. The results showed that downstream of the crossing shock intersection, the stagnation pressure losses were significant and the stagnation pressure profiles were highly nonuniform. Despite the severe shock disturbances, the law of the wall and the law of the wake were found to give relatively good agreement with the experimental data.

Lachowicz, Jason T.↗

NASA NTRS1993

A computational and experimental investigation of a delta wing with vertical tails

The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 deg and a Reynolds number of 500,000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

Krist, Sherrie L.↗

NASA NTRS1993

Lateral control at high angles of attack using pneumatic blowing through a chined forebody

Directional control through the use of pneumatic blowing was investigated on a generic subscale model with a chined forebody with blowing through a chine slot in a direction normal to the forebody surface. Comparisons are made with a vertical tail on and off, and with control through rudder deflection. Force and moment data were obtained for various blowing coefficients over a 0-75 deg alpha range, and flow visualization was also conducted in order to see qualitative effects on the flowfield. Blowing through a chined forebody generates yaw moments at large alpha where control surfaces lose their effectiveness; these moments are much larger than obtained by jet thrust alone, since the forebody flowfield is modified through the interaction of the jet with the chine vortices. Directional control increased with angle of attack for a given blowing coefficient until a maximum was reached. Further increases in angle of attack results in a rapid loss of effectiveness. For angles of attack above 60 deg, yaw moments are generated by simple jet thrust effect. The effectiveness of the pneumatic system depended on tail configuration.

Arena, A. S., Jr.↗

NASA NTRS1989

Numerical study of a delta planform with multiple jets in ground effect

The flow past a 60-deg delta wing equipped with two thrust-reverser jets near the inboard trailing edge has been analyzed by numerical solution of the 3D thin-layer Navier-Stokes equations. An implicit, partially flux-split, approximately-factored Navier-Stokes solver coupled with a multiple grid embedding scheme has been adapted to this problem. Studies of the impact of numerical parameters (e.g., grid refinement and dissipation levels), and flow-field parameters such as the height of the delta wing above the ground plane and the jet size on the solution, were performed. Results of these numerical studies indicate some challenges in the accurate resolution of complex 3D free shear layers and jets. Nevertheless, flow features such as jet deformation and ground vortex formation observed in experimental flow visualizations are captured. Further, comparisons with experimental data confirm the ability to simulate the loss of wing-borne lift, commonly referred to 'suckdown, as the delta planform flies at slow speeds in close proximity to the ground. Detailed analysis of the numerical results has also given additional insight into the structure of the ground vortex and the mechanisms of lift loss.

Chawla, K.↗