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At least 847 records · Page 47
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Experimental investigation of the inlet detector configuration variation in the flow field at Mach 1.9

In recent years, active research has been conducted to study the technological feasibility of supersonic laminar flow control on the wing of the High Speed Civil Transport (HSCT). For this study, the F-16XL has been chosen due to its highly swept crank wing planform that closely resembles the HSCT configurations. During flights, it is discovered that the shock wave generated from the aircraft inlet introduces disturbances on the wing where the data acquisition is conducted. The flow field about a supersonic inlet is characterized by a complex three dimensional pattern of shock waves generated by the geometrical configuration of a deflector and a cowl lip. Hence, in this study, experimental method is employed to investigate the effects of the variation of deflector configuration on the flow field, and consequently, the possibility of diverting the incoming shock-disturbances away from the test section. In the present experiments, a model composed of a simple circular tube with a triangular deflector is designed to study the deflector length and the deflector base width variation in the flow field. Experimental results indicate that the lowest external pressure ratio is observed at the junction where the deflector lip and the inlet cowl lip merge. Also, it is noted that the external pressure ratio, the internal pressure ratio, the coefficient of spillage drag, and the shock standoff distance decrease as the deflector length increases. In addition, the Redefined Total Pressure Recovery Ratio (RTPRR) increases with an increase in the deflector length. Results from the study of the effect of the deflector's base width variation on the flow field indicate that the lowest external pressure ratio is observed at the junction between the inlet cowl lip and the deflector lip. As the base width of the deflector increases, the external pressure ratio at 0 rotation increases, whereas the external pressure ratio at 180 rotation decreases. In addition, the internal pressure ratio and the coefficient of spillage drag decrease as the base width of the deflector increases. However, RTPRR and shock standoff distance increase as the base width increases. In conclusion, as deflector dimensions vary, distinctive patterns in the pressure variation around the inlet deflector are observed. With an increase in the deflector length and base width, the magnitude of shock-disturbances are weakened due to a decrease in the external pressure ratio. Also, as the deflector length and base width increase, a smaller bow shock angle is formed. Therefore, the inlet shock wave formation would be significantly altered, and consequently, shock disturbances on the wing test section can be avoided through appropriately designing the deflector.

Hwang, Kyu C.↗
NASA NTRS1995

Experimental Results from the Thermal Energy Storage-1 (TES-1) Flight Experiment

The Thermal Energy Storage-1 (TES-1) is a flight experiment that flew on the Space Shuttle Columbia (STS-62), in March 1994, as part of the OAST-2 mission. TES-1 is the first experiment in a four experiment suite designed to provide data for understanding the long duration microgravity behavior of thermal energy storage fluoride salts that undergo repeated melting and freezing. Such data have never been obtained before and have direct application for the development of space-based solar dynamic (SD) power systems. These power systems will store solar energy in a thermal energy salt such as lithium fluoride or calcium fluoride. The stored energy is extracted during the shade portion of the orbit. This enables the solar dynamic power system to provide constant electrical power over the entire orbit. Analytical computer codes have been developed for predicting performance of a spaced-based solar dynamic power system. Experimental verification of the analytical predictions is needed prior to using the analytical results for future space power design applications. The four TES flight experiments will be used to obtain the needed experimental data. This paper will focus on the flight results from the first experiment, TES-1, in comparison to the predicted results from the Thermal Energy Storage Simulation (TESSIM) analytical computer code. The TES-1 conceptual development, hardware design, final development, and system verification testing were accomplished at the NASA lewis Research Center (LeRC). TES-1 was developed under the In-Space Technology Experiment Program (IN-STEP), which sponsors NASA, industry, and university flight experiments designed to enable and enhance space flight technology. The IN-STEP Program is sponsored by the Office of Space Access and Technology (OSAT).

Wald, Lawrence W.↗
NASA NTRS1996

Verification of Experimental Techniques for Flow Surface Determination

The concept of a yield surface is central to the mathematical formulation of a classical plasticity theory. However, at elevated temperatures, material response can be highly time-dependent, which is beyond the realm of classical plasticity. Viscoplastic theories have been developed for just such conditions. In viscoplastic theories, the flow law is given in terms of inelastic strain rate rather than the inelastic strain increment used in time-independent plasticity. Thus, surfaces of constant inelastic strain rate or flow surfaces are to viscoplastic theories what yield surfaces are to classical plasticity. The purpose of the work reported herein was to validate experimental procedures for determining flow surfaces at elevated temperatures. Since experimental procedures for determining yield surfaces in axial/torsional stress space are well established, they were employed -- except inelastic strain rates were used rather than total inelastic strains. In yield-surface determinations, the use of small-offset definitions of yield minimizes the change of material state and allows multiple loadings to be applied to a single specimen. The key to the experiments reported here was precise, decoupled measurement of axial and torsional strain. With this requirement in mind, the performance of a high-temperature multi-axial extensometer was evaluated by comparing its results with strain gauge results at room temperature. Both the extensometer and strain gauges gave nearly identical yield surfaces (both initial and subsequent) for type 316 stainless steel (316 SS). The extensometer also successfully determined flow surfaces for 316 SS at 650 C. Furthermore, to judge the applicability of the technique for composite materials, yield surfaces were determined for unidirectional tungsten/Kanthal (Fe-Cr-Al).

Lissenden, Cliff J.↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges: Sharp Leading Edge - Volume 1

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Experimental study of vortex breakdown in a cylindrical, swirling flow

The stability of a steady, vortical flow in a cylindrical container with one rotating endwall has been experimentally examined to gain insight into the process of vortex breakdowwn. The dynamics of the flow are governed by the Reynolds number (Re) and the aspect ratio of the cylinder. Re is given by Omega R(sup 2)/nu, where Omega is the speed of rotation of the endwall, R is the cylinder radius, and nu is the kinematic viscosity of the fluid filling the cylinder. The aspect ratio is H/R, where H is the height of the cylinder. Numerical simulation studies disagree whether or not the steady breakdown is stable beyond a critical Reynolds number, Re(sub c). Previous experimental researches have considered the steady and unsteady flows near Re(sub c), but have not explored the stability of the steady breakdown structures beyond this value. In this investigation, laser induced fluorescence was utilized to observe both steady and unsteady vortex breakdown at a fixed H/R of 2.5 with Re varying around Re(sub c). When the Re of a steady flow was slowly increased beyond Re(sub c), the breakdown structure remained steady even though unsteadiness was possible. In addition, a number of hysteresis events involving the oscillation periods of the unsteady flow were noted. The results show that both steady and unsteady vortex breakdown occur for a limited range of Re above Re(sub c). Also, with increasing Re, complex flow transformations take place that alter the period at which the unsteady flow oscillates.

Stevens, J. L.↗
NASA NTRS1996

Experimental and computational investigation of lift-enhancing tabs on a multi-element airfoil

An experimental and computational investigation of the effect of lift enhancing tabs on a two-element airfoil was conducted. The objective of the study was to develop an understanding of the flow physics associated with lift enhancing tabs on a multi-element airfoil. A NACA 63(sub 2)-215 ModB airfoil with a 30 percent chord Fowler flap was tested in the NASA Ames 7 by 10 foot wind tunnel. Lift enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computer results predict all of the trends in the experimental data quite well. When the flow over the flap upper surface is attached, tabs mounted at the main element trailing edge (cove tabs) produce very little change in lift. At high flap deflections. however, the flow over the flap is separated and cove tabs produce large increases in lift and corresponding reductions in drag by eliminating the separated flow. Cove tabs permit high flap deflection angles to be achieved and reduce the sensitivity of the airfoil lift to the size of the flap gap. Tabs attached to the flap training edge (flap tabs) are effective at increasing lift without significantly increasing drag. A combination of a cove tab and a flap tab increased the airfoil lift coefficient by 11 percent relative to the highest lift tab coefficient achieved by any baseline configuration at an angle of attack of zero percent and the maximum lift coefficient was increased by more than 3 percent. A simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift enhancing tabs work. The tabs were modeled by a point vortex at the training edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift enhancing tabs on a multi-element airfoil. Results of the modeling effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

Ashby, Dale↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges: Small-Radius Leading Edge - Volume 2

An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

An Experimental Study of Richtmyer-Meshkov Instability

Richtmyer-Meshkov (RM) instability occurs when a planar interface separating two fluids of different density is impulsively accelerated In the direction of its normal. It is one of the most fundamental of fluid instabilities and is of importance in fields ranging from astrophysics to materials processing. Because RM instability experiments are normally carried out in shock tubes, where the generation of a sharp well controlled interface between gases is difficult, there is a scarcity of good experimental results. The experiments presented here utilize a novel technique which circumvents many of the experimental difficulties that have previously limited the study of RM instability. In this system, the instability is generated by bouncing a thin rectangular tank containing two liquids off of a fixed spring. Results obtained from these experiments yield particularly well visualized images of the nonlinear development of the instability. However, because the run time in these experiments is limited, new experiments capable of achieving longer run times are planned.

Jacobs, Jeffrey W.↗
NASA NTRS1996

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions. The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row. The blockage is evaluated in the core flow area as well as in the casing endwall region. Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed. Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed. In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis. The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer. The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer. However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss.

Suder, Kenneth L.↗
NASA NTRS1995

Active Flow Control: Instrumentation Automation and Experimental Technique

In investigating the potential of a new actuator for use in an active flow control system, several objectives had to be accomplished, the largest of which was the experimental setup. The work was conducted at the NASA Langley 20x28 Shear Flow Control Tunnel. The actuator named Thunder, is a high deflection piezo device recently developed at Langley Research Center. This research involved setting up the instrumentation, the lighting, the smoke, and the recording devices. The instrumentation was automated by means of a Power Macintosh running LabVIEW, a graphical instrumentation package developed by National Instruments. Routines were written to allow the tunnel conditions to be determined at a given instant at the push of a button. This included determination of tunnel pressures, speed, density, temperature, and viscosity. Other aspects of the experimental equipment included the set up of a CCD video camera with a video frame grabber, monitor, and VCR to capture the motion. A strobe light was used to highlight the smoke that was used to visualize the flow. Additional effort was put into creating a scale drawing of another tunnel on site and a limited literature search in the area of active flow control.

Gimbert, N. Wes↗
NASA NTRS1995

Experimental Verification of a Progressive Damage Model for IM7/5260 Laminates Subjected to Tension-Tension Fatigue

The durability and damage tolerance of laminated composites are critical design considerations for airframe composite structures. Therefore, the ability to model damage initiation and growth and predict the life of laminated composites is necessary to achieve structurally efficient and economical designs. The purpose of this research is to experimentally verify the application of a continuum damage model to predict progressive damage development in a toughened material system. Damage due to monotonic and tension-tension fatigue was documented for IM7/5260 graphite/bismaleimide laminates. Crack density and delamination surface area were used to calculate matrix cracking and delamination internal state variables to predict stiffness loss in unnotched laminates. A damage dependent finite element code predicted the stiffness loss for notched laminates with good agreement to experimental data. It was concluded that the continuum damage model can adequately predict matrix damage progression in notched and unnotched laminates as a function of loading history and laminate stacking sequence.

Coats, Timothy W.↗
NASA NTRS1995

Experimental Studies on Mass Transport of Cadmium-Zinc Telluride by Physical Vapor Transport

Experimental studies on mass transport of ternary compound, Cd(1-x)Zn(x)Te by physical vapor transport (PVT) for source compositions up to X = 0.21 are presented. The effect of thermochemical (temperatures, vapor composition) and other factors (preparation of the source, crystal growth rate, temperature gradient) on composition and composition profiles of the grown crystals were investigated. A steep decrease in the mass flux with an increase in X(crystal) for X less than 0.1, and a difference in composition between the source and the deposited material have been observed. The composition profiles of the crystals were found to depend on the density and pretreatment of the source, and on the temperature gradient in the source zone. The homogeneity of the crystals improves at low undercoolings and/or when an appropriate excess of metal constituents is present in the vapor phase. The experimental results are in good agreement with our thermochemical model of this system.

Palosz, W.↗
NASA NTRS1996

A Computational-Experimental Development of Vortex Generator Use for a Transitioning S-Diffuser

The development of an effective design strategy for surface-mounted vortex generator arrays in a subsonic diffuser is described in this report. This strategy uses the strengths of both computational and experimental analyses to determine beneficial vortex generator locations and sizes. A parabolized Navier-Stokes solver, RNS3D, was used to establish proper placement of the vortex generators for reduction in circumferential total pressure distortion. Experimental measurements were used to determine proper vortex generator sizing to minimize total pressure recovery losses associated with vortex generator device drag. The best result achieved a 59% reduction in the distortion index DC60, with a 0.3% reduction in total pressure recovery.

Wendt, Bruce J.↗
NASA NTRS1996

Experimental and Computational Investigation of Lift-Enhancing Tabs on a Multi-Element Airfoil

An experimental and computational investigation of the effect of lift-enhancing tabs on a two-element airfoil has been conducted. The objective of the study was to develop an understanding of the flow physics associated with lift-enhancing tabs on a multi-element airfoil. An NACA 63(2)-215 ModB airfoil with a 30% chord fowler flap was tested in the NASA Ames 7- by 10-Foot Wind Tunnel. Lift-enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. A combination of tabs located at the main element and flap trailing edges increased the airfoil lift coefficient by 11% relative to the highest lift coefficient achieved by any baseline configuration at an angle of attack of 0 deg, and C(sub 1max) was increased by 3%. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computed results predicted all of the trends observed in the experimental data quite well. In addition, a simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift-enhancing tabs work. The tabs were modeled by a point vortex at the air-foil or flap trailing edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift-enhancing tabs on a multi-element airfoil. Results of the modeling effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

Ashby, Dale L.↗
NASA NTRS1997

An Experimental Study of Swirling Flows as Applied to Annular Combustors

This thesis presents an experimental study of swirling flows with direct applications to gas turbine combustors. Two separate flowfields were investigated: a round, swirling jet and a non-combusting annular combustor model. These studies were intended to allow both a further understanding of the behavior of general swirling flow characteristics, such as the recirculation zone, as well as to provide a base for the development of computational models. In order to determine the characteristics of swirling flows the concentration fields of a round, swirling jet were analyzed for varying amount of swirl. The experimental method used was a light scattering concentration measurement technique known as marker nephelometry. Results indicated the formation of a zone of recirculating fluid for swirl ratios (rotational speed x jet radius over mass average axial velocity) above a certain critical value. The size of this recirculation zone, as well as the spread angle of the jet, was found to increase with increase in the amount of applied swirl. The annular combustor model flowfield simulated the cold-flow characteristics of typical current annular combustors: swirl, recirculation, primary air cross jets and high levels of turbulence. The measurements in the combustor model made by the Laser Doppler Velocimetry technique, allowed the evaluation of the mean and rms velocities in the three coordinate directions, one Reynold's shear stress component and the turbulence kinetic energy: The primary cross jets were found to have a very strong effect on both the mean and turbulence flowfields. These cross jets, along with a large step change in area and wall jet inlet flow pattern, reduced the overall swirl in the test section to negligible levels. The formation of the strong recirculation zone is due mainly to the cross jets and the large step change in area. The cross jets were also found to drive a four-celled vortex-type motion (parallel to the combustor longitudinal axis) near the cross jet injection plane.

Seal, Michael Damian, II↗
NASA NTRS1997

Experimental Observations of PMMA Spheres Burning at Reduced Gravity

Polymer combustion is a very complicated process which involves the coupling of gas-phase combustion, melting, pyrolysis, and possible charring of the condensed phase. Although only a few studies on the combustion of a spherically shaped polymer have been conducted, there is renewed interest in the subject. At the Third International Microgravity Combustion Workshop, we presented a preliminary experimental plan and proposed an apparatus to study the combustion of a PolyMethylMethAcrylate (PMMA) sphere at reduced gravity. In this paper, we describe the experimental hardware in detail, summarize our observations since the last workshop, and describe future studies. The main objective is to determine the burning rate of PMMA spheres at reduced gravity under different ambient oxygen concentrations and total pressures. The dependence of the burning rate on the initial sphere diameter will be examined. It is anticipated that the simple spherical geometry in conjunction with reduced gravity will facilitate an assessment of the effect of condensed phase behavior on polymer burning processes and can be used as a means to test polymer combustion models.

Yang, Ji-Ann C.↗