SEARCH · Engineering Papers
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
A comprehensive analytical, numerical and experimental study of convective flow in multilayer immiscible liquids has been performed. Studies include transparent high Prandtl number of liquid as well as opaque low Prandtl metallic melts. A new radioscopic flow visualization has been developed for the latter studies.
The transitional flight characteristics of a geometrically simplified Short Take-Off Vertical Landing (STOVL) aircraft configuration have been measured in the NASA Ames 7- by 10-Foot Wind Tunnel. The experiment is the first in a sequence of tests designed to provide detailed data for evaluating the capability of computational fluid dynamics methods to predict the important flow parameters for powered lift. The model consists of a 60 deg cropped delta wing platform, blended fuselage and two circular in-line jets that exit perpendicularly from the flat lower surface. The measured flows have a maximum freestream Mach number of 0.2. Model angle of attack is varied between -10 and +20 deg. The flow is ambient temperature in both jet exits and the nozzle pressure ratios are varied between 1 and 3. The data presented includes forces and moments, pressures measured at 281 surface pressure ports and the pressures of the jets. Measurements of the flow are also made in the tunnel test section upstream and downstream of the model and at the jet exits to guide boundary condition selection for the planned computations. Flow visualization and total pressure measurements in the jet plumes provide a description of the three-dimensional jet efflux flowfield.
The principal objective of this research effort was to develop a multicomponent strain gauge balance to measure forces and moments on models tested in flow visualization water tunnels. An internal balance was designed that allows measuring normal and side forces, and pitching, yawing and rolling moments (no axial force). The five-components to applied loads, low interactions between the sections and no hysteresis. Static experiments (which are discussed in this Volume) were conducted in the Eidetics water tunnel with delta wings and a model of the F/A-18. Experiments with the F/A-18 model included a thorough baseline study and investigations of the effect of control surface deflections and of several Forebody Vortex Control (FVC) techniques. Results were compared to wind tunnel data and, in general, the agreement is very satisfactory. The results of the static tests provide confidence that loads can be measured accurately in the water tunnel with a relatively simple multicomponent internal balance. Dynamic experiments were also performed using the balance, and the results are discussed in detail in Volume 2 of this report.
The use of sub-scale wind tunnel test data to predict the behavior of commercial transport high lift systems at in-flight Reynolds number is limited by the so-called 'inverse Reynolds number effect'. This involves an actual deterioration in the performance of a high lift device with increasing Reynolds number. A lack of understanding of the relevant flow field physics associated with numerous complicated viscous flow interactions that characterize flow over high-lift devices prohibits computational fluid dynamics from addressing Reynolds number effects. Clearly there is a need for research that has as its objective the clarification of the fundamental flow field physics associated with viscous effects in high lift systems. In this investigation, a detailed experimental investigation is being performed to study the interaction between the slat wake and the boundary layer on the primary airfoil which is known as a confluent boundary layer. This little-studied aspect of the multi-element airfoil problem deserves special attention due to its importance in the lift augmentation process. The goal of this research is is to provide an improved understanding of the flow physics associated with high lift generation. This process report will discuss the status of the research being conducted at the Hessert Center for Aerospace Research at the University of Notre Dame. The research is sponsored by NASA Ames Research Center under NASA grant NAG2-905. The report will include a discussion of the models that have been built or that are under construction, a description of the planned experiments, a description of a flow visualization apparatus that has been developed for generating colored smoke for confluent boundary layer studies and some preliminary measurements made using our new 3-component fiber optic LDV system.
Detailed three-dimensional three-component phase averaged measurements of the spanwise and streamwise vorticity formation and evolution in acoustically forced plane free-shear flows have been obtained. For the first time, phase-averaged measurements of all three velocity components have been obtained in both a mixing layer and a wake on three-dimensional grids, yielding the spanwise and streamwise vorticity distributions without invoking Taylor's hypothesis. Initially, two-frequency forcing was used to phase-lock the roll-up and first pairing of the spanwise vortical structures in a plane mixing layer. The objective of this study was to measure the near-field vortical structure morphology in a mixing layer with 'natural' laminar initial boundary layers. For the second experiment the second and third subharmonics of the fundamental roll-up frequency were added to the previous two-frequency forcing in order to phase-lock the roll-up and first three pairings of the spanwise rollers in the mixing layer. The objective of this study was to determine the details of spanwise scale changes observed in previous time-averaged measurements and flow visualization of unforced mixing layers. For the final experiment, single-frequency forcing was used to phase-lock the Karman vortex street in a plane wake developing from nominally two-dimensional laminar initial boundary layers. The objective of this study was to compare measurements of the three-dimensional structure in a wake developing from 'natural' initial boundary layers to existing models of wake vortical structure.
A low-speed multistage axial compressor built by the NASA Lewis Research Center is described. The purpose of this compressor is to increase the understanding of the complex flow phenomena within multistage axial compressors and to obtain detailed data from a multistage compressor environment for use in developing and verifying models for computational fluid dynamic code assessment. The compressor has extensive pressure instrumentation in both stationary and rotating frames of reference, and has provisions for flow visualization and laser velocimetry. The compressor will accommodate rotational speeds to 1050 rpm and is rated at a pressure ratio of 1.042.
Results from computational fluid dynamic analyses for complex three-dimensional internal flows in the Advanced Solid Rocket Motor (ASRM) are presented. In particular, flow visualization and tabulated results from a fine-grid model consisting of 1.5 M grid points for the ASRM Aft Segment at the l9-second burn time with an 8-degree nozzle gimbal angle are shown. The results from this model will enable characterization of various aspects of the ASRM internal environment, and in particular will allow an assessment of the heat transfer and stresses exerted on the submerged nozzle, casing insulation, and nozzle case joint.
This report includes copies of relevant publications, including conference proceedings, resulting from studies funded by NAG3-1010. Topics include flow visualization, LDV (laser doppler velocimeter), straight and curved simulations, and kinetics modeling of nitrogen oxides, including a turbulent mixing model for O2 and a maximum mixedness model.
Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.
A study of a full-scale, semi-span business jet wing has been conducted to investigate the potential of two types of high-lift devices for improving aircraft high-lift performance. The research effort involved low-speed wind-tunnel tests of micro-vortex generators and Gurney flaps applied to the flap system of the business jet wing and included force and moment measurements, surface pressure surveys and flow visualization on the wing and flap. Results showed that the micro-vortex generators tested had no beneficial effects on the longitudinal force characteristics in this particular application, while the Gurney flaps were an effective means of increasing lift. However, the Gurney flaps also caused an increase in drag in most circumstances.
Low-speed wind-tunnel tests were conducted in the Langley 12-Foot Low-Speed Tunnel on a model of the Boeing Multirole Fighter (BMRF) aircraft. This single-seat, single-engine configuration was intended to be an F-16 replacement that would incorporate many of the design goals and advanced technologies of the F-22. Its mission requirements included supersonic cruise without afterburner, reduced observability, and the ability to attack both air-to-air and air-to-ground targets. So that it would be effective in all phases of air combat, the ability to maneuver at angles of attack up to and beyond maximum lift was also desired. Traditional aerodynamic yaw controls, such as rudders, are typically ineffective at these higher angles of attack because they are usually located in the wake from the wings and fuselage. For this reason, this study focused on investigating forebody-mounted controls that produces yawing moments by modifying the strong vortex flowfield being shed from the forebody at high angles of attack. Two forebody strakes were tested that varied in planform and chordwise location. Various patterns of porosity in the forebody skin were also tested that differed in their radial coverage and chordwise location. The tests were performed at a dynamic pressure of 4 lb/ft(exp 2) over an angle-of-attack range of -4 deg to 72 deg and a sideslip range of -10 deg to 10 deg. Static force data, static pressures on the surface of the forebody, and videotapes of flow-visualization using laser-illuminated smoke were obtained.
To identify planform characteristics which have promise for a highly maneuverable vehicle, an investigation was conducted in the Langley Subsonic Basic Research Tunnel to determine the low-speed longitudinal aerodynamics of 21 planform geometries. Concepts studied included twin bodies, double wings, cutout wings, and serrated forebodies. The planform models tested were all 1/4-in.-thick flat plates with beveled edges on the lower surface to ensure uniform flow separation at angle of attack. A 1.0-in.-diameter cylindrical metric body with a hemispherical nose was used to house the six-component strain gauge balance for each configuration. Aerodynamic force and moment data were obtained across an angle-of-attack range of 0 to 70 deg with zero sideslip at a free-stream dynamic pressure of 30 psf. Surface flow visualization studies were also conducted on selected configurations using fluorescent minitufts. Results from the investigation indicate that a cutout wing planform can improve lift characteristics; however, cutout size, shape, and position and wing leading-edge sweep will all influence the effectiveness of the cutout configuration. Tests of serrated forebodies identified this concept as an extremely effective means of improving configuration lift characteristics; increases of up to 25 percent in the value of maximum lift coefficient were obtained.
The Boeing-Vertol VR-7 airfoil was experimentally studied with steady and pulsed upper-surface blowing for sinusoidal pitching oscillations described by alpha = alpha(sub m) + 10 deg sin(omega t). The tests were conducted in the U.S. Army Aeroflightdynamics Directorate's Water Tunnel at NASA Ames Research Center. The experiment was performed at a Reynolds number of 100,000. Pitch oscillations with alpha(sub m) = 10 deg and 15 deg and with reduced frequencies ranging from k = 0.005 to 0.15 were examined. Blowing conditions ranged from C(sub mu) = 0.03 to 0.66 and F(+) = 0 to 3. Unsteady lift, drag, and pitching-moment loads were measured, and fluorescent-dye flow visualizations were obtained. Steady, upper-surface blowing was found to be capable of trapping a separation bubble near the leading edge during a portion of the airfoil's upward rotation. When this occurred, the lift was increased significantly and stall was averted. In all cases, steady blowing reduced the hysteresis amplitudes present in the loads and produced a large thrust force. The benefits of steady blowing diminished as the reduced frequency and mean angle of oscillation increased. Pulsed blowing showed only marginal benefits for the conditions tested. The greatest gains from pulsed blowing were achieved at F(+) = 0.9.
A series of studies were undertaken to measure various parameters such as velocity, species concentration, and temperature downstream of a shear coaxial injector in an optically accessible uni-element rocket chamber. Techniques applied include: laser Doppler velocimetry (LDV) for velocity; laser light scattering (LLS) for flow visualization and estimating mixture fraction and density; laser induced fluorescence (LIF) of hydroxyl radials (OH) to determine the characteristics and extent of the reaction zone; and Raman spectroscopy to measure major species concentration and temperature.
The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.
The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.
The test results briefly described in this report were obtained on the three-dimensional 1:48 scale tunnel modeled on the design proposed by Messrs. D.S.M.A. Corporation. More particularly, while the test chamber dimensions were indeed scaled down in the ration of 1:48, including the contraction and the collector as well, the duct system itself leading to and from the chamber was adapted to suit laboratory conditions and space limitations. Earlier tests with the two-dimensional model showed that blowing mode was preferred as against the suction mode, hence all tests were performed with blowing only. At the exit of the contraction the maximum airspeed attained with the 1 HP blower unit was about 200 ft/sec. This airspeed may be increased in future if desired. The test results show that pressure recovery in the diffuser was about 34 percent due to the large blockage at its entrance. Velocity traverses taken across the diffuser entrance explain the reason for this blockage. Recirculation, studied with both, hot-wire anemometry and flow-visualization techniques, was largely affected by the design of the test chamber itself and the amount of vent-air admitted to the chamber. Vent-air helped to decrease the level of turbulence.
Two major regimes where laser light scattered by molecules or particles in a flow contains significant information about the flow are Mie scattering and Rayleigh scattering. Mie scattering is used to obtain only velocity information, while Rayleigh scattering can be used to measure both the velocity and the thermodynamic properties of the flow. Now, recently introduced (1990, 1991) absorption filter based diagnostic techniques have started a new era in flow visualization, simultaneous velocity and thermodynamic measurements, and planar velocity measurements. Using a filtered planar velocimetry (FPV) technique, we have modified the optically thick iodine filter profile of Miles, et al., and used it in the pressure-broaden regime which accommodates measurements in a wide range of velocity applications. Measuring velocity and thermodynamic properties simultaneously, using absorption filtered based Rayleigh scattering, involves not only the measurement of the Doppler shift, but also the spectral profile of the Rayleigh scattering signal. Using multiple observation angles, simultaneous measurement of one component velocity and thermodynamic properties in a supersonic jet were measured. Presently, the technique is being extended for simultaneous measurements of all three components of velocity and thermodynamic properties.