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At least 73 records · Page 4

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NASA NTRS1976

Flow visualization in long neck Helmholtz resonators with grazing flow

Both oscillating and steady flows were applied to a single plexiglass resonator cavity with colored dyes injected in both the orifice and grazing flow field to record the motion of the fluid. For oscillatory flow, the instantaneous dye streamlines were similar for both the short and long-neck orifices. The orifice flow blockage appears to be independent of orifice length for a fixed amplitude of flow oscillation and magnitude of the grazing flow. The steady flow dye studies showed that the acoustic and steady flow resistances do not necessarily correspond for long neck orifices.

Baumeister, K. J.↗

NASA NTRS1980

Analytical and experimental study of mean flow and turbulence characteristics inside the passages of an axial flow inducer

The inviscid and viscid effects existing within the passages of a three bladed axial flow inducer operating at a flow coefficient of 0.065 are investigated. The blade static pressure and blade limiting streamline angle distributions were determined and the three components of mean velocity, turbulence intensities, and turbulence stresses were measured at locations inside the inducer blade passage utilizing a rotating three sensor hotwire probe. Applicable equations were derived for the hotwire data reduction analysis and solved numerically to obtain the appropriate flow parameters. The three dimensional inviscid flow in the inducer was predicted by numerically solving the exact equations of motion, and the three dimensional viscid flow was predicted by incorporating the dominant viscous terms into the exact equations. The analytical results are compared with the experimental measurements and design values where appropriate. Radial velocities are found to be of the same order as axial velocities within the inducer passage, confirming the highly three dimensional characteristic of inducer flow. Total relative velocity distribution indicate a substantial velocity deficiency near the tip at mid-passage which expands significantly as the flow proceeds toward the inducer trailing edge. High turbulence intensities and turbulence stresses are concentrated within this core region. Considerable wake diffusion occurs immediately downstream of the inducer trailing edge to decay this loss core. Evidence of boundary layer interactions, blade blockage effects, radially inward flows, annulus wall effects, and backflows are all found to exist within the long, narrow passages of the inducer.

Gorton, C. A.↗

NASA NTRS1984

Dynamics of lava flow - Thickness growth characteristics of steady two-dimensional flow

The thickness growth characteristics of flowing lava are investigated using a heat balance model and a two-dimensional model for flow of a Bingham plastic fluid down an inclined plane. It is found that yield strength plays a crucial role in the thickening of a lava flow of given flow rate. To illustrate this point, downstream thickness profiles and yield strength distributions were calculated for flows with mass flow rates of 10,000 and 100,000 kg/m-sec. Higher flow rates led to slow cooling rates which resulted in slow rate of increase of yield strength and thus greater flow lengths.

Park, S.↗

NASA NTRS1988

Compressibility effects on flow friction in a fully developed pipe flow

A method was evolved to account for the effects of compressibility in flows through long circular pipes of uniform diameter. Velocity and density variations across the flow cross section were approximated by profile shapes recommended for a flow past flat plates. The wall of the pipe was assumed to have attained an adiabatic condition. An integral method was used to obtain solutions for the governing equations of flow. The method was then used to analyze the detailed pressure measurements made in a long pipe for flow of air at different flow rate settings. Flow friction values derived from these pressure measurements were compared with the standard incompressible pipe flow data available in existing literature.

Murthy, Sreedhara V.↗

NASA NTRS1991

Low Flow Vortex Shedding Flow Meter for Hypergolics/all Media

A family of vortex shedding flow meters, for measurement of hypergol flows, was designed and fabricated. The test loops to evaluate the flow meters for water flow, as well as Freon -113 flow which simulates the hypergolic fluids, were modified and constructed to utilize a pump system which has an output capacity of 200 gpm. Test runs were conducted on the small 1/2 inch model with Freon 113 and on the larger models with water. Results showed that the linearity between the frequency of the vortices and the flow rate of the fluids was very close to that of the turbine flow meter. It is suggested that the vortex shedding flow meter is a possible replacement for the existing turbine type.

Thinh, Ngo Dinh↗

NASA NTRS1993

Effects of flow-path variations on internal reversing flow in a tailpipe offtake configuration for ASTOVL aircraft

A one-third-scale model of a generic tailpipe offtake system for an advanced short takeoff, vertical landing (ASTOVL) aircraft was tested at the NASA Lewis Research Center Powered Lift Facility. The basic model consisted of a tailpipe with a center body to form an annulus simulating turbine outflow with no swirl; twin offtake ducts with elbows at the ends to turn the flow to a downward direction; flow control nozzles at the ends of the elbows; and a blind flange at the end of the tailpipe to simulate a closed cruise nozzle. The offtake duct-to-tailpipe diameter ratio was 0.74. Modifications of a generic nature were then made to this basic configuration to measure the effects of flow-path changes on the flow and pressure-loss characteristics. The modifications included adding rounded entrances at the forward edges of the offtake openings, blocking the tailpipe just aft the openings instead of at the cruise nozzle, changing the location of the openings along the tailpipe, removing the center body, and varying the Mach number (flow rate) over a wide range in the tailpipe ahead of the openings by changing the size of the flow control nozzles. The tests were made with unheated air at tailpipe-to-ambient pressure ratios from 1.4 to 5. Results are presented and compared with performance graphs, total-pressure contour plots, paint streak flow visualization photographs, and a flow-angle probe traverse at the offtake entrance.

Mcardle, Jack G.↗

NASA NTRS1993

Effects of flow-path variations on internal reversing flow in a tailpipe offtake configuration for ASTOVL aircraft

Mcardle, Jack G.↗

NASA NTRS1998

Flow Simulation in Secondary Flow Passages of a Rocket Engine Turbopump

This paper presents application of a Generalized Fluid System Simulation Program (GFSSP) to model fluid flow in a very complex secondary flow circuit of a rocket engine turbopump, SIMPLEX. SIMPLEX was a low cost turbopump designed, manufactured and tested to demonstrate the ability to reduce the overall cost and design cycle time of turbomachinery. GFSSP is a general purpose computer program for analyzing flowrate, pressure, temperature and concentration distribution in a complex flow network. The program employs a finite volume formulation of mass, momentum and energy conservation equations in conjunction with thermodynamic equation of state of a real fluid. The secondary flow passages modeled include flow between the impeller shroud and housing, impeller back face and housing, flow through labyrinth seal and bearing, return line flow to inducer and flow adjacent to front and back face of the turbine. Axial load on the bearing are calculated from the predicted pressure on all axial faces of the rotor. The predicted pressure and temperature distributions compared well with the test data.

Schallhorn, Paul↗

NASA NTRS1994

Low-Disturbance Flow Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel

A unique, low-disturbance (quiet) supersonic wind tunnel has been commissioned at the NASA-Ames Fluid Mechanics Laboratory (FML) to support Supersonic Laminar Flow Control (SLFC) research. Known as the Laminar Flow Supersonic Wind Tunnel (LFSWT), this tunnel is designed to operate at potential cruise Mach numbers and unit Reynolds numbers (Re) of the High Speed Civil Transport (HSCT). The need to better understand the receptivity of the transition phenomena on swept (HSCT) wings to attachment-line contamination and cross-flows has provided the impetus for building the LFSWT. Low-disturbance or "quiet" wind tunnels are known to be an essential part of any meaningful boundary layer transition research. In particular, the receptivity of supersonic boundary layers to wind tunnel disturbances can significantly alter the transition phenomena under investigation on a test model. Consequently, considerable effort has gone into the design of the LFSWT to provide quiet flow. The paper describes efforts to quantify the low-disturbance flows in the LFSWT operating at Mach 1.6, as a precursor to transition research on wing models. The research includes: (1) Flow measurements in both the test section and settling chamber of the LFSWT, using a full range of measurement techniques; (2) Study of the state of the test section boundary layer so far by using a single hot-wire mounted above the floor centerline, with and without boundary layer trips fitted at the test section entrance; (3) The effect of flow quality of unsteady supersonic diffuser flow, joint steps and gaps, and wall vibration.

Wolf, Stephen W. D.↗

NASA NTRS1995

Investigation Flow Uniformity in a Supersonic Duct with High Enthalpy Flows

Flow uniformity in a high enthalpy facility is investigated. The scramjet research facility is composed of a rectangular combustor duct connected to a 100 MW electric arc air heater. The Mach 3.3 flow is accelerated through a two-dimensional contoured nozzle. Instream measurements were made with water-cooled Pitot probes and stagnation point heat flux gages at stream enthalpy levels ranging from 4 to 7 Mj/kg. Flow surveys were made on the flow centerline and off centerline in order to measure the three dimensional uniformity of the flow in the rectangular duct. Measurements indicated that although the flow in the aspect ratio 6:1 duct was relatively uniform on the centerline, three dimensional viscous effects were apparent near the corners. Flow through the nozzle and constant area duct was modeled computationally using a two dimensional, Navier-Stokes, reacting gas code. The computations predict that the flow in the test section is in vibrational equilibrium. The computed and measured Pitot pressure and heat flux profiles are in reasonable agreement with the experimental data.

Balboni, John↗

NASA NTRS1949

Lecture Series "Boundary Layer Theory". Part I - Laminar Flows: Laminar Flows - Part 1

In the lecture series starting today author want to give a survey of a field of aerodynamics which has for a number of years been attracting an ever growing interest. The subject is the theory of flows with friction, and, within that field, particularly the theory of friction layers, or boundary layers. A great many considerations of aerodynamics are based on the ideal fluid, that is the frictionless incompressibility and fluid. By neglect of compressibility and friction the extensive mathematical theory of the ideal fluid, (potential theory) has been made possible. Actual liquids and gases satisfy the condition of incomressibility rather well if the velocities are not extremely high or, more accurately, if they are small in comparison with sonic velocity. For air, for instance, the change in volume due to compressibility amounts to about 1 percent for a velocity of 60 meters per second. The hypothesis of absence of friction is not satisfied by any actual fluid; however, it is true that most technically important fluids, for instance air and water, have a very small friction coefficient and therefore behave in many cases almost like the ideal frictionless fluid. Many flow phenomena, in particular most cases of lift, can be treated satisfactorily, - that is, the calculations are in good agreement with the test results, -under the assumption of frictionless fluid. However, the calculations with frictionless flow show a very serious deficiency; namely, the fact, known as d'Alembert's paradox, that in frictionless flow each body has zero drag whereas in actual flow each body experiences a drag of greater or smaller magnitude. For a long time the theory has been unable to bridge this gap between the theory of frictionless flow and the experimental findings about actual flow. The cause of this fundamental discrepancy is the viscosity which is neglected in the theory of ideal fluid; however, in spite of its extraordinary smallness it is decisive for the course of the flow phenomena.

Schlichting, H.↗

NASA NTRS2005

Simulations of a Liquid Hydrogen Inducer at Low-Flow Off-Design Flow Conditions

The ability to accurately model details of inlet back flow for inducers operating a t low-flow, off-design conditions is evaluated. A sub-scale version of a three-bladed liquid hydrogen inducer tested in water with detailed velocity and pressure measurements is used as a numerical test bed. Under low-flow, off-design conditions the length of the separation zone as well as the swirl velocity magnitude was under predicted with a standard k-E model. When the turbulent viscosity coefficient was reduced good comparison was obtained a t all the flow conditions examined with both the magnitude and shape of the profile matching well with the experimental data taken half a diameter upstream of the leading edge. The velocity profiles and incidence angles a t the leading edge itself were less sensitive to the back flow length predictions indicating that single-phase performance predictions may be well predicted even if the details of flow separation modeled are incorrect. However, for cavitating flow situations the prediction of the correct swirl in the back flow and the pressure depression in the core becomes critical since it leads to vapor formation. The simulations have been performed using the CRUNCH CFD(Registered Trademark) code that has a generalized multi-element unstructured framework and a n advanced multi-phase formulation for cryogenic fluids. The framework has been validated rigorously for predictions of temperature and pressure depression in cryogenic fluid cavities and has also been shown to predict the cavitation breakdown point for inducers a t design conditions.

Hosangadi, A.↗

NASA NTRS2014

Reverse Flow Engine Core Having a Ducted Fan with Integrated Secondary Flow Blades

Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Kisska, Michael K.↗

NASA NTRS2022

Flow Visualization, Heat Transfer, and Critical Heat Flux of Flow Boiling in Earth Gravity with Saturated Liquid-Vapor Mixture Inlet Conditions – In Preparation for Experiments Onboard the International Space Station

This study investigates flow boiling of n-Perfluorohexane with saturated two-phase mixture inlet in a rectangular channel of dimensions 114.6-mm heated length, 2.5-mm width, and 5 mm-height. The experiments were performed as part of the Mission Sequence Testing of the Flow Boiling and Condensation Experiment’s (FBCE) Flow Boiling Module (FBM) in the vertical upflow configuration in Earth gravity using the same experimental system that was launched to the International Space Station (ISS) in August 2021. The operating parameters varied are heating configuration (single- and double-sided), mass velocity (380 – 2400 kg/m2s), inlet quality (0.011 – 0.519), and inlet pressure (120 – 179 kPa). High-speed video photographs are presented to explain the two-phase flow patterns within the channel’s heated length. Flow patterns are constituted by low-density and high-density fronts moving along the channel, with the high-density fronts gradually reducing in length due to evaporation. Heat transfer results in terms of flow boiling curves, streamwise wall temperature profiles, streamwise heat transfer coefficient profiles, and average heat transfer coefficients are presented and trends discussed. CHF data from the present experiments are combined with prior databases to compile a consolidated FBCE-CHF database for saturated inlet to expand the ranges of operating conditions and include other flow orientations in Earth gravity. Experimental CHF trends are also discussed. The interfacial lift-off model shows a good CHF predictive accuracy evidenced by a mean absolute error of 11.97% for this consolidated database after constraining it to mass velocities greater than or equal to 500 kg/m2s. Finally, this study confirmed reliability of the upcoming ISS experiments for saturated inlet conditions and the collected Earth-gravity data will be compared to ISS microgravity data.

Flow boiling↗

NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper presents and discusses the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of the test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model angle of incidence. Laser Doppler velocimetry (LDV) measurements were made at three model angles of incidence (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data were obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data on the test article and Reynolds-averaged Navier-Stokes CFD results are presented and discussed.

Juncture Flow↗

NASA NTRS2023

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗

NASA NTRSDate not supplied

Measurements and Computations of the Turbulent Corner Flow on the NASA Juncture-Flow Model with a Symmetric Wing

The NASA Juncture Flow experiment is designed to acquire high-quality flowfield data deep in the corner of a wing-fuselage junction specifically for the purpose of computational fluid dynamics (CFD) validation and turbulence model improvement. This paper will present and discuss the results of a recent experiment with the juncture-flow model in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The main objective of that test was to expand the existing juncture-flow dataset with a symmetric wing case that displays fully attached, incipient separation, and separated flow in the corner of the wing-fuselage junction, depending on the model pitch angle. Laser Doppler velocimetry (LDV) measurements were made at three model pitch angles (0 deg: fully attached, 1 deg: incipient separation, and 5 deg: separated flow) and for each one, mean-flow and Reynolds-stress data was obtained on the fuselage and at several streamwise locations along the corner of the wing-fuselage junction. Supporting measurements were made during the test campaign and included model and tunnel wall static pressures, tunnel wall and ceiling boundary-layer rake data, oil-flow visualizations, and laser-based measurements of the as-built model geometry and model position in the test section. Comparisons between the experimental data and Reynolds averaged Navier-Stokes CFD results will also be presented and discussed.

Juncture Flow↗

NASA NTRSDate not supplied

Visible and Infrared Imaging of Various Flow Regimes in the Flow Boiling and Condensation Experiment Transfer Line

The Flow Boiling and Condensation Experiment (FBCE) Transfer Line (TL) test module is being developed to study fluid physics of cryogenic transfer lines in microgravity. The TL module will reside in the FBCE facility located within the Fluids Integration Rack (FIR) aboard the International Space Station (ISS). TL will use the same normal perfluoro hexane (nPFH) used in the previous FBCE test sections. To simulate cryogenic line transfer in the visualization section, a sapphire tube coated with indium tin oxide (ITO) is heated via Joule heated while a bypass line is used to flow liquid nPFH at the desired inlet conditions. Once the visualization line reaches steady-state temperature and pressure conditions, flow is diverted through the heated line and chilldown begins. All flow regimes observed within a cryogenic TL are created and studied using visible imaging with diffuse backlighting to image inverted annular flow, transition, and nucleate boiling. A visible laser light sheet provided by FIR will provide a means to observe and track droplets in the dispersed flow regime. IR imaging of the visualization tube outer surface with the emissivity of the ITO coating provides the surface chilldown temperature profile. Preliminary imaging system development includes some of the components used in the Flow Boiling Module (FBM) and some new components that are easily integrated into the PC-based system but will require some refinement to operate properly on the FIR PCIe/104 imaging system. Optical design and packaging within the limited volume of the module enclosure is also presented. This paper presents FBCE-TL visible and infrared laboratory-based imaging studies and anticipated visualization capabilities of the TL module under preliminary breadboard development and under the utilization constraints within the FBCE/FIR host facility.

transfer line↗