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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.

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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Flow-Field Surveys for Rectangular Nozzles

Flow field survey results for three rectangular nozzles are presented for a low subsonic condition obtained primarily by hot-wire anemometry. The three nozzles have aspect ratios of 2:1, 4:1 and 8:1. A fourth case included has 2:1 aspect ratio with chevrons added to the long edges. Data on mean velocity, turbulent normal and shear stresses as well as streamwise vorticity are presented covering a streamwise distance up to sixteen equivalent diameters from the nozzle exit. These detailed flow properties, including initial boundary layer characteristics, are usually difficult to measure in high speed flows and the primary objective of the study is to aid ongoing and future computational and noise modeling efforts. This supplement contains data files, charts and source code.

K B M Q Zaman↗
NASA NTRS2012

Boundary Layer Flow Control by an Array of Ramp-Shaped Vortex Generators

Flow field survey results for the effect of ramp-shaped vortex generators (VG) on a turbulent boundary layer are presented. The experiments are carried out in a low-speed wind tunnel and the data are acquired primarily by hot-wire anemometry. Distributions of mean velocity and turbulent stresses as well as streamwise vorticity, on cross-sectional planes at various downstream locations, are obtained. These detailed flow field properties, including the boundary layer characteristics, are documented with the primary objective of aiding possible computational investigations. The results show that VG orientation with apex upstream, that produces a downwash directly behind it, yields a stronger pair of streamwise vortices. This is in contrast to the case with apex downstream that produces a pair of vortices of opposite sense. Thus, an array of VG s with the former orientation, usually considered for film-cooling application, may also be superior for mixing enhancement and boundary layer separation control. The data files can be found on a supplemental CD.

Zaman, K. B. M. Q.↗
NASA NTRS2012

Experimental Studies of Low-Pressure Turbine Flows and Flow Control

This report summarizes research performed in support of the NASA Glenn Research Center (GRC) Low-Pressure Turbine (LPT) Flow Physics Program. The work was performed experimentally at the U.S. Naval Academy faculties. The geometry corresponded to "Pak B" LPT airfoil. The test section simulated LPT flow in a passage. Three experimental studies were performed: (a) Boundary layer measurements for ten baseline cases under high and low freestream turbulence conditions at five Reynolds numbers of 25,000, 50,000, 100,000, 200,000, and 300,000, based on passage exit velocity and suction surface wetted length; (b) Passive flow control studies with three thicknesses of two-dimensional bars, and two heights of three-dimensional circular cylinders with different spanwise separations, at same flow conditions as the 10 baseline cases; (c) Active flow control with oscillating synthetic (zero net mass flow) vortex generator jets, for one case with low freestream turbulence and a low Reynolds number of 25,000. The Passive flow control was successful at controlling the separation problem at low Reynolds numbers, with varying degrees of success from case to case and varying levels of impact at higher Reynolds numbers. The active flow control successfully eliminated the large separation problem for the low Reynolds number case. Very detailed data was acquired using hot-wire anemometry, including single and two velocity components, integral boundary layer quantities, turbulence statistics and spectra, turbulent shear stresses and their spectra, and intermittency, documenting transition, separation and reattachment. Models were constructed to correlate the results. The report includes a summary of the work performed and reprints of the publications describing the various studies.

Volino, Ralph J.↗
NASA NTRS2014

An Experimental Study of Roughness-Induced Instabilities in a Supersonic Boundary Layer

Progress on an experimental study of laminar-to-turbulent transition induced by an isolated roughness element in a supersonic laminar boundary layer is reported in this paper. Here, the primary focus is on the effects of roughness planform shape on the instability and transition characteristics. Four different roughness planform shapes were considered (a diamond, a circle, a right triangle, and a 45 degree fence) and the height and width of each one was held fixed so that a consistent frontal area was presented to the oncoming boundary layer. The nominal roughness Reynolds number was 462 and the ratio of the roughness height to the boundary layer thickness was 0.48. Detailed flow- field surveys in the wake of each geometry were performed via hot-wire anemometry. High- and low-speed streaks were observed in the wake of each roughness geometry, and the modified mean flow associated with these streak structures was found to support a single dominant convective instability mode. For the symmetric planform shapes - the diamond and circular planforms - the instability characteristics (mode shapes, growth rates, and frequencies) were found to be similar. For the asymmetric planform shapes - the right-triangle and 45 degree fence planforms - the mode shapes were asymmetrically distributed about the roughness-wake centerline. The instability growth rates for the asymmetric planforms were lower than those for the symmetric planforms and therefore, transition onset was delayed relative to the symmetric planforms.

Kegerise, Michael A.↗
NASA NTRS2013

Interaction of a Rectangular Jet with a Flat-Plate Placed Parallel to the Flow

An experimental study is carried out addressing the flowfield and radiated noise from the interaction of a large aspect ratio rectangular jet with a flat plate placed parallel to but away from the direct path of the jet. Sound pressure level spectra exhibit an increase in the noise levels for both the 'reflected' and 'shielded' sides of the plate relative to the free-jet case. Detailed cross-sectional distributions of flowfield properties obtained by hot-wire anemometry are documented for a low subsonic condition. Corresponding mean Mach number distributions obtained by Pitot-probe surveys are presented for high subsonic conditions. In the latter flow regime and for certain relative locations of the plate, a flow resonance accompanied by audible tones is encountered. Under the resonant condition the jet cross-section experiences an 'axis-switching' and flow visualization indicates the presence of an organized 'vortex street'. The trends of the resonant frequency variation with flow parameters exhibit some similarities to, but also marked differences with, corresponding trends of the well-known edgetone phenomenon.

Zaman, K. B. M. Q.↗
NASA NTRS2016

Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

An experimental investigation of a tip vortex from a NACA0012 airfoil is conducted in a low-speed wind tunnel at a chord Reynolds number of 4x10(exp 4). Initially, data for a stationary airfoil held at various angles-of-attack (alpha) are gathered. Detailed surveys are done for two cases: alpha=10 deg with attached flow and alpha=25 deg with massive flow separation on the upper surface. Distributions of various properties are obtained using hot-wire anemometry. Data include mean velocity, streamwise vorticity and turbulent stresses at various streamwise locations. For all cases, the vortex core is seen to involve a mean velocity deficit. The deficit apparently traces to the airfoil wake, part of which gets wrapped by the tip vortex. At small alpha, the vortex is laminar within the measurement domain. The strength of the vortex increases with increasing alpha but undergoes a sudden drop around alpha (is) greater than 16 deg. The drop in peak vorticity level is accompanied by transition and a sharp rise in turbulence within the core. Data are also acquired with the airfoil pitched sinusoidally. All oscillation cases pertain to a mean alpha=15 deg while the amplitude and frequency are varied. An example of phase-averaged data for an amplitude of +/-10 deg and a reduced frequency of k=0.2 is discussed. All results are compared with available data from the literature shedding further light on the complex dynamics of the tip vortex.

vortex flows↗
NASA NTRS2016

Femtosecond Laser Tagging Characterization of a Sweeping Jet Actuator Operating in the Compressible Regime

A sweeping jet (SWJ) actuator operating over a range of nozzle pressure ratios (NPRs) was characterized with femtosecond laser electronic excitation tagging (FLEET), single hot-wire anemometry (HWA) and high-speed/phase-averaged schlieren. FLEET velocimetry was successfully demonstrated in a highly unsteady, oscillatory flow containing subsonic through supersonic velocities. Qualitative comparisons between FLEET and HWA (which measured mass flux since the flow was compressible) showed relatively good agreement in the external flow profiles. The spreading rate was found to vary from 0.5 to 1.2 depending on the pressure ratio. The precision of FLEET velocity measurements in the external flow field was poorer (is approximately equal to 25 m/s) than reported in a previous study due to the use of relatively low laser fluences, impacting the velocity fluctuation measurements. FLEET enabled velocity measurements inside the device and showed that choking likely occurred for NPR ≥ 2.0, and no internal shockwaves were present. Qualitative oxygen concentration measurements using FLEET were explored in an effort to gauge the jet's mixing with the ambient. The jet was shown to mix well within roughly four throat diameters and mix fully within roughly eight throat diameters. Schlieren provided visualization of the internal and external flow fields and showed that the qualitative structure of the internal flow does not vary with pressure ratio and the sweeping mechanism observed for incompressible NPRs also probably holds for compressible NPRs.

Peters, Christopher J.↗
NASA NTRS2016

Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

An experimental investigation of tip vortex flow from a NACA0012 airfoil, pitched periodically at various frequencies, is conducted in a low-speed wind tunnel. Initially, data for stationary airfoil held fixed at various angles-of-attack are gathered. Flow visualization pictures as well as detailed cross-sectional properties areobtained at various streamwise locations using hot-wire anemometry. Data include mean velocity, streamwise vorticity as well as various turbulent stresses. Preliminary data are also acquired for periodically pitched airfoil. These results are briefly presented in this extended abstract.

wing aerodynamics↗
NASA NTRS2017

Transition Induced by a Streamwise Array of Roughness Elements on a Supersonic Flat Plate

Roughness is unavoidable on practical high-speed vehicles, so it is critical to determine its impact on boundary layer transition. The flow field downstream of a streamwise array of cylindrical roughness elements is probed with hot-wire anemometry in this experiment. Mean flow distortion is examined in several measurement planes in the wake of the cylindrical roughness using the streak strength profiles and contour plots of the mass flux and total temperature. The roughness element heights and spacings were varied and their instability modes were examined. Cylindrical roughness elements approximately 140 micron tall produce an odd instability mode that grows weakly with downstream distance in the measurement range of this experiment. Cylindrical roughness elements approximately 280 micron tall produce an even instability mode that grows, becomes nonlinear, and then breaks down. Transition onset remains constant relative to the most downstream roughness in the streamwise array when the 280 micron roughness elements are spaced 2 diameters apart. Transition onset occurs at an earlier upstream location relative to the most downstream roughness in the streamwise array when the roughness elements are spaced 4 diameters appear to recover before the next downstream roughness element, so the location of transition shifts with the location of the most downstream roughness element in the array. When the rough- apart. The wake behind roughness elements spaced 2 diameters apart do not ness elements are spaced 4 diameters apart, the flow behind the first roughness element has enough space to recover before feeding into the second roughness element, and thus, moves transition forward.

Chou, Amanda↗
NASA NTRS2019

Exit Boundary Layer Data for a Round Convergent Nozzle in Support of Numerical Simulation Efforts

An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves ‘Blasius-like’ mean velocity profiles but a ‘highly disturbed laminar state’ with large turbulence intensities in a range of subsonic Mach numbers.

Transition↗
NASA NTRS2019

Considerations for Femtosecond Laser Electronic Excitation Tagging in High-Speed Flows

Femtosecond laser electronic excitation tagging (FLEET) is an unseeded method for molecular tagging which offers valuable opportunities for measurement of high-speed (transonic, supersonic or hypersonic) flows. The unique nature of high-speed testing demands certain performance from FLEET such as satisfactory signal-to-noise ratio (SNR) at depressed static conditions (i.e., low temperatures, pressures and densities), wide dynamic range for velocity determination (especially single-shot), and measurements with acceptable accuracy and precision. This dissertation strives to evaluate FLEET in those regards and provide strategies to maximize the method's capabilities. A zero-dimensional kinetics model in nitrogen explains FLEET signal changes with pressure/density and/or temperature in terms of plasma-chemical reactions. Poorly known rate coefficients are tuned by comparing model output to measurements, with temporal agreement up to several hundred nanoseconds. Modeling reveals that initial signal peaks at reduced density because of slowed temporal evolution (and decay) of excited populations. Low temperatures enhance signal by enlarging cluster ion populations which contribute to excited species via electron-ion dissociative recombination. A purpose-built free jet facility provides experimental validation of the kinetics model and assesses FLEET velocimetry in low temperature and pressure/density conditions. Signal, lifetime, accuracy and precision results are obtained from unheated subsonic through Mach 4.0 operation of the facility, with best results noted. FLEET measurements of a sweeping jet (SWJ) actuator in compressible operation showcase its advantages in a highly unsteady jet containing subsonic through supersonic velocities. FLEET velocimetry is performed in the device's internal and external flow fields, with the latter compared to hot-wire anemometry. Internal measurements reveal the absence of shockwaves theorized to occur at high pressure ratios. Simultaneous qualitative measurements of compressible jet mixing are shown as a proof-of-concept. Overall, the work demonstrates that previous understanding of SWJ incompressible operation readily extends into the compressible realm. Practical aspects of performing FLEET velocimetry are detailed, along with strategies for improving measurement quality. Determination of a fundamental precision in nitrogen and air is attempted. Experiments show that increasing time delay and/or SNR improves velocimetry precision. A comparison of five camera systems indicates sensors with larger pixels capture higher SNR data and produce more precise results.

FLEET↗
NASA NTRS2019

Exit Boundary Layer Data for a Round Convergent Nozzle in Support of Numerical Simulation Efforts

An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves ‘Blasius-like’ mean velocity profiles but a ‘highly disturbed laminar state’ with large turbulence intensities in a range of subsonic Mach numbers.

Transition↗
NASA NTRS1995

Turbulence measurements in hypersonic boundary layers using constant-temperature anemometry and Reynolds stress measurements in hypersonic boundary layers

The primary objective in the two research investigations performed under NASA Langley sponsorship (Turbulence measurements in hypersonic boundary layers using constant temperature anemometry and Reynolds stress measurements in hypersonic boundary layers) has been to increase the understanding of the physics of hypersonic turbulent boundary layers. The study began with an extension of constant-temperature thermal anemometry techniques to a Mach 11 helium flow, including careful examinations of hot-wire construction techniques, system response, and system calibration. This was followed by the application of these techniques to the exploration of a Mach 11 helium turbulent boundary layer (To approximately 290 K). The data that was acquired over the course of more than two years consists of instantaneous streamwise mass flux measurements at a frequency response of about 500 kHz. The data are of exceptional quality in both the time and frequency domain and possess a high degree of repeatability. The data analysis that has been performed to date has added significantly to the body of knowledge on hypersonic turbulence, and the data reduction is continuing. An attempt was then made to extend these thermal anemometry techniques to higher enthalpy flows, starting with a Mach 6 air flow with a stagnation temperature just above that needed to prevent liquefaction (To approximately 475 F). Conventional hot-wire anemometry proved to be inadequate for the selected high-temperature, high dynamic pressure flow, with frequent wire breakage and poor system frequency response. The use of hot-film anemometry has since been investigated for these higher-enthalpy, severe environment flows. The difficulty with using hot-film probes for dynamic (turbulence) measurements is associated with construction limitations and conduction of heat into the film substrate. Work continues under a NASA GSRP grant on the development of a hot film probe that overcomes these shortcomings for hypersonic flows. Each of the research tasks performed during the NASA Langley research grants is discussed separately below.

Spina, Eric F.↗
NASA NTRS1988

Features of separating turbulent boundary layers

In the present study of two strong adverse pressure gradient flows, mean flow and turbulence characteristics are measured, together with frequency spectra, using hot-wire and laser anemometry. In these separating flows, reattachment occurs over a shorter distance than separation. It is noted that the outer flow variables form a unique set of scaling parameters for streamwise power spectra in adverse pressure gradient flows. The inner flow scaling of Perry et al. (1985) for streamwise spectra does not hold in the backflow region unless the value of the downstream-upstream intermittency in the flow is unity.

Nagabushana, K. A.↗
NASA NTRS1991

Fiber optic laser Doppler anemometry in swirling jets

Time-averaged and fluctuating quantities are measured in a free turbulent swirling jet. Data from a two-component laser Doppler anemometry (LDA) are compared to the measurements via hot-wire and 5-hole pitot probes. To acquire the proper seeding density near the axis of a swirling jet for LDA measurements proved difficult. This is due to an imbalance of the centrifugal force and radial pressure gradient, which throws the seeding material off the axis. Despite this problem, close agreement between various measurement techniques is obtained.

Taghavi, R.↗
NASA NTRS2015

Flexible Micropost Arrays for Shear Stress Measurement

Increased fuel costs, heightened environmental protection requirements, and noise abatement continue to place drag reduction at the forefront of aerospace research priorities. Unfortunately, shortfalls still exist in the fundamental understanding of boundary-layer airflow over aerodynamic surfaces, especially regarding drag arising from skin friction. For example, there is insufficient availability of instrumentation to adequately characterize complex flows with strong pressure gradients, heat transfer, wall mass flux, three-dimensionality, separation, shock waves, and transient phenomena. One example is the acoustic liner efficacy on aircraft engine nacelle walls. Active measurement of shear stress in boundary layer airflow would enable a better understanding of how aircraft structure and flight dynamics affect skin friction. Current shear stress measurement techniques suffer from reliability, complexity, and airflow disruption, thereby compromising resultant shear stress data. The state-of-the-art for shear stress sensing uses indirect or direct measurement techniques. Indirect measurements (e.g., hot-wire, heat flux gages, oil interferometry, laser Doppler anemometry, small scale pressure drag surfaces, i.e., fences) require intricate knowledge of the studied flow, restrictive instrument arrangements, large surface areas, flow disruption, or seeding material; with smaller, higher bandwidth probes under development. Direct measurements involve strain displacement of a sensor element and require no prior knowledge of the flow. Unfortunately, conventional "floating" recessed components for direct measurements are mm to cm in size. Whispering gallery mode devices and Fiber Bragg Gratings are examples of recent additions to this type of sensor with much smaller (m) sensor components. Direct detection techniques are often single point measurements and difficult to calibrate and implement in wind tunnel experiments. In addition, the wiring, packaging, and installation of delicate micro-electromechanical devices impede the use of most direct shear sensors. Similarly, the cavity required for sensing element displacement is sensitive to particulate obstruction. This work was focused on developing a shear stress sensor for use in subsonic wind tunnel test facilities applicable to an array of test configurations. The non-displacement shear sensors described here have minimal packaging requirements resulting in minimal or no disturbance of boundary layer flow. Compared to previous concepts, device installation could be simple with reduced cost and down-time. The novelty lies in the creation of low profile (nanoscale to 100 μm) micropost arrays that stay within the viscous sub-layer of the airflow. Aerodynamic forces, which are related to the surface shear stress, cause post deflection and optical property changes. Ultimately, a reliable, accurate shear stress sensor that does not disrupt the airflow has the potential to provide high value data for flow physics researchers, aerodynamicists, and aircraft manufacturers leading to greater flight efficiency arising from more in-depth knowledge on how aircraft design impacts near surface properties.

Wohl, Christopher J.↗
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.↗