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At least 451 records · Page 25
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements

In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.

Hah, Chunill↗
NASA NTRS2004

Glenn Heat Transfer Simulation and Solver Graphical User Interface: Development and Testing

In the Tui ine Branch of the Turbomachinery and Propulsion Systems Division, researching and developing efficient turbine aerothermodynamics technologies is the main objective. Creating effective turbines for jet engines is a process which, if based purely on physical experimental testing, would be extremely expensive. It is for this reason, and also for the reasons of speed and ease, that the Turbine Branch spends a large amount of effort working with simulations of turbines. Specifically, they focus their work on two main fields: Computational Field Dynamics (CFD), and Experimental data analysis. The experimental field involves comparing experimental results to simulated results, whereas the CFD field involves running these simulations. The simulations are applied to aerodynamics and heat transfer cases, for both steady and unsteady flow conditions. By and large this work is applied to the domain of flow and heat transfer in axial turbines. The main application used to run these heat flow simulations is GlennHT. This program, recently rewritten in FORTRAN 90, allows the user to input a job file which specifies all the necessary parameters needed to simulate flow through a user-defined grid. There are several other executables used as well, ranging in application from converting grid files to and from particular formats, to merging blocks in a connectivity file, to converting connectivity files to a GlennHT compatible format. All of these executables are run from the command line in a terminal; some of them have interactive prompts where the user must specify the files to be manipulated after the program starts, while others take all of their parameters from the command line. With this amount of variation comes a good deal of commands and formats to memorize, which can cause slower and less efficient work, as users may forget how to execute a certain program, or not remember the pathnames of the files they wish to use. Two years ago, steps were made to expedite this process with a graphical user interface (GUI) that combines the functionality of all the executables along with adding some new functionality, such as residuals graphing and boundary conditions creation. Upon my beginning here at Glenn, many parts of the GUI, which was developed in Java, were nonfunctional. There were also issues with cross-platforming, as systems in the branch were transitioning from Silicon Graphics (SGI) machines to Linux machines. My goals this summer are to finish the parts of the GUI that are not yet completed, fix parts that did not work correctly, expand the functionality to include other useful features, such as grid surface highlighting, and make the system compatible with both Linux and SGI. I will also be heavily testing the system and providing sufficient documentation on how to use the GUI, as no such documentation existed previously.

Kardamis, Joseph R.↗
NASA NTRSDate not supplied

Progress on Transonic Flutter and Shock Buffet Computationsin Support of the Third Aeroelastic Prediction Workshop

This paper reports on the progress of the NASA Langley team contributions to the third Aeroelastic Pre-diction Workshop’s (AePW-3) High Angle Working Group (HAWG). The primary objectives of HAWG is to predict the fluter dynamic pressure of the NASA Benchmark Supercritical Wing (BSCW) configuration at Mach 0.8 and 5◦angle of attack. The secondary objective is to determine if a shock-buffet onset is present at or near that flow condition. The computational results are obtained using FUN3D, an unstructured grid Reynolds-averaged Navier-Stokes solver developed at the NASA Langley Research Center. The preliminary analysis results show a computationally-obtained flutter dynamic pressure of approximately 120 psf. Initial results describing unforced BSCW unsteady flow environment at flutter condition are also presented.

Pawel Chwalowski↗
NASA NTRS1976

Noise generated by impingement of turbulent flow on airfoils of varied chord, cylinders, and other flow obstructions

Noise spectra were measured in three dimensions for several surfaces immersed in turbulent flow from a jet and over a range of flow conditions. The data are free field and were corrected to remove the small contributions of jet noise, atmospheric attenuation and feedback tones. These broadband data were compared with the results of available theories which are only strictly applicable to simple geometries over a limited range of conditions. The available theories proved to be accurate over the range of flow, chord length, thickness, angle of attack, and surface geometries defined by the experiments. These results apply to the noise generated by fixed surfaces in engine passages, the lifting surfaces of aircraft and also to fan noise.

Olsen, W. A.↗
NASA NTRS1976

Noise generated by impingement of turbulent flow on airfoils of varied chord, cylinders, and other flow obstructions

Noise spectra were measured in three dimensions for several surfaces immersed in turbulent flow from a jet and over a range of flow conditions. The data are free field and were corrected to remove the small contributions of jet noise, atmospheric attenuation and feedback tones. These broadband data were compared with the results of available theories which are only strictly applicable to simple geometries over a limited range of conditions. The available theories proved to be accurate over the range of flow, chord length, thickness, angle of attack, and surface geometries defined by the experiments. These results apply to the noise generated by fixed surfaces in engine passages, the lifting surfaces of aircraft and also to fan noise.

Olsen, W. A.↗
NASA NTRS1982

Numerical simulation of viscous-inviscid interactions on indented nose tips

An implicit numerical algorithm to solve the unsteady thin-layer Navier-Stokes equations in a strong conservative form has been used to compute the viscous flow over indented nose tips placed in a supersonic free stream. Numerical solutions are presented for axisymmetric and three-dimensional indented configurations for laminar flow conditions. Results demonstrate the capability of the present numerical procedure to predict flow fields that contain strong viscous-inviscid interactions, including boundary-layer separation, reattachment, and embedded discontinuities. Comparisons with available experimental data for the shock shape and surface pressure are also given.

Rizk, Y. M.↗
NASA NTRS1993

Developmental testing resulting in a simplified liquid oxygen check valve for the Space Shuttle Main Propulsion System

The coil spring in a Space Shuttle liquid oxygen check valve failed due to cyclic fatigue in September, 1991. The dual-flapper, swing check valve is used to prevent reverse flow to the Space Shuttle Main Engines. Upon inspection of the failed component, the spring tangs were found to be missing and heavy wear was observed on the inner diameter of the spring coils. The fracture surfaces revealed that the metal had been steadily worn away until a simple overload caused the final fracture. A series of flow tests using water and a water/gas mixture was conducted to identify the flow phenomenon which produced the cyclic wear. A Plexiglas outlet housing was utilized to view the flapper behavior under different flow conditions and to aid in high speed photography. The tests revealed that flow instabilities induced two oscillatory flapper responses: a rocking mode and a chattering mode. Initially, attempts were made to reduce the spring-flapper oscillations. However, the final solution to the problem was a springless configuration which satisfied the valve's design requirements and eliminated the oscillations. The springless design relied on the inherent ability of the reverse flow momentum to close the flappers.

Aber, Gregory S.↗
NASA NTRS1994

A workstation based simulator for teaching compressible aerodynamics

A workstation-based interactive flow simulator has been developed to aid in the teaching of undergraduate compressible aerodynamics. By solving the equations found in NACA 1135, the simulator models three basic fluids problems encountered in supersonic flow: flow past a compression corner, flow past two wedges in series, and flow past two opposed wedges. The study can vary the geometry or flow conditions through a graphical user interface and the new conditions are calculated immediately. Various graphical formats present the results of the flow calculations to the student. The simulator includes interactive questions and answers to aid in both the use of the tool and to develop an understanding of some of the complexities of compressible aerodynamics. A series of help screens make the simulator easy to learn and use.

Benson, Thomas J.↗
NASA NTRS2013

Development of Flow Boiling and Condensation Experiment on the International Space Station- Normal and Low Gravity Flow Boiling Experiment Development and Test Results

Flow boiling and condensation have been identified as two key mechanisms for heat transport that are vital for achieving weight and volume reduction as well as performance enhancement in future space systems. Since inertia driven flows are demanding on power usage, lower flows are desirable. However, in microgravity, lower flows are dominated by forces other than inertia (like the capillary force). It is of paramount interest to investigate limits of low flows beyond which the flow is inertial enough to be gravity independent. One of the objectives of the Flow Boiling and Condensation Flight Experiment sets to investigate these limits for flow boiling and condensation. A two-phase flow loop consisting of a Flow Boiling Module and two Condensation Modules has been developed to experimentally study flow boiling condensation heat transfer in the reduced gravity environment provided by the reduced gravity platform. This effort supports the development of a flow boiling and condensation facility for the International Space Station (ISS). The closed loop test facility is designed to deliver the test fluid, FC-72 to the inlet of any one of the test modules at specified thermodynamic and flow conditions. The zero-g-aircraft tests will provide subcooled and saturated flow boiling critical heat flux and flow condensation heat transfer data over wide range of flow velocities. Additionally, these tests will verify the performance of all gravity sensitive components, such as evaporator, condenser and accumulator associated with the two-phase flow loop. We will present in this paper the breadboard development and testing results which consist of detailed performance evaluation of the heater and condenser combination in reduced and normal gravity. We will also present the design of the reduced gravity aircraft rack and the results of the ground flow boiling heat transfer testing performed with the Flow Boiling Module that is designed to investigate flow boiling heat transfer and Critical Heat Flux (CHF) phenomena.

microgravity↗
NASA NTRS2016

Experiment Design and Characterization for the Study of Afterbody Radiation During Mars Entry

Recent work has shown that a significant contributor to the afterbody aeroheating during Mars entry is radiation. However, relevant ground test data is not available to help assess the uncertainty associated with prediction of the radiation when designing the thermal protection system for the aeroshell afterbody. The present work is aimed at designing an experiment which allows the study of the afterbody radiation experienced during Mars entry. The X2 expansion tube at the University of Queensland is used to generate the relevant experimental freestream flow conditions. Analysis is carried out to accurately characterize the generated experimental freestream conditions. A two dimensional wedge model is used to produce the expanding flow which simulates aspects of the afterbody flow around Mars entry vehicles. Preliminary analysis of the generated expanding flow shows that it produces significant radiation in the mid-infrared region and has a steady duration of about 50-110 μs. This allows emission spectroscopy to be conducted in the future.

Gu, S.↗
NASA NTRS2019

Influence of Source Propagation Direction and Shear Flow Profile in Impedance Eduction of Acoustic Liners

The acoustic impedance of liners is a key parameter for their design, and depends on the flow conditions, i.e., the sound pressure level and the presence of a grazing flow. The surface impedance of a locally reacting liner is defined as a local intrinsic property relating the acoustic pressure to the normal acoustic particle velocity at the liner surface. Impedance eduction techniques are now widely used to retrieve the impedance of liners in aeroacoustic facilities in the presence of a shear grazing flow. While surface impedance is intrinsic by definition, the educed impedance has recently been shown to depend on the direction of the incident waves relative to the mean flow. Different studies have investigated this issue by considering different acoustic propagation models used in the education process in the hope of matching the educed values. The purpose of the present work is to continue the previous investigations by evaluating the influence of the shear flow profile on the educed impedance, while considering a Bayesian inference process in order to evaluate the uncertainty on the educed values. The identified uncertainties were not able to totally account for the observed discrepancies between educed impedances.

Roncen, R.↗
NASA NTRS2012

Large Eddy Simulation (LES) of Particle-Laden Temporal Mixing Layers

High-fidelity models of plume-regolith interaction are difficult to develop because of the widely disparate flow conditions that exist in this process. The gas in the core of a rocket plume can often be modeled as a time-dependent, high-temperature, turbulent, reacting continuum flow. However, due to the vacuum conditions on the lunar surface, the mean molecular path in the outer parts of the plume is too long for the continuum assumption to remain valid. Molecular methods are better suited to model this region of the flow. Finally, granular and multiphase flow models must be employed to describe the dust and debris that are displaced from the surface, as well as how a crater is formed in the regolith. At present, standard commercial CFD (computational fluid dynamics) software is not capable of coupling each of these flow regimes to provide an accurate representation of this flow process, necessitating the development of custom software. This software solves the fluid-flow-governing equations in an Eulerian framework, coupled with the particle transport equations that are solved in a Lagrangian framework. It uses a fourth-order explicit Runge-Kutta scheme for temporal integration, an eighth-order central finite differencing scheme for spatial discretization. The non-linear terms in the governing equations are recast in cubic skew symmetric form to reduce aliasing error. The second derivative viscous terms are computed using eighth-order narrow stencils that provide better diffusion for the highest resolved wave numbers. A fourth-order Lagrange interpolation procedure is used to obtain gas-phase variable values at the particle locations.

Bellan, Josette↗
NASA NTRS1976

Simulation of flight-type engine fan noise in the NASA-Lewis 9 x 15 anechoic wind tunnel

A major problem in the measurement of aircraft engine fan noise is the difficulty of simulating, in a ground-based facility, the noise that occurs during flight. Flight-type noise as contrasted to the usual ground-static test noise exhibits substantial reductions in both (1) the time unsteadiness of tone noise and (2) the mean level of tones calculated to be nonpropagating or cut-off. A model fan designed with cut-off of the fundamental tone was acoustically tested in the anechoic wind tunnel under both static and tunnel flow conditions. The properties that characterize flight-type noise were progressively simulated with increasing tunnel flow. The distinctly lobed directivity pattern of propagating rotor/stator interaction modes was also observed. The results imply that the excess noise attributed to the ingestion of the flow disturbances that prevail near most static test facilities was substantially reduced with tunnel flow. The anechoic wind tunnel appears to be a useful facility for applied research on aircraft engine fan noise under conditions of simulated flight.

Heidmann, M. F.↗
NASA NTRS2018

Transition Delay via Vortex Generators in a Hypersonic Boundary Layer at Flight Conditions

The potential of realizable, stationary streaks undergoing non-modal growth to stabilize a hypersonic boundary-layer flow and, subsequently, delay the laminar-turbulent transition onset, is studied via numerical computations. The geometry and flow conditions are selected to match a relevant trajectory location from the ascent phase of the HIFiRE-1 flight experiment, namely, a 7-degree half-angle cone with 2.5 mm nose radius, freestream Mach number of 5.30, freestream unit Reynolds number equal to 13.42 x 10(exp 6)/m, and wall-to-adiabatic temperature ratio of approximately 0.35 over most of the test article. This paper investigates flow modifications induced by wall-mounted vortex generators (VGs), followed by an analysis of the modal instability of the perturbed, streaky boundary-layer flow. Results are presented both for a single array of VGs that was designed on the basis of optimal growth theory and for a VG configuration involving two separate arrays with opposite orientations that ware designed to provide staged control of flow instabilities while simultaneously reducing the amplification of streak instabilities resulting from the control devices. Earlier research had shown that the onset of transition during the HIFiRE-1 flight experiment, which did not include any control devices, correlated with an amplification factor of N = 14.7 for the planar Mack modes. If one assumes that the transition N -factor is not affected by the introduction of the VGs, then the control configurations based on a single array of VGs and two separate arrays would result in a transition delay of 17% and 40%, respectively. These findings suggest a passive flow control s to induce streaks that would delay transition in hypersonic boundary dominated by Mack-mode instabilities.

Incompressible boundary layer↗
NASA NTRS2018

Instability Wave–Streak Interactions in a High Mach Number Boundary Layer at Flight Conditions

The interaction of stationary streaks undergoing nonmodal growth with modally unstable instability waves in a hypersonic boundary-layer flow is studied using numerical computations. The geometry and flow conditions are selected to match a relevant trajectory location from the ascent phase of the HIFiRE-1 flight experiment; namely, a 7 degree half-angle, circular cone with 2.5 mm nose radius, freestream Mach number equal to 5.30, unit Reynolds number equal to 13.42 m -1 , and wall-to-adiabatic temperature ratio of approximately 0.35 over most of the vehicle. This paper investigates the nonlinear evolution of initially linear optimal disturbances that evolve into finite-amplitude streaks, followed by an analysis of the modal instability characteristics of the perturbed, streaky boundary-layer flow. The investigation is performed with stationary direct numerical simulations (DNS) and plane-marching parabolized stability equations (PSE), in conjunction with partial-differential-equation-based planar eigenvalue analysis. The overall effect of streaks is to reduce the peak amplification factors of instability waves, indicating a possible downstream shift in the onset of laminar-turbulent transition. The present study confirms previous findings that the mean flow distortion of the nonlinear streak perturbation reduces the amplification rates of the Mack-mode instability. More importantly, however, the present results demonstrate that the spanwise varying component of the streak can produce a larger effect on the Mack-mode amplification. The study with selected azimuthal wavenumbers for the stationary streaks reveals that a wavenumber of approximately 1.4 times larger than the optimal wavenumber is more effective in stabilizing the planar Mack-mode instabilities. In the absence of unstable first-mode waves for the present cold-wall condition, transition onset is expected to be delayed until the peak streak amplitude increases to nearly 35 percent of the freestream velocity, when intrinsic instabilities of the boundary-layer streaks begin to dominate the transition process. For streak amplitudes below that limit a significant net stabilization is achieved, yielding a potential transition delay that can exceed 100 percent of the length of the laminar region in the uncontrolled case.

Boundary layer stability↗
NASA NTRS1972

Applications of ERTS-A Data Collection System (DCS) in the Arizona Regional Ecological Test Site (ARETS)

The author has identified the following significant results. Preliminary analysis of DCS data from the USGS Verde River stream flow measuring site indicates the DCS system is furnishing high quality data more frequently than had been expected. During the 43-day period between Nov. 3, and Dec. 15, 1972, 552 DCS transmissions were received during 193 data passes. The amount of data received far exceeded the single high quality transmission per 12-hour period expected from the DCS system. The digital-parallel ERTS-1 data has furnished sufficient to accurately compute mean daily gage heights. These in turn, are used to compute average daily streamflow rates during periods of stable or slowly changing flow conditions. The digital-parallel data has also furnished useful information during peak flow periods. However, the serial-digital DCS capability, currently under development for transmitting streamflow data, should provide data of greater utility for determining times of flood peaks.

Schumann, H. H.↗