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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 55 records · Page 3
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NASA NTRS2018

Assessment of Axial Wave Number and Mean Flow Uncertainty on Acoustic Liner Impedance Education

A key parameter in designing and assessing advanced broadband acoustic liners to achieve the current and future noise reduction goals is the acoustic impedance presented by the liner. This parameter, intrinsic to a specific liner configuration, is dependent on sound pressure level and grazing flow velocity. Current impedance eduction approaches have, in general, provided excellent results and continue to be employed throughout the acoustic liner community. However, some recent applications have indicated a possible dependence of the educed impedance on the direction of incident waves relative to the mean flow. The purpose of the current study is to investigate this unexpected behavior for various impedance eduction methods based on the Pridmore-Brown and convected Helmholtz equations. Specifically, the effects of flow profile and axial wavenumber uncertainties on educed impedances for upstream and downstream sources are investigated. The uniform flow results demonstrate the importance of setting a correct Mach number value in obtaining consistent educed impedances for upstream and downstream sources. In fact, the consistency of results over the two source locations was greatly improved by a slight modification of the uniform flow Mach number. In addition, uncertainty in educed axial wavenumber was also illustrated to correlate well with differences in the educed impedances, even with modified uniform flow Mach number. Finally, while less straightforward than in the uniform flow case, it appears that modification of the mean flow profile may also improve consistency of results for upstream and downstream results when shear flow is included.

Nark, Douglas M.↗
NASA NTRS2020

Full Aerothermal Characterization of the 6- by 9-Foot Icing Research Tunnel (2019 Test)

A full aerothermal characterization of the NASA Glenn Icing Research Tunnel (IRT) was completed in August 2019. A mandatory maintenance period required the test entry to be separated into two phases: the first occurring in April 2019 and the second in August 2019. The hardware used during this test entry consisted of the 6-foot quick-check rake, 6-foot traversing rake, RTD array, and hot-wire rake. The 6-foot quick check rake collects static pressure, total pressure, total temperature, and flow angularity data along the centerline of the test section. The data from the 6-foot quick-check rake were used to generate test section calibration relationships for total and static pressure as well as implement statistical process control. The 6-foot traversing rake collects flow uniformity data across the test section for various flow parameters, such as static pressure, total pressure, and flow angularity. Total temperature flow uniformity data is mapped by the RTD array and is used to generate the test section calibration relationship for total temperature. The hot-wire rake measures turbulence intensity at three lateral locations and characterizes airspeed unsteadiness in the test section. The test entry resulted in updated calibration relationships and a full characterization of test section flow quality and flow uniformity in the IRT.

Wind↗
NASA NTRS2000

Generation and Radiation of Acoustic Waves from a 2-D Shear Layer

A parallel numerical simulation of the radiation of sound from an acoustic source inside a 2-D jet is presented in this paper. This basic benchmark problem is used as a test case for scattering problems that are presently being solved by using the Impedance Mismatch Method (IMM). In this technique, a solid body in the domain is represented by setting the acoustic impedance of each medium, encountered by a wave, to a different value. This impedance discrepancy results in reflected and scattered waves with appropriate amplitudes. The great advantage of the use of this method is that no modifications to a simple Cartesian grid need to be made for complicated geometry bodies. Thus, high order finite difference schemes may be applied simply to all parts of the domain. In the IMM, the total perturbation field is split into incident and scattered fields. The incident pressure is assumed to be known and the equivalent sources for the scattered field are associated with the presence of the scattering body (through the impedance mismatch) and the propagation of the incident field through a non-uniform flow. An earlier version of the technique could only handle uniform flow in the vicinity of the source and at the outflow boundary. Scattering problems in non-uniform mean flow are of great practical importance (for example, scattering from a high lift device in a non-uniform mean flow or the effects of a fuselage boundary layer). The solution to this benchmark problem, which has an acoustic wave propagating through a non-uniform mean flow, serves as a test case for the extensions of the IMM technique.

Agarwal, Anurag↗
NASA NTRS1994

Optimized design of a hypersonic nozzle

Conventional procedures for designing nozzles involve the design of an inviscid contour (using the method of characteristics) that is corrected with a displacement thickness calculated from boundary-layer theory. However, nozzles designed using this classical procedure have been shown to exhibit poor flow quality at Mach numbers characteristic of hypersonic applications. The nozzle to be designed will be a part of the NASA HYPULSE facility which is being used for hypervelocity flight research. Thus, the flow quality of the nozzle is a critical question that needs to be addressed. Design of nozzles for hypersonic applications requires a proper assessment of the effects of the thick boundary layer on the inviscid flowfield. Since the flow field is largely supersonic, the parabolized form of the Navier-Stokes (PNS) equations can be used. The requirement of a uniform flow at the exit plane of the nozzle can be used to define an objective function as part of an optimization procedure. The design procedure used in this study involves the coupling of a nonlinear (least-squares) optimization algorithm with an efficient, explicit PNS solver. The thick boundary layers growing on the walls of the nozzle limit the extent of the usable core region (region with uniform flow) for testing models (especially rectangular). In order to maximize the region of uniform flow, it was decided to have the exit plane of this nozzle to be (nearly) rectangular. Thus, an additional constraint on the nozzle shape resulted, namely the nozzle will have a shape transitioning from a circular one at the inlet to that of a rectangle at the exit. In order to provide for a smooth shape transition, the cross sectional contour of the nozzle is defined by a superellipse. The nozzle is taken to be a meter in length. The axial variations of the major and minor radii of the superellipse are governed by cubic splines. The design parameters are the coefficients of the splines associated with the local nozzle wall slopes. Extensive calculations have been made (with a three-dimensional Euler code) to understand the effects of various parameters such as location of the knot points of the spline function, different ways of characterizing the uniformity of the flow in the exit plane, as well as the effect of constraining the area of the nozzle to be invariant. Turbulent flow (measurements indicate that the flow at the nozzle inlet is turbulent) calculations are now being performed (with the inviscidly designed nozzle contours) to assess the flow quality.

Krishnamurthy, Ramesh↗
NASA NTRS2005

A Quasi-3-D Theory for Impedance Eduction in Uniform Grazing Flow

A 2-D impedance eduction methodology is extended to quasi-3-D sound fields in uniform or shearing mean flow. We introduce a nonlocal, nonreflecting boundary condition to terminate the duct and then educe the impedance by minimizing an objective function. The introduction of a parallel, sparse, equation solver significantly reduces the wall clock time for educing the impedance when compared to that of the sequential band solver used in the 2-D methodology. The accuracy, efficiency, and robustness of the methodology is demonstrated using two examples. In the first example, we show that the method reproduces the known impedance of a ceramic tubular test liner. In the second example, we illustrate that the approach educes the impedance of a four-segment liner where the first, second, and fourth segments consist of a perforated face sheet bonded to honeycomb, and the third segment is a cut from the ceramic tubular test liner. The ability of the method to educe the impedances of multisegmented liners has the potential to significantly reduce the amount of time and cost required to determine the impedance of several uniform liners by allowing them to be placed in series in the test section and to educe the impedance of each segment using a single numerical experiment. Finally, we probe the objective function in great detail and show that it contains a single minimum. Thus, our objective function is ideal for use with local, inexpensive, gradient-based optimizers.

Watson, W. R.↗
NASA NTRS2020

Investigation of Atmospheric Boundary-Layer Effects on Launch-Vehicle Ground Wind Loads

A launch vehicle ground-wind-loads program is underway at the NASA Langley Transonic Dynamics Tunnel. The objectives are to quantify key aerodynamic and structural characteristics that impact the occurrence of large wind-induced oscillations of a launch vehicle when exposed to ground winds prior to launch. Of particular interest is the dynamic response of a launch vehicle when a von Kármán vortex street forms in the wake of the vehicle resulting in quasiperiodic lift and drag forces. Vehicle response to these quasiperiodic forces can become quite large when the frequency of vortex shedding nears that of a lowly-damped structural mode thereby exciting a resonant response. Wind approaching the vehicle can be characterized by a varying speed with height and turbulence content. The combination of both the varying speed and turbulence content is referred to herein as the atmospheric boundary-layer. The importance of the atmospheric boundary-layer upon launch vehicle wind-induced oscillation response has long been questioned, and its effects are not well understood. Although there are several facilities around the world dedicated to replicating atmospheric boundary layers, the development of such a boundary layer in a wind tunnel capable of producing flight-representative Reynolds numbers for aeroelastically-scaled launch vehicle models has only recently been accomplished. The NASA Langley Transonic Dynamics Tunnel is capable of simulating flight-representative Reynolds numbers of launch vehicles on the pad and is uniquely capable of replicating many fluid-structure scaling parameters typical of aeroelastic tests. Recent test efforts successfully developed representative atmospheric boundary-layers for three launch sites in the Transonic Dynamics Tunnel, thereby allowing all known aerodynamic and fluid-structure coupling parameters to be simultaneously simulated for those sites. Dynamic aeroelastically-scaled models representative of typical large launch vehicles were constructed for testing. Aeroelastic scaling includes matching geometry, mode shapes, reduced frequencies, damping, running mass ratios, and running stiffness ratios. The models were tested in smooth uniform flow and then immersed in the atmospheric boundary-layer for comparison of these effects. Dynamic data were acquired measuring unsteady pressure, acceleration, and base bending moment. It was discovered that peak dynamic loads resulting from resonant wind-induced oscillation response are similar when acquired in either smooth uniform flow or an atmospheric boundary-layer. This indicates that resonant lock-in events are minimally impacted by representative turbulence and/or wind profile. Alternately, nonresonant wind-induced oscillation response events are stronger when acquired in an atmospheric boundary-layer. This indicates that a lowly-damped structural response will increase when exposed to an increased magnitude of random excitation, which is consistent with historical comparisons. Loads created by the resonant response events were substantially stronger than those from the nonresonant response events. Therefore, if testing is done to simply identify worst-case conditions and load magnitude, then smooth uniform flow is likely an adequate test technique. However, if nonresonant response loads are of primary interest, then atmospheric boundary-layer simulation is required.

Thomas G Ivanco↗
NASA NTRS1974

Numerical solution of the time-dependent compressible Navier-Stokes equations in inlet regions

The results of a study to determine the effects of compressibility on the viscous flow through channels that have straight, parallel walls are presented. Two channel configurations are considered, the flow between two semi-infinite flat plates with uniform flow prescribed at the inlet plane and a cascade of semi-infinite flat plates with uniform flow introduced upstream. The flow field is modeled by using the time dependent, compressible Navier-Stokes equations. Time dependent solutions are obtained by using an explicit finite difference technique which advances the pressure on near field subsonic boundaries such that accurate steady state solutions are obtained. Steady state results at Reynolds number 20 and 150 are presented for Mach numbers between 0.09 and 0.36 and compared with the incompressible solutions of previous studies.

Olson, L. E.↗
NASA NTRS1983

Aeroacoustic computation of cylinder wake flow

Sound generation by uniform flow over a cylinder at the Reynolds number of 200 is calculated. The incompressible, time dependent flow field is first computed by standard computational fluid dynamics techniques using a stream function/vorticity formulation. The sound radiation is then obtained by integrating over the flow field with the Coriolis acceleration as the source term and a low frequency Green's function technique. Time histories, spectra, and directivity of the radiated sound are obtained. Comparisons with experimental data, where possible, indicate the feasibility of this computational aeroacoustics approach.

Hardin, J. C.↗
NASA NTRS1975

The transmission of sound in nonuniform ducts

The method of weighted residuals in the form of a modified Galerkin method with boundary residuals was developed for the study of the transmission of sound in nonuniform ducts carrying a steady, compressible flow. In this development, the steady flow was modeled as essentially one dimensional but with a kinematic modification to force tangency of the flow at the duct walls. Three forms of the computational scheme were developed using for basis functions (1) the no-flow uniform duct modes, (2) positive running uniform duct modes, with flow, and (3) positive and negative running uniform duct modes, with flow. The formulation using the no-flow modes was the most highly developed, and has advantages primarily due to relative computational simplicity. Results using the three methods are shown to converge to known solutions for several special cases, and the most significant check case is against low frequency, one dimensional results over the complete subsonic Mach number range. Development of the method is continuing, with emphasis on assessing the relative accuracy and efficiency of the three implementations.

Eversman, W.↗
NASA NTRS1995

Effects of gravity on sheared and nonsheared turbulent nonpremixed flames

The present numerical study is concerned with the fundamental physics of the multiway interaction between turbulence, chemical reaction, and buoyancy in a nonpremixed flame. The method of direct numerical simulation (DNS) is used to solve the instantaneous, three-dimensional governing equations. Because of the present supercomputer limitations, we consider two simple flow geometries, namely an initially uniform flow without shear (equivalent to grid-generated turbulence) and an initially uniform shear flow. In each flow, the fuel and oxidant initially exist as two separate streams. As the reactants mix, chemical reaction takes place and exothermic energy is released causing variations in density. In the presence of a gravity field, the spatial and temporal distributions of the induced buoyancy forces depend on the local density gradients and the direction of the gravitational acceleration. The effects of buoyancy include the generation of local shear, baroclinic production or destruction of vorticity, and countergradient heat and mass transport. Increased vorticity and small-scale turbulence promote further mixing and reaction. However, if the strain-rates become too high, local flame extinction can occur. Our objective is to gain an understanding of the complex interactions between the physical phenomena involved, with particular attention to the effects of buoyancy on the turbulence structure, flame behavior, and factors influencing flame extinction.

Elghobashi, Said↗
NASA NTRS1997

Effects of Gravity on Sheared and Nonsheared Turbulent Nonpremixed Flames

The present numerical study is concerned with the fundamental physics of the multi-way interaction between turbulence, chemical reaction and buoyancy in a nonpremixed flame. The method of direct numerical simulation (DNS) is used to solve the instantaneous three-dimensional governing equations (continuity, Navier-Stokes, species mass fractions and energy) under the zero-Mach number assumption. Because of the present supercomputer limitations, we consider two simple flow geometries, namely an initially uniform flow without shear (equivalent to grid-generated turbulence) and an initially uniform shear flow. In each flow, the fuel and oxidant initially exist as two separate streams. As the reactants mix, chemical reaction takes place and exothermic energy is released causing variations in density. In the presence of a gravity field, the spatial and temporal distributions of the induced buoyancy forces depend on the local density gradients and the direction of the gravitational acceleration.

Elghobashi, Said↗
NASA NTRS1986

Confinement of thermocapillary floating zone flow by uniform rotation

Smith and Greenspan (1984) examined theoretically the idea of using a uniform rotation of the floating zone system to confine the thermocapillary flow in crystal growth experiments to the melt sidewall, leaving the interior of the melt passive. Here, that model is extended to a full zone with a more realistic temperature distribution imposed on the sidewall, and both linear and nonlinear thermocapillary flows are theoretically studied. Linearized, analytical solutions are found using singular perturbation theory and the various sidewall boundary layers described by Greenspan (1969) for rotating fluids. The analytical and linearized numerical results are compared, and the linear and nonlinear flows are discussed. The results demonstrate that the thermocapillary flow is strong and that rotation cannot confine the flow. Temperature advection by strongly nonlinear flow is significant even for the small Prandtl number of silicon.

Fowlis, W. W.↗
NASA NTRS1999

Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies

Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined.

Cobleigh, Brent R.↗
NASA NTRS1999

Flush Airdata Sensing (FADS) System Calibration Procedures and Results for Blunt Forebodies

Blunt-forebody pressure data are used to study the behavior of the NASA Dryden Flight Research Center flush airdata sensing (FADS) pressure model and solution algorithm. The model relates surface pressure measurements to the airdata state. Spliced from the potential flow solution for uniform flow over a sphere and the modified Newtonian impact theory, the model was shown to apply to a wide range of blunt-forebody shapes and Mach numbers. Calibrations of a sphere, spherical cones, a Rankine half body, and the F-14, F/A-18, X-33, X-34, and X-38 configurations are shown. The three calibration parameters are well-behaved from Mach 0.25 to Mach 5.0, an angle-of-attack range extending to greater than 30 deg, and an angle-of-sideslip range extending to greater than 15 deg. Contrary to the sharp calibration changes found on traditional pitot-static systems at transonic speeds, the FADS calibrations are smooth, monotonic functions of Mach number and effective angles of attack and sideslip. Because the FADS calibration is sensitive to pressure port location, detailed measurements of the actual pressure port locations on the flight vehicle are required and the wind-tunnel calibration model should have pressure ports in similar locations. The procedure for calibrating a FADS system is outlined.

Cobleigh, Brent R.↗
NASA NTRS1973

A difference theory for noise propagation in an acoustically lined duct with mean flow

A finite difference formulation is presented for sound propagation in a two-dimensional straight soft-walled duct with uniform flow. The difference analysis is developed in terms of complex notation. The governing acoustic difference equations and the appropriate displacement boundary conditions associated with uniform flow are presented for the sound attenuation in straight hard and soft-walled ducts. At present the finite Mach number case is solved only for the one-dimensional hard walled duct.

Baumeister, K. J.↗