SEARCH · Engineering Papers

Results for “unsteady flow”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 865 records · Page 48

All sources NASA NTRS DOE OSTI

NASA NTRS1997

CFD Analysis and Design Optimization Using Parallel Computers

A versatile and efficient multi-block method is presented for the simulation of both steady and unsteady flow, as well as aerodynamic design optimization of complete aircraft configurations. The compressible Euler and Reynolds Averaged Navier-Stokes (RANS) equations are discretized using a high resolution scheme on body-fitted structured meshes. An efficient multigrid implicit scheme is implemented for time-accurate flow calculations. Optimum aerodynamic shape design is achieved at very low cost using an adjoint formulation. The method is implemented on parallel computing systems using the MPI message passing interface standard to ensure portability. The results demonstrate that, by combining highly efficient algorithms with parallel computing, it is possible to perform detailed steady and unsteady analysis as well as automatic design for complex configurations using the present generation of parallel computers.

Martinelli, Luigi↗

NASA NTRS1994

Simulation of Unsteady Combustion in a Ramjet Engine Using a Highly Parallel Computer

Combustion instability in ramjets is a complex phenomenon that involve nonlinear interaction between acoustic waves, vortex motion and unsteady heat release in the combustor. To numerically simulate this 3-D, transient phenomenon, enormous computer resources (time, memory and disk storage) are required. Although current generation vector supercomputers are capable of providing adequate resources for simulations of this nature, their high cost and limited availability, makes such machines less than satisfactory for routine use. The primary focus of this study is to assess the feasibility of using highly parallel computer systems as a cost-effective alternative for conducting such unsteady flow simulations. Towards this end, a large-eddy simulation model for combustion instability was implemented on the Intel iPSC/860 and a careful study was conducted to determine the benefits and the problems associated with the use of such machines for transient simulations. Details of this study along with the results obtained from the unsteady combustion simulations carried out on the iPSC/860 are discussed in this paper.

Menon, Suresh↗

NASA NTRS2019

Investigation of Transitional Shock-Wave/Boundary Layer Interactions Using Direct Numerical Simulations

Interaction of a transitional boundary layer with a normal-shock is investigated using unstructured tetrahedral meshes under the numerical framework of the space-time conservation element, solution element (CESE) method. The computations mimic recent experimental efforts at the University of Tennessee Space Institute, where a Mach 2.0 flow interacts with a tall cylinder attached to a flat plate. The location of the cylinder with respect to the flat plate leading edge determines if the incoming boundary layer is laminar, transitional or fully turbulent. Four representative flow conditions exemplifying laminar and transitional boundary layers are analyzed by direct numerical simulations. Similar to what was observed in the experiments for the case of transitional interaction, the computations reveal an intermittent upstream influence (UI) shock that repeatedly travels upstream from the lambda-foot toward the leading edge before vanishing. Through detailed unsteady flow analysis obtained using Fourier analysis and dynamic mode decomposition techniques, the presence of disturbances with similar frequencies as those measured in experiments were identified in the flow along with locations that appear to influence the dynamics of the flow.

Venkatachari, Balaji Shankar↗

NASA NTRSDate not supplied

Modification of Mach 6 Freestream Flow by Pitot Probe Bow Shock as Measured by Line FLDI and SAFS

Measurements of the flow in front of flat-face Pitot probes in NASA Langley ResearchCenter’s 20-Inch Mach 6 Air Tunnel were made using a 35-point line focused laser differentialinterferometer (FLDI), a two-point FLDI, and a self-aligned focusing schlieren system. Thesesystems were used to make measurements of flow unsteadiness between the bow shock andPitot probe face, while also probing the freestream region just upstream of the bow shock.Shock standoff distances from the non-intrusive measurements match closely with analyticalsolutions using the probe face diameter. Acoustic wave angles and velocities in the freestreamcompare favorably with DNS simulations of a Mach 6 freestream flow, and wave velocities in thepost-shock region slow as they approach the probe face. Peak frequencies in the post-shockregion as measured by FLDI compare favorably to the estimates computed from the shockstandoff distance for most probe diameters. With increasing unit Reynolds number, fluctuationintensities in the post-shock region were attenuated below the intensities of the freestream, witha spectral, spatial, and probe size dependence. Pitot probe power spectral density estimatesare presented from the test campaign, which can be compared with FLDI and SAFS data toevaluate the applicability of the transfer functions derived from computational simulations.

Joshua M Weisberger↗

NASA NTRS1987

Computation of transonic potential flow on helicopter rotor blades

Two computer codes, the full-potential three-dimensional (FP3D) code and the full-potential rotor (FPR) code have recently been developed. Both of these codes solve the three-dimensional conservative formulation of the full potential equation. The FPR code was developed at the U.S. Army Aeroflightdynamics Directorate (AFDD) while the FP3D code was a joint development by ONERA and AFDD. Both of these codes were used to predict the nonlifting, unsteady flow over a rotor operating at high advance ratio and tip speed. Three different rotor tip planform shapes were studied: a rectangular tip, a 30 deg aft swept tip and a 30 deg forward swept tip. Results of these computations are compared to results obtained using an earlier small-disturbances code. Also, the lifting flow over a rectangular tip operating at a slightly different condition was computed. These results are also compared with the small disturbances computations and with experimental results.

Costes, M.↗

NASA NTRS1985

Recent transonic unsteady pressure measurements at the NASA Langley Research Center

Four semispan wing model configurations were studied in the Transonic Dynamics Tunnel (TDT). The first model had a clipped delta planform with a circular arc airfoil, the second model had a high aspect ratio planform with a supercritical airfoil, the third model has a rectangular planform with a supercritical airfoil and the fourth model had a high aspect ratio planform with a supercritical airfoil. To generate unsteady flow, the first and third models were equipped with pitch oscillation mechanisms and the first, second and fourth models were equipped with control surface oscillation mechanisms. The fourth model was similar in planform and airfoil shape to the second model, but it is the only one of the four models that has an elastic wing structure. The unsteady pressure studies of the four models are described and some typical results for each model are presented. Comparison of selected experimental data with analytical results also are included.

Sandford, M. C.↗

NASA NTRS1985

Recent transonic unsteady pressure measurements at the NASA Langley Research Center

Sandford, M. C.↗

NASA NTRS1992

Natural convection in binary gases driven by combined horizontal thermal and vertical solutal gradients

An investigation of natural convection is presented to examine the influence of a horizontal temperature gradient and a concentration gradient occurring from the bottom to the cold wall in a cavity. As the solutal buoyancy force changes from augmenting to opposing the thermal buoyancy force, the fluid motion switches from unicellular to multicellular flow (fluid motion is up the cold wall and down the hot wall for the bottom counterrotating flow cell). Qualitatively, the agreement between predicted streamlines and smoke flow patterns is generally good. In contrast, agreement between measured and predicted temperature and concentration distributions ranges from fair to poor. Part of the discrepancy can be attributed to experimental error. However, there remains considerable discrepancy between data and predictions due to the idealizations of the mathematical model, which examines only first-order physical effects. An unsteady flow, variable thermophysical properties, conjugate effects, species interdiffusion, and radiation were not accounted for in the model.

Weaver, J. A.↗

NASA NTRS1979

The lambda-scheme

A method for integrating the Euler equations of gas dynamics for compressible flows in any hyperbolic case is presented. This method is applied to the Mach number distribution over a stretch of an infinite duct having a variable cross section, and to the distribution in a channel opening into a vacuum with the Mach number equalling 1.04. An example of the ability of this method to handle two-dimensional unsteady flows is shown using the steady shock-and-isobars pattern reached asymptotically about an ablated blunt body with a free stream Mach number equalling 12. A final example is presented where the technique is applied to a three-dimensional steady supersonic flow, with a Mach number of 2 and an angle of attack of 5 deg.

Moretti, G.↗

NASA NTRS1989

An efficient method for computing unsteady transonic aerodynamics of swept wings with control surfaces

A transonic equivalent strip (TES) method was further developed for unsteady flow computations of arbitrary wing planforms. The TES method consists of two consecutive correction steps to a given nonlinear code such as LTRAN2; namely, the chordwise mean flow correction and the spanwise phase correction. The computation procedure requires direct pressure input from other computed or measured data. Otherwise, it does not require airfoil shape or grid generation for given planforms. To validate the computed results, four swept wings of various aspect ratios, including those with control surfaces, are selected as computational examples. Overall trends in unsteady pressures are established with those obtained by XTRAN3S codes, Isogai's full potential code and measured data by NLR and RAE. In comparison with these methods, the TES has achieved considerable saving in computer time and reasonable accuracy which suggests immediate industrial applications.

Liu, D. D.↗

NASA NTRS1980

Techniques for aerodynamic and turbulence measurements in turbomachinery rotors

Nonoptical techniques currently used to measure flows in turbomachine rotors are reviewed, with particular reference to the measurement of the three-dimensional mean velocity, turbulence intensity, Reynolds stresses, static and stagnation pressures, and blade surface measurements. The high level of sophistication of rotor flow measurement techniques is demonstrated. Rotating probe measurements are well suited for low- and moderate-speed turbomachinery, while stationary quick-response probes (hot-wire and pressure probes) are applicable both in low- and high-speed turbomachines. Hot-wire probes, blade pressure transducers, and telemetry systems will provide the urgently needed basic information on steady and unsteady flow through a blade passage, including blade boundary layer, shock configuration, leakage and secondary flows.

Lakshminarayana, B.↗

NASA NTRS1995

Auto-adaptive finite element meshes

Accurate capturing of discontinuities within compressible flow computations is achieved by coupling a suitable solver with an automatic adaptive mesh algorithm for unstructured triangular meshes. The mesh adaptation procedures developed rely on non-hierarchical dynamical local refinement/derefinement techniques, which hence enable structural optimization as well as geometrical optimization. The methods described are applied for a number of the ICASE test cases are particularly interesting for unsteady flow simulations.

Richter, Roland↗

NASA NTRS2002

Rocket Engine Oscillation Diagnostics

Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

Nesman, Tom↗

NASA NTRS1981

Stall flutter experiment in a transonic oscillating linear cascade

Two dimensional biconvex airfoils were oscillated at reduced frequencies up to 0.5 based on semi-chord and a free stream Mach number of 0.80 to simulate transonic stall flutter in rotors. Steady-state periodicity was confirmed through end-wall pressure measurements, exit flow traverses, and flow visualization. The initial flow visualization results from flutter tests indicated that the oscillating shock on the airfoils lagged the airfoil motion by as much as 80 deg. These initial data exhibited an appreciable amount of scatter; however, a linear fit of the results indicated that the greatest shock phase lag occurred at a positive interblade phase angle. Photographs of the steady-state and unsteady flow fields reveal some of the features of the lambda shock wave on the suction surface of the airfoils.

Boldman, D. R.↗

NASA NTRS1993

Development of an engineering level prediction method for high angle of attack aerodynamics

The present work is concerned with predicting the unsteady flow considered to be the cause of the structural failure of twin vertical tail aircraft. An engineering tool has been produced for high angle of attack aerodynamics using the simplest physical models. The main innovation behind this work is its emphasis on the modeling of two key aspects of the dominant physics associated with high angle-of-attack airflows, namely unsteady separation and vortex breakdown.

Reisenthel, Patrick H.↗

NASA NTRS1987

A computational analysis of flow separation over five different airfoil geometries at high angles-of-attack

The flux splitting method of Van Leer (1982) and an implicit, upwind-biased, finite-volume scheme of Rumsey (1985) for solving the two-dimensional Reynolds-averaged Navier-Stokes equations were applied to analyze five airfoil geometries (the NACA 0012, 12-percent Joukowski, NACA 4412, NACA 65-015, and LS1-0417 sections), at high angles of attack. The effect of grid density and grid extent on the Strouhal number of the unsteady flow was computed, as well as the effect of variations in Mach number, Reynolds number, and angle of attack. It was found, that past an angle of attack of 20 deg, Strouhal numbers were independent of angle of attack (at a constant value of 0.115), agreeing well with the experimental findings of Tyler (1931).

Rumsey, Christopher L.↗

NASA NTRS1996

The Effect of Vane-Blade Spacing on Transonic Turbine Stage Performance

This paper presents the results of a computational study on the effect of axial spacing between the vane and blade rows of a transonic turbine stage. The study was performed on the mid-span section of a high-pressure turbine stage using a quasi-3D, unsteady Navier-Stokes solver that provides a fully interactive vane-blade unsteady flow solution. Three different cases were considered, corresponding to axial spacings of 20%, 40%, and 60% of the vane axial chord. The calculated vane and blade pressure distributions for the 40 percent case were found to compare favorably with experimental measurements acquired in a short-duration shock tunnel. In addition, the analysis shows a marked increase in the amplitude of the unsteady pressure fluctuations on the vane and blade surfaces as the spacing decreases. Time-averaged stage adiabatic efficiency predictions for each case are presented to show the effect of spacing on aerodynamic performance.

F J Kelecy↗

NASA NTRS1959

Performance of Typical Rear-Stage Axial-Flow Compressor Rotor Blade Row at Three Different Blade Setting Angles

A comparison of the performance of a single-stage rotor run at three different blade setting angles is presented. The rotor was of a design typical for a last stage of a multistage compressor. At each setting angle, the rotor blade row was operated from 53 to 100 percent of equivalent maximum speed (850 ft/sec tip speed) at constant inlet pressure. Hot-wire anemometry was used to observe rotating-stall and surge patterns in time unsteady flow. Results indicated that an increase in peak pressure ratio and an increase in maximum equivalent weight flow were obtained at each speed investigated when the blade setting angle was decreased. An increase in peak efficiency was achieved with decrease in blade setting angle for part of the range of speeds investigated. However, the peak efficiencies for the three blade setting angles were approximately the same at the maximum speed investigated. The flow ranges for all three configurations were about the same at minimum speed and decreased at almost the same rate when the rotative speed was increased through part of the range of speeds investigated. At maximum speed, the flow range for the smallest setting angle was considerably less than the flow range for the other two configurations. A decrease in efficiency and flow range for the smallest blade setting angle at maximum speed can be attributed primarily to a Mach number effect. In addition, because of the difference in projected axial chord lengths at the casing wall, some effect on performance could be expected from the change in three-dimensional flow occurring at the tip. Rotating-stall characteristics for the two smaller blade setting angles were essentially the same. Only surge could be detected for the largest blade setting angle in the unstable-flow region of operation.

Kussoy, Marvin I.↗