SEARCH · Engineering Papers

Results for “Structured Grids”

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 577 records · Page 32

All sources NASA NTRS DOE OSTI

NASA NTRS2003

Active Control of F/A-18 Vertical Tail Buffeting using Piezoelectric Actuators

Vertical tail buffeting is a serious multidisciplinary problem that limits the performance of twin-tail fighter aircraft. The buffet problem occurs at high angles of attack when the vortical flow breaks down ahead of the vertical tails resulting in unsteady and unbalanced pressure loads on the vertical tails. This paper describes a multidisciplinary computational investigation for buffet load alleviation of full F/A-18 aircraft using distributed piezoelectric actuators. The inboard and outboard surfaces of the vertical tail are equipped with piezoelectric actuators to control the buffet responses in the first bending and torsion modes. The electrodynamics of the smart structure are expressed with a three-dimensional finite element model. A single-input-single-output controller is designed to drive the active piezoelectric actuators. High-fidelity multidisciplinary analysis modules for the fluid dynamics, structure dynamics, electrodynamics of the piezoelectric actuators, fluid-structure interfacing, and grid motion are integrated into a multidisciplinary computing environment that controls the temporal synchronization of the analysis modules. Peak values of the power spectral density of tail tip acceleration are reduced by as much as 22% in the first bending mode and by as much as 82% in the first torsion mode. RMS values of tip acceleration are reduced by as much as 12%.

Sheta, Essam F.↗

NASA NTRS2003

An Active Smart Material Control System for F/A-18 Buffet Alleviation

The vertical tail buffet problem of fighter aircraft occurs at high angles of attack when the vortical flow breaks down ahead of the vertical tails resulting in unsteady and unbalanced pressure loads on the vertical tails. The buffet loads imposed upon the vertical tails resulted in a premature fatigue failure of the tails, and consequently limits the performance and super maneuverability of twin-tail fighter aircraft. An active smart material control system using distributed piezoelectric actuators has been developed for buffet alleviation and is presented. The inboard and outboard surfaces of the vertical tail are equipped with piezoelectric actuators to control the buffet responses in the first bending and torsion modes. The electrodynamics of the piezoelectric actuators are expressed with a three-dimensional finite-element model. A single-input-single-output controller is designed to drive the active piezoelectric actuators. High-fidelity multidisciplinary analysis modules for the fluid dynamics, structure dynamics, electrodynamics of the piezoelectric actuators, control law, fluid structure interfacing, and grid motion are integrated into a multidisciplinary computing environment that controls the temporal synchronization of the analysis modules. At 30 degree angle of attack, RMS values of tip acceleration are reduced by as much as 12%. The peak values of the power spectral density of tail-tip acceleration are reduced by as much as 22% in the first bending mode and by as much as 82% in the first torsion mode. The actively controlled piezoelectric actuators were also effective in adding damping at wide range of angles of attack.

Sheta, Essam F.↗

NASA NTRS1998

Aeroelastic Deflection of NURBS Geometry

The purpose of this paper is to present an algorithm for using NonUniform Rational B-Spline (NURBS) representation in an aeroelastic loop. The algorithm is based on creating a least-squares NURBS surface representing the aeroelastic defection. The resulting NURBS surfaces are used to update either the original Computer- Aided Design (CAD) model, Computational Structural Mechanics (CSM) grid or the Computational Fluid Dynamics (CFD) grid. Results are presented for a generic High-Speed Civil Transport (HSCT).

Samareh, Jamshid A.↗

NASA NTRS2005

Progress Towards Computational Method for Circulation Control Airfoils

The compressible Reynolds-averaged Navier-Stokes equations are solved for circulation control airfoil flows. Numerical solutions are computed with both structured and unstructured grid solvers. Several turbulence models are considered, including the Spalart-Allmaras model with and without curvature corrections, the shear stress transport model of Menter, and the k-enstrophy model. Circulation control flows with jet momentum coefficients of 0.03, 0.10, and 0.226 are considered. Comparisons are made between computed and experimental pressure distributions, velocity profiles, Reynolds stress profiles, and streamline patterns. Including curvature effects yields the closest agreement with the measured data.

Swanson, R. C.↗

NASA NTRS2004

Adjoint Algorithm for CAD-Based Shape Optimization Using a Cartesian Method

Adjoint solutions of the governing flow equations are becoming increasingly important for the development of efficient analysis and optimization algorithms. A well-known use of the adjoint method is gradient-based shape optimization. Given an objective function that defines some measure of performance, such as the lift and drag functionals, its gradient is computed at a cost that is essentially independent of the number of design variables (geometric parameters that control the shape). More recently, emerging adjoint applications focus on the analysis problem, where the adjoint solution is used to drive mesh adaptation, as well as to provide estimates of functional error bounds and corrections. The attractive feature of this approach is that the mesh-adaptation procedure targets a specific functional, thereby localizing the mesh refinement and reducing computational cost. Our focus is on the development of adjoint-based optimization techniques for a Cartesian method with embedded boundaries.12 In contrast t o implementations on structured and unstructured grids, Cartesian methods decouple the surface discretization from the volume mesh. This feature makes Cartesian methods well suited for the automated analysis of complex geometry problems, and consequently a promising approach to aerodynamic optimization. Melvin et developed an adjoint formulation for the TRANAIR code, which is based on the full-potential equation with viscous corrections. More recently, Dadone and Grossman presented an adjoint formulation for the Euler equations. In both approaches, a boundary condition is introduced to approximate the effects of the evolving surface shape that results in accurate gradient computation. Central to automated shape optimization algorithms is the issue of geometry modeling and control. The need to optimize complex, "real-life" geometry provides a strong incentive for the use of parametric-CAD systems within the optimization procedure. In previous work, we presented an effective optimization framework that incorporates a direct-CAD interface. In this work, we enhance the capabilities of this framework with efficient gradient computations using the discrete adjoint method. We present details of the adjoint numerical implementation, which reuses the domain decomposition, multigrid, and time-marching schemes of the flow solver. Furthermore, we explain and demonstrate the use of CAD in conjunction with the Cartesian adjoint approach. The final paper will contain a number of complex geometry, industrially relevant examples with many design variables to demonstrate the effectiveness of the adjoint method on Cartesian meshes.

Nemec, Marian↗

NASA NTRS2009

Lessons Learned from Numerical Simulations of the F-16XL Aircraft at Flight Conditions

Nine groups participating in the Cranked Arrow Wing Aerodynamics Project International (CAWAPI) project have contributed steady and unsteady viscous simulations of a full-scale, semi-span model of the F-16XL aircraft. Three different categories of flight Reynolds/Mach number combinations were computed and compared with flight-test measurements for the purpose of code validation and improved understanding of the flight physics. Steady-state simulations are done with several turbulence models of different complexity with no topology information required and which overcome Boussinesq-assumption problems in vortical flows. Detached-eddy simulation (DES) and its successor delayed detached-eddy simulation (DDES) have been used to compute the time accurate flow development. Common structured and unstructured grids as well as individually-adapted unstructured grids were used. Although discrepancies are observed in the comparisons, overall reasonable agreement is demonstrated for surface pressure distribution, local skin friction and boundary velocity profiles at subsonic speeds. The physical modeling, steady or unsteady, and the grid resolution both contribute to the discrepancies observed in the comparisons with flight data, but at this time it cannot be determined how much each part contributes to the whole. Overall it can be said that the technology readiness of CFD-simulation technology for the study of vehicle performance has matured since 2001 such that it can be used today with a reasonable level of confidence for complex configurations.

Rizzi, Arthur↗

NASA NTRS2009

What was Learned from Numerical Simulations of F-16XL (CAWAPI) at Flight Conditions

Nine groups participating in the CAWAPI project have contributed steady and unsteady viscous simulations of a full-scale, semi-span model of the F-16XL aircraft. Three different categories of flight Reynolds/Mach number combinations were computed and compared with flight-test measurements for the purpose of code validation and improved understanding of the flight physics. Steady-state simulations are done with several turbulence models of different complexity with no topology information required and which overcome Boussinesq-assumption problems in vortical flows. Detached-eddy simulation (DES) and its successor delayed detached-eddy simulation (DDES) have been used to compute the time accurate flow development. Common structured and unstructured grids as well as individually-adapted unstructured grids were used. Although discrepancies are observed in the comparisons, overall reasonable agreement is demonstrated for surface pressure distribution, local skin friction and boundary velocity profiles at subsonic speeds. The physical modeling, be it steady or unsteady flow, and the grid resolution both contribute to the discrepancies observed in the comparisons with flight data, but at this time it cannot be determined how much each part contributes to the whole. Overall it can be said that the technology readiness of CFD-simulation technology for the study of vehicle performance has matured since 2001 such that it can be used today with a reasonable level of confidence for complex configurations.

Rizzi, Arthur↗

NASA NTRS2010

Micro-Ramps for External Compression Low-Boom Inlets

The application of vortex generators for flow control in an external compression, axisymmetric, low-boom concept inlet was investigated using RANS simulations with three-dimensional (3-D), structured, chimera (overset) grids and the WIND-US code. The low-boom inlet design is based on previous scale model 1- by 1-ft wind tunnel tests and features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. Validation of the methodology was first performed for micro-ramps in supersonic flow on a flat plate with and without oblique shocks. For the inlet configuration, simulations with several types of vortex generators were conducted for positions both upstream and downstream of the terminating normal shock. The performance parameters included incompressible axisymmetric shape factor, separation area, inlet pressure recovery, and massflow ratio. The design of experiments (DOE) methodology was used to select device size and location, analyze the resulting data, and determine the optimal choice of device geometry. The optimum upstream configuration was found to substantially reduce the post-shock separation area but did not significantly impact recovery at the aerodynamic interface plane (AIP). Downstream device placement allowed for fuller boundary layer velocity profiles and reduced distortion. This resulted in an improved pressure recovery and massflow ratio at the AIP compared to the baseline solid-wall configuration.

Rybalko, Michael↗

NASA NTRS2016

Unsteady Aerodynamic Force Sensing from Measured Strain

A simple approach for computing unsteady aerodynamic forces from simulated measured strain data is proposed in this study. First, the deflection and slope of the structure are computed from the unsteady strain using the two-step approach. Velocities and accelerations of the structure are computed using the autoregressive moving average model, on-line parameter estimator, low-pass filter, and a least-squares curve fitting method together with analytical derivatives with respect to time. Finally, aerodynamic forces over the wing are computed using modal aerodynamic influence coefficient matrices, a rational function approximation, and a time-marching algorithm. A cantilevered rectangular wing built and tested at the NASA Langley Research Center (Hampton, Virginia, USA) in 1959 is used to validate the simple approach. Unsteady aerodynamic forces as well as wing deflections, velocities, accelerations, and strains are computed using the CFL3D computational fluid dynamics (CFD) code and an MSC/NASTRAN code (MSC Software Corporation, Newport Beach, California, USA), and these CFL3D-based results are assumed as measured quantities. Based on the measured strains, wing deflections, velocities, accelerations, and aerodynamic forces are computed using the proposed approach. These computed deflections, velocities, accelerations, and unsteady aerodynamic forces are compared with the CFL3D/NASTRAN-based results. In general, computed aerodynamic forces based on the lifting surface theory in subsonic speeds are in good agreement with the target aerodynamic forces generated using CFL3D code with the Euler equation. Excellent aeroelastic responses are obtained even with unsteady strain data under the signal to noise ratio of -9.8dB. The deflections, velocities, and accelerations at each sensor location are independent of structural and aerodynamic models. Therefore, the distributed strain data together with the current proposed approaches can be used as distributed deflection, velocity, and acceleration sensors. This research demonstrates the feasibility of obtaining induced drag and lift forces through the use of distributed sensor technology with measured strain data. An active induced drag control system thus can be designed using the two computed aerodynamic forces, induced drag and lift, to improve the fuel efficiency of an aircraft. Interpolation elements between structural finite element grids and the CFD grids and centroids are successfully incorporated with the unsteady aeroelastic computation scheme. The most critical technology for the success of the proposed approach is the robust on-line parameter estimator, since the least-squares curve fitting method depends heavily on aeroelastic system frequencies and damping factors.

shape sensing↗

NASA NTRSDate not supplied

Scale resolving simulations of the NASA Juncture Flow Model using the LAVA solver

Two distinct simulation methodologies: Delayed Detached Eddy Simulation (DDES) and stress based Wall-modelled Large Eddy Simulation (WMLES) are evaluated using structured overset curvilinear grids for the NASA juncture flow model. While both methodologies are shown to mitigate the primary shortcomings of steady state Reynolds Averaged Navier Stokes(RANS) simulations, several unresolved aspects are identified. Strong sensitivity to RANS-typegrid refinement is observed in the DDES with a substantial deterioration of the solution quality with increasing spatial resolution associated with deficiencies in the shielding function. Mean profiles for attached boundary layers on the fuselage show spurious inflections suggesting modelled stress depletion on finer grids. Lower numerical dissipation in terms of spatial discretization and time step size is seen to improve the solution quality, and the advantage of using RC and QCR2000 corrections in the underlying RANS closure is demonstrated for DDES on RANS-type meshes. Equilibrium Wall-Modelled LES used grids that resolved the tripping dots over the fuselage nose and the wing leading edge, consistent with the experiment setup. These simulations resulted in a cost-competitive approach compared to DDES on RANS-type grids. Although the agreement between WMLES predictions of first and second order single point statistics with experimental measurements is promising, a fundamental shortcoming is noted in terms of an overshoot in streamwise momentum in corner regions of the wing-fuselage juncture. This excess momentum subsequently delays the onset of separation,thereby resulting in an underprediction in length of the separation bubble. Some quantitative sensitivity to numerical discretization is observed; lowering of numerical dissipation shows better agreement with the experiment.

T^3↗

NASA NTRSDate not supplied

Wall-Modeled Large-Eddy Simulations of Jet Noise in Flight Conditions

A campaign of wall-modeled large-eddy simulations (WMLES) using structured curvilinear overlapping grids has been performed with the Launch Ascent and Vehicle Aerodynamics (LAVA) computational fluid dynamics (CFD) software to predict jet noise for single-stream axisymmetric round jets. The simulations address the new Prediction Uncertainty Reduction (PUR) technical challenge within the context of NASA’s Commercial Supersonic Technology (CST) project. The focus of this effort is to generate a simulation database for single-stream axisymmetric round nozzles at several operating points both for static (no ambient co-flow), and in-flight (M ͚ =0.3 co-flow) conditions. The operating conditions range in jet exit Mach number from 0.38 to 1.1 with nozzle temperature ratios (NTR) from 0.84 to 2.7. The effect of the flight-stream on far-field noise sound spectra is assessed. Comparison of LES predictions to microphone array measurements demonstrate good agreement within the resolved frequency range. A dip in the predicted low frequency noise spectra for observers between 120° and 145° is observed. This dip is smaller for lower Mach numbers and seems to be correlated to the Mach wave radiation angle. The Mach 1.1 jet shows broadband-shock associated noise. While the onset of BBSN appears to be captured correctly in WMLES, some differences in its magnitude and the prominent frequency at which it occurs persist between the experiment and the simulations. The effect of the outer nozzle boundary layer state created by the co-flow is assessed and shows to be important for accurate comparisons with experiments. A change of 2.5dB between a slip-wall condition and a artificially thickened turbulent boundary layer was observed. In addition, simulations were performed with an alternative nozzle geometry that includes a internal plug and has twice the nozzle exit diameter. These two configurations resulted in very comparable spectra which is consistent with experimental observations. A generally stronger deviation from experimental results is observed for in-flight cases compared to static conditions. The applicability and correct usage of acoustic analogies used for far-field propagation with strong turbulent co-flows needs to be investigated more systematically using canonical problems to improve comparisons between experiments and WMLES. This is especially true for coherent noise sources seen in BBSN.

CST↗

NASA NTRS2023

Unsteady RANS and Scale Resolving Simulations of Open Rotor Noise

Computational simulations using structured curvilinear overset grids with the Launch Ascent and Vehicle Aerodynamics (LAVA) framework are presented for predicting the aerodynamic and aeroacoustic behavior of a contra-rotating open rotor (CROR) propulsion system using a newly developed and computationally efficient sliding mesh technique. A comprehensive evaluation of two turbulence closure treatments is carried out as part of the scope of this work: Unsteady Reynolds-Averaged Navier-Stokes Simulation (URANS), and Hybrid RANS/LES (HRLES). The computed results are then validated against existing experimental measurements at both the nominal take-off and cruise flight conditions. To propagate the noise generated by the near-field CROR system, aeroacoustic analysis based on a frequency-domain permeable surface Ffowcs Williams-Hawkings (FW-H) approach is applied.

ARMD↗

NASA NTRS2023

Unsteady RANS and Scale Resolving Simulations of Open Rotor Noise

ARMD↗

NASA NTRSDate not supplied

Scale Resolving Simulations of Contra Rotating Open Rotor Noise Prediction

The development of breakthrough open-rotor propulsion systems that meet performance, fuel efficiency, and noise requirements for commercial viability relies heavily on the availability of reliable and computationally efficient aerodynamic performance and noise prediction tools.This investigation showcases the proficiency of the structured curvilinear overset grid Computational Fluid Dynamics solver implemented in the Launch, Ascent, and Vehicle Aerodynamics framework for predicting the aerodynamic and aeroacoustic characteristics of a Contra-RotatingOpen Rotor propulsion system. Additionally, the study includes comprehensive sensitivity analyses employing a Hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation - based scale-resolving turbulence closure treatment. The computed results are validated through comparison with existing experimental measurements.

ARMD↗

NASA NTRSDate not supplied

Simulating Transonic Buffet Aerodynamics for the Boeing Transonic Truss-Braced Wing Aircraft

The flow past the Boeing Transonic Truss-Braced Wing (TTBW) aircraft is simulated with a Hybrid Reynolds-Averaged Navier Stokes Large-Eddy Simulations (HRLES) turbulence modeling approach to study transonic buffet onset and evaluate the predictive capability of the numerical approach. All simulations used structured overset curvilinear grids for the free-air configuration, with all simulations performed using the Launch, Ascent, and Vehicle Aerodynamics (LAVA) computational framework. Aerodynamic loads and surface pressure obtained from the HRLES solutions are compared to results from unsteady RANS (URANS) and experiments conducted at NASA Ames 11-by 11-Foot Transonic Wind Tunnel. Overall good agreement is obtained in the predicted loads and surface pressure. Unsteady pressure data from HRLES and URANS simulations is utilized to compute power spectral density (PSD) to predict transonic buffet onset and compared to the Kulite data obtained from the experiment. Comparisons of the PSD spectrum also show reasonable agreement with the wind tunnel data, with HRLES and URANS predicting buffet onset at a slightly earlier angle-of-attack than the experiment.

AATT↗

NASA NTRS1999

Data Structure and Parallel Decomposition Considerations on a Fibonacci Grid

The Fibonacci grid, proposed by Swinbank and Purser (see companion abstract), provides attractive properties for global numerical atmospheric prediction by offering an optimally homogeneous, geometrically regular, and approximately isotropic discretization, with only the polar regions requiring special numerical treatment. It is a mathematical idealization, applied to the sphere, of the multi-spiral patterns often found in botanical structures, such as in pine cones and sunflower heads. Computationally, it is natural to organize the domain, into zones, in each of which the same pair, or triple, of "Fibonacci spirals" dominate. But the further subdivision of such zones into "tiles" of a shape and size suitable for distribution to the processors of a massively parallel computer requires very careful consideration if the subsequent spatial computations along the respective spirals, especially those computations (such as compact differencing schemes) that involve recursion, can be implemented in an efficient "load-balanced "manner without requiring excessive amounts of inter-processor communications. In this paper we show how certain "number theoretic" properties of the Fibonacci sequence (whose numbers prescribe the multiplicity of successive spirals) may be exploited in the decomposition of grid zones into tidy arrangements of triangular grid tiles, each tile possessing one side approximately parallel to the constant-latitude zone boundary. We also describe how the spatially recursive processes may be decomposed across such a tiling, and the directionality of the recursions reversed on alternate grid lines, to ensure a very high degree of load balancing throughout the execution of the computations required for one time step of a global model.

Michalakes, John↗

NASA NTRS1984

A statistical evaluation and comparison of VISSR Atmospheric Sounder (VAS) data

In order to account for the temporal and spatial discrepancies between the VAS and rawinsonde soundings, the rawinsonde data were adjusted to a common hour of release where the new observation time corresponded to the satellite scan time. Both the satellite and rawinsonde observations of the basic atmospheric parameters (T Td, and Z) were objectively analyzed to a uniform grid maintaining the same mesoscale structure in each data set. The performance of each retrieval algorithm in producing accurate and representative soundings was evaluated using statistical parameters such as the mean, standard deviation, and root mean square of the difference fields for each parameter and grid level. Horizontal structure was also qualitatively evaluated by examining atmospheric features on constant pressure surfaces. An analysis of the vertical structure of the atmosphere were also performed by looking at colocated and grid mean vertical profiles of both the satellite and rawinsonde data sets. Highlights of these results are presented.

Jedlovec, G. J.↗

NASA NTRS1989

Application of a lower-upper implicit scheme and an interactive grid generation for turbomachinery flow field simulations

A finite-volume lower-upper (LU) implicit scheme is used to simulate an inviscid flow in a tubine cascade. This approximate factorization scheme requires only the inversion of sparse lower and upper triangular matrices, which can be done efficiently without extensive storage. As an implicit scheme it allows a large time step to reach the steady state. An interactive grid generation program (TURBO), which is being developed, is used to generate grids. This program uses the control point form of algebraic grid generation which uses a sparse collection of control points from which the shape and position of coordinate curves can be adjusted. A distinct advantage of TURBO compared with other grid generation programs is that it allows the easy change of local mesh structure without affecting the grid outside the domain of independence. Sample grids are generated by TURBO for a compressor rotor blade and a turbine cascade. The turbine cascade flow is simulated by using the LU implicit scheme on the grid generated by TURBO.

Choo, Yung K.↗