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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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NASA NTRS1999

A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

A three-dimensional Navier-Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Glenn Research Center experiment and has both circular cross-section and shaped film cooling holes. This complex geometry is modeled using a multi-block grid which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges.

Heidmann, James D.↗

NASA NTRS1999

N-S Simulations of Crow-Type Instabilities in Vortex Wake

Three-dimensional Navier-Stokes simulations of the Crow instability of wake vortices are conducted using large-eddy simulations without background turbulence. Sinusoidal displacement has been specified as the initial perturbation for the vortex system. The results have shown that the minimum Crow instability wavelength is about one vortex spacing shorter than predicted by Crow's linear stability theory. The planar- standing-wave-angle value and the amplitude indifference behavior agree with Crow's analysis. Simulations with periodicity in the axial direction have indicated minor influence of axial flow on the Crow instability.

Zheng, Z. C.↗

NASA NTRS1993

CFD zonal modeling of leading-edge ice effects for a complete aircraft

A simplified, uncoupled zonal procedure was utilized to assess the capability of numerically simulating icing effects on a Boeing 727-200 aircraft. The computational approach combines potential flow, plus boundary layer simulations by VSAERO for the un-iced aircraft forces and moments, with Navier-Stokes simulations by ARC3D for the incremental forces and moments due to iced components. These are compared with wind tunnel longitudinal force and moment data. Although the computational results compared favorably with the test data in the linear angle of attack range, it is clear that for general aircraft icing calculations, a multiblock Navier-Stokes code will be required for the viscous component of this zonal method.

Summa, J. M.↗

NASA NTRS2004

Modeling of High Speed Reacting Flows: Established Practices and Future Challenges

Computational fluid dynamics (CFD) has proven to be an invaluable tool for the design and analysis of high- speed propulsion devices. Massively parallel computing, together with the maturation of robust CFD codes, has made it possible to perform simulations of complete engine flowpaths. Steady-state Reynolds-Averaged Navier-Stokes simulations are now routinely used in the scramjet engine development cycle to determine optimal fuel injector arrangements, investigate trends noted during testing, and extract various measures of engine efficiency. Unfortunately, the turbulence and combustion models used in these codes have not changed significantly over the past decade. Hence, the CFD practitioner must often rely heavily on existing measurements (at similar flow conditions) to calibrate model coefficients on a case- by-case basis. This paper provides an overview of the modeled equations typically employed by commercial- quality CFD codes for high-speed combustion applications. Careful attention is given to the approximations employed for each of the unclosed terms in the averaged equation set. The salient features (and shortcomings) of common models used to close these terms are covered in detail, and several academic efforts aimed at addressing these shortcomings are discussed.

Baurle, R. A.↗

NASA NTRSDate not supplied

Progress Toward Evaluating Roughness-Resolved Aerothermodynamic Simulations for the Mars Sample Return Earth Entry System

This work pursues the efficient Navier-Stokes simulation, using NASA’s standard LAURA and DPLR simulation tools, of the flowfield and convective heating resulting from the patterned-roughness surface of the Mars Sample Return (MSR) Earth Entry System (EES). This type of simulation has been infeasible for past NASA missions. However, for EES, the combination of a patterned roughness heatshield approximated with a sinusoidal pattern, the zero degrees angle of attack nominal flow, and the supersonic boundary layer edge Mach number, all combine to make these roughness resolved grid simulations feasible. The patterned roughness and zero degrees angle of attack nominal flow allow the simulation domain to be reduced to a thin axial flowfield slice that captures a minimum of single period of the roughness pattern. With a periodic boundary condition applied along the boundaries of this slice, simulation on the reduced domain is identical to the full domain but with significant cost saving. The supersonic boundary layer edge Mach number prevents the upstream influence, which allows the smooth stagnation region to be frozen. This freezing of the stagnation region flow is required because of the numerical instability resulting from the tight spacing of the grid in the axial direction as the stagnation point is approached. Applying the developed analysis approach to a peak heating EES case results in a roughness augmentation factor for the convective heating of 1.32, which agrees well with correlations that predict values between 1.30 and 1.37. Analysis approach developed in this work provides the framework for future studies of turbulence modeling impacts and the influence of an ablating surface.

Earth Entry System↗

NASA NTRS2007

Simulations of Boundary-Layer Transition

For incompressible benchmark flows, we have demonstrated the capability of the parabolized stability equations (PSE) to simulate the transition process in excellent agreement with microscopic experiments and direct Navier-Stokes simulations at modest computational cost. Encouraged by these results, we have developed the PSE methodology of three-dimensional boundary-layers in general curvilinear coordinates for the range from low to hypersonic speeds, and for both linear and nonlinear problems. For given initial and boundary conditions, the approach permits simulations from receptivity through linear and secondary instabilities into the late stages of transition where significant changes in skin friction and heat transfer coefficients occur. We have performed transition simulations for a variety of two- and three-dimensional similarity solutions and for realistic flows over swept wings at subsonic and supersonic speeds, the pressure ans suction side of turbine blades at low and medium turbulence levels, and over a blunt cone at Mach number Ma = 8. We present selected results for different transition mechanisms with emphasis on the late stage of transition and the evolution of wall-shear stress and heat transfer.

Herbert, Thorwald↗

NASA NTRS2015

An Investigation of Transonic Resonance in a Mach 2.2 Round Convergent-Divergent Nozzle

Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.

DES↗

NASA NTRS1996

Development of Three-Dimensional Flow Code Package to Predict Performance and Stability of Aircraft with Leading Edge Ice Contamination

In the work reported herein, a simplified, uncoupled, zonal procedure is utilized to assess the capability of numerically simulating icing effects on a Boeing 727-200 aircraft. The computational approach combines potential flow plus boundary layer simulations by VSAERO for the un-iced aircraft forces and moments with Navier-Stokes simulations by NPARC for the incremental forces and moments due to iced components. These are compared with wind tunnel force and moment data, supplied by the Boeing Company, examining longitudinal flight characteristics. Grid refinement improved the local flow features over previously reported work with no appreciable difference in the incremental ice effect. The computed lift curve slope with and without empennage ice matches the experimental value to within 1%, and the zero lift angle agrees to within 0.2 of a degree. The computed slope of the un-iced and iced aircraft longitudinal stability curve is within about 2% of the test data. This work demonstrates the feasibility of a zonal method for the icing analysis of complete aircraft or isolated components within the linear angle of attack range. In fact, this zonal technique has allowed for the viscous analysis of a complete aircraft with ice which is currently not otherwise considered tractable.

Aircraft icing↗

NASA NTRS2018

Numerical Simulation of Dynamic Stall Using Near-Body Adaptive Mesh Refinement

Time-dependent Reynolds-averaged Navier-Stokes simulations have been carried out for a NACA 0012 airfoil and a flexible UH-60A rotor undergoing dynamic stall. The OVERFLOW computational fluid dynamics (CFD) code is used to explore the use of near-body adaptive mesh refinement (NB-AMR) for the first time on a flexible helicopter rotor in forward flight. Emphasis is placed on understanding the temporal and spatial convergence of the solutions, their numerical efficiency and stability, and establishing a grid-converged solution. Some of the remaining differences between CFD and flight-test measurements are discussed. Time-dependent flow visualization is used to provide an improved understanding of the physical mechanisms involved with two-dimensional and three-dimensional dynamic stall, and the NB-AMR process.

Near-Body↗

NASA NTRS2006

Blunt Body Aerodynamics for Hypersonic Low Density Flows

Numerical simulations are performed for the Apollo capsule from the hypersonic rarefied to the continuum regimes. The focus is on flow conditions similar to those experienced by the Apollo 6 Command Module during the high altitude portion of its reentry. The present focus is to highlight some of the current activities that serve as a precursor for computational tool assessments that will be used to support the development of aerodynamic data bases for future capsule flight environments, particularly those for the Crew Exploration Vehicle (CEV). Results for aerodynamic forces and moments are presented that demonstrate their sensitivity to rarefaction; that is, free molecular to continuum conditions. Also, aerodynamic data are presented that shows their sensitivity to a range of reentry velocities, encompassing conditions that include reentry from low Earth orbit, lunar return, and Mars return velocities (7.7 to 15 km/s). The rarefied results obtained with direct simulation Monte Carlo (DSMC) codes are anchored in the continuum regime with data from Navier-Stokes simulations.

Moss, James N.↗

NASA NTRS1989

Stereo image visualization of numerically simulated turbulence

Stereo imaging techniques are used in the current study to analyze the results of a direct Navier-Stokes simulation of a turbulent boundary layer. Coherent structures embedded within the unsteady, highly three-dimensional turbulent flow fields are significantly enhanced through the use of stereo computer graphic renderings. In the paper, the geometrical foundations of two different methods for generating stereo images are reviewed. Sufficient detail is provided to enable readers to code stereo display algorithms from scratch. The effects of varying the geometric parameters of the stereo algorithm are demonstrated. Sample image pairs from animations of the turbulence numerical simulation are presented in color. Stereo imagery, especially when used with time-evolving data, provides new insight into the physical processes responsible for the generation an maintenance of turbulence within boundary layers.

Robinson, Stephen K.↗

NASA NTRS2002

Unsteady Simulation of a Landing-Gear Flow Field

This paper presents results of an unsteady Reynolds-averaged Navier-Stokes simulation of a landing-gear flow field. The geometry of the four-wheel landing gear assembly consists of several of the fine details including the oleo-strut, two diagonal struts, a door, yokes/pin and a flat-plate simulating the wing surface. The computational results, obtained by using 13.3 million grid points, are presented with an emphasis on the characteristics of the unsteadiness ensuing from different parts of the landing-gear assembly, including vortex shedding patterns and frequencies of dominant oscillations. The results show that the presence of the diagonal struts and the door significantly influence the flow field. Owing to the induced asymmetry, vortices are shed only from one of the rear wheels and not the other. Present computations also capture streamwise vortices originating from the upstream corners of the door.

Li, Fei↗

NASA NTRS1996

Unsteady Simulation of the Viscous Flow About a V-22 Rotor and Wing in Hover

Results of an unsteady thin-layer Navier-Stokes simulation of a 0.658-scale V-22 rotor and wing configuration in hover are presented. All geometric components of the flapped-wing and rotor test rig, including rotor blades, are accurately modeled. Rotor motion and rotor/airframe interference effects are simulated directly using moving body overset grid methods. Tiltrotor hover aerodynamics are visualized via unsteady particle trace images. Wing download predictive ability is demonstrated. Simulation results are compared with experimental data.

Meakin, Robert L.↗

NASA NTRS2010

Automatic Tip Vortex Core Profiling for Numerical Flow Simulations of Rotorcraft in Hover

An automated approach is presented that extracts visual and quantitative data from vortex cores produced by Navier-Stokes simulations of rotorcraft in hover mode. This approach extracts contiguous rotor tip vortex-core trajectories, cross-flow velocity profiles, and vortex-core diameter variation with wake age (azimuth angle). This automated approach is faster and more accurate than a conventional manual approach. Moreover, this new approach allows for an efficient way to quantitatively compare vortex-core profiles from different flow simulations, e.g., grid resolution studies, and validate computed results with experimental data

Kao, David L.↗

NASA NTRSDate not supplied

Wall-modeled LES of Flow Over a Gaussian Bump

We perform wall-modeled large eddy simulations (WMLES) of turbulent flow over a Gaussian-shaped bump geometry, to assess its performance in the accelerating and separation regions of the flow. The flow conditions are based on the ongoing CFD validation experiments of Slotnick [NATO STO-MP-AVT-307, 2019]. The oncoming flow Mach number is 0.176, and two Reynolds numbers are simulated that are about 10000 and 36000 based on boundary layer thickness upstream of the bump. Preliminary Reynolds-averaged Navier-Stokes simulations are first performed to assess the effects of Mach number, Reynolds number, tunnel top and sidewall effects. Finally, WMLES results with an equilibrium wall model will be presented at two Reynolds numbers to assess their performance for this flow by making detailed comparisons with available experimental and higher-fidelity numerical data.

turbulence, wall model, large eddy simulation↗

NASA NTRS1990

Supercomputer applications in gas turbine flowfield simulation

The numerical simulation of the unsteady three-dimensional viscous flow in a gas turbine stage is considered. Results from a three-dimensional time-accurate Navier-Stokes simulation of rotor-stator interaction in an axial turbine stage are presented. The present study uses a fine grid in the spanwise direction to better resolve the complex three-dimensional flowfield, and complements earlier reported coarse-grid calculations. Several different features of the flowfield are analyzed and compared to earlier calculations and to experimental data whenever possible. Computer animation techniques are used to visualize various unsteady three-dimensional features of the flow. The results demonstrate the capabilities of current computing hardware in obtaining accurate simulations of unsteady flows in turbomachines.

Madavan, Nateri K.↗

NASA NTRS2003

A Computational Study of the Abrupt Wing Stall (AWS) Characteristics for Various Fighter Jets: F/A-18E and F-16C - Part 1

Steady state Computational Fluid Dynamic (CFD) simulations are used to gain an understanding of the physics behind the Abrupt Wing Stall (AWS) phenomenon and to arrive at static Figures of Merit (FOMs). Navier-Stokes simulations are conducted using the NASA Langley developed TetrUSS simulation suite which is based on tetrahedral, unstructured grids. The physics of the AWS phenomenon is understood by comparing CFD simulation results on two aircraft; a pre-production F/A-18E configuration which exhibits AWS phenomenon under certain geometric and flow conditions, and an F-16C aircraft configuration that does not. The CFD code is first validated against two sets of experimental data to build confidence in its use for the problem of AWS. An attempt is then made to understand the possible causes of AWS by analyzing and comparing the detailed flow fields between the two configurations under a variety of flow conditions. Based on this approach, a number of static Figures of Merit are developed to predict the potential existence of AWS. The FOMs include the break in the lift and wing root bending moment versus angle-of-attack (AoA) curves and the rate of change of sectional lift with respect to AoA. A companion paper, in Part II, describes a similar CFD study on two other aircraft, the AV-8B Harrier and F/A-18C. Results from both these studies, as well as other CFD studies conducted as part of the AWS program are used to recommend a CFD procedure for predicting the existence of AWS in future aerospace designs.

Parikh, Paresh↗

NASA NTRS1999

Development and Validation of a Fast, Accurate and Cost-Effective Aeroservoelastic Method on Advanced Parallel Computing Systems

Progress to date towards the development and validation of a fast, accurate and cost-effective aeroelastic method for advanced parallel computing platforms such as the IBM SP2 and the SGI Origin 2000 is presented in this paper. The ENSAERO code, developed at the NASA-Ames Research Center has been selected for this effort. The code allows for the computation of aeroelastic responses by simultaneously integrating the Euler or Navier-Stokes equations and the modal structural equations of motion. To assess the computational performance and accuracy of the ENSAERO code, this paper reports the results of the Navier-Stokes simulations of the transonic flow over a flexible aeroelastic wing body configuration. In addition, a forced harmonic oscillation analysis in the frequency domain and an analysis in the time domain are done on a wing undergoing a rigid pitch and plunge motion. Finally, to demonstrate the ENSAERO flutter-analysis capability, aeroelastic Euler and Navier-Stokes computations on an L-1011 wind tunnel model including pylon, nacelle and empennage are underway. All computational solutions are compared with experimental data to assess the level of accuracy of ENSAERO. As the computations described above are performed, a meticulous log of computational performance in terms of wall clock time, execution speed, memory and disk storage is kept. Code scalability is also demonstrated by studying the impact of varying the number of processors on computational performance on the IBM SP2 and the Origin 2000 systems.

Goodwin, Sabine A.↗