SEARCH · Engineering Papers

Results for “CFD modeling”

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 1,495 records · Page 83

All sources NASA NTRS DOE OSTI

NASA NTRSDate not supplied

Leveraging Large Eddy Simulations to Characterize Aeroacoustic Environments on Launch Vehicles

NASA's Space Launch System (SLS) program has relied on multiple wind tunnel campaigns to predict expected aeroacoustic environments. Time, cost, Reynolds number matching, contamination from test section acoustics, and limitations in spatial resolution are all factors that impact the use experimental data for development of launch vehicle environments. Methodological and computational advances over the last decade are enabling scale resolving computational fluid dynamics (CFD) simulations to be completed that can address some the challenges with conventional datasets, enabling more complete/accurate modeling of dynamic loading events on ascent. These simulations are classified as so-called "Large-Eddy Simulations" (LES). As the name implies, these simulations seek to directly resolve the turbulence in the largest, most-energetic turbulent structures, or eddies, and the dissipation of the smallest eddies in the flow field, is modeled. Like anything else, these simulations still must undergo rigorous testing and validation for the results to ultimately be trusted in the analysis of a launch vehicle. EV33 has begun to conduct some of this early-stage testing and validation to better support NASA missions in the future.

Jonathan Boustani↗

NASA NTRSDate not supplied

LOFTID Surface Heating Reconstruction

On November 10, 2022, the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) reentry vehicle launched to low-Earth orbit aboard a United Launch Alliance Atlas V rocket out of Vandenberg Air Force Base. The aeroshell, the largest Hypersonic Inflatable Aerodynamic Decelerator (HIAD) ever flown, was inflated to its full 6-meter diameter before successful re-entry into the atmosphere. The aeroshell was heavily instrumented in order to understand its behavior during entry. There were 82 thermocouples (TCs) distributed across the aeroshell, with 22 integrated into the flexible thermal protection system (FTPS) on the rigid nose, 36 in the FTPS on the deployable structure, and 24 on the inflatable structure. TCs were placed at different depths throughout the FTPS. Those nearest to the surface were located just beneath the two SiC outer fabric layers. The near-surface TCs on the rigid nose were Type R with flame spray alumina insulation, while those on the flank were Type N with mica/ceramic insulation. Additionally, a radiometer was placed at the center of the nose surrounded by four total heat flux gauges in a cruciform configuration at a radius of 0.41 m. The nose instrumentation is shown in Fig. 1 and a cross-section of the aeroshell with all TC locations is shown in Fig. 2. The objective of this work was to use the temperatures measured by the TCs during flight to estimate the surface heat rate across the aeroshell throughout the period of re-entry by inverse analysis methodology. The results were used to evaluate the fidelity of measurements from the total heat flux gauges on the nose, determine the surface heat flux at aeroshell locations where gauges were not present, and compare to pre-flight CFD-based heating predictions. Inversely estimated surface heat flux continues to be used to correlate FTPS thermal models to reconstruct in-flight thermal response.

LOFTID↗

NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code, are used with a blowing boundary condition in the DPLR hypersonic CFD code. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver, is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Ablation↗

NASA NTRSDate not supplied

HLPW-4/GMGW-3: Overview and Workshop Summary

The Fourth AIAA CFD High Lift Prediction Workshop and the Third Geometry and Mesh Generation Workshop were held collaboratively with the common goal of assessing the numerical prediction capability of current-generation computational fluid dynamics (CFD) technology for swept, medium/high-aspect-ratio wings in high-lift configurations. A key aspect of this joint endeavor was the use of Technology Focus Groups, an innovative new approach for workshops involving close collaboration between participants. These groups, which included both mesh generation and flow solver experts, worked to accelerate advancements for their particular methodologies by addressing key questions of importance {\em prior} to the workshop. The high-lift version of the NASA Common Research Model (CRM-HL) configuration was the focus of this workshop. Measured experimental wind tunnel data were available for comparison. The workshop also included a two-dimensional turbulence model verification exercise based on the CRM-HL wing shape. Altogether, 44 participants submitted a total of 184 data sets of CFD results. This paper provides a high-level summary of the results and conclusions from the workshop. Like at past workshops, fixed-grid Reynolds-averaged Navier-Stokes continued to be inaccurate and inconsistent for high lift. However, mesh adaptation definitively brought more consistency. Scale-resolving methods appeared most promising for predicting high-lift flow physics.

Christopher L Rumsey↗

NASA NTRS1993

Computational aeroelasticity

Recent large-scale studies of computational unsteady aerodynamics for aeroelastic applications are reviewed. The variety of fluid dynamic flow models available to address such computations are illustrated in two cases: (1) 2D Navier-Stokes (NS) equations illustrating the highest level of modeling usually employed and (2) the transonic small-disturbance potential equation representing the entry level for nonlinear flow modeling. Estimates of computer resources necessary to produce accurate converged results are given. Application of potential and NS equation codes for flows which are generally at lower angles and high speeds are addressed, including (1) the treatment of wings and configuration details using potential equation codes and (2) recent NS code calculations of complete vehicle configurations. Attention is given to high-angle conditions involving separated vortex-dominated flows. Accuracy requirements for vortex shedding over airfoils is discussed and the calculation of vorticity convected over significant distances is addressed. Steady CFD computations about delta wings at high angles are given, including a discussion of the required level of fluid dynamic flow modeling. Recent results on unsteady 'buffetlike flow' about complete vehicle models are reviewed.

Edwards, John W.↗

NASA NTRS1993

Three-D Flow Analysis of the Alternate SSME HPOT TAD

This paper describes the results of numerical flow analyses performed in support of design development of the Space Shuttle Main Engine Alternate High Pressure Oxidizer Turbine Turn-around duct (TAD). The flow domain has been modeled using a 3D, Navier-Stokes, general purpose flow solver. The goal of this effort is to achieve an alternate TAD exit flow distribution which closely matches that of the baseline configuration. 3D Navier Stokes CFD analyses were employed to evaluate numerous candidate geometry modifications to the TAD flowpath in order to achieve this goal. The design iterations are summarized, as well as a description of the computational model, numerical results and the conclusions based on these calculations.

Kubinski, Cheryl A.↗

NASA NTRSDate not supplied

Coupling CFD and Material Response for Analysis of Mars Entry

In computing the response of an ablating thermal protection system during atmospheric entry, the aerothermal environment and material response are generally computed separately with a blowing correction term in the material response model to account for the blowing of char and pyrolysis gases [1]. In this work, we apply a coupled approach in which pyrolysis blowing gases, computed in the PATO material response code [2], are used with a blowing boundary condition in the DPLR hypersonic CFD code [3]. This leads to an iterative method in which blowing products from PATO are input into DPLR to update surface heating estimates. The full iterative method, with the addition of radiative heating estimates using the NEQAIR radiation solver [4], is shown in Fig. 1. The method is demonstrated on a sphere case with the environment and material properties based on the Mars Science Laboratory entry. Future work includes utilizing this method in computing full 3D material response during the Mars 2020 entry.

Heat Transfer↗

NASA NTRS1986

On recent advances and future research directions for computational fluid dynamics

This paper highlights some recent accomplishments regarding CFD numerical algorithm constructions for generation of discrete approximate solutions to classes of Reynolds-averaged Navier-Stokes equations. Following an overview of turbulent closure modeling, and development of appropriate conservation law systems, a Taylor weak-statement semi-discrete approximate solution algorithm is developed. Various forms for completion to the final linear algebra statement are cited, as are a range of candidate numerical linear algebra solution procedures. This development sequence emphasizes the key building blocks of a CFD RNS algorithm, including solution trial and test spaces, integration procedure and added numerical stability mechanisms. A range of numerical results are discussed focusing on key topics guiding future research directions.

Baker, A. J.↗

NASA NTRS1992

Automatic differentiation of advanced CFD codes for multidisciplinary design

Automated multidisciplinary design of aircraft and other flight vehicles requires the optimization of complex performance objectives with respect to a number of design parameters and constraints. The effect of these independent design variables on the system performance criteria can be quantified in terms of sensitivity derivatives which must be calculated and propagated by the individual discipline simulation codes. Typical advanced CFD analysis codes do not provide such derivatives as part of a flow solution; these derivatives are very expensive to obtain by divided (finite) differences from perturbed solutions. It is shown that sensitivity derivatives can be obtained accurately and efficiently using the ADIFOR source translator for automatic differentiation. In particular, it is demonstrated that the 3-D, thin-layer Navier-Stokes, multigrid flow solver called TLNS3D is amenable to automatic differentiation in the forward mode even with its implicit iterative solution algorithm and complex turbulence modeling. It is significant that by using computational differentiation, consistent discrete nongeometric sensitivity derivatives have been obtained from an aerodynamic 3-D CFD code in a relatively short time, e.g., O(man-week) not O(man-year).

Bischof, C.↗

NASA NTRS2019

Overflow Juncture Flow Computations Compared with Experimental Data

NASA’s Transformational Tools and Technologies Project (T(cu.) is supporting a substantial effort to investigate the formation and origin of separation bubbles found on wing-body juncture zones: the Juncture Flow experiment. The first phase of the Juncture Flow experiment, performed in NASA Langley’s 14- by 22-Foot Subsonic Tunnel, has been completed. This paper documents the CFD analysis done in conjunction with the experiment. Comparisons between CFD simulations and wind tunnel experimental results will be shown. Oil flow results, surface pressure cuts, velocity profiles, and Reynolds stress profiles will be compared. Preliminary results analyzing the effect of grid resolution, wind tunnel walls, and turbulence models on the above data will be presented. The results are not meant to be a validation study, but more as an evaluation of the current status of Reynolds averaged Navier Stokes (RANS) CFD simulations.

Lee, Henry C.↗

NASA NTRS2000

A Validation Summary of the NCC Turbulent Reacting/non-reacting Spray Computations

This pper provides a validation summary of the spray computations performed as a part of the NCC (National Combustion Code) development activity. NCC is being developed with the aim of advancing the current prediction tools used in the design of advanced technology combustors based on the multidimensional computational methods. The solution procedure combines the novelty of the application of the scalar Monte Carlo PDF (Probability Density Function) method to the modeling of turbulent spray flames with the ability to perform the computations on unstructured grids with parallel computing. The calculation procedure was applied to predict the flow properties of three different spray cases. One is a nonswirling unconfined reacting spray, the second is a nonswirling unconfined nonreacting spray, and the third is a confined swirl-stabilized spray flame. The comparisons involving both gas-phase and droplet velocities, droplet size distributions, and gas-phase temperatures show reasonable agreement with the available experimental data. The comparisons involve both the results obtained from the use of the Monte Carlo PDF method as well as those obtained from the conventional computational fluid dynamics (CFD) solution. Detailed comparisons in the case of a reacting nonswirling spray clearly highlight the importance of chemistry/turbulence interactions in the modeling of reacting sprays. The results from the PDF and non-PDF methods were found to be markedly different and the PDF solution is closer to the reported experimental data. The PDF computations predict that most of the combustion occurs in a predominantly diffusion-flame environment. However, the non-PDF solution predicts incorrectly that the combustion occurs in a predominantly vaporization-controlled regime. The Monte Carlo temperature distribution shows that the functional form of the PDF for the temperature fluctuations varies substantially from point to point. The results also bring to the fore some of the deficiencies associated with the use of assumed-shape PDF methods in spray computations.

Raju, M. S.↗

NASA NTRSDate not supplied

Comparison of OVERFLOW Computational and Experimental Results of the CobraMRV Mars Entry Vehicle Concept during Supersonic Retropropulsion

The CFD solver Overflow was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a CFD and wind tunnel integration effort. Selection of numerical methods and resulting comparisons to selected tunnel conditions of interest are detailed. Comparisons of mean pressures with static taps and pressure-sensitive paint are provided for multiple SST-based turbulence models. URANS predicitons of heat shield pressures at high thrust conditions are poor compared to TMRC and DES predictions. Comparisons to experimental measurements of pressure fluctuations indicate complex relationships between heat shield unsteadiness, flight Mach number, and engine thrust coefficient. Bow shock topology and shock cell structure between experimental and OVERFLOW data show good agreement in Schlieren imaging. Recommendations for future investigation into flow field unsteadiness and turbulence models are suggested.

CFD↗

NASA NTRS2017

Modeling Xenon Tank Pressurization using One-Dimensional Thermodynamic and Heat Transfer Equations

As a first step in understanding what ground support equipment (GSE) is required to provide external cooling during the loading of 5,000 kg of xenon into 4 aluminum lined composite overwrapped pressure vessels (COPVs), a modeling analysis was performed using Microsoft Excel. The goals of the analysis were to predict xenon temperature and pressure throughout loading at the launch facility, estimate the time required to load one tank, and to get an early estimate of what provisions for cooling xenon might be needed while the tanks are being filled. The model uses the governing thermodynamic and heat transfer equations to achieve these goals. Results indicate that a single tank can be loaded in about 15 hours with reasonable external coolant requirements. The model developed in this study was successfully validated against flight and test data. The first data set is from the Dawn mission which also utilizes solar electric propulsion with xenon propellant, and the second is test data from the rapid loading of a hydrogen cylindrical COPV. The main benefit of this type of model is that the governing physical equations using bulk fluid solid temperatures can provide a quick and accurate estimate of the state of the propellant throughout loading which is much cheaper in terms of computational time and licensing costs than a Computation Fluid Dynamics (CFD) analysis while capturing the majority of the thermodynamics and heat transfer.

Thermodynamic tank modeling↗

NASA NTRS2014

NASA Engineering and Safety Center Technical Bulletin No. 14-02: Aerodynamic Reaction Control System (RCS) Orientation and Jet Interaction (JI) Model Validation

Careful consideration should be given to placement and orientation of RCS jet thrusters on hypersonic entry vehicles in order to minimize adverse RCS JI. Modern state-of-the-art computational fluid dynamics (CFD) tools should be able to predict the RCS effectiveness and JI accurately when coupled with verification and validation studies, including the use of appropriately designed experimental wind tunnel testing.

Aerodynamic Reaction Control System↗

NASA NTRS2016

Computational Investigation of a Boundary-Layer Ingesting Propulsion System for the Common Research Model

The present paper examines potential propulsive and aerodynamic benefits of integrating a Boundary-Layer Ingestion (BLI) propulsion system into a typical commercial aircraft using the Common Research Model (CRM) geometry and the NASA Tetrahedral Unstructured Software System (TetrUSS). The Numerical Propulsion System Simulation (NPSS) environment is used to generate engine conditions for CFD analysis. Improvements to the BLI geometry are made using the Constrained Direct Iterative Surface Curvature (CDISC) design method. Previous studies have shown reductions of up to 25% in terms of propulsive power required for cruise for other axisymmetric geometries using the BLI concept. An analysis of engine power requirements, drag, and lift coefficients using the baseline and BLI geometries coupled with the NPSS model are shown. Potential benefits of the BLI system relating to cruise propulsive power are quantified using a power balance method, and a comparison to the baseline case is made. Iterations of the BLI geometric design are shown and any improvements between subsequent BLI designs presented. Simulations are conducted for a cruise flight condition of Mach 0.85 at an altitude of 38,500 feet and an angle of attack of 2 deg for all geometries. A comparison between available wind tunnel data, previous computational results, and the original CRM model is presented for model verification purposes along with full results for BLI power savings. Results indicate a 14.4% reduction in engine power requirements at cruise for the BLI configuration over the baseline geometry. Minor shaping of the aft portion of the fuselage using CDISC has been shown to increase the benefit from Boundary-Layer Ingestion further, resulting in a 15.6% reduction in power requirements for cruise as well as a drag reduction of eighteen counts over the baseline geometry.

Blumenthal, Brennan T.↗

NASA NTRS1985

Duct flow nonuniformities study for space shuttle main engine

To improve the Space Shuttle Main Engine (SSME) design and for future use in the development of generation rocket engines, a combined experimental/analytical study was undertaken with the goals of first, establishing an experimental data base for the flow conditions in the SSME high pressure fuel turbopump (HPFTP) hot gas manifold (HGM) and, second, setting up a computer model of the SSME HGM flow field. Using the test data to verify the computer model it should be possible in the future to computationally scan contemplated advanced design configurations and limit costly testing to the most promising design. The effort of establishing and using the computer model is detailed. The comparison of computational results and experimental data observed clearly demonstrate that computational fluid mechanics (CFD) techniques can be used successfully to predict the gross features of three dimensional fluid flow through configurations as intricate as the SSME turbopump hot gas manifold.

Thoenes, J.↗

NASA NTRS1991

Development of a CFD Code for Analysis of Fluid Dynamic Forces in Seals

The aim is to develop a 3-D computational fluid dynamics (CFD) code for the analysis of fluid flow in cylindrical seals and evaluation of the dynamic forces on the seals. This code is expected to serve as a scientific tool for detailed flow analysis as well as a check for the accuracy of the 2D industrial codes. The features necessary in the CFD code are outlined. The initial focus was to develop or modify and implement new techniques and physical models. These include collocated grid formulation, rotating coordinate frames and moving grid formulation. Other advanced numerical techniques include higher order spatial and temporal differencing and an efficient linear equation solver. These techniques were implemented in a 2D flow solver for initial testing. Several benchmark test cases were computed using the 2D code, and the results of these were compared to analytical solutions or experimental data to check the accuracy. Tests presented here include planar wedge flow, flow due to an enclosed rotor, and flow in a 2D seal with a whirling rotor. Comparisons between numerical and experimental results for an annular seal and a 7-cavity labyrinth seal are also included.

Athavale, Mahesh M.↗

NASA NTRS1995

Experimental results for a hypersonic nozzle/afterbody flow field

This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion ramp-nozzle (SERN) flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5 Foot Hypersonic Wind Tunnel at the NASA Ames Research Center, in a cooperative experimental program involving Ames and McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aerospace Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadow graph flow-visualization photographs, afterbody surface-pressure distributions, rake boundary-layer measurements, Preston-tube skin-friction measurements, and flow field surveys with five-hole and thermocouple probes. The probe data consist of impact pressure, flow direction, and total temperature profiles in the interaction flow field.

Spaid, Frank W.↗