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NASA NTRSDate not supplied

Validation of Artemis I Aerothermal Design Models Using Developmental Flight Instrumentation

The inaugural flight of the Space Launch System (SLS) Block 1 launch vehicle, Artemis I, occurred on November 16, 2022, and featured a full suite of Developmental Flight Instrumentation (DFI) that provided aerothermodynamic measurements to assess thermal design and substantiate aerothermodynamic models. The Block 1 launch vehicle aerothermal instrumentation consisted of approximately 277 aerothermal gauges mounted throughout the Orion Multi-Purpose Crew Vehicle (MPCV), Integrated Spacecraft and Payload Element (ISPE), Core Stage (CS) and Solid Rocket Boosters (SRB) and an additional 179 thermal gauges on the Orion Crew Module (CM). Instrumentation included calorimeters, radiometers, pressure transducers, gas temperature probes, and thermocouples. Data was collected from lift-off through CS Main Engine Cut-Off (MECO). The flight data was invaluable for determining aerothermal model performance and developing flight-derived aerothermal environments for flight reconstruction thermal analysis and future SLS aerothermal models. The data offered critical insights into the aerothermodynamic conditions experienced during the launch and ascent of the SLS vehicle. This study compares the flight derived environments to pre-existing design models. The aerothermal models were constructed using MINIVER, the aerothermal engineering code which predicts aerodynamic heating and acts as an integration tool for incorporating databases from computational fluid dynamics (CFD) simulations and wind tunnel test data. The comparisons reveal the fidelity of the design models, highlighting areas where the design models accurately predicted flight conditions and instances where deviations were observed. Preliminary results suggest that while the design models largely aligned with the observed flight data, there were unique observations that reflected needed areas of model refinement. Aerothermal flight data from Artemis I for the SLS Block 1 vehicle will be further utilized to enhance the accuracy of Block 1B and Block 2 aerothermal models, ensuring improved safety and performance for subsequent Artemis missions.

aerothermodynamics↗

NASA NTRSDate not supplied

Validation of Artemis I Aerothermal Design Models Using Developmental Flight Instrumentation

aerothermodynamics↗

NASA NTRSDate not supplied

Recent Advancements in the PATO Material Response Code

Introduction: Predicting the complicated multiphysics phenomena during atmospheric entry requires high-fidelity modeling tools to refine estimates of mission risks during entry. To this end, new capabilities are being added to the Porous-material Analysis Toolbox based on OpenFOAM (PATO). PATO is an open-source software for Computational Material Response (CMR) of reactive porous materials submitted to high-temperature environments. The objective of this work is to highlight current efforts to add to and improve upon the modeling capabilities of PATO. These include efforts to loosely couple PATO with other discipline specialized codes including hypersonic Computational Fluid Dynamics (CFD), to assess the interaction effects between pyrolysis gas blowing and the boundary layer, and Computational Solid Mechanics (CSM), to address modeling of mechanical erosion. Other refinements include surface phenomena modeling capabilities to address the effects of silicone-based coatings applied to the TPS during flight preparation, and a unified multiphase solver for a mixed porous-material and plain-fluid domain. Coupling CMR with CFD (CMR/CFD): A loose coupling between PATO and the Data Parallel Line Relaxation (DPLR) CFD code has been achieved by making use of a blowing boundary condition at the heatshield surface available in DPLR. Starting with heat flux estimates with no pyrolysis gas blowing at the surface, blowing gases are computed by the CMR and passed to the CFD such that aerothermal properties of the environment can be recomputed for a new CMR computation. This leads to an iterative process which is supplemented with an estimate of the radiative heat flux using the Nonequilibrium air radiation (NEQAIR) program. The entire iterative process is illustrated in Figure 1. This coupling strategy has been utilized in computing the MSL material response. The goal is to compare the coupled CMR/CFD results with material response results obtained using traditional blowing corrections. Coupling CMS with CMR: A mechanical erosion model is currently being implemented in PATO to account for the additional mass removal induced by high shear conditions. The modeling process at each timestep consists of updating the mechanical properties as a function of temperature and computing the stress tensor and displacement fields of the material. Then, a failure criteria model determines the regions in which the stress exceeds the ultimate strength values resulting in mesh movement to account for mass removal. This model allows the material response simulation to compute the recession due to both oxidation and shear-induced erosion. The model is demonstrated by computing material response of sphere-cone arc jet samples. Surface Modeling Capabilities: NuSil, a silicone-based coating, was sprayed onto the MSL and Mars 2020 heatshields to mitigate shedding of phenolic dust. To better understand the effects of the NuSil coating on the material response, a novel model has been implemented in PATO. In this model, the equilibrium of the charred NuSil surface is modeled as pure silica, and a constant offset, inspired by the classical spallation model, is added to the the char blowing rate and wall enthalpy to reproduce HyMETS experimental results. The model has also been used to estimate the 3D material response of the MSL heatshield. Unified Solver: In addition to the iterative loose coupling approach mentioned above, a multiphase unified solver is being developed to couple the environment (plain-fluid phase) and the porous-material phase. The solver is based on the volume averaged conservation of mass, momentum, and energy for the macroscale with closure models which include microscale effects through effective physicochemical properties. The unified solver has been used to compute flow through a porous plug and solve the Beavers and Joseph problem. Since the strong coupling between phases is inherent to this solver, modeling assumptions present in other coupling methods of material response are mitigated. This strategy also makes it feasible to capture the competition between surface and volume ablation in the same computational domain, which is usually not possible with other coupling approaches.

Material Response↗

NASA NTRSDate not supplied

Recent Advancements in the PATO Material Response Code

Introduction: Predicting the complicated multiphysics phenomena during atmospheric entry requires high-fidelity modeling tools to refine estimates of mission risks during entry. To this end, new capabilities are being added to the Porous-material Analysis Toolbox based on OpenFOAM (PATO) [1,2,3]. PATO is an open-source software for Computational Material Response (CMR) of reactive porous materials submitted to high-temperature environments. The objective of this work is to highlight current efforts to add to and improve upon the modeling capabilities of PATO. These include efforts to loosely couple PATO with other discipline specialized codes including hypersonic Computational Fluid Dynamics (CFD), to assess the interaction effects between pyrolysis gas blowing and the boundary layer, and Computational Solid Mechanics (CSM), to address modeling of mechanical erosion. Other refinements include surface phenomena modeling capabilities to address the effects of silicone-based coatings applied to the TPS during flight preparation, and a unified multiphase solver for a mixed porous-material and plain-fluid domain. Coupling CMR with CFD (CMR/CFD): A loose coupling between PATO and the Data Parallel Line Relaxation (DPLR) [4] CFD code has been achieved by making use of a blowing boundary condition at the heatshield surface available in DPLR. Starting with heat flux estimates with no pyrolysis gas blowing at the surface, blowing gases are computed by the CMR and passed to the CFD such that aerothermal properties of the environment can be recomputed for a new CMR computation. This leads to an iterative process which is supplemented with an estimate of the radiative heat flux using the Nonequilibrium air radiation (NEQAIR) [5] program. The entire iterative process is illustrated in Figure 1. This coupling strategy has been utilized in computing the MSL material response. The goal is to compare the coupled CMR/CFD results with material response results obtained using traditional blowing corrections. Coupling CMS with CMR: A mechanical erosion model is currently being implemented in PATO to account for the additional mass removal induced by high shear conditions. The modeling process at each timestep consists of updating the mechanical properties as a function of temperature and computing the stress tensor and displacement fields of the material. Then, a failure criteria model determines the regions in which the stress exceeds the ultimate strength values resulting in mesh movement to account for mass removal. This model allows the material response simulation to compute the recession due to both oxidation and shear-induced erosion. The model is demonstrated by computing material response of sphere-cone arc jet samples. Surface Modeling Capabilities: NuSil, a silicone-based coating, was sprayed onto the MSL and Mars 2020 heatshields to mitigate shedding of phenolic dust. To better understand the effects of the NuSil coating on the material response, a novel model has been implemented in PATO. In this model, the equilibrium of the charred NuSil surface is modeled as pure silica, and a constant offset, inspired by the classical spallation model, is added to the the char blowing rate and wall enthalpy to reproduce HyMETS experimental results. The model has also been used to estimate the 3D material response of the MSL heatshield [6]. Unified Solver: In addition to the iterative loose coupling approach mentioned above, a multiphase unified solver is being developed to couple the environment (plain-fluid phase) and the porous-material phase. The solver is based on the volume averaged conservation of mass, momentum, and energy for the macroscale with closure models which include microscale effects through effective physicochemical properties. The unified solver has been used to compute flow through a porous plug and solve the Beavers and Joseph problem [7]. Since the strong coupling between phases is inherent to this solver, modeling assumptions present in other coupling methods of material response are mitigated. This strategy also makes it feasible to capture the competition between surface and volume ablation in the same computational domain, which is usually not possible with other coupling approaches.

Thermal Protection Systems↗

NASA NTRS2016

Design of the Hybrid Wing Body with Nacelle: N3-X Propulsion-Airframe Configuration

The Hybrid Wing Body (HWB) aircraft is of great interest for future transport concepts due to itspromises of reduced aircraft noise, nitrous-oxide emissions, and fuel consumption. A design parameterizationmethod for HWB configurations with mail slot nacelle has been developed for a fast exploration of designspace in conceptual and preliminary design phases of a HWB configuration. A HWB planform model byLaughlin [11] was implemented, and the Class Shape Transformation (CST) airfoil generation method byKulfan [10] was utilized to construct the needed geometry for computational high fidelity aerodynamicsimulations. Geometric constraints for the parameterization such as internal cabin and cargo hold layoutswere imposed on the geometry generation. A CFD simulation was performed for a HWB configurationgenerated by the current geometric modeler, clearly showing a significant effect of the installed nacelle on theflowfield.

Nacelle Installation↗

NASA NTRS2020

Computational Simulations of a Mach 0.745 Transonic Truss-Braced Wing Design

A joint effort between the NASA Ames and Langley Research Centers was undertaken to analyze the Mach 0.745 variant of the Boeing Transonic Truss-Braced Wing (TTBW) Design. Two different flow solvers, LAVA and USM3D, were used to predict the TTBW flight performance. Sensitivity studies related to mesh resolution and numerical schemes were conducted to define best practices for this type of geometry and flow regime. Validation efforts compared the numerical simulation results of various modeling methods against experimental data taken from the NASA Ames 11-foot Unitary Wind Tunnel experimental data. The fidelity of the computational representation of the wind tunnel experiment, such as utilizing a porous wall boundary condition to model the ventilated test section, was varied to examine how different tunnel effects influence CFD predictions. LAVA and USM3D results both show an approximate 0.5o angle of attack shift from experimental lift curve data. This drove an investigation that revealed that the trailing edge of the experimental model was rounded in comparison to the CAD model, due to manufacturing tolerances, which had not been accounted for in the initial simulations of the experiment. Simulating the TTBW with an approximation of this rounded trailing-edge reduces error by approximately 60%. An accurate representation of the tested TTBW geometry, ideally including any wing twists and deflections experienced during the test under various loading conditions, will be necessary for more thorough validation of the CFD.

Maldonado, Daniel↗

NASA NTRS1995

Verification and Validation of Simulation Models for High-Speed Flow Fields

This paper is largely based on the paper titled "Guide to Credible Computational Fluid Dynamics Simulations (Invited)" (AIAA Paper 95-2225). The significance of computational fluid dynamics (CFD) simulations depends solely on their credibility. A customer of CFD products -- simulations and software -- expects that these products are credible for the intended use. Simulation model verification and validation are critical in establishing the credibility of simulations and in certifying simulation software. The only thing that matters in establishing the credibility is uncertainty, not veracity or validity. The sensitivity-uncertainty analysis is the key to the establishment of this credibility. Assessing the credibility of complex simulation results poses a significant challenge. Terminology, concepts, framework, and guidelines are presented for addressing this challenge. Verification assesses whether the problem is solved correctly and estimates the level of computational accuracy of simulations; validation determines whether the right problem is solved and assesses the level of validity of the simulation model by estimating the degree to which simulations accurately represent reality. These concepts and the related guidelines are discussed with examples from high-speed flow fields.

Mehta, Unmeel B.↗

NASA NTRS2003

Effect of Turbulence Models on Two Massively-Separated Benchmark Flow Cases

Two massively-separated flow cases (the 2-D hill and the 3-D Ahmed body) were computed with several different turbulence models in the Reynolds-averaged Navier-Stokes code CFL3D as part of participation in a turbulence modeling workshop held in Poitiers, France in October, 2002. Overall, results were disappointing, but were consistent with results from other RANS codes and other turbulence models at the workshop. For the 2-D hill case, those turbulence models that predicted separation location accurately ended up yielding a too-long separation extent downstream. The one model that predicted a shorter separation extent in better agreement with LES data did so only by coincidence: its prediction of earlier reattachment was due to a too-late prediction of the separation location. For the Ahmed body, two slant angles were computed, and CFD performed fairly well for one of the cases (the larger slant angle). Both turbulence models tested in this case were very similar to each other. For the smaller slant angle, CFD predicted massive separation, whereas the experiment showed reattachment about half-way down the center of the face. These test cases serve as reminders that state- of-the-art CFD is currently not a reliable predictor of massively-separated flow physics, and that further validation studies in this area would be beneficial.

Rumsey, Christopher L.↗

NASA NTRS2013

Wind-US Code Contributions to the First AIAA Shock Boundary Layer Interaction Prediction Workshop

This report discusses the computations of a set of shock wave/turbulent boundary layer interaction (SWTBLI) test cases using the Wind-US code, as part of the 2010 American Institute of Aeronautics and Astronautics (AIAA) shock/boundary layer interaction workshop. The experiments involve supersonic flows in wind tunnels with a shock generator that directs an oblique shock wave toward the boundary layer along one of the walls of the wind tunnel. The Wind-US calculations utilized structured grid computations performed in Reynolds-averaged Navier-Stokes mode. Four turbulence models were investigated: the Spalart-Allmaras one-equation model, the Menter Baseline and Shear Stress Transport k-omega two-equation models, and an explicit algebraic stress k-omega formulation. Effects of grid resolution and upwinding scheme were also considered. The results from the CFD calculations are compared to particle image velocimetry (PIV) data from the experiments. As expected, turbulence model effects dominated the accuracy of the solutions with upwinding scheme selection indicating minimal effects.

Georgiadis, Nicholas J.↗

NASA NTRS2019

Unsteady Model Estimation for Generic T-Tail Transport Aircraft Using Computational Data

Models including nonlinear and unsteady behaviors are developed for the longitudinal axis of the NASA Generic T-Tail Aircraft over a large range of angle of attack. These models are based on computational simulations of forced-oscillation tests in a wind tunnel. This work continues a recent study and an ongoing effort by NASA to improve aircraft simulations for pilot training in loss-of-control and stalled conditions. The objective of this work is to develop appropriate aerodynamic models that provide representative responses in simulation for a given class of aircraft. In the stall region, nonlinear unsteady responses are often present and may require an extended aerodynamic model compared to that used in the conventional flight envelope. In this study, two objectives are addressed. The first is to obtain representative models for the NASA Generic T-Tail aircraft over a wide range of angle of attack and the second is to continue development of a specialized CFD test technique that uses Schroeder sweeps to create information rich responses for unsteady aerodynamic model identification.

Murphy, Patrick C.↗

NASA NTRS2018

Measurements to Understand the Flow Mechanisms Contributing to Tandem-Rotor Outwash

Downwash and outwash characteristics of a model-scale tandem-rotor system in the presence of the ground were analyzed by identifying and understanding the physical mechanisms contributing to the observed flow field behavior. A building block approach was followed in simplifying the problem, separating the effects of the fuselage, effects of one rotor on the other, etc. Flow field velocities were acquired in a vertical plane at four aircraft azimuths of a small-scale tandem rotor system using the particle image velocimetry (PIV) technique for radial distances up to 4 times the rotor diameter. Results were compared against full-scale CH-47D measurements. Excellent correlation was found between the small- and full-scale mean flow fields (after appropriate normalization using rotor and wall jet parameters). Following the scalability analysis, the effect of rotor height on the outwash was also studied. Close to the aircraft, an increase in rotor height above ground decreased the outwash velocity at all aircraft azimuths. However, farther away, the longitudinal and lateral axes of the aircraft showed increasing and decreasing outwash velocities, respectively, with increasing rotor height. Measurements also indicated the presence of large-scale (of the size of the rotor height) shear-layer vortical structures along the ground that could be the source of low-frequency (approximately 1 Hz) flow variation observed in the full-scale measurements. Flow visualization studies and PIV measurements were also made on jets of different sizes to complement the observations made on rotors wherever possible. Baseline rotor measurements were made out-of-ground effect to understand the nature of inflow distribution for realistic rotor configurations and their modified characteristics in the presence of ground. Lastly, a feasibility study on applying high-fidelity CFD simulations for outwash study was conducted using Helios to model an isolated rotor configuration IGE at full scale Reynolds number. The results were encouraging and demonstrated the practical challenges associated with predicting rotor outwash.

Measurements↗

NASA NTRS2015

Measurements to Understand the Flow Mechanisms Contributing to Tandem-Rotor Outwash

Downwash and outwash characteristics of a model-scale tandem-rotor system in the presence of the ground were analyzed by identifying and understanding the physical mechanisms contributing to the observed flow field behavior. A building block approach was followed in simplifying the problem, separating the effects of the fuselage, effects of one rotor on the other, etc. Flow field velocities were acquired in a vertical plane at four aircraft azimuths of a small-scale tandem rotor system using the particle image velocimetry (PIV) technique for radial distances up to 4 times the rotor diameter. Results were compared against full-scale CH-47D measurements. Excellent correlation was found between the small- and full-scale mean flow fields (after appropriate normalization using rotor and wall jet parameters). Following the scalability analysis, the effect of rotor height on the outwash was also studied. Close to the aircraft, an increase in rotor height above ground decreased the outwash velocity at all aircraft azimuths. However, farther away, the longitudinal and lateral axes of the aircraft showed increasing and decreasing outwash velocities, respectively, with increasing rotor height. Measurements also indicated the presence of large-scale (of the size of the rotor height) shear-layer vortical structures along the ground that could be the source of low-frequency (approximately 1 Hz) flow variation observed in the full-scale measurements. Flow visualization studies and PIV measurements were also made on jets of different sizes to complement the observations made on rotors wherever possible. Baseline rotor measurements were made out-of-ground effect to understand the nature of inflow distribution for realistic rotor configurations and their modified characteristics in the presence of ground. Lastly, a feasibility study on applying high-fidelity CFD simulations for outwash study was conducted using Helios to model an isolated rotor configuration IGE at full-scale Reynolds number. The results were encouraging and demonstrated the practical challenges associated with predicting rotor outwash.

Ramasamy, Manikandan↗

NASA NTRSDate not supplied

Conceptual Design of Propulsors for the SUSAN Electrofan Transport Aircraft

Conceptual designs of the propulsor modules for the SUSAN electro-fan aircraft are sought after. Shaft power requirement is reduced by using boundary layer ingesting propulsion technology. There are several obstacles to designing feasible BLI propulsion systems such as inlet distortion, flow blockage from ingesting low-momentum flow, strong coupling between airframe and propulsion systems. Thus, the high fidelity CFD tool is indispensable to assess the performance of the propulsion systems and evaluate the inlet profiles during the design process. Consequently, the CFD data is used to update the inputs for the NPSS model from the initial stage of the system design. Hence, the inlet sizing, fan design, and estimation of the power saving are carried out. Various installation concepts of the mail-slot nacelle, such as under-/over-wing and trailing edge configurations, are investigated for the wing-mounted turbo-electric distributed propulsor module, and their power-saving is evaluated. As for the tail-mounted turbofan engine, the fan diameter and inlet captured area is determined based on the CFD profiles. An appropriate area ratio of the bypass and core ducts is derived from meeting the target bypass ratio from the system design. The baseline fan stage is analyzed by full annulus URANS CFD to assess the efficiency penalty due to the ingested pressure and swirl distortions.

SUSAN Electro-Fan Aircraft↗

NASA NTRSDate not supplied

Conceptual Design of Propulsors for the SUSAN Electrofan Transport Aircrafts

Electric Aircraft↗

NASA NTRS2006

Solver and Turbulence Model Upgrades to OVERFLOW 2 for Unsteady and High-Speed Applications

An implicit unfactored SSOR algorithm has been added to the overset Navier-Stokes CFD code OVERFLOW 2 for unsteady and moving body applications. The HLLEM and HLLC third-order spatial upwind convective flux models have been added for high-speed flow applications. A generalized upwind transport equation has been added for solution of the two-equation turbulence models and the species equations. The generalized transport equation is solved using an unfactored SSOR implicit algorithm. Three hybrid RANS/DES turbulence models have been added for unsteady flow applications. Wall function boundary conditions that include compressibility and heat transfer effects have been also been added to OVERFLOW 2.

Nichols, Robert H.↗

NASA NTRS2010

RANS Modeling of Benchmark Shockwave / Boundary Layer Interaction Experiments

This presentation summarizes the computations of a set of shock wave / turbulent boundary layer interaction (SWTBLI) test cases using the Wind-US code, as part of the 2010 American Institute of Aeronautics and Astronautics (AIAA) shock / boundary layer interaction workshop. The experiments involve supersonic flows in wind tunnels with a shock generator that directs an oblique shock wave toward the boundary layer along one of the walls of the wind tunnel. The Wind-US calculations utilized structured grid computations performed in Reynolds-averaged Navier-Stokes mode. Three turbulence models were investigated: the Spalart-Allmaras one-equation model, the Menter Shear Stress Transport wavenumber-angular frequency two-equation model, and an explicit algebraic stress wavenumber-angular frequency formulation. Effects of grid resolution and upwinding scheme were also considered. The results from the CFD calculations are compared to particle image velocimetry (PIV) data from the experiments. As expected, turbulence model effects dominated the accuracy of the solutions with upwinding scheme selection indicating minimal effects.!

Georgiadis, Nick↗

NASA NTRS2017

A CFD Case Study of a Fan Stage with Split Flow Path Subject to Total Pressure Distortion Inflow

This report is the documentation of the work performed under the Hypersonic Project of the NASA's Fundamental Aeronautics Program. It was funded through Task Number NNC10E444T under GESS-2 Contract NNC06BA07B. The objective of the task is to develop advanced computational tools for the simulation of multi-stage turbomachinery in support of aeropropulsion. This includes work elements in extending the TURBO code and validating the multi-stage URANS (Unsteady Reynolds Averaged Navier Stokes) simulation results with the experimental data. The unsteady CFD (Computation Fluid Dynamics) calculations were performed in full wheel mode with and without screen generated total pressure distortion at the computational inflow boundary, as well as in single passage phase lag mode for uniform inflow. The experimental data were provided by NASA from the single stage RTA (Revolutionary Turbine Accelerator) fan test program.Significant non-uniform flow condition at the fan-face of the aeropropulsion system is frequentlyencountered in many of the advanced aerospace vehicles. These propulsion systems can be eithera podded or an embedded design employed in HWB (Hybrid Wing Body) airframe concept. It isalso a topic of interest in military applications, in which advanced air vehicles have already deployedsome form of embedded propulsion systems in their design because of the requirementsof compact and low observable inlets. Even in the conventional airframe/engine design, the fancould operate under such condition when the air vehicle is undergoing rapid maneuvering action.It is believed that a better understanding of the fan’s aerodynamic and aeromechanical responseto this type of operating condition or off design operation would be beneficial to designing distortiontolerant blades for improved engine operability.The objective for this research is to assess the capability of turbomachinery code as an analysistool in understanding the effects and evaluating the impact of flow distortion on the aerodynamicand aeromechanical performance of the fan in advanced propulsion systems. Results from thetesting of an advanced fan stage released by NASA are available and will be used here for CFDcode validation. The experiment was performed at NASA’s high speed compressor facility aspart of the RTA (Revolutionary Turbine Accelerator) demonstration project, a joint effort ofNASA Glenn Research Center and GE Aircraft Engines in developing an advanced Mach 4TBCC (Turbine Based Combined Cycle) turbofan/ramjet engine for access to space. Part of thetest was to assess the aerodynamic performance and operability of the fan stage under nonuniforminflow condition. Various flow distortion patterns were created at the fan-face by manipulatingsets of screens placed upstream of the wind tunnel. Measurements at the fan-face willprovide the necessary distortion flow information as the inflow boundary condition for the CFDin a full wheel simulation. Therefore the purpose of this work is to demonstrate the NASA supportedmulti-stage turbomachinery code, TURBO [1-5], in the aerodynamic performance analysisof a modern fan design operating under off design condition, and in particular to validate theCFD results with the RTA fan test data.A brief description of the RTA fan rig configuration is given in the next section, explaining onhow flow distortion were measured in the test and constructed for the CFD at the fan-face. It isfollowed by a section summarizing previous CFD work performed at NASA relevant to the currentfan configuration. A short description of the TURBO code is given next, followed by detailsin the computational model of the fan rig, the required computing resources, and the numericalprocedure for the simulations. The CFD results are presented in the discussion section and finallyconcluding remarks are summarized.

Fan Stage↗

NASA NTRS1990

Atomization modeling in a multiphase flow environment and comparison with experiments

An atomization model based on Reitz's instability wave analysis has been implemented into the ARICC3D multiphase CFD combustion code. Preliminary test runs with cold non-evaporating liquid jet and coaxial gas-liquid atomization cases appeared to have verified basic performance of the model, generating realistic-looking sprays. Furthermore, the extended liquid jet is explicitly resolved, and predicted jet lengths agree well with classical correlations. Fair agreement with test data is obtained for predicted spray tip penetrations and liquid mass flux radial distributions, with obvious room for improvement. Some numerical problems also appear to have resulted with the current implementation when low gas Mach number and high liquid velocities are involved.

Liang, P. Y.↗