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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Comparisons of CFD Simulations of Icing Wind Tunnel Clouds with Experiments Conducted at the NASA Propulsions Systems Laboratory

This paper evaluates simulation predictions against experimental test data of icing clouds that were produced during 2018 ice crystal icing physics tests conducted at the NASA Propulsion Systems Laboratory icing wind tunnel. Aero-thermal and cloud parameters are set and known upstream at the tunnel inlet and spray system, but change as the cloud and air thermodynamically interact as the flowing masses reach the tunnel test section. Utilizing the ANSYS Fluent Discrete Phase Model function, 3D computational fluid dynamics (CFD) simulations were performed, capturing the thermodynamic interactions between the test parameters, and providing predictions of the aero thermal and cloud conditions at the tunnel test section. Simulation predictions were compared with test data that were measured at the tunnel exit plane. Evaluations focused on the cloud concentration (total water content), humidity content, and air temperature. CFD simulation predictions showed areas of agreement and disagreement. Simulations showed that cloud concentration profiles at the test section are strongly related to the initial spray nozzle pattern used at the tunnel inlet. Experimental data suggest that greater dispersion of the cloud occurred as the simulated cloud predicted areas of high and low cloud concentration compared to test data profiles. Simulations, however, captured the magnitude and location of the change in humidity content and the change in air temperature due to the presence of the cloud reasonably well, when compared to test data. This result would suggest that while the ANSYS simulation did not fully predict the spreading of the cloud as measured during experiment, it did capture evaporation and the molecular movements of air and vapor relatively well.

CFD↗
NASA NTRS1991

Viscous flow simulations of internal store carriage and separation

The internal carriage of stores by the military aircraft is an option for possible reductions in the aerodynamic drag and the observability. Trade studies of this option require considering the aircraft and the stores together. In an effort to develop a computational fluid dynamic (CFD) code for such studies, an investigation was conducted from 1986 to 1990. The study was divided into five building-block steps. First, a full Navier-Stokes code was developed to simulate the unsteady, three-dimensional cavity flow. As the second step, this code was then used to simulate the flows past various missile configurations at angles of attack up to 44 deg. The effects of incidence as well as the turbulence on the leeside flows were computationally captured. The objective of this study has involved the interference flows of rather complex configurations with multiple, joint or disjoint, components of nonsimilar geometries. Hence, a hybrid domain decomposition (HDD) method was developed as the third step of the investigation. The strengths of the multiblock, zonal, and overlapped grids were judiciously combined and employed for the present problem. In the fourth step, the interference flow past a missile near a flat-plate wing was simulated using the HDD method. Finally, the fifth step involved the simulation of the internal store carriage and separation. Four different cases for two different configurations were simulated. The computational results of all five steps were successfully compared with the available wind tunnel test data. The unsteady aerodynamic forces on the separating store were computationally predicted. The CFD code developed for this project is called Viscous Internal Store Carriage Code (VISCC).

Baysal, Oktay↗
NASA NTRS2002

Flowfield Comparisons from Three Navier-Stokes Solvers for an Axisymmetric Separate Flow Jet

To meet new noise reduction goals, many concepts to enhance mixing in the exhaust jets of turbofan engines are being studied. Accurate steady state flowfield predictions from state-of-the-art computational fluid dynamics (CFD) solvers are needed as input to the latest noise prediction codes. The main intent of this paper was to ascertain that similar Navier-Stokes solvers run at different sites would yield comparable results for an axisymmetric two-stream nozzle case. Predictions from the WIND and the NPARC codes are compared to previously reported experimental data and results from the CRAFT Navier-Stokes solver. Similar k-epsilon turbulence models were employed in each solver, and identical computational grids were used. Agreement between experimental data and predictions from each code was generally good for mean values. All three codes underpredict the maximum value of turbulent kinetic energy. The predicted locations of the maximum turbulent kinetic energy were farther downstream than seen in the data. A grid study was conducted using the WIND code, and comments about convergence criteria and grid requirements for CFD solutions to be used as input for noise prediction computations are given. Additionally, noise predictions from the MGBK code, using the CFD results from the CRAFT code, NPARC, and WIND as input are compared to data.

Koch, L. Danielle↗
NASA NTRS2000

Numerical Simulation of Non-Rotating and Rotating Coolant Channel Flow Fields

Future generations of ultra high bypass-ratio jet engines will require far higher pressure ratios and operating temperatures than those of current engines. For the foreseeable future, engine materials will not be able to withstand the high temperatures without some form of cooling. In particular the turbine blades, which are under high thermal as well as mechanical loads, must be cooled. Cooling of turbine blades is achieved by bleeding air from the compressor stage of the engine through complicated internal passages in the turbine blades (internal cooling, including jet-impingement cooling) and by bleeding small amounts of air into the boundary layer of the external flow through small discrete holes on the surface of the blade (film cooling and transpiration cooling). The cooling must be done using a minimum amount of air or any increases in efficiency gained through higher operating temperature will be lost due to added load on the compressor stage. Turbine cooling schemes have traditionally been based on extensive empirical data bases, quasi-one-dimensional computational fluid dynamics (CFD) analysis, and trial and error. With improved capabilities of CFD, these traditional methods can be augmented by full three-dimensional simulations of the coolant flow to predict in detail the heat transfer and metal temperatures. Several aspects of turbine coolant flows make such application of CFD difficult, thus a highly effective CFD methodology must be used. First, high resolution of the flow field is required to attain the needed accuracy for heat transfer predictions, making highly efficient flow solvers essential for such computations. Second, the geometries of the flow passages are complicated but must be modeled accurately in order to capture all important details of the flow. This makes grid generation and grid quality important issues. Finally, since coolant flows are turbulent and separated the effects of turbulence must be modeled with a low Reynolds number turbulence model to accurately predict details of heat transfer.

Rigby, David L.↗
NASA NTRSDate not supplied

Computational Fluid Dynamics Simulations of the Transonic Dynamics Tunnel Airstream Oscillator System

This paper presents a Computational Fluid Dynamics (CFD) model of the flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) with the Airstream Oscillator System (AOS) in operation. The TDT is a continuous-flow, closed circuit wind tunnel with a 16- by 16-foot slotted test section with cropped corners. The tunnel was originally built as the 19-ft Pressure Tunnel in 1938, but it was converted to the current transonic tunnel in the 1950s, with capabilities to use either air or heavy gas (R-134a) as the test medium. The TDT was also fitted with an AOS which can be used to create a gust field in the test section. To date, no computational analyses of the TDT involving the AOS have been performed. In this study, experimental data acquired of an oscillating airstream in the tunnel will be used to calibrate the computational analyses. A significant motivation of this work is to attempt to compare computational gust velocities to those recorded in the TDT literature. In addition to a validation of this model with experimental data, it may also be possible to supplement the experimental data with computational data. The experimental data is rather sparse and at only a few Mach numbers. Computational results may be able to expand the AOS data set. Another motivation for this work is that the Integrated Adaptive Wing Technology Maturation (IAWTM) semi-span model of the high-aspect-ratio CRM with 10 trailing edge control surfaces is currently being fabricated and will be delivered to the NASA Langley Transonic Dynamics Tunnel (TDT) for testing in late 2020. Among tests to be conducted, gust load alleviation (GLA) will be demonstrated at transonic conditions using the AOS.

Computational Fluid Dynamics↗
NASA NTRS1994

An analysis code for the Rapid Engineering Estimation of Momentum and Energy Losses (REMEL)

Nonideal behavior has traditionally been modeled by defining efficiency (a comparison between actual and isentropic processes), and subsequent specification by empirical or heuristic methods. With the increasing complexity of aeropropulsion system designs, the reliability of these more traditional methods is uncertain. Computational fluid dynamics (CFD) and experimental methods can provide this information but are expensive in terms of human resources, cost, and time. This report discusses an alternative to empirical and CFD methods by applying classical analytical techniques and a simplified flow model to provide rapid engineering estimates of these losses based on steady, quasi-one-dimensional governing equations including viscous and heat transfer terms (estimated by Reynold's analogy). A preliminary verification of REMEL has been compared with full Navier-Stokes (FNS) and CFD boundary layer computations for several high-speed inlet and forebody designs. Current methods compare quite well with more complex method results and solutions compare very well with simple degenerate and asymptotic results such as Fanno flow, isentropic variable area flow, and a newly developed, combined variable area duct with friction flow solution. These solution comparisons may offer an alternative to transitional and CFD-intense methods for the rapid estimation of viscous and heat transfer losses in aeropropulsion systems.

Dechant, Lawrence J.↗
NASA NTRS2016

Numerical Simulation of a High-Lift Configuration Embedded with High Momentum Fluidic Actuators

Numerical simulations have been performed for a vertical tail configuration with deflected rudder. The suction surface of the main element of this configuration, just upstream of the hinge line, is embedded with an array of 32 fluidic actuators that produce oscillating sweeping jets. Such oscillating jets have been found to be very effective for flow control applications in the past. In the current paper, a high-fidelity computational fluid dynamics (CFD) code known as the PowerFLOW R code is used to simulate the entire flow field associated with this configuration, including the flow inside the actuators. A fully compressible version of the PowerFLOW R code valid for high speed flows is used for the present simulations to accurately represent the transonic flow regimes encountered in the flow field due to the actuators operating at higher mass flow (momentum) rates required to mitigate reverse flow regions on a highly-deflected rudder surface. The computed results for the surface pressure and integrated forces compare favorably with measured data. In addition, numerical solutions predict the correct trends in forces with active flow control compared to the no control case. The effect of varying the rudder deflection angle on integrated forces and surface pressures is also presented.

Vatsa, Veer N.↗
NASA NTRS1993

An experimental study of characteristic combustion-driven flows for CFD validation

The application of laser-based diagnostic techniques has become commonplace to a wide variety of combustion problems. New insights into combustion phenomena at a level previously unattainable has been made possible by non-intrusive measurements of velocity, temperature, and species. However, due to the adverse conditions which exist inside rocket engines, relatively few studies have addressed these combustion environments. The high pressure, high speed, combusting environment in a rocket engine prohibits the application of several measurement techniques. However, in the rocket community, there is a critical need for rocket flow field data to validate computational fluid dynamic (CFD) codes. Currently at Penn State, there is an effort to obtain flowfield measurements inside a rocket engine. Velocity measurements have been made inside the combustion chamber of a uni-element (shear coaxial injector) optically accessible rocket chamber at several axial locations downstream of the injector. These measurements, combined with future measurements, will provide benchmark data for CFD code validation.

Pal, S.↗
NASA NTRS2017

Computational Modeling Approaches to Multiscale Design of Icephobic Surfaces

To aid in the design of surfaces that prevent icing, a model and computational simulation of impact ice formation at the single droplet scale was implemented. The nucleation of a single supercooled droplet impacting on a substrate, in rime ice conditions, was simulated. Open source computational fluid dynamics (CFD) software was used for the simulation. To aid in the design of surfaces that prevent icing, a model of impact ice formation at the single droplet scale was proposed•No existing model simulates simultaneous impact and freezing of a single super-cooled water droplet•For the 10-week project, a low-fidelity feasibility study was the goal.

adhesion↗
NASA NTRS1988

The numerical simulation of transonic separated flow about the complete F-16A

The thin-layer, Reynolds-averaged, Navier-Stokes equations are used to simulate the transonic viscous flow about the complete F-16A fighter aircraft. These computations demonstrate how computational fluid dynamics (CFD) can be used to simulate turbulent viscous flow about realistic aircraft geometries. A zonal grid approach is used to provide adequate viscous grid clustering on all aircraft surfaces. Zonal grids extend inside the F-16A inlet and up to the compressor face while power on conditions are modeled by employing a zonal grid extending from the exhaust nozzle to the far field. A simple solution adaptive grid procedure is used on the wing surface and good agreement with experimental data is obtained. Computations for the F-16A in side slip are also presented.

Flores, Jolen↗
NASA NTRS1996

A Structured Grid Based Solution-Adaptive Technique for Complex Separated Flows

The objective of this work was to enhance the predictive capability of widely used computational fluid dynamic (CFD) codes through the use of solution adaptive gridding. Most problems of engineering interest involve multi-block grids and widely disparate length scales. Hence, it is desirable that the adaptive grid feature detection algorithm be developed to recognize flow structures of different type as well as differing intensity, and adequately address scaling and normalization across blocks. In order to study the accuracy and efficiency improvements due to the grid adaptation, it is necessary to quantify grid size and distribution requirements as well as computational times of non-adapted solutions. Flow fields about launch vehicles of practical interest often involve supersonic freestream conditions at angle of attack exhibiting large scale separate vortical flow, vortex-vortex and vortex-surface interactions, separated shear layers and multiple shocks of different intensity. In this work, a weight function and an associated mesh redistribution procedure is presented which detects and resolves these features without user intervention. Particular emphasis has been placed upon accurate resolution of expansion regions and boundary layers. Flow past a wedge at Mach=2.0 is used to illustrate the enhanced detection capabilities of this newly developed weight function.

Thornburg, Hugh↗
NASA NTRS2000

Testing New Programming Paradigms with NAS Parallel Benchmarks

Over the past decade, high performance computing has evolved rapidly, not only in hardware architectures but also with increasing complexity of real applications. Technologies have been developing to aim at scaling up to thousands of processors on both distributed and shared memory systems. Development of parallel programs on these computers is always a challenging task. Today, writing parallel programs with message passing (e.g. MPI) is the most popular way of achieving scalability and high performance. However, writing message passing programs is difficult and error prone. Recent years new effort has been made in defining new parallel programming paradigms. The best examples are: HPF (based on data parallelism) and OpenMP (based on shared memory parallelism). Both provide simple and clear extensions to sequential programs, thus greatly simplify the tedious tasks encountered in writing message passing programs. HPF is independent of memory hierarchy, however, due to the immaturity of compiler technology its performance is still questionable. Although use of parallel compiler directives is not new, OpenMP offers a portable solution in the shared-memory domain. Another important development involves the tremendous progress in the internet and its associated technology. Although still in its infancy, Java promisses portability in a heterogeneous environment and offers possibility to "compile once and run anywhere." In light of testing these new technologies, we implemented new parallel versions of the NAS Parallel Benchmarks (NPBs) with HPF and OpenMP directives, and extended the work with Java and Java-threads. The purpose of this study is to examine the effectiveness of alternative programming paradigms. NPBs consist of five kernels and three simulated applications that mimic the computation and data movement of large scale computational fluid dynamics (CFD) applications. We started with the serial version included in NPB2.3. Optimization of memory and cache usage was applied to several benchmarks, noticeably BT and SP, resulting in better sequential performance. In order to overcome the lack of an HPF performance model and guide the development of the HPF codes, we employed an empirical performance model for several primitives found in the benchmarks. We encountered a few limitations of HPF, such as lack of supporting the "REDISTRIBUTION" directive and no easy way to handle irregular computation. The parallelization with OpenMP directives was done at the outer-most loop level to achieve the largest granularity. The performance of six HPF and OpenMP benchmarks is compared with their MPI counterparts for the Class-A problem size in the figure in next page. These results were obtained on an SGI Origin2000 (195MHz) with MIPSpro-f77 compiler 7.2.1 for OpenMP and MPI codes and PGI pghpf-2.4.3 compiler with MPI interface for HPF programs.

Jin, H.↗
NASA NTRS2003

Enhancement and Extension of Porosity Model in the FDNS-500 Code to Provide Enhanced Simulations of Rocket Engine Components

In the past, the design of rocket engines has primarily relied on the cold flow/hot fire test, and the empirical correlations developed based on the database from previous designs. However, it is very costly to fabricate and test various hardware designs during the design cycle, whereas the empirical model becomes unreliable in designing the advanced rocket engine where its operating conditions exceed the range of the database. The main goal of the 2nd Generation Reusable Launching Vehicle (GEN-II RLV) is to reduce the cost per payload and to extend the life of the hardware, which poses a great challenge to the rocket engine design. Hence, understanding the flow characteristics in each engine components is thus critical to the engine design. In the last few decades, the methodology of computational fluid dynamics (CFD) has been advanced to be a mature tool of analyzing various engine components. Therefore, it is important for the CFD design tool to be able to properly simulate the hot flow environment near the liquid injector, and thus to accurately predict the heat load to the injector faceplate. However, to date it is still not feasible to conduct CFD simulations of the detailed flowfield with very complicated geometries such as fluid flow and heat transfer in an injector assembly and through a porous plate, which requires gigantic computer memories and power to resolve the detailed geometry. The rigimesh (a sintered metal material), utilized to reduce the heat load to the faceplate, is one of the design concepts for the injector faceplate of the GEN-II RLV. In addition, the injector assembly is designed to distribute propellants into the combustion chamber of the liquid rocket engine. A porosity mode thus becomes a necessity for the CFD code in order to efficiently simulate the flow and heat transfer in these porous media, and maintain good accuracy in describing the flow fields. Currently, the FDNS (Finite Difference Navier-Stakes) code is one of the CFD codes which are most widely used by research engineers at NASA Marshall Space Flight Center (MSFC) to simulate various flow problems related to rocket engines. The objective of this research work during the 10-week summer faculty fellowship program was to 1) debug the framework of the porosity model in the current FDNS code, and 2) validate the porosity model by simulating flows through various porous media such as tube banks and porous plate.

Cheng, Gary↗
NASA NTRS2011

Comparison of PLIF and CFD Results for the Orion CEV RCS Jets

Nitric-oxide planar laser-induced fluorescence (NO PLIF) was used to visualize and measure centerline streamwise velocity of the Orion Crew Exploration Vehicle (CEV) Reaction Control System (RCS) Jets at NASA Langley Research Center's 31-Inch Mach 10 Air wind tunnel. Fluorescence flow visualizations of pitch, roll, and yaw RCS jets were obtained using different plenum pressures and wind tunnel operating stagnation pressures. For two yaw RCS jet test cases, the PLIF visualizations were compared to computational flow imaging (CFI) images based on Langley Aerothermal Upwind Relaxation Algorithm (LAURA) computational fluid dynamics (CFD) simulations of the flowfield. For the same test cases, the streamwise velocity measurements were compared to CFD. The CFD solution, while showing some unphysical artifacts, generally agree with the experimental measurements.

Ivey, Christopher B.↗
NASA NTRS2011

A CFD Analysis of Hydrogen Leakage During On-Pad Purge in the ORION/ARES I Shared Volume

A common open volume is created by the stacking of the Orion vehicle onto the Ares I Upper Stage. Called the Shared Volume, both vehicles contribute to its gas, fluid, and thermal environment. One of these environments is related to hazardous hydrogen gas. While both vehicles use inert purge gas to mitigate any hazardous gas buildup, there are concerns that hydrogen gas may still accumulate and that the Ares I Hazardous Gas Detection System will not be sufficient for monitoring the integrated volume. This Computational Fluid Dynamics (CFD) analysis has been performed to examine these topics. Results of the analysis conclude that the Ares I Hazardous Gas Detection System will be able to sample the vent effluent containing the highest hydrogen concentrations. A second conclusion is that hydrogen does not accumulate under the Orion Service Module (SM) avionics ring as diffusion and purge flow mixing sufficiently dilute the hydrogen to safe concentrations. Finally the hydrogen concentrations within the Orion SM engine nozzle may slightly exceed the 1 percent volume fraction when the entire worse case maximum full leak is directed vertically into the engine nozzle.

Ajmani, Kumud↗
NASA NTRS1989

Cold Flow Determination of the Internal Flow Environment Around the Submerged TVC Nozzle for the Space Shuttle SRM

A series of subscale cold flow tests was performed to quantify the gas flow characteristics at the aft end of the Space Shuttle Solid Rocket Motor. This information was used to support the analyses of the redesigned nozzle/case joint. A portion of the thermal loads at the joint are due to the circumferential velocities and pressure gradients caused primarily by the gimbaling of the submerged nose TVC nozzle. When the nozzle centerline is vectored with respect to the motor centerline, asymmetries are set up in the flow field under the submerged nozzle and immediately adjacent to the nozzle/case joint. Specific program objectives included: determination of the effects of nozzle gimbal angle and propellant geometry on the circumferential flow field; measurement of the static pressure and gas velocities in the vicinity of the nozzle/case joint; use of scaling laws to apply the subscale cold flow data to the full scale SRM; and generation of data for use in validation of 3-D computational fluid dynamic, CFD, models of the SRM flow field. These tests were conducted in the NASA Marshall Space Flight Center Airflow Facility with a 7.5 percent scale model of the aft segment of the SRM. Static and dynamic pressures were measured in the model to quantify the flow field. Oil flow data was also acquired to obtain qualitative visual descriptions of the flow field. Nozzle gimbal angles of 0, 3.5, and 7 deg were used with propellant grain configurations corresponding to motor burn times of 0, 9, 19, and 114 seconds. This experimental program was successful in generating velocity and pressure gradient data for the flow field around the submerged nose nozzle of the Space Shuttle SRM at various burn times and gimbal angles. The nature of the flow field adjacent to the nozzle/case joint was determined with oil droplet streaks, and the velocity and pressure gradients were quantified with pitot probes and wall static pressure measurements. The data was applied to the full scale SRM thru a scaling analysis and the results compared well with the 3-D computational fluid dynamics computer model.

Whitesides, R. H.↗
NASA NTRS1991

Hypersonic Shock/Boundary-Layer Interaction Database

Turbulence modeling is generally recognized as the major problem obstructing further advances in computational fluid dynamics (CFD). A closed solution of the governing Navier-Stokes equations for turbulent flows of practical consequence is still far beyond grasp. At the same time, the simplified models of turbulence which are used to achieve closure of the Navier-Stokes equations are known to be rigorously incorrect. While these models serve a definite purpose, they are inadequate for the general prediction of hypersonic viscous/inviscid interactions, mixing problems, chemical nonequilibria, and a range of other phenomena which must be predicted in order to design a hypersonic vehicle computationally. Due to the complexity of turbulence, useful new turbulence models are synthesized only when great expertise is brought to bear and considerable intellectual energy is expended. Although this process is fundamentally theoretical, crucial guidance may be gained from carefully-executed basic experiments. Following the birth of a new model, its testing and validation once again demand comparisons with data of unimpeachable quality. This report concerns these issues which arise from the experimental aspects of hypersonic modeling and represents the results of the first phase of an effort to develop compressible turbulence models.

Settles, G. S.↗
NASA NTRS1992

CFD Techniques for Propulsion Applications

The symposium was composed of the following sessions: turbomachinery computations and validations; flow in ducts, intakes, and nozzles; and reacting flows. Forty papers were presented, and they covered full 3-D code validation and numerical techniques; multidimensional reacting flow; and unsteady viscous flow for the entire spectrum of propulsion system components. The capabilities of the various numerical techniques were assessed and significant new developments were identified. The technical evaluation spells out where progress has been made and concludes that the present state of the art has almost reached the level necessary to tackle the comprehensive topic of computational fluid dynamics (CFD) validation for propulsion.

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