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At least 163 records · Page 9

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NASA NTRSDate not supplied

EMU Helmet Free Water Transport Assessment for the HAB in Support of Eva 80

After water was reported in the EMU helmet during ISS US EVA-80, mitigation strategies were created to attempt to arrest the motion of any droplets that enter the helmet for future Extravehicular Activities (EVAs). This included adding absorbent materials into the interior of the helmet. To assess the effectiveness of this strategy, a computational fluid dynamics (CFD) model of a human mannikin head in the EMU helmet was used to track water droplets and quantify how much water is likely to be caught by the absorbent material. A combination of engineering judgment, tests, and CFD results were used to develop the expected path of droplets in the helmet, to account for the simplifications necessary in modeling two-phase flow.

Abigail Baukus↗

NASA NTRSDate not supplied

Liquid Hydrogen Tank Chill and No-Vent Fill Prediction Using Computational Fluid Dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management (CFM) technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents a computational fluid dynamics (CFD) model with conjugate heat transfer that is used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34 liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, fluid temperatures, and wall temperatures. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface. Several sensitivity studies are considered due to notable uncertainties in the experiment.

Computational Fluid Dynamics↗

NASA NTRSDate not supplied

EMU Helmet Free Water Transport Assessment for the HAB in Support of Eva 80

Abigail Baukus↗

NASA NTRSDate not supplied

Liquid Hydrogen Tank Chill and No-Vent Fill Prediction using Computational Fluid Dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents computational fluid dynamics (CFD) models with conjugate heat transfer that are used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34-liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, and wall temperature. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface.

Computational Fluid Dynamics↗

NASA NTRS2023

Thermal Management System Modeling in the Heat Transport System Simulation (HeaTSSPy) Package

This paper describes the development of a thermal management system (TMS) concept design and analysis software package called Heat Transport System Simulation (HeaTSSPy). Built within Python using the OpenMDAO framework, HeaTSSPy can be used to size and optimize an active (using liquid/air heat exchangers) or passive (using finned heat sinks) TMS. The package makes use of modular TMS elements that allow for the creation of different system architectures and includes components such as heat sinks, heat exchangers, liquid pumps, fans, ducts, air inlets, air nozzles, and liquid pipes. Modeling methods for these components include a combination of physics-based analytical and empirical equations that relate component sizing criteria to system performance. The HeatSSPy heat sink methods are fully detailed within this paper, while the heat exchanger methods are described and referenced from previous work. This paper also uses high-fidelity simulation to validate two different methods of calculating thermal resistance using CFD results. Once the methods are fully described, the code is exercised to compare an active TMS with that of a passive TMS. Design criteria for this study include rejected heat, system altitude, Mach number, and ambient temperature. These criteria are used to develop a TMS system with estimated performance metrics such as weight, drag, and operational power. Results of this paper show the crossover point when a passive system begins to weigh more than an active system for a given heat rejection.

Thermal management system↗

NASA NTRSDate not supplied

HLPW-5: Overview and Workshop Summary

The Fifth AIAA CFD High-Lift Prediction Workshop was held with the 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 endeavor was the use of Technology Focus Groups (TFG), which included both mesh generation and flow-solver experts working together to accelerate advancements for their particular CFD methodology, by addressing key questions of importance prior to the workshop. The high-lift version of the NASA Common Research Model (CRM-HL) configuration was the focus of this workshop, and was used for three unique test cases. Wind-tunnel data were available for comparison for one of the test cases. Altogether, 365 datasets of CFD results were submitted by 47 teams, with 41 teams contributing to the multiple configurations of Case 1, 40 to Case 2, and 18 to Case 3. This paper provides a high-level summary of the results and conclusions from the workshop. As concluded from past workshops, fixed-grid Reynolds-averaged Navier-Stokes methods continued to be inaccurate and inconsistent for high-lift flows near maximum lift. However, application of mesh-adaptation technology helped to achieve improved consistency. Scale-resolving methods appeared most promising for predicting high-lift flow physics, particularly at maximum lift. Best practices for these methods were refined over the course of the workshop and new challenges were identified.

Adam M Clark↗

NASA NTRS2010

Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio: Comparison of CFD and Experimental Results - Part 2

NASA and GE teamed to design and build a 57 percent engine scaled fan stage for a Mach 4 variable cycle turbofan/ramjet engine for access to space with multipoint operations. This fan stage was tested in NASA's transonic compressor facility. The objectives of this test were to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off; 2) transition over large swings in fan bypass ratio; 3) transition from turbofan to ramjet; and 4) fan wind-milling operation at high Mach flight conditions. This paper will focus on an assessment of APNASA, a multistage turbomachinery analysis code developed by NASA, to predict the fan stage performance and operability over a wide range of speeds (37 to 100 percent) and bypass ratios.

Celestina, Mark L.↗

NASA NTRSDate not supplied

Review of IRT CFD & Diffuser Study Results - February 2024 Test Entries

In February 2024, multiple test entries were conducted in the NASA Glenn Research Center Icing Research Tunnel (IRT) to collect CFD validation data, particularly near the facility heat exchanger inlet and outlet planes, measure the boundary layer characteristics in the test section, and survey the diffuser flow field. The CFD validation testing utilized ultrasonic anemometers to acquire 3-components of flow velocity at the inlet and outlet planes across a row of heat exchanger blocks. The test section boundary layer surveys used 18-inch boundary layer rakes centered on each of the four test section surfaces with pitot tubes at turntable centerline. The diffuser study included surveys at three streamwise positions within the diffuser with a 5-hole flow angularity probe at between 3 and 7 lateral locations at a given streamwise position. All IRT diffuser data were acquired with the flow angularity probe at the centerline of the test section. An ultrasonic anemometer was also used in the diffuser to survey near-wall velocity profiles.

Icing↗

NASA NTRS1997

Comparative Results from a CFD Challenge Over a 2D Three-Element High-Lift Airfoil

A high-lift workshop was held in May of 1993 at NASA Langley Research Center. A major part of the workshop centered on a blind test of various computational fluid dynamics (CFD) methods in which the flow about a two- dimensional (2D) three-element airfoil was computed without prior knowledge of the experimental data. The results of this 'blind' test revealed: (1) The Reynolds Averaged Navier-Stokes (RANS) methods generally showed less variability among codes than did potential/Euler solvers coupled with boundary-layer solution techniques. However, some of the coupled methods still provided excellent predictions. (2) Drag prediction using coupled methods agreed more closely with experiment than the RANS methods. Lift was more accurately predicted than drag for both methods. (3) The CFD methods did well in predicting lift and drag changes due to changes in Reynolds number, however, they did not perform as well when predicting lift and drag increments due to changing flap gap, (4) Pressures and skin friction compared favorably with experiment for most of the codes. (5) There was a large variability in most of the velocity profile predictions. Computational results predict a stronger siat wake than measured suggesting a missing component in turbulence modeling, perhaps curvature effects.

Klausmeyer, Steven M.↗

NASA NTRSDate not supplied

1st AIAA CFD Transition Modeling and Prediction Workshop: OVERFLOW Results for the SST and Langtry-Menter Models

This paper reports the results of CFD simulations of the test cases for the AIAA 1stCFDTransition Modeling and Prediction Workshop held at the AIAA Scitech Forum on 21-22January, 2021 using the NASA solver OVERFLOW. The Langtry-Menter transition model was used in combination with the Shear Stress Transport (SST) turbulence model. The testcases include the ERCOFTAC zero-pressure-gradient flat plate cases T3A and T3B with by-pass transition, the natural laminar flow NLF(1)-0416 airfoil, the DLR inclined 6:1 prolate spheroid, and the NASA Common Research Model with wing design optimized for natural laminar flow transition (CRM-NLF). OVERFLOW load and transition predictions obtained using best practices are compared with experiments and prior simulations. Additional results are provided for modified turbulence model boundary specifications to improve alignment to reference data and assess model sensitivity.

ARMD↗

NASA NTRS2005

CFD Validation with LDV Test Data for Payload/Fairing Internal Flow

Flowfield testing of a 1/5th scale model of a payload/fairing configuration, typical of an expendable launch vehicle, has been performed. Two-dimensional (planar) velocity measurements were carried out in four planes with the aid of Laser Doppler Velocimetry (LDV). Computational Fluid Dynamics (CFD) analysis results for the scale model flowfleld are compared with the test data. The CFD results are in general agreement with the test data. The ability of the CFD methodology in identifying the global flow features (including critical points such as vortex, saddle point, etc.) has been demonstrated. Practical problems and difficulties associated with the LDV method applied to the complex geometry under consideration have been summarized.

Kandula, max↗

NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 x 10 ft supersonic wind tunnel. The 2-D inlet model was designed to study the Mach 3.0 to 5.0 speed range for an over-under turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both 3-D parabolized Navier-Stokes (PNS) and 3-D full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗

NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 ft. by 10 ft. supersonic wind tunnel. The two-dimensional inlet model was designed to study the Mach 3.0 to 5.0 speed range for an 'over-under' turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both three-dimensional parabolized Navier-Stokes (PNS) and three-dimensional full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗

NASA NTRS2000

Benchmark of FDNS CFD Code For Direct Connect RBCC Test Data

Computational Fluid Dynamics (CFD) analysis results are compared with experimental data from the Pennsylvania State University's (PSU) Propulsion Engineering Research Center (PERC) rocket based combined cycle (RBCC) rocket-ejector experiments. The PERC RBCC experimental hardware was in a direct-connect configuration in diffusion and afterburning (DAB) operation. The objective of the present work was to validate the Finite Difference Navier Stokes (FDNS) CFD code for the rocket-ejector mode internal fluid mechanics and combustion phenomena. A second objective was determine the best application procedures to use FDNS as a predictive/engineering tool. Three-dimensional CFD analysis was performed. Solution methodology and grid requirements are discussed. CFD results are compared to experimental data for static pressure, Raman Spectroscopy species distribution data and RBCC net thrust and specified impulse.

Ruf, J. H.↗

NASA NTRS1997

Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions. Freejet tests of a candidate flowpath for this RBCC engine were conducted at the NASA Lewis Research Center's Hypersonic Tunnel Facility between July and September 1996. This paper describes the engine flowpath and installation, outlines the primary objectives of the program, and describes the overall results of this activity. Through this program 15 full duration tests, including 13 fueled tests were made. The first major achievement was the further demonstration of the HTF capability. The facility operated at conditions up to 1950 K and 7.34 MPa, simulating approximately Mach 6.6 flight. The initial tests were unfueled and focused on verifying both facility and engine starting. During these runs additional aerodynamic appliances were incorporated onto the facility diffuser to enhance starting. Both facility and engine starting were achieved. Further, the static pressure distributions compared well with the results previously obtained in a 40% subscale flowpath study conducted in the LERC 1X1 supersonic wind tunnel (SWT), as well as the results of CFD analysis. Fueled performance results were obtained for the engine at both simulated Mach 6 (1670 K) and Mach 6.6 (1950 K) conditions. For all these tests the primary fuel was liquid JP-10 with gaseous silane (a mixture of 20% SiH4 and 80% H2 by volume) as an ignitor/pilot. These tests verified performance of this engine flowpath in a freejet mode. High combustor pressures were reached and significant changes in axial force were achieved due to combustion. Future test plans include redistributing the fuel to improve mixing, and consequently performance, at higher equivalence ratios.

Thomas, Scott R.↗

NASA NTRS1989

NASA Computational Fluid Dynamics Conference. Volume 1: Sessions 1-6

This publication is a collection of the presentations given at the NASA Computational Fluid Dynamics (CFD) Conference held at NASA Ames Research Center, Moffett Field, California, March 7-9, 1989. The objectives of the conference were to disseminate CFD research results to industry and university CFD researchers, to promote synergy among NASA CFD researchers, and to permit feedback from researchers outside of NASA on issues pacing the discipline of CFD. The focus of the conference was on the application of CFD technology but also included fundamental activities. The conference was sponsored by the Aerodynamics Division, Office of Aeronautics and Space Technology (OAST), NASA Headquarters, Washington, D.C. 20546. The conference consisted of twelve sessions of papers representative of CFD research conducted within NASA and three non-NASA panel sessions. For each panel session, the panel membership consisted of industry and university CFD researchers. A summary of the comments made during the panel sessions has been included in this publication. Volume 1 contains the papers given in Sessions I-VT. Volume 2 covers Sessions VII-XII.

Computational fluid dynamics↗

NASA NTRS1989

NASA Computational Fluid Dynamics Conference. Volume 2: Sessions 7-12

The objectives of the conference were to disseminate CFD research results to industry and university CFD researchers, to promote synergy among NASA CFD researchers, and to permit feedback from researchers outside of NASA on issues pacing the discipline of CFD. The focus of the conference was on the application of CFD technology but also included fundamental activities.

Source record↗

NASA NTRS2011

The Crucial Role of Error Correlation for Uncertainty Modeling of CFD-Based Aerodynamics Increments

The Ares I ascent aerodynamics database for Design Cycle 3 (DAC-3) was built from wind-tunnel test results and CFD solutions. The wind tunnel results were used to build the baseline response surfaces for wind-tunnel Reynolds numbers at power-off conditions. The CFD solutions were used to build increments to account for Reynolds number effects. We calculate the validation errors for the primary CFD code results at wind tunnel Reynolds number power-off conditions and would like to be able to use those errors to predict the validation errors for the CFD increments. However, the validation errors are large compared to the increments. We suggest a way forward that is consistent with common practice in wind tunnel testing which is to assume that systematic errors in the measurement process and/or the environment will subtract out when increments are calculated, thus making increments more reliable with smaller uncertainty than absolute values of the aerodynamic coefficients. A similar practice has arisen for the use of CFD to generate aerodynamic database increments. The basis of this practice is the assumption of strong correlation of the systematic errors inherent in each of the results used to generate an increment. The assumption of strong correlation is the inferential link between the observed validation uncertainties at wind-tunnel Reynolds numbers and the uncertainties to be predicted for flight. In this paper, we suggest a way to estimate the correlation coefficient and demonstrate the approach using code-to-code differences that were obtained for quality control purposes during the Ares I CFD campaign. Finally, since we can expect the increments to be relatively small compared to the baseline response surface and to be typically of the order of the baseline uncertainty, we find that it is necessary to be able to show that the correlation coefficients are close to unity to avoid overinflating the overall database uncertainty with the addition of the increments.

Hemsch, Michael J.↗