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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Development of code evaluation criteria for assessing predictive capability and performance

Computational Fluid Dynamics (CFD), because of its unique ability to predict complex three-dimensional flows, is being applied with increasing frequency in the aerospace industry. Currently, no consistent code validation procedure is applied within the industry. Such a procedure is needed to increase confidence in CFD and reduce risk in the use of these codes as a design and analysis tool. This final contract report defines classifications for three levels of code validation, directly relating the use of CFD codes to the engineering design cycle. Evaluation criteria by which codes are measured and classified are recommended and discussed. Criteria for selecting experimental data against which CFD results can be compared are outlined. A four phase CFD code validation procedure is described in detail. Finally, the code validation procedure is demonstrated through application of the REACT CFD code to a series of cases culminating in a code to data comparison on the Space Shuttle Main Engine High Pressure Fuel Turbopump Impeller.

Lin, Shyi-Jang↗
NASA NTRSDate not supplied

Computational Icing Analysis on NASA’s SIDRM Geometry to Investigate Collection Efficiency

Computational icing analysis results were compared to experimental icing tunnel data including aerothermal (e.g. dry air) and supercooled water droplet rime-ice conditions from tests conducted in early 2022 at the NASA Icing Research Tunnel (IRT). The Simulated Intercompressor Duct Research Model (SIDRM) test article was used in this study, and its geometry represents the inter-compressor duct region of a turbofan engine. The test article’s purpose is to study the physics of supercooled water icing and ice crystal icing. This study compared three different icing codes: FENSAP-ICE (Eulerian approach), LEWICE3D (Lagrangian approach), and GlennICE (Lagrangian approach). All three icing codes were conducted on SIDRM’s complex body flow-field and compared to different experimental supercooled water rime runs. The test article instrumentation (pressure taps, thermocouples, etc.) and 3D laser scans of final ice shapes were used to compare against the different icing code simulations. The overall objectives are to understand how the icing codes handle capturing collection efficiency on the complex test article’s unheated surfaces. In the aerothermal cases, pressure tap readings matched the CFD results, but dry air CFD underpredicted thermocouple readings. Collection efficiency results from all three icing codes matched well together on the main body leading edge, main body slope, and the strut leading edges. All three icing codes underpredicted the final ice shape using a single-shot constant ice density approach, with more difficulty coming from the strut leading edge ice shape due to the swept wing like flow field. NASA’s overall goal for this effort is to develop computational icing tools to assist in the design and certification of engines for flight in icing conditions.

Supercooled icing, SIDRM, aerodynamics, computatio↗
NASA NTRS1997

Aerodynamic Synthesis of a Centrifugal Impeller Using CFD and Measurements

The performance and flow structure in an unshrouded impeller of approximately 4:1 pressure ratio is synthesized on the basis of a detailed analysis of 3D viscous CFD results and aerodynamic measurements. A good data match was obtained between CFD and measurements using laser anemometry and pneumatic probes. This solidified the role of the CFD model as a reliable representation of the impeller internal flow structure and integrated performance. Results are presented showing the loss production and secondary flow structure in the impeller. The results indicate that while the overall impeller efficiency is high, the impeller shroud static pressure recovery potential is underdeveloped leading to a performance degradation in the downstream diffusing element. Thus, a case is made for a follow-on impeller parametric design study to improve the flow quality. A strategy for aerodynamic performance enhancement is outlined and an estimate of the gain in overall impeller efficiency that might be realized through improvements to the relative diffusion process is provided.

Larosiliere, L. M.↗
NASA NTRS1992

Comparison between computational and experimental data for a hypersonic laser propelled vehicle

A 3-D axisymmetric hypersonic engine inlet was investigated using PARC2D, an ideal gas Computational Fluid Dynamics code. The code was used to predict the results of tests conducted in the Rensselaer Polytechnic Institute Hypersonic Shock Tunnel which measured surface and pitot pressures, and shock positions (through Schlieren photography) at freestream Mach numbers of 10, 13, and 15. A strong viscous/shock interaction was observed in both the experiment and the CFD results, due to the model's parabolic compression ramp. Good agreement was found between the experimental results and the CFD solution both for surface pressures and shock positions. Agreement between pitot pressures was less reliable.

Messitt, D. G.↗
NASA NTRS2015

Advancing the State-of-the-Practice for Liquid Rocket Engine Injector Design

Current shortcomings in both the overall injector design process and its underlying combustion stability assessment methodology are rooted in the use of empirically based or low fidelity representations of complex physical phenomena and geometry details that have first order effects on performance, thermal environments and combustion stability. The result is a design and analysis capability that is often inadequate to reliably arrive at a suitable injector design in an efficient manner. Specifically, combustion instability has been particularly difficult to predict and mitigate. Large hydrocarbon-fueled booster engines have been especially problematic in this regard. Where combustion instability has been a problem, costly and time-consuming redesign efforts have often been an unfortunate consequence. This paper presents an overview of a recently completed effort at NASA Marshall Space Flight Center to advance the state-of-the-practice for liquid rocket engine injector design. Multiple perturbations of a gas-centered swirl coaxial (GCSC) element that burned gaseous oxygen and RP-1 were designed, assessed for combustion stability, and tested. Three designs, one stable, one marginally unstable and one unstable, were used to demonstrate both an enhanced overall injector design process and an improved combustion stability assessment process. High-fidelity results from state-of-the-art computational fluid dynamics CFD simulations were used to substantially augment and improve the injector design methodology. The CFD results were used to inform and guide the overall injector design process. They were also used to upgrade selected empirical or low-dimensional quantities in the ROCket Combustor Interactive Design (ROCCID) stability assessment tool. Hot fire single element injector testing was used to verify both the overall injector designs and the stability assessments. Testing was conducted at the Air Force Research Laboratory and at Purdue University. Companion papers provide details of the overall injector design process, full- and sub-scale testing, ROCCID-based stability assessments and the CFD simulations.

Tucker, P. K.↗
NASA NTRSDate not supplied

The NASA Turbulent Heat Flux Experiments: Summary and Lessons Learned

The Turbulent Heat Flux (THX) experiments were conducted at NASA Glenn Research Center (GRC) in order to collect measurements of velocities and temperatures for computational fluid dynamics (CFD) validation of heated flows, with a focus on propulsion system components. The experiments spanned 5 phases; four of which were conducted in the GRC AeroAcoustic Propulsion Laboratory (AAPL) using the Small Hot Jet Flow Rig (SHJAR). In addition to making velocity measurements with Particle Image Velocimetry (PIV), the THX experiments introduced a new Raman-scattering based capability to measure temperatures. Computational studies were also conducted for each of the experimental configurations, in order to provide a baseline of expected CFD results and conduct an assessment of the capability of various CFD approaches for calculating flows where the turbulent transport of heat was important. Two of the collected sets of data were used for American Institute of Aeronautics and Astronautics (AIAA) Propulsion Aerodynamic Workshops (PAWs). The data set from the 5th phase, collected for heated supersonic jets, was also used to construct new validation cases for the NASA Turbulence Model Resource (TMR). This paper provides an overview of the experiments and associated computations for each of the 5 test phases. Key experimental findings are presented. Lessons learned are provided concerning the effect of computational modeling choice on accuracy of predicting turbulent flows where thermal transport is important. Emphasis is placed on comparing Reynolds- averaged Navier-Stokes approaches with large-eddy simulation approaches. The benefits of utilizing a conjugate heat transfer method in conjunction with CFD solver for film cooling is demonstrated.

RANS↗
NASA NTRSDate not supplied

The NASA Turbulent Heat Flux (THX) Experiments: Summary and Lessons Learned

The Turbulent Heat Flux (THX) experiments were conducted at NASA Glenn Research Center (GRC) in order to collect measurements of velocities and temperatures for computational fluid dynamics (CFD) validation of heated flows, with a focus on propulsion system components. The experiments spanned 5 phases; four of which were conducted in the GRC AeroAcoustic Propulsion Laboratory (AAPL) using the Small Hot Jet Flow Rig (SHJAR). In addition to making velocity measurements with Particle Image Velocimetry (PIV), the THX experiments introduced a new Raman-scattering based capability to measure temperatures. Computational studies were also conducted for each of the experimental configurations, in order to provide a baseline of expected CFD results and conduct an assessment of the capability of various CFD approaches for calculating flows where the turbulent transport of heat was important. Two of the collected sets of data were used for American Institute of Aeronautics and Astronautics (AIAA) Propulsion Aerodynamic Workshops (PAWs). The data set from the 5th phase, collected for heated supersonic jets, was also used to construct new validation cases for the NASA Turbulence Model Resource (TMR). This paper provides an overview of the experiments and associated computations for each of the 5 test phases. Key experimental findings are presented. Lessons learned are provided concerning the effect of computational modeling choice on accuracy of predicting turbulent flows where thermal transport is important. Emphasis is placed on comparing Reynolds-averaged Navier-Stokes approaches with large-eddy simulation approaches. The benefits of utilizing a conjugate heat transfer method in conjunction with CFD solver for film cooling is demonstrated.

RANS↗
NASA NTRS2012

Aerothermal Ground Testing of Flexible Thermal Protection Systems for Hypersonic Inflatable Aerodynamic Decelerators

Flexible TPS development involves ground testing and analysis necessary to characterize performance of the FTPS candidates prior to flight testing. This paper provides an overview of the analysis and ground testing efforts performed over the last year at the NASA Langley Research Center and in the Boeing Large-Core Arc Tunnel (LCAT). In the LCAT test series, material layups were subjected to aerothermal loads commensurate with peak re-entry conditions enveloping a range of HIAD mission trajectories. The FTPS layups were tested over a heat flux range from 20 to 50 W/cm with associated surface pressures of 3 to 8 kPa. To support the testing effort a significant redesign of the existing shear (wedge) model holder from previous testing efforts was undertaken to develop a new test technique for supporting and evaluating the FTPS in the high-temperature, arc jet flow. Since the FTPS test samples typically experience a geometry change during testing, computational fluid dynamic (CFD) models of the arc jet flow field and test model were developed to support the testing effort. The CFD results were used to help determine the test conditions experienced by the test samples as the surface geometry changes. This paper includes an overview of the Boeing LCAT facility, the general approach for testing FTPS, CFD analysis methodology and results, model holder design and test methodology, and selected thermal results of several FTPS layups.

Bruce, Walter E., III↗
NASA NTRS1999

Measurements of the Basic SR-71 Airplane Near-Field Signature

Airplane design studies have developed configuration concepts that may produce lower sonic boom annoyance levels. Since lower noise designs differ significantly from other HSCT designs, it is necessary to accurately assess their potential before HSCT final configuration decisions are made. Flight tests to demonstrate lower noise design capability by modifying an existing airframe have been proposed for the Mach 3 SR-71 reconnaissance airplane. To support the modified SR-71 proposal, baseline in-flight measurements were made of the unmodified aircraft. These measurements of SR-71 near-field sonic boom signatures were obtained by an F-16XL probe airplane at flightpath separation distances ranging from approximately 740 to 40 ft. This paper discusses the methods used to gather and analyze the flight data, and makes comparisons of these flight data with CFD results from Douglas Aircraft Corporation and NASA Langley Research Center. The CFD solutions were obtained for the near-field flow about the SR-71, and then propagated to the flight test measurement location using the program MDBOOM.

Haering, Edward A., Jr.↗
NASA NTRS2006

Three-Dimensional Thermal Boundary Layer Corrections for Circular Heat Flux Gauges Mounted in a Flat Plate with a Surface Temperature Discontinuity

Three-dimensional Navier-Stokes computational fluid dynamics (CFD) analysis has been performed in an effort to determine thermal boundary layer correction factors for circular convective heat flux gauges (such as Schmidt-Boelter and plug type)mounted flush in a flat plate subjected to a stepwise surface temperature discontinuity. Turbulent flow solutions with temperature-dependent properties are obtained for a free stream Reynolds number of 1E6, and freestream Mach numbers of 2 and 4. The effect of gauge diameter and the plate surface temperature have been investigated. The 3-D CFD results for the heat flux correction factors are compared to quasi-21) results deduced from constant property integral solutions and also 2-D CFD analysis with both constant and variable properties. The role of three-dimensionality and of property variations on the heat flux correction factors has been demonstrated.

Kandula, M.↗
NASA NTRS1999

Forced and Moment Measurements with Pressure-Sensitive Paint

The potential of pressure-sensitive paint (PSP) to provide aerodynamic loads measurements has been a driving force behind the development of this measurement technique. To demonstrate the suitability of PSP for this purpose, it is necessary to show that PSP-derived pressures can be accurately integrated over the model surface. This cannot be done simply by demonstrating the accuracy of PSP as compared to pressure taps. PSP errors due to misregistration or temperature sensitivity may be high near model edges, where they will have a strong effect on moment measurements, but where pressure taps are rarely installed. A more suitable technique is to compare integrated PSP data over the entire model surface with balance and/or CFD results. This paper presents results from three experiments in which integrated PSP data is compared with balance and/or CFD data. This allows the usefulness of PSP for force and moment measurements, and by implication for loads measurements, to be assessed.

Bell, James H.↗
NASA NTRS2015

Computational Fluid Dynamics Simulation of Dual Bell Nozzle Film Cooling

Marshall Space Flight Center (MSFC) and Armstrong Flight Research Center (AFRC) are working together to advance the technology readiness level (TRL) of the dual bell nozzle concept. Dual bell nozzles are a form of altitude compensating nozzle that consists of two connecting bell contours. At low altitude the nozzle flows fully in the first, relatively lower area ratio, nozzle. The nozzle flow separates from the wall at the inflection point which joins the two bell contours. This relatively low expansion results in higher nozzle efficiency during the low altitude portion of the launch. As ambient pressure decreases with increasing altitude, the nozzle flow will expand to fill the relatively large area ratio second nozzle. The larger area ratio of the second bell enables higher Isp during the high altitude and vacuum portions of the launch. Despite a long history of theoretical consideration and promise towards improving rocket performance, dual bell nozzles have yet to be developed for practical use and have seen only limited testing. One barrier to use of dual bell nozzles is the lack of control over the nozzle flow transition from the first bell to the second bell during operation. A method that this team is pursuing to enhance the controllability of the nozzle flow transition is manipulation of the film coolant that is injected near the inflection between the two bell contours. Computational fluid dynamics (CFD) analysis is being run to assess the degree of control over nozzle flow transition generated via manipulation of the film injection. A cold flow dual bell nozzle, without film coolant, was tested over a range of simulated altitudes in 2004 in MSFC's nozzle test facility. Both NASA centers have performed a series of simulations of that dual bell to validate their computational models. Those CFD results are compared to the experimental results within this paper. MSFC then proceeded to add film injection to the CFD grid of the dual bell nozzle. A series of nozzle pressure ratios and film coolant flow rates are investigated to determine the effect of the film injection on the nozzle flow transition behavior. The results of this CFD study of a dual bell with film injection are presented in this paper.

Braman, Kalen↗
NASA NTRSDate not supplied

Evaluation of CFD as a Surrogate for Wind-Tunnel Testing: Experimental Uncertainty Quantification for the UPWT Flow Survey Test

A series of wind tunnel tests are being performed at the Unitary Plan Wind Tunnel (UPWT)at NASA Langley Research Center to assess the validity of using computational fluid dynamics(CFD) as a surrogate for wind tunnel testing. In order to make proper comparisons, uncertainties in CFD results and experimental data must be well understood. The material presented highlights the methods, assumptions, and inputs used to achieve experimental uncertainty estimates. Results for a small subset of tunnel conditions and variables of interest from the first test in the series, the Flow Survey Test, are highlighted and sample CFD comparisons are shown. The bulk of the results from this work are used for comparisons in other AIAA conference papers related to this test series.

Erin P Hubbard↗
DOE OSTI2022

OC6 Phase Ia: CFD Simulations of the Free-Decay Motion of the DeepCwind Semisubmersible

Currently, the design of floating offshore wind systems is primarily based on mid-fidelity models with empirical drag forces. The tuning of the model coefficients requires data from either experiments or high-fidelity simulations. As part of the OC6 (Offshore Code Comparison Collaboration, Continued, with Correlation, and unCertainty (OC6) is a project under the International Energy Agency Wind Task 30 framework) project, the present investigation explores the latter option. A verification and validation study of computational fluid dynamics (CFD) models of the DeepCwind semisubmersible undergoing free-decay motion is performed. Several institutions provided CFD results for validation against the OC6 experimental campaign. The objective is to evaluate whether the CFD setups of the participants can provide valid estimates of the hydrodynamic damping coefficients needed by mid-fidelity models. The linear and quadratic damping coefficients and the equivalent damping ratio are chosen as metrics for validation. Large numerical uncertainties are estimated for the linear and quadratic damping coefficients; however, the equivalent damping ratios are more consistently predicted with lower uncertainty. Some difference is observed between the experimental and CFD surge-decay motion, which is caused by mechanical damping not considered in the simulations that likely originated from the mooring setup, including a Coulomb-friction-type force. Overall, the simulations and the experiment show reasonable agreement, thus demonstrating the feasibility of using CFD simulations to tune mid-fidelity models.

17 WIND ENERGY↗
DOE OSTI2023

Convolutional Neural Network–Aided Temperature Field Reconstruction: An Innovative Method for Advanced Reactor Monitoring

In this study, the capabilities of a physics-informed convolutional neural network (CNN) for reconstructing the temperature field from a limited set of measurements taken at the boundaries of internal flows are demonstrated. Such an approach enables the development of less invasive monitoring methods for real-time plant diagnostics. As a test case, a Molten Salt Fast Reactor (MSFR) design was selected. This circulating fuel reactor has received interest from both scientific and industrial communities due to its intrinsic safety and sustainability. Molten salt flows in such reactors, however, can present highly localized temperature peaks that can induce significant thermal stresses onto the vessel walls. At these local maxima, the salt temperature may exceed a thousand kelvins, which makes a direct measurement challenging or even unfeasible. The proposed CNN algorithm allows one to detect indirectly such discontinuities through an accurate, albeit indirect, temperature measurement method during reactor operation. The datasets employed to train and test the machine learning models in the present work were generated with Nek5000, a computational fluid dynamics (CFD) code developed at Argonne National Laboratory. The CNN algorithm is trained with CFD results that span a set of MSFR operational power and flow ranges. Here, to demonstrate the efficacy of the algorithm, predictions are made for test cases contained within the training range but for which the CFD data were not used when training. Results demonstrate that the proposed technique properly characterizes temperature peaks and distributions within the domain for a broad range of scenarios.

21 SPECIFIC NUCLEAR REACTORS AND ASSOCIATED PLANTS↗
DOE OSTI2022

Development of A Microsampler for Aerial Platform on Collection of Aerosol Particles

A microsampler for aerosol particles was designed and evaluated in the laboratory for dry (no condensed water) and wet (foggy, misty, or rainy) air conditions. Computational Fluid Dynamics (CFD) results suggested a design that provided a higher sampling efficiency with minimal inlet loss. The microsampler was experimentally evaluated using fluorescent Polystyrene Latex (PSL) particles at four nominal sizes. The experimental results at the four PSL particle sizes agree reasonably well with the collection efficiencies predicted by CFD. We found no impacts on the collection efficiency of the microsampler at conditions under 80% Relative Humidity (RH). However, when condensed water exists, significant biases on the collection efficiencies were found suggesting this microsampler should not be used in the condition where there was condensed water in the atmosphere, e.g., foggy, misty, and rainy conditions.

54 ENVIRONMENTAL SCIENCES↗
NASA NTRS2000

Experiment/Analytical Characterization of the RBCC Rocket-Ejector Mode

Experimental and complementary CFD results from the study of the rocket-ejector mode of a Rocket Based Combined Cycle (RBCC) engine are presented and discussed. The experiments involved systematic flowfield measurements in a two-dimensional, variable geometry rocket-ejector system. The rocket-ejector system utilizes a single two-dimensional, gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a thorough understanding of the rocket-ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions. Overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (oxygen, hydrogen, nitrogen and water vapor). The experimental results for both the direct-connect and sea-level static configurations are compared with CFD predictions of the flowfield.

Ruf, J. H.↗
NASA NTRS2003

Dynamic Stall Measurements and Computations for a VR-12 Airfoil with a Variable Droop Leading Edge

High density-altitude operations of helicopters with advanced performance and maneuver capabilities have lead to fundamental research on active high-lift system concepts for rotor blades. The requirement for this type of system was to improve the sectional lift-to-drag ratio by alleviating dynamic stall on the retreating blade while simultaneously reducing the transonic drag rise of the advancing blade. Both measured and computational results showed that a Variable Droop Leading Edge (VDLE) airfoil is a viable concept for application to a rotor high-lift system. Results are presented for a series of 2D compressible dynamic stall wind tunnel tests with supporting CFD results for selected test cases. These measurements and computations show a dramatic decrease in the drag and pitching moment associated with severe dynamic stall when the VDLE concept is applied to the Boeing VR-12 airfoil. Test results also show an elimination of the negative pitch damping observed in the baseline moment hysteresis curves.

Martin, P. B.↗