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DOE OSTI2025

Assessing Clouds in GFDL's AM4.0 With Different Microphysical Parameterizations Using the Satellite Simulator Package COSP

We evaluate cloud simulations using satellite simulators against multiple observational data sets. These simulators have been run within the Geophysical Fluid Dynamics Laboratory's Atmosphere Model version 4.0 (AM4.0), as well as an alternative configuration where a fully two‐moment Morrison‐Gettelman cloud microphysical parameterization with prognostic precipitation (MG2) is applied, denoted as AM4‐MG2. The modeled cloud spatial distributions, vertical profiles, phase partitioning, cloud‐to‐precipitation transitions, and radiative effects compare reasonably well with satellite observations. Model biases include the under‐prediction of total and low‐level clouds, especially optically thin/intermediate clouds with cloud optical depth of less than 23, but the over‐prediction of thick clouds, indicating “too few, too bright” biases. These biases counteract each other, and give rise to reasonable estimates of cloud radiative effects. The underestimate of low‐level clouds is associated with too early and too frequent drizzle/precipitation formation. The precipitation bias is improved in AM4‐MG2, where the autoconversion scheme initiates the precipitation more realistically. There also exist discrepancies between models and observations for midlevel and high‐level clouds. Additional biases include the underestimate of liquid cloud fraction and the overestimate of ice cloud fraction.

54 ENVIRONMENTAL SCIENCES↗

NASA NTRS1992

Improved Depiction Of Computed Unsteady Flows

PLOT4D and STREAKER programs increase ease of production and utility of animated depictions of unsteady flows simulated numerically by advanced computational fluid dynamics. Programs developed to maximize information displayed while minimizing amount stored. PLOT4D, an extended version of PLOT3D, used to depict steady three-dimensional flows, incorporates features making it more computationally efficient than to visualize unsteady flow as succession of steady flows, animating frame by frame. PLOT4D postprocesses numerical output of fluid-dynamical computation into maps of flow quantities on two-coordinate surfaces depicting cut through flow field or conform to body surface. STREAKER generates streaklines, close approximations of patterns traced by streams of dye or smoke released into simulated flows. Streaklines defined via spatial succession of endpoints of pathlines of massless particles released into flow. STREAKER computes endpoints by fourth-order Runge-Kutta integration of velocity vectors.

Smith, Merritt H.↗

NASA NTRS2015

Application of Pinniped Vibrissae to Aeropropulsion

Vibrissae of Phoca Vitulina (Harbor Seal) and Mirounga Angustirostris (Elephant Seal) possessundulations along their length. Harbor Seal Vibrissae were shown to reduce vortex induced vibrations and reduce dragcompared to appropriately scaled cylinders and ellipses. Samples of Harbor Seal vibrissae, Elephant Seal vibrissae andCalifornia Sea Lion vibrissae were collected from the Marine Mammal Center in California. CT scanning, microscopy and3D scanning techniques were utilized to characterize the whiskers. Computational fluid dynamics simulations of thewhiskers were carried out to compare them to an ellipse and a cylinder. Leading edge parameters from the whiskerswere used to create a 3D profile based on a modern power turbine blade. The NASA SW-2 facility was used to performwind tunnel cascade testing on the 'Seal Blades'. Computational Fluid Dynamics simulations were used to studyincidence angles from -37 to +10 degrees on the aerodynamic performance of the Seal Blade. The tests and simulationswere conducted at a Reynolds number of 100,000. The Seal Blades showed consistent performance improvements overthe baseline configuration. It was determined that a fuel burn reduction of approximately 5 could be achieved for a fixedwing aircraft. Noise reduction potential is also explored

vortex induced vibration↗

NASA NTRS1991

NASA low-speed centrifugal compressor for 3-D viscous code assessment and fundamental flow physics research

A low speed centrifugal compressor facility recently built by the NASA Lewis Research Center is described. The purpose of this facility is to obtain detailed flow field measurements for computational fluid dynamic code assessment and flow physics modeling in support of Army and NASA efforts to advance small gas turbine engine technology. The facility is heavily instrumented with pressure and temperature probes, both in the stationary and rotating frames of reference, and has provisions for flow visualization and laser velocimetry. The facility will accommodate rotational speeds to 2400 rpm and is rated at pressures to 1.25 atm. The initial compressor stage being tested is geometrically and dynamically representative of modern high-performance centrifugal compressor stages with the exception of Mach number levels. Preliminary experimental investigations of inlet and exit flow uniformly and measurement repeatability are presented. These results demonstrate the high quality of the data which may be expected from this facility. The significance of synergism between computational fluid dynamic analysis and experimentation throughout the development of the low speed centrifugal compressor facility is demonstrated.

Hathaway, M. D.↗

NASA NTRS1991

NASA low-speed centrifugal compressor for 3-D viscous code assessment and fundamental flow physics research

Hathaway, M. D.↗

NASA NTRS2015

Application of Pinniped Vibrissae to Aeropropulsion

Vibrissae of Phoca Vitulina (Harbor Seal) and Mirounga Angustirostris (Elephant Seal) possess undulations along their length. Harbor Seal Vibrissae were shown to reduce vortex induced vibrations and reduce drag compared to appropriately scaled cylinders and ellipses. Samples of Harbor Seal vibrissae, Elephant Seal vibrissae and California Sea Lion vibrissae were collected from the Marine Mammal Center in California. CT scanning, microscopy and 3D scanning techniques were utilized to characterize the whiskers. Computational fluid dynamics simulations of the whiskers were carried out to compare them to an ellipse and a cylinder. Leading edge parameters from the whiskers were used to create a 3D profile based on a modern power turbine blade. The NASA SW-2 facility was used to perform wind tunnel cascade testing on the 'Seal Blades'. Computational Fluid Dynamics simulations were used to study incidence angles from -37 to +10 degrees on the aerodynamic performance of the Seal Blade. The tests and simulations were conducted at a Reynolds number of 100,000. The Seal Blades showed consistent performance improvements over the baseline configuration. It was determined that a fuel burn reduction of approximately 5 could be achieved for a fixed wing aircraft. Noise reduction potential is also explored.

wakes↗

NASA NTRS2017

Uncertainty Quantification of CFD Data Generated for a Model Scramjet Isolator Flowfield

Computational fluid dynamics is now considered to be an indispensable tool for the design and development of scramjet engine components. Unfortunately, the quantification of uncertainties is rarely addressed with anything other than sensitivity studies, so the degree of confidence associated with the numerical results remains exclusively with the subject matter expert that generated them. This practice must be replaced with a formal uncertainty quantification process for computational fluid dynamics to play an expanded role in the system design, development, and flight certification process. Given the limitations of current hypersonic ground test facilities, this expanded role is believed to be a requirement by some in the hypersonics community if scramjet engines are to be given serious consideration as a viable propulsion system. The present effort describes a simple, relatively low cost, nonintrusive approach to uncertainty quantification that includes the basic ingredients required to handle both aleatoric (random) and epistemic (lack of knowledge) sources of uncertainty. The nonintrusive nature of the approach allows the computational fluid dynamicist to perform the uncertainty quantification with the flow solver treated as a "black box". Moreover, a large fraction of the process can be automated, allowing the uncertainty assessment to be readily adapted into the engineering design and development workflow. In the present work, the approach is applied to a model scramjet isolator problem where the desire is to validate turbulence closure models in the presence of uncertainty. In this context, the relevant uncertainty sources are determined and accounted for to allow the analyst to delineate turbulence model-form errors from other sources of uncertainty associated with the simulation of the facility flow.

Baurle, R. A.↗

NASA NTRS2005

Heart Pump Design for Cleveland Clinic Foundation

Through a Lewis CommTech Program project with the Cleveland Clinic Foundation, the NASA Lewis Research Center is playing a key role in the design and development of a permanently implantable, artificial heart pump assist device. Known as the Innovative Ventricular Assist System (IVAS), this device will take on the pumping role of the damaged left ventricle of the heart. The key part of the IVAS is a nonpulsatile (continuous flow) artificial heart pump with centrifugal impeller blades, driven by an electric motor. Lewis is part of an industry and academia team, led by the Ohio Aerospace Institute (OAI), that is working with the Cleveland Clinic Foundation to make IVAS a reality. This device has the potential to save tens of thousands of lives each year, since 80 percent of heart attack victims suffer irreversible damage to the left ventricle, the part of the heart that does most of the pumping. Impeller blade design codes and flow-modeling analytical codes will be used in the project. These codes were developed at Lewis for the aerospace industry but will be applicable to the IVAS design project. The analytical codes, which currently simulate the flow through the compressor and pump systems, will be used to simulate the flow within the blood pump in the artificial heart assist device. The Interdisciplinary Technology Office heads up Lewis' efforts in the IVAS project. With the aid of numerical modeling, the blood pump will address many design issues, including some fluid-dynamic design considerations that are unique to the properties of blood. Some of the issues that will be addressed in the design process include hemolysis, deposition, recirculation, pump efficiency, rotor thrust balance, and bearing lubrication. Optimum pumping system performance will be achieved by modeling all the interactions between the pump components. The interactions can be multidisciplinary and, therefore, are influenced not only by the fluid dynamics of adjacent components but also by thermal and structural effects. Lewis-developed flow-modeling codes to be used in the pump simulations will include a one-dimensional code and an incompressible three-dimensional Navier-Stokes flow code. These codes will analyze the prototype pump designed by the Cleveland Clinic Foundation. With an improved understanding of the flow phenomena within the prototype pump, design changes to improve the performance of the pump system can be verified by computer prior to fabrication in order to reduce risks. The use of Lewis flow modeling codes during the design and development process will improve pump system performance and reduce the number of prototypes built in the development phase. The first phase of the IVAS project is to fully develop the prototype in a laboratory environment that uses a water/glycerin mixture as the surrogate fluid to simulate blood. A later phase of the project will include testing in animals for final validation. Lewis will be involved in the IVAS project for 3 to 5 years.

Source record↗

NASA NTRS1997

Surface Modeling and Grid Generation of Orbital Sciences X34 Vehicle

The surface modeling and grid generation requirements, motivations, and methods used to develop Computational Fluid Dynamic volume grids for the X34-Phase 1 are presented. The requirements set forth by the Aerothermodynamics Branch at the NASA Langley Research Center serve as the basis for the final techniques used in the construction of all volume grids, including grids for parametric studies of the X34. The Integrated Computer Engineering and Manufacturing code for Computational Fluid Dynamics (ICEM/CFD), the Grid Generation code (GRIDGEN), the Three-Dimensional Multi-block Advanced Grid Generation System (3DMAGGS) code, and Volume Grid Manipulator (VGM) code are used to enable the necessary surface modeling, surface grid generation, volume grid generation, and grid alterations, respectively. All volume grids generated for the X34, as outlined in this paper, were used for CFD simulations within the Aerothermodynamics Branch.

Alter, Stephen J.↗

NASA NTRS2004

COMSAC: Computational Methods for Stability and Control

Work on stability and control included the following reports:Introductory Remarks; Introduction to Computational Methods for Stability and Control (COMSAC); Stability & Control Challenges for COMSAC: a NASA Langley Perspective; Emerging CFD Capabilities and Outlook A NASA Langley Perspective; The Role for Computational Fluid Dynamics for Stability and Control:Is it Time?; Northrop Grumman Perspective on COMSAC; Boeing Integrated Defense Systems Perspective on COMSAC; Computational Methods in Stability and Control:WPAFB Perspective; Perspective: Raytheon Aircraft Company; A Greybeard's View of the State of Aerodynamic Prediction; Computational Methods for Stability and Control: A Perspective; Boeing TacAir Stability and Control Issues for Computational Fluid Dynamics; NAVAIR S&C Issues for CFD; An S&C Perspective on CFD; Issues, Challenges & Payoffs: A Boeing User s Perspective on CFD for S&C; and Stability and Control in Computational Simulations for Conceptual and Preliminary Design: the Past, Today, and Future?

Fremaux, C. Michael↗

NASA NTRS1993

Simulation of tail buffet using delta wing-vertical tail configuration

Computational simulation of the vertical tail buffet problem is accomplished using a delta wing-vertical tail configuration. Flow conditions are selected such that the wing primary-vortex cores experience vortex breakdown and the resulting flow interacts with the vertical tail. This multidisciplinary problem is solved successively using three sets of equations for the fluid flow, aeroelastic deflections and grid displacements. For the fluid dynamics part, the unsteady, compressible, full Navier-Stokes equations are solved accurately in time using an implicit, upwind, flux-difference splitting, finite-volume scheme. For the aeroelastic part, the aeroelastic equation for bending vibrations is solved accurately in time using the Galerkin method and the four-stage Runge-Kutta scheme. The grid for the fluid dynamics computations is updated every few time steps using a third set of interpolation equations. The computational application includes a delta wing of aspect ratio 1 and a rectangular vertical tail of aspect ratio 2, which is placed at 0.5 root-chord length downstream of the wing trailing edge. The wing angle of attack is 35 deg and the flow Mach number and Reynolds number are 0.4 and 10,000, respectively.

Kandil, Osama A.↗

NASA NTRS2016

Toward Improved Predictions of Slender Airframe Aerodynamics Using the F-16XL Aircraft

A coordinated project has been underway to improve computational fluid dynamics predictions of slender airframe aerodynamics. The work is focused on two flow conditions and leverages a unique flight data set obtained with an F-16XL aircraft. These conditions, a low-speed high angle-of-attack case and a transonic low angle-of-attack case, were selected from a prior prediction campaign wherein the computational fluid dynamics failed to provide acceptable results. In this paper, the background, objectives, and approach to the current project are presented. The work embodies predictions from multiple numerical formulations that are contributed from multiple organizations, and the context of this campaign to other multicode, multi-organizational efforts is included. The relevance of this body of work toward future supersonic commercial transport concepts is also briefly addressed.

Luckring, James M.↗

NASA NTRS2018

A Numerical and Experimental Investigation of Flow Separation Control over a Wall-Mounted Hump Model

Numerical simulations and wind tunnel experiments were performed to investigate flow separation and its control over a wall-mounted hump model. Three-dimensional, unsteady fluid dynamic simulations were used to capture the relevant flow physics of unsteady flow separation and its control via active and passive flow control methods. Surface pressure measurements and surface oilflow visualization were obtained in the experiments. Current measurements documented a considerably thinner incoming boundary layer, but similar pressure distribution, compared to the reference data reported in the CFD (Computational Fluid Dynamics) Validation (CFDVAL 2004) Experiments. Consistent with the data available in the literature, numerical simulations predicted a longer separation bubble, which manifested itself as a shift in the pressure distributions. Streamwise vortices generated by passive vortex generators increased the suction pressure and provided substantial pressure recovery. Numerical simulations of vortex generators qualitatively predicted the effects of control, and revealed patterns of attached/separated flow regions. On the other hand, active flow control using steady discrete jets reduced, but was not able to eliminate, flow separation for the tested configuration. Although the major disagreement between the numerical and experimental results was in the pressure recovery region due to the inaccurate prediction of separation bubbles, the relative pressure recoveries from the baselines were comparable. Surface oilflow visualization and simulated surface streamlines were qualitatively in good agreement revealing the key flow features.

Koklu, Mehti↗

NASA NTRS2014

Uncertainty Quantification and Certification Prediction of Low-Boom Supersonic Aircraft Configurations

The primary objective of this work was to develop and demonstrate a process for accurate and efficient uncertainty quantification and certification prediction of low-boom, supersonic, transport aircraft. High-fidelity computational fluid dynamics models of multiple low-boom configurations were investigated including the Lockheed Martin SEEB-ALR body of revolution, the NASA 69 Delta Wing, and the Lockheed Martin 1021-01 configuration. A nonintrusive polynomial chaos surrogate modeling approach was used for reduced computational cost of propagating mixed, inherent (aleatory) and model-form (epistemic) uncertainty from both the computation fluid dynamics model and the near-field to ground level propagation model. A methodology has also been introduced to quantify the plausibility of a design to pass a certification under uncertainty. Results of this study include the analysis of each of the three configurations of interest under inviscid and fully turbulent flow assumptions. A comparison of the uncertainty outputs and sensitivity analyses between the configurations is also given. The results of this study illustrate the flexibility and robustness of the developed framework as a tool for uncertainty quantification and certification prediction of low-boom, supersonic aircraft.

West, Thomas K., IV↗

NASA NTRS2011

CFD Study of the Hydrocarbon Boost Low-Pressure Inducer and Kicker in the Presence of a Circumferential Groove

Results are presented of a computational fluid dynamics (CFD) study done in support of Marshall Space Flight Center's (MSFC) sub-scale water flow experiments of the Hydrocarbon Boost (HCB) Oxidizer Turbopump (OTP) being developed by the Air Force Research Laboratory (AFRL) and Aerojet. A circumferential groove may be added to the pump to reduce synchronous cavitation and subsequent bearing loads at a minimal performance cost. However, the energy may reappear as high order cavitation (HOC) that spans a relatively large frequency range. Thus, HOC may have implications for the full-scale OTP inducer in terms of reduced structural margin at higher mode frequencies. Simulations using the LOCI/Stream CFD program were conducted in order to explore the fluid dynamical impact of the groove on the low-pressure inducer and kicker. It was found that the circumferential groove has minimal head performance impact, but causes back-flowing high-swirl fluid to interact with the nearly-axial incoming fluid just above the inducer blades. The high-shear interface between the fluids is Kelvin-Helmholtz unstable, resulting in trains of low pressure regions or 'pearls' forming near the upstream edge of the groove. When the static pressure in these regions becomes low enough and they get cut by the blade leading edge, HOC is thought to occur. Although further work is required, the numerical models indicate that HOC will occur in the runbox of the AFRL/Aerojet HCB OTP. Comparisons to the ongoing water flow experiments will be discussed, as well as possible designs that may mitigate HOC while continuing to reduce synchronous cavitation. December 2011 MSS/LPS/SPS Joint Subcommittee Meeting ABSTRACT SUBMITTAL FORM

Coker, Robert↗

NASA NTRS1995

Prospects and expectations for unstructured methods

The last decade has witnessed a vigorous and sustained research effort on unstructured methods for computational fluid dynamics. Unstructured mesh generators and flow solvers have evolved to the point where they are now in use for design purposes throughout the aerospace industry. In this paper we survey the various mesh types, structured as well as unstructured, and examine their relative strengths and weaknesses. We argue that unstructured methodology does offer the best prospect for the next generation of computational fluid dynamics algorithms.

Baker, Timothy J.↗

NASA NTRS2016

What Was Learned in Predicting Slender Airframe Aerodynamics with the F-16XL Aircraft

The second Cranked-Arrow Wing Aerodynamics Project, International, coordinated project has been underway to improve high-fidelity computational-fluid-dynamics predictions of slender airframe aerodynamics. The work is focused on two flow conditions and leverages a unique flight data set obtained with the F-16XL aircraft for comparison and validation. These conditions, a low-speed high-angle-of-attack case and a transonic low-angle-of-attack case, were selected from a prior prediction campaign wherein the computational fluid dynamics failed to provide acceptable results. In revisiting these two cases, approaches for improved results include better, denser grids using more grid adaptation to local flow features as well as unsteady higher-fidelity physical modeling like hybrid Reynolds-averaged Navier-Stokes/unsteady Reynolds-averaged Navier-Stokes/large-eddy simulation methods. The work embodies predictions from multiple numerical formulations that are contributed from multiple organizations where some authors investigate other possible factors that could explain the discrepancies in agreement (e.g., effects due to deflected control surfaces during the flight tests as well as static aeroelastic deflection of the outer wing). This paper presents the synthesis of all the results and findings and draws some conclusions that lead to an improved understanding of the underlying flow physics, finally making the connections between the physics and aircraft features.

Rizzi, Arthur↗

NASA NTRS2003

Comparison of Methods for Determining Boundary Layer Edge Conditions for Transition Correlations

Data previously obtained for the X-33 in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel have been reanalyzed to compare methods for determining boundary layer edge conditions for use in transition correlations. The experimental results were previously obtained utilizing the phosphor thermography technique to monitor the status of the boundary layer downstream of discrete roughness elements via global heat transfer images of the X-33 windward surface. A boundary layer transition correlation was previously developed for this data set using boundary layer edge conditions calculated using an inviscid/integral boundary layer approach. An algorithm was written in the present study to extract boundary layer edge quantities from higher fidelity viscous computational fluid dynamic solutions to develop transition correlations that account for viscous effects on vehicles of arbitrary complexity. The boundary layer transition correlation developed for the X-33 from the viscous solutions are compared to the previous boundary layer transition correlations. It is shown that the boundary layer edge conditions calculated using an inviscid/integral boundary layer approach are significantly different than those extracted from viscous computational fluid dynamic solutions. The present results demonstrate the differences obtained in correlating transition data using different computational methods.

Liechty, Derek S.↗