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NASA NTRSDate not supplied

Validation of a Computational Fluid Dynamics Model of Axial Jet Mixing for Cryogenic Propellant Tank Pressure Control

The Fluid Dynamics Branch at the NASA Marshall Space Flight Center is preparing to support flight programs through analysis of a variety of cryogenic fluid management (CFM) applications. Many vehicles being considered for future manned missions to the moon and beyond use chemical or nuclear thermal propulsion systems that rely on cryogenic propellants. Storing cryogenic propellants for later use is a challenge, though. Many areas of active CFM research, testing, and design involve propellant conditioning to ensure propellant remains a usable liquid for propulsion. Multiple technologies may be used to achieve adequate conditioning including the subject of this paper, a jet-based mixer. Mixing serves to homogenize fluid temperatures and decrease ullage pressure. Development and validation of a modeling methodology for jet-based mixing was conducted to prepare for in-line design work.

Jacob M. Brodnick↗

NASA NTRS2011

Assessment of Computational Fluid Dynamics (CFD) Models for Shock Boundary-Layer Interaction

A workshop on the computational fluid dynamics (CFD) prediction of shock boundary-layer interactions (SBLIs) was held at the 48th AIAA Aerospace Sciences Meeting. As part of the workshop numerous CFD analysts submitted solutions to four experimentally measured SBLIs. This paper describes the assessment of the CFD predictions. The assessment includes an uncertainty analysis of the experimental data, the definition of an error metric and the application of that metric to the CFD solutions. The CFD solutions provided very similar levels of error and in general it was difficult to discern clear trends in the data. For the Reynolds Averaged Navier-Stokes methods the choice of turbulence model appeared to be the largest factor in solution accuracy. Large-eddy simulation methods produced error levels similar to RANS methods but provided superior predictions of normal stresses.

DeBonis, James R.↗

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 NTRSDate not supplied

Validation of Cryogenic Propellant Tank Self-Pressurization by Leveraging Reduced Order Modeling within Computational Fluid Dynamics Simulation

Validation of cryogenic propellant tank self-pressurization was performed using a hybrid Computational Fluid Dynamics (CFD) and reduced order modeling methodology. Data from a liquid hydrogen ground test conducted at the K-site facility at the National Aeronautics and Space Administration (NASA) Glenn Research Center was used for the validation effort. Liquid phase dynamics were explicitly resolved with a CFD tool. Vapor phase dynamics were modeled as a point mass that communicated heat from the tank wall to the liquid phase via a boundary condition used at the gas-liquid interface. The method proved to be more accurate, robust, and efficient than explicit resolution of the dynamics using a standard Volume of Fluid (VOF) methodology. The subject pressurization process was found to be heavily dependent upon both the relatively high liquid temperature gradient near the gas-liquid interface and the natural convection flow path. Modeling the gas-liquid interface as an immovable surface eliminated temperature gradient destroying gas-liquid interface velocities observed in VOF simulations, and correspondingly enabled more rapid simulation since interface advection was not allowed. The single phase computational domain also facilitated the ability to demonstrate spatial resolution convergence of natural convection cells within the liquid which significantly impacted the tank pressurization rate. This work was used to demonstrate the critical physics for tank self-pressurization and numerical methodologies that may be used to best resolve those physics. The findings informed development and operation of production level CFD tools used in the Fluid Dynamics Branch at NASA Marshall Space Flight Center.

J. M. Brodnick↗

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 NTRSDate not supplied

Coupling a Computational Fluid Dynamics (CFD) Model to a Spacecraft Thermal System Model for the DraMS Instrument Thermal Analysis

The Dragonfly Mass Spectrometer (DraMS) is an instrument on the Dragonfly mission, which will spend 7 years in deep space cruise before landing and operating on the surface of Titan. Vacuum thermal analyses are required for deep space cruise, and convection analyses are required for the Titan surface operations. Model exchanges across multiple thermal teams are needed for all phases of the mission. For DraMS, Thermal Desktop (TD) has been the main thermal analytical tool of choice due to its capability in modeling complex thermal systems with relatively low computational power and for its availability across thermal teams. However, TD does not have computational fluid dynamics (CFD) capability and struggles to accurately capture complex convective behavior. DraMS has fans operating in tandem and gas flow behaviors are not easily predicted due to its complex flow paths. CFD software, such as Fluent, can model and predict such complex flow behaviors, but CFD models are computationally expensive, and its workflow processes are not tailored towards simulating large and complex systems. Therefore, a coupled modeling approach was chosen for DraMS: A TD model was used for simulating all the conductive, radiative, and source terms, while a Fluent CFD model was added on, as needed, to the TD model to provide the convective boundary conditions using the System Coupling software. The coupling software allows the TD and Fluent models to communicate data and arrive at a co-solved and co-converged solution. Furthermore, Thermal Iso-value Exchange (TIE) method was developed to facilitate and improve the TD-Fluent data exchange process. This paper will discuss the analytical studies that were done to verify the accuracy and usability of the coupled approach and the challenges associated, which lead to the development of the TIE approach. DraMS thermal design and co-solved analysis results will also be discussed.

heat transfer↗

NASA NTRS2015

Modeling of Non-Isothermal Cryogenic Fluid Sloshing

A computational fluid dynamic model was used to simulate the thermal destratification in an upright self-pressurized cryostat approximately half-filled with liquid nitrogen and subjected to forced sinusoidal lateral shaking. A full three-dimensional computational grid was used to model the tank dynamics, fluid flow and thermodynamics using the ANSYS Fluent code. A non-inertial grid was used which required the addition of momentum and energy source terms to account for the inertial forces, energy transfer and wall reaction forces produced by the shaken tank. The kinetics-based Schrage mass transfer model provided the interfacial mass transfer due to evaporation and condensation at the sloshing interface. The dynamic behavior of the sloshing interface, its amplitude and transition to different wave modes, provided insight into the fluid process at the interface. The tank pressure evolution and temperature profiles compared relatively well with the shaken cryostat experimental test data provided by the Centre National D'Etudes Spatiales.

Cryogenics↗

NASA NTRS1984

Finite-difference fluid dynamics computer mathematical models for the design and interpretation of experiments for space flight

Numerical methods are used to design a spherical baroclinic flow model experiment of the large scale atmosphere flow for Spacelab. The dielectric simulation of radial gravity is only dominant in a low gravity environment. Computer codes are developed to study the processes at work in crystal growing systems which are also candidates for space flight. Crystalline materials rarely achieve their potential properties because of imperfections and component concentration variations. Thermosolutal convection in the liquid melt can be the cause of these imperfections. Such convection is suppressed in a low gravity environment. Two and three dimensional finite difference codes are being used for this work. Nonuniform meshes and implicit iterative methods are used. The iterative method for steady solutions is based on time stepping but has the options of different time steps for velocity and temperature and of a time step varying smoothly with position according to specified powers of the mesh spacings. This allows for more rapid convergence. The code being developed for the crystal growth studies allows for growth of the crystal as the solid-liquid interface. The moving interface is followed using finite differences; shape variations are permitted. For convenience in applying finite differences in the solid and liquid, a time dependent coordinate transformation is used to make this interface a coordinate surface.

Roberts, G. O.↗

NASA NTRS1992

Integrating aerodynamic surface modeling for computational fluid dynamics with computer aided structural analysis, design, and manufacturing

This presentation will discuss the development of a NASA Geometry Exchange Specification for transferring aerodynamic surface geometry between LeRC systems and grid generation software used for computational fluid dynamics research. The proposed specification is based on a subset of the Initial Graphics Exchange Specification (IGES). The presentation will include discussion of how the NASA-IGES standard will accommodate improved computer aided design inspection methods and reverse engineering techniques currently being developed. The presentation is in viewgraph format.

Thorp, Scott A.↗

NASA NTRSDate not supplied

Pre-Flight Hazardous Gas Assessment Methodology for the Space Launch System

The complexity of the Artemis I liquid rocket propulsion systems produces many sources of hazards during the operation of the vehicle. Pre-flight hazardous gas leakages pose fire and explosion risks to rocket that could potentially lead to loss of the vehicle, mission, and crew. The National Aeronautics and Space Administration (NASA) uses a probability risk assessment strategy to assess the likelihood of the hazard and the consequence should the hazardous situation occur. For high probability hazards that pose a potentially catastrophic consequence to the vehicle, the most effective strategy for reducing the risk is accomplished by engineering and implementing systems that actively mitigate the hazard. A hazards analysis is completed to determine the likelihood and severity of each hazard. These hazard analyses are informed using relevant empirical data and physics-based analytical models. Time-accurate Computational Fluid Dynamics models used to quantify and understand the risks posed to Artemis I due to expulsions of hazardous gas near the vehicle. Several sources of hazardous gas that posed a risk to Artemis I during nominal pre-flight operations are analyzed in this work. The results of the analyses were used to develop flight rationale and make risk acceptance decisions for the Artemis I launch.

Brian R. Richardson↗

NASA NTRSDate not supplied

Pre-Flight Hazardous Gas Assessment Methodology for the Space Launch System

Brian R. Richardson↗

NASA NTRS1991

Modeling of InP metalorganic chemical vapor deposition

The growth of InP by metalorganic chemical vapor deposition (MOCVD) in a horizontal reactor is being modeled with a commercially available computational fluid dynamics modeling code. The mathematical treatment of the MOCVD process has four primary areas of concern: 1) transport phenomena, 2) chemistry, 3) boundary conditions, and 4) numerical solution methods. The transport processes involved in CVD are described by conservation of total mass, momentum, energy, and atomic species. Momentum conservation is described by a generalized form of the Navier-Stokes equation for a Newtonian fluid and laminar flow. The effect of Soret diffusion on the transport of particular chemical species and on the predicted deposition rate is examined. Both gas-phase and surface chemical reactions are employed in the model. Boundary conditions are specified at the inlet and walls of the reactor for temperature, fluid flow and chemical species. The coupled set of equations described above is solved by a finite difference method over a nonuniform rectilinear grid in both two and three dimensions. The results of the 2-D computational model is presented for gravity levels of zero- and one-g. The predicted growth rates at one-g are compared to measured growth rates on fused silica substrates.

Black, Linda R.↗

NASA NTRS2018

Aerodynamic Conceptual Design of Boundary Layer Ingestion Propulsor Systems: A Quasi-2D Through Flow Analysis Method and Multi-Fidelity Propulsor Design Framework

A propulsor design framework for maximizing the benefits of boundary layer ingestion is presented. The performance of BLI is strongly affected by all propulsor components, including the boundary layer characteristics (displacement thickness and form factor) of the ingested boundary layer at the inlet, the radical loading characteristics if the fan and exit guide vane (EGV), the area contraction from inlet to nozzle, and flow expansion at the exhaust cone. A strategy and it's associated multi-fidelity design framework are proposed for an efficient conceptual design of the BLI proplusor which inherently differs from the conventional engine. In the frame work, a quasi-2D through flow model served as the underlying fidelity model is introduced to incorporate the radial effect of the boundary layer entering the propulsor. Multi-fidelity design work is conducted to maximize the predefined performance metrics. On top of this efficient quasi-2D model. computational fluid dynamics based 3-D propulsor models are implemented to refine and validate the design. A BLI propulsion system integrated with fuselage is designed to showcase the framework. The performance of the resulting BLI propulsor system is evaluated via body-force model in unstructured Reynolds-Averaged Navier-Stoke (RANS) CFD and improvement is presented.

Lee, B. J.↗

NASA NTRS2020

Generation and Calibration of Linear Models of Aircraft with Highly Coupled Aeroelastic and Flight Dynamics

The lightweight structures and unconventional configurations being considered for the next generation of aircraft mean that any effort to predict or control the flight dynamics is impacted by the structural dynamics. One of the most severe forms of coupling between aeroelasticity and flight dynamics is an instability called body freedom flutter. The existing tools often assume a relatively weak effect of structural dynamics on the flight dynamics, and are therefore incapable of modeling strong interactions like body freedom flutter. A method of combining different sources of data traditionally used for aeroelasticity and flight dynamics is described by reconciling many of the differences between these models. By building upon past modeling efforts, a level of familiarity in the approach is achieved. Generally the differences from the traditional approaches are subtle but significant. The traditional frequency domain flutter model in a modal coordinate system is converted to a form consistent with a time domain flight dynamics model. The time domain rational function approximation about a non-inertial coordinate system and the unique constraints for the conversion between the inertial and non-inertial coordinate systems are discussed. A consistent transformation of the states of aeroelastic models to flight dynamics models is derived, which enables the integration of data from higher fidelity computational fluid dynamics models or wind-tunnel testing. The present method of integrating multidisciplinary data was used to create models that compare well with X-56A flight-test data, including conditions past the flutter speed.

Ouellette, Jeffrey↗

NASA NTRS1990

Tank Pressure Control Experiment - A low-g mixing investigation

The Tank Pressure Control Experiment (TPCE) was conceived to meet the need for an aspect of cryogenic fluid management technology that is critical to several future national space missions: control of cryogenic storage tank pressures by active mixing. In-space testing is the only means of obtaining the fluid dynamic data necessary to develop and test predictive models of mixing. These models, when validated, will allow future cryogenic systems to be designed with more efficient and reliable pressure control systems. The objectives of the TPCE project are to characterize the fluid dynamics of jet-induced mixing in low gravity, evaluate the validity of empirical models and correlations, and provide data for use in developing and validating computational fluid dynamic models. This paper discusses prior studies of low-g mixing, the objectives and benefits of TPCE, the design and status of the payload, and preliminary results obtained from low-g aircraft testing.

Bentz, M. D.↗

NASA NTRS2015

Nonequilibrium Supersonic Freestream Studied Using Coherent Anti-Stokes Raman Spectroscopy

Measurements were conducted at the University of Virginia Supersonic Combustion Facility of the flow in a constant-area duct downstream of a Mach 2 nozzle. The airflow was heated to approximately 1200 K in the facility heater upstream of the nozzle. Dual-pump coherent anti-Stokes Raman spectroscopy was used to measure the rotational and vibrational temperatures of N2 and O2 at two planes in the duct. The expectation was that the vibrational temperature would be in equilibrium, because most scramjet facilities are vitiated air facilities and are in vibrational equilibrium. However, with a flow of clean air, the vibrational temperature of N2 along a streamline remains approximately constant between the measurement plane and the facility heater, the vibrational temperature of O2 in the duct is about 1000 K, and the rotational temperature is consistent with the isentropic flow. The measurements of N2 vibrational temperature enabled cross-stream nonuniformities in the temperature exiting the facility heater to be documented. The measurements are in agreement with computational fluid dynamics models employing separate lumped vibrational and translational/rotational temperatures. Measurements and computations are also reported for a few percent steam addition to the air. The effect of the steam is to bring the flow to thermal equilibrium, also in agreement with the computational fluid dynamics.

Cutler, Andrew D.↗

NASA NTRS2004

NASA Multidimensional Stirling Convertor Code Developed

A high-efficiency Stirling Radioisotope Generator (SRG) for use on potential NASA Space Science missions is being developed by the Department of Energy, Lockheed Martin, Stirling Technology Company, and the NASA Glenn Research Center. These missions may include providing spacecraft onboard electric power for deep space missions or power for unmanned Mars rovers. Glenn is also developing advanced technology for Stirling convertors, aimed at substantially improving the specific power and efficiency of the convertor and the overall power system. Performance and mass improvement goals have been established for second- and third-generation Stirling radioisotope power systems. Multiple efforts are underway to achieve these goals, both in house at Glenn and under various grants and contracts. These efforts include the development of a multidimensional Stirling computational fluid dynamics code, high-temperature materials, advanced controllers, an end-to-end system dynamics model, low-vibration techniques, advanced regenerators, and a lightweight convertor. Under a NASA grant, Cleveland State University (CSU) and its subcontractors, the University of Minnesota (UMN) and Gedeon Associates, have developed a twodimensional computer simulation of a CSUmod Stirling convertor. The CFD-ACE commercial software developed by CFD Research Corp. of Huntsville, Alabama, is being used. The CSUmod is a scaled version of the Stirling Technology Demonstrator Convertor (TDC), which was designed and fabricated by the Stirling Technology Company and is being tested by NASA. The schematic illustrates the structure of this model. Modeled are the fluid-flow and heat-transfer phenomena that occur in the expansion space, the heater, the regenerator, the cooler, the compression space, the surrounding walls, and the moving piston and displacer. In addition, the overall heat transfer, the indicated power, and the efficiency can be calculated. The CSUmod model is being converted to a two-dimensional model of the TDC at NASA Glenn. Validation of the multidimensional Stirling code is an important part of the grant effort. UMN has been generating data in an oscillating-flow test facility using two different test sections: a 90 turn and a cooler/regenerator/heater test section. CSU has created computational fluid dynamics models of both these test sections and has been making comparisons with the data, then improving their models to improve the agreement with the test data. CSU has also been using data available in the literature for code validation. UMN is now preparing to begin fabrication of a new 180 turn test section that will be more representative of certain portions of the Stirling engine geometry. Simulations to almost periodic steady state with the two-dimensional CSUmod model indicate that, to reach periodic steady state on a single 2-GHz desktop computer, 75 to 100 complete simulation cycles would be required and between 1 and 2 months of computer time. Therefore, Glenn has purchased the first 8 computers, of a 64-computer cluster, to be run in parallel to accelerate the simulation. On the basis of CFD Research Corp.'s experience with running the parallelized version of CFD-ACE on their clusters, we estimate that the complete 64-computer cluster will reduce simulation computing time by a factor of about 40. Plans are to continue development of these multidimensional Stirling codes and to use them to study the fluid-flow and heat-transfer phenomena that occur inside Stirling convertors. This is expected to lead to improved thermodynamic loss understanding, onedimensional design and performance codes, and engine performance.

Tew, Roy C.↗

NASA NTRS2012

NASA Trapezoidal Wing Computations Including Transition and Advanced Turbulence Modeling

Flow about the NASA Trapezoidal Wing is computed with several turbulence models by using grids from the first High Lift Prediction Workshop in an effort to advance understanding of computational fluid dynamics modeling for this type of flowfield. Transition is accounted for in many of the computations. In particular, a recently-developed 4-equation transition model is utilized and works well overall. Accounting for transition tends to increase lift and decrease moment, which improves the agreement with experiment. Upper surface flap separation is reduced, and agreement with experimental surface pressures and velocity profiles is improved. The predicted shape of wakes from upstream elements is strongly influenced by grid resolution in regions above the main and flap elements. Turbulence model enhancements to account for rotation and curvature have the general effect of increasing lift and improving the resolution of the wing tip vortex as it convects downstream. However, none of the models improve the prediction of surface pressures near the wing tip, where more grid resolution is needed.

Rumsey, C. L.↗