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NASA NTRSDate not supplied

Multiphase Simulations of the SLS Launch Environment

NASA’s Space Launch System (SLS), which will send astronauts back to the Moon in the next few years, is powered by four RS-25 engines and two RSRMV solid rocket boosters (SRBs). During launch the SLS propulsion system generates intense acoustics and other powerful waves, such as ignition overpressure (IOP) which, if unmitigated, have the potential to damage the vehicle and possibly cause loss of mission or crew. To protect the vehicle from these powerful waves, the SLS launch pad design includes an ignition overpressure/sound suppression (IOP/SS) system which sprays 270,000 gallons per minute of water very close to the SRB and RS-25 nozzles. The SRB and RS-25 engine plumes, and the proximity of the IOP/SS water, create a complex multiphase (gas and liquid) environment during the SLS ignition sequence. The interplay among these systems creates challenges related to water spray into/onto engine nozzles, potential debris transport, and additional transient loads due to strong plume-water interactions - all of which the SLS vehicle must be able to withstand. Prior to the Artemis I launch, the SLS multiphase liftoff environment was largely unknown due to differences from the Space Shuttle and other programs. Some data was available from tests of individual systems, but no integrated testing or analysis was available. Even post-launch analysis of Artemis I cannot provide a full understanding of the complex physics involved due to limited (or obstructed) camera views and instrumentation. Computational fluid dynamics (CFD) is being used to investigate the details of the multiphase environment which could not be measured, help comprehend the data gathered from the launch, and ultimately identify phenomena that are a concern for future flights. Project Details Engineers at NASA’s Marshall Space Flight Center (MSFC) have executed simulations using the Loci/STREAM-Volume of Fluid (VoF) multiphase CFD solver to understand this environment. Initial efforts successfully validated the CFD solver on various tests, giving confidence to simulate the SLS multiphase liftoff environment prior to the Artemis I launch. The CFD simulation of the SLS ignition sequence was conducted in three phases. First the IOP/SS water system was simulated for approximately 6 seconds to reach a quasi-steady state. Next, the RS-25 engine plumes were activated and held at full power for 1 second. Lastly, the SRB booster was activated and the simulation was carried out until just prior to vehicle motion. This simulation process mimics the conditions that exist at launch. Results and Impact The SLS ignition sequence simulation results provide deep understanding of the underlying physics occuring during launch. Observations from the simulation include reduction of water splashing into/onto the engine nozzles, change in angling of the dense water sheets, and the origin of the powerful ignition overpressure (IOP) wave. These observations directly inform the SLS program on subjects including plume-water induced side loads, debris transport, and the acoustic launch environment. Additionally, with post launch comparison of CFD observations to flight data, these tools can be applied to launch vehicles and environments other than SLS with confidence. Why HPC Matters The SLS ignition sequence CFD simulations are conducted on meshes up to hundreds of millions of cells on thousands of processors for weeks at a time. These simulations generate terabytes of data that must also be stored and archived for future use on HPC systems. Simply put, the CFD simulations would not be possible without NASA HPC resources. What’s Next Comparisons between the Artemis I flight data and the CFD simulations will be continued to both improve confidence in the CFD results and provide deeper understanding into the SLS multiphase launch environment. This will be used to provide insight for decision making for the first manned SLS flight, Artemis II. Future simulations will target new configurations of the SLS IOP/SS water required to support the more powerful variants of the SLS vehicle, such as Block 1B. Additionally, this capability provides NASA the ability to investigate launch environments for vehicles other than SLS to support other missions.

Travis Rivord↗

NASA NTRS1991

Algorithm Computes Hypersonic Flow Of Air

Algorithm simulates three-dimensional hypersonic flow of air or another mixture of gases around blunt body. Accounts for nonequilibrium thermochemical effects via two-temperature mathematical model of molecular excitations and equations for finite-rate chemical reactions. These equations fully coupled to fluid dynamical equations.

Palmer, Grant↗

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 NTRS1991

Navier-Stokes computations with finite-rate chemistry for LO2/LH2 rocket engine plume flow studies

Computational fluid dynamics methods have been developed and applied to Space Shuttle Main Engine LO2/LH2 plume flow simulation/analysis of airloading and convective base heating effects on the vehicle at high flight velocities and altitudes. New methods are described which were applied to the simulation of a Return-to-Launch-Site abort where the vehicle would fly briefly at negative angles of attack into its own plume. A simplified two-perfect-gases-mixing approach is used where one gas is the plume and the other is air at 180-deg and 135-deg flight angle of attack. Related research has resulted in real gas multiple-plume interaction methods with finite-rate chemistry described herein which are applied to the same high-altitude-flight conditions of 0 deg angle of attack. Continuing research plans are to study Orbiter wake/plume flows at several Mach numbers and altitudes during ascent and then to merge this model with the Shuttle 'nose-to-tail' aerodynamic and SRB plume models for an overall 'nose-to-plume' capability. These new methods are also applicable to future launch vehicles using clustered-engine LO2/LH2 propulsion.

Dougherty, N. Sam↗

NASA NTRS1993

A computational investigation of fuel mixing in a hypersonic scramjet

A parabolized, Navier-Stokes code, SHIP3D, is used to numerically investigate the mixing between air injection and hydrogen injection from a swept ramp injector configuration into either a mainstream low-enthalpy flow or a hypervelocity test flow. The mixing comparisons between air and hydrogen injection reveal the importance of matching injectant-to-mainstream mass flow ratios. In flows with the same injectant-to-mainstream dynamic pressure ratio, the mixing definition was altered for the air injection cases. Comparisons of the computed results indicate that the air injection cases overestimate the mixing performance associated with hydrogen injection simulation. A lifting length parameter, to account for the time a fluid particle transverses through the mixing region, is defined and used to establish a connection of injectant mixing in hypervelocity flows, based on nonreactive, low-enthalpy flows.

Fathauer, Brett W.↗

NASA NTRS1998

Subsonic Maneuvering Effectiveness of High Performance Aircraft Which Employ Quasi-Static Shape Change Devices

This paper represents an initial study on the use of quasi-static shape change devices in aircraft maneuvering. The macroscopic effects and requirements for these devices in flight control are the focus of this study. Groups of devices are postulated to replace the conventional leading-edge flap (LEF) and the all-moving wing tip (AMT) on the tailless LMTAS-ICE (Lockheed Martin Tactical Aircraft Systems - Innovative Control Effectors) configuration. The maximum quasi-static shape changes are 13.8% and 7.7% of the wing section thickness for the LEF and AMT replacement devices, respectively. A Computational Fluid Dynamics (CFD) panel code is used to determine the control effectiveness of groups of these devices. A preliminary design of a wings-leveler autopilot is presented. Initial evaluation at 0.6 Mach at 15,000 ft. altitude is made through batch simulation. Results show small disturbance stability is achieved, however, an increase in maximum distortion is needed to statically offset five degrees of sideslip. This only applies to the specific device groups studied, encouraging future research on optimal device placement.

Montgomery, Raymond C.↗

NASA NTRS1990

The Importance of the Measurement of the Unsteady Wake of the Aeroassist Flight Experiment Vehicle: A White Paper

The High Frequency Radiometer (HFR) is the only instrument on the Aeroassist Flight Experiment (AFE) with sufficient temporal resolution to discern the frequency of unsteady wake oscillations. Determining both the frequency and amplitude of wake unsteadiness during AFE atmospheric entry is essential for reliably predicting the geometry and motion of the wake of future Aeroassisted Space Transfer Vehicles (ASTV). These parameters directly affect the location and size of the pay- load and the weight of the required afterbody heat protection. The purpose of the AFE is to validate the technologies required for the design of ASTVs, which will be used to exploit Earth-lunar space. This validation will be conducted at a combination of vehicle size, altitude, and velocity not obtainable in ground-based facilities. The AFE will provide the experimental flight data needed to improve our understanding of hypersonic-wake physics and to validate computational predictions of the aerodynamic and heating loads, including afterbody radiative heating loads, on an ASTV. Reliable prediction of ASTV wake flows will ensure that payloads are located within the shear-layer envelope and will determine the amount of thermal protection the payloads require. Specifically, understanding the temporal nature of the wake unsteadiness is important for two reasons. Most importantly, analysis of ground-based experiments suggests that wake unsteadiness results in a variation of as much as +/- 5 deg in the shear-flow turning angle. This angle must be reliably predicted to avoid shear-layer impingement on the vehicle afterbody, which would result in heating rates of about 10 W/sq cm, of the same order as on the forebody stagnation point. Secondly, the energy associated with wake unsteadiness will reduce the static enthalpy of the wake fluid and cause an error of as much as 30% in the amount of predicted wake radiative heating. Therefore, the HFR flight data, which will quantify the frequency and amplitude of the wake unsteadiness, are required for the verification of computational models of ASTV flowfields. Measurement of wake unsteadiness has been endorsed by the Peer Science Steering Group and the AFE Computational Fluid Dynamics (CFD) Working Group. The Peer Science Steering Group has stressed that a reliable measurement of the wake unsteadiness is fundamental because incorporating unsteadiness represents a substantial challenge to the CFD community, and a reliable measurement will raise confidence in the computer simulation. The AFE CFD Working Group has stated that 44 any data which would prove the existence and effects of unsteady flow would be extremely valuable.

Strawa, A. W.↗

NASA NTRS2001

Thermal Fault Tolerance Analysis of Carbon Fiber Rope Barrier Systems for Use in the Reusable Solid Rocket Motor ( RSRM) Nozzle Joints

Carbon Fiber Rope (CFR) thermal barrier systems are being considered for use in several RSRM (Reusable Solid Rocket Motor) nozzle joints as a replacement for the current assembly gap close-out process/design. This study provides for development and test verification of analysis methods used for flow-thermal modeling of a CFR thermal barrier subject to fault conditions such as rope combustion gas blow-by and CFR splice failure. Global model development is based on a 1-D (one dimensional) transient volume filling approach where the flow conditions are calculated as a function of internal 'pipe' and porous media 'Darcy' flow correlations. Combustion gas flow rates are calculated for the CFR on a per-linear inch basis and solved simultaneously with a detailed thermal-gas dynamic model of a local region of gas blow by (or splice fault). Effects of gas compressibility, friction and heat transfer are accounted for the model. Computational Fluid Dynamic (CFD) solutions of the fault regions are used to characterize the local flow field, quantify the amount of free jet spreading and assist in the determination of impingement film coefficients on the nozzle housings. Gas to wall heat transfer is simulated by a large thermal finite element grid of the local structure. The employed numerical technique loosely couples the FE (Finite Element) solution with the gas dynamics solution of the faulted region. All free constants that appear in the governing equations are calibrated by hot fire sub-scale test. The calibrated model is used to make flight predictions using motor aft end environments and timelines. Model results indicate that CFR barrier systems provide a near 'vented joint' style of pressurization. Hypothetical fault conditions considered in this study (blow by, splice defect) are relatively benign in terms of overall heating to nozzle metal housing structures.

Clayton, J. Louie↗

NASA NTRS2019

SINDA/FLUINT and Thermal Desktop Multi-Node Settled and Unsettled Propellant Tank Modeling of Zero Boil Off Test

Cryogenic propellant storage tank self-pressurization involves complex physical phenomena which are usually analytically modelled via complex multidimensional CFD (Computational Fluid Dynamics) codes. Unfortunately these codes, even when modelling axisymmetric domains, may take weeks or longer to obtain transient pressure and temperature information for relatively short periods of time (several seconds to several hours). Propellant tank storage end-to-end mission simulations can last a duration of days to weeks to months. Multi-node modelling of propellant tanks is a viable alternative to traditional CFD modelling and presents the advantage of greatly reduced run times on the order of hours and days compared to the weeks or longer for CFD codes. A multi-node model represents the fluid within the storage tank, as well as the storage tank itself, as a fluid-thermal network. This type of setup is not necessarily geometrically based. This can be accomplished using a commercial generalized fluid-thermal network code, such as SINDA/FLUINT (SF). The advantage of using a fluid-thermal network code like SF lies in its extensive ability to model the external environment of the storage tank through the graphical user interface, Thermal Desktop (TD). The total heat load into the tank may be a function of heaters and a complex radiative environment as well. Thermal Desktop may be used to address the detailed radiative environment of the tank as well as building a geometrically accurate depiction of the storage tank itself.

Thermal Desktop↗

NASA NTRS1988

Flowfield visualization for SSME hot gas manifold

The objective of this research, as defined by NASA-Marshall Space Flight Center, was two-fold: (1) to numerically simulate viscous subsonic flow in a proposed elliptical two-duct version of the fuel side Hot Gas Manifold (HGM) for the Space Shuttle Main Engine (SSME), and (2) to provide analytical support for SSME related numerical computational experiments, being performed by the Computational Fluid Dynamics staff in the Aerophysics Division of the Structures and Dynamics Laboratory at NASA-MSFC. Numerical results of HGM were calculations to complement both water flow visualization experiments and air flow visualization experiments and air experiments in two-duct geometries performed at NASA-MSFC and Rocketdyne. In addition, code modification and improvement efforts were to strengthen the CFD capabilities of NASA-MSFC for producing reliable predictions of flow environments within the SSME.

Roger, Robert P.↗

NASA NTRS1992

NAS technical summaries: Numerical aerodynamic simulation program, March 1991 - February 1992

NASA created the Numerical Aerodynamic Simulation (NAS) Program in 1987 to focus resources on solving critical problems in aeroscience and related disciplines by utilizing the power of the most advanced supercomputers available. The NAS Program provides scientists with the necessary computing power to solve today's most demanding computational fluid dynamics problems and serves as a pathfinder in integrating leading-edge supercomputing technologies, thus benefiting other supercomputer centers in Government and industry. This report contains selected scientific results from the 1991-92 NAS Operational Year, March 4, 1991 to March 3, 1992, which is the fifth year of operation. During this year, the scientific community was given access to a Cray-2 and a Cray Y-MP. The Cray-2, the first generation supercomputer, has four processors, 256 megawords of central memory, and a total sustained speed of 250 million floating point operations per second. The Cray Y-MP, the second generation supercomputer, has eight processors and a total sustained speed of one billion floating point operations per second. Additional memory was installed this year, doubling capacity from 128 to 256 megawords of solid-state storage-device memory. Because of its higher performance, the Cray Y-MP delivered approximately 77 percent of the total number of supercomputer hours used during this year.

Source record↗

NASA NTRS1993

Research activities at the Center for Modeling of Turbulence and Transition

The main research activities at the Center for Modeling of Turbulence and Transition (CMOTT) are described. The research objective of CMOTT is to improve and/or develop turbulence and transition models for propulsion systems. The flows of interest in propulsion systems can be both compressible and incompressible, three dimensional, bounded by complex wall geometries, chemically reacting, and involve 'bypass' transition. The most relevant turbulence and transition models for the above flows are one- and two-equation eddy viscosity models, Reynolds stress algebraic- and transport-equation models, pdf models, and multiple-scale models. All these models are classified as one-point closure schemes since only one-point (in time and space) turbulent correlations, such as second moments (Reynolds stresses and turbulent heat fluxes) and third moments, are involved. In computational fluid dynamics, all turbulent quantities are one-point correlations. Therefore, the study of one-point turbulent closure schemes is the focus of our turbulence research. However, other research, such as the renormalization group theory, the direct interaction approximation method, and numerical simulations are also pursued to support the development of turbulence modeling.

Shih, Tsan-Hsing↗

NASA NTRS1994

NAS technical summaries. Numerical aerodynamic simulation program, March 1992 - February 1993

Source record↗

NASA NTRS1995

NAS Technical Summaries, March 1993 - February 1994

Source record↗

NASA NTRS2015

Development of an Aerodynamic Analysis Method and Database for the SLS Service Module Panel Jettison Event Utilizing Inviscid CFD and MATLAB

This paper describes the development, testing, and utilization of an aerodynamic force and moment database for the Space Launch System (SLS) Service Module (SM) panel jettison event. The database is a combination of inviscid Computational Fluid Dynamic (CFD) data and MATLAB code written to query the data at input values of vehicle/SM panel parameters and return the aerodynamic force and moment coefficients of the panels as they are jettisoned from the vehicle. The database encompasses over 5000 CFD simulations with the panels either in the initial stages of separation where they are hinged to the vehicle, in close proximity to the vehicle, or far enough from the vehicle that body interference effects are neglected. A series of viscous CFD check cases were performed to assess the accuracy of the Euler solutions for this class of problem and good agreement was obtained. The ultimate goal of the panel jettison database was to create a tool that could be coupled with any 6-Degree-Of-Freedom (DOF) dynamics model to rapidly predict SM panel separation from the SLS vehicle in a quasi-unsteady manner. Results are presented for panel jettison simulations that utilize the database at various SLS flight conditions. These results compare favorably to an approach that directly couples a 6-DOF model with the Cart3D Euler flow solver and obtains solutions for the panels at exact locations. This paper demonstrates a method of using inviscid CFD simulations coupled with a 6-DOF model that provides adequate fidelity to capture the physics of this complex multiple moving-body panel separation event.

Applebaum, Michael P.↗

NASA NTRS2019

Pilot Sensitivity to Simulator Flight Dynamics Model Formulation for Stall Training

A piloted simulation study was performed in the Cockpit Motion Facility at the National Aeronautics and Space Administration Langley Research Center. The research was motivated by the desire to reduce the commercial transport airplane fatal accident rate due to in-flight loss of control. The purpose of this study, which focused on a generic T-tail transport airplane, was to assess pilot sensitivity to flight dynamics model formulation used during a simulator stall recognition and recovery training/demonstration profile. To accomplish this, the flight dynamics model was designed with many configuration options. The model options were based on recently acquired static and dynamic stability and control data from sources that included wind tunnel, water tunnel, and computational fluid dynamics. The results, which are specific to a transport airplane stall recognition and recovery guided demonstration scenario, showed the two most important aerodynamic effects (other than stick pusher) to model were stall roll- off and the longitudinal static stability characteristic associated with the pitch break.

Cunningham, Kevin↗

NASA NTRSDate not supplied

Plume Impingement Studies in Space Environments for NASA Deep Space Logistics

NASA’s Artemis program is preparing to return Americans to the moon, in preparation for future deep space missions to Mars and beyond. The Lunar Gateway – a space station to be placed in orbit around the moon – serves as a critical component of the Artemis program, and will be the launchpad for future astronauts to visit the lunar surface and for humanity to learn how to live in a deep space environment. Much like the ISS, during operations and visiting vehicles approaches, the Lunar Gateway structures will be impinged upon from various plume sources: reaction control system rocket engine plumes, wastewater venting plumes, and plasma plumes from electric thrusters for station keeping. This presentation will describe the engineering challenges associated with each of these plume impingement scenarios and will detail the physical models and simulation approaches to predict the thermomechanical loading, contamination accretion, and erosion of station components due to plume impingement in space environments.

computational fluid dynamics↗

NASA NTRS2022

Prediction of Turbulent Diffusing Flows Using FUN3D

The FUN3D code was used to perform Reynolds-averaged Navier-Stokes (RANS) simulations to compute subsonic flow in an S-duct diffuser and transonic flow in a two-dimensional diffuser using the Speziale-Sarkar-Gatski/Launder-Rodi-Reece (SSG/LRR) Reynolds stress model (RSM). For comparison purposes, additional simulations were run with the one-equation Spalart-Allmaras (SA) and the two-equation Menter Shear-Stress Transport (SST) turbulence models. Each model was run with and without the quadratic constitutive relation (QCR) for computing the turbulent stresses. It was shown that the simulations that utilized the RSM had better overall predictions of the diffusive flow fields compared to the simulations that utilized the one and two equation turbulence models.

Inlet Distortion↗