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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Ice-Accretion Test Results for Three Large-Scale Swept-Wing Models in the NASA Icing Research Tunnel

Icing simulation tools and computational fluid dynamics codes are reaching levels of maturity such that they are being proposed by manufacturers for use in certification of aircraft for flight in icing conditions with increasingly less reliance on natural-icing flight testing and icing-wind-tunnel testing. Sufficient high-quality data to evaluate the performance of these tools is not currently available. The objective of this work was to generate a database of ice-accretion geometry that can be used for development and validation of icing simulation tools as well as for aerodynamic testing. Three large-scale swept wing models were built and tested at the NASA Glenn Icing Research Tunnel (IRT). The models represented the Inboard (20% semispan), Midspan (64% semispan) and Outboard stations (83% semispan) of a wing based upon a 65% scale version of the Common Research Model (CRM). The IRT models utilized a hybrid design that maintained the full-scale leading-edge geometry with a truncated afterbody and flap. The models were instrumented with surface pressure taps in order to acquire sufficient aerodynamic data to verify the hybrid model design capability to simulate the full-scale wing section. A series of ice-accretion tests were conducted over a range of total temperatures from -23.8 deg C to -1.4 deg C with all other conditions held constant. The results showed the changing ice-accretion morphology from rime ice at the colder temperatures to highly 3-D scallop ice in the range of -11.2 deg C to -6.3 deg C. Warmer temperatures generated highly 3-D ice accretion with glaze ice characteristics. The results indicated that the general scallop ice morphology was similar for all three models. Icing results were documented for limited parametric variations in angle of attack, drop size and cloud liquid-water content (LWC). The effect of velocity on ice accretion was documented for the Midspan and Outboard models for a limited number of test cases. The data suggest that there are morphological characteristics of glaze and scallop ice accretion on these swept-wing models that are dependent upon the velocity. This work has resulted in a large database of ice-accretion geometry on large-scale, swept-wing models.

Wind-Tunnel Tests↗
NASA NTRS2018

Aeromechanics Analysis of a Distortion-Tolerant Fan with Boundary Layer Ingestion

A propulsion system with Boundary Layer Ingestion (BLI) has the potential to significantly reduce aircraft engine fuel burn. But a critical challenge is to design a fan that can operate continuously with a persistent BLI distortion without aeromechanical failure -- flutter or high cycle fatigue due to forced response. High-fidelity computational aeromechanics analysis can be very valuable to support the design of a fan that has satisfactory aeromechanic characteristics and good aerodynamic performance and operability. Detailed aeromechanics analyses together with careful monitoring of the test article is necessary to avoid unexpected problems or failures during testing. In the present work, an aeromechanics analysis based on a three-dimensional, time-accurate, Reynolds-averaged Navier-Stokes computational fluid dynamics code is used to study the performance and aeromechanical characteristics of the fan in both circumferentially-uniform and circumferentially-varying distorted flows. Pre-test aeromechanics analyses are used to prepare for the wind tunnel test and comparisons are made with measured blade vibration data after the test. The analysis shows that the fan has low levels of aerodynamic damping at various operating conditions examined. In the test, the fan remained free of flutter except at one near-stall operating condition. Analysis could not be performed at this low mass flow rate operating condition since it fell beyond the limit of numerical stability of the analysis code. The measured resonant forced response at a specific low-response crossing indicated that the analysis under-predicted this response and work is in progress to understand possible sources of differences and to analyze other larger resonant responses. Follow-on work is also planned with a coupled inlet-fan aeromechanics analysis that will more accurately represent the interactions between the fan and BLI distortion.

Propulsion Aeroelasticity↗
NASA NTRS2018

Sensitivity Analysis for Multidisciplinary Systems (SAMS)

This report describes the research conducted under an interagency collaboration agreement between the Aerospace Systems Directorate of the Air Force Research Laboratory (AFRL/RQ) and the Computational AeroSciences Branch of NASA Langley (NASA LaRC). Both organizations have a long-term goal of developing a modular computational system for coupling fluids and structures to enable both analysis and optimization of aerospace vehicles. Ultimately, the system should support multiple solvers within the fluid and structure domains to allow the best combination for the task at hand, as well as to allow for institutional preferences of specific software components. Towards this goal, the current research was focused on enhancing the existing modal aeroelastic analysis in the NASA FUN3D (Fully-UNstructured three-dimensional CFD (Computational Fluid Dynamics) code) software (Biedron et al. 2018), as well as developing new aeroelastic analysis and optimization capabilities based on a non-linear finite-element method. The methods and enhancements described in this document pertain to FUN3D Version 13.4.

Biedron, Robert T.↗
NASA NTRS2017

Implementation of a Body Force Model in OVERFLOW for Propulsor Simulations

We present an implementation of a propulsor model based on body force method into the OVERFLOW computational fluid dynamics code to model turbofan engines and propulsors of similar type. The model estimates the forces imparted on the fluid by the blade camber surfaces as a body force source terms in the momentum and energy equations over grids that represent the rotor and stator of a fan stage. We tested the implementation on three cases: 1) Source Diagnostics Test (SDT) fan with R4 rotors, which has extensive test data on aerodynamic performance and rotor wake surveys, 2) A stand-alone Aeronaut TF8000 propulsor which is an off-the-shelf propulsor used on model aircraft and 3) The D8 aircraft model with TF8000 propulsors placed in a wind tunnel. Despite missing some of the features in the velocity profiles near the endwalls (i.e. hub and casing surfaces), the OVERFLOW simulations with body force model predicted area-averaged flow speed and total pressure rise through the SDT propulsor within a few percent of the LDV (Laser Doppler Velocimetry) measurements. In the case of TF8000 propulsor on the D8 airframe, the model under-predicted mechanical power coefficient by several percent of the wind tunnel test results when the horizontal force balance condition over the airframe is targeted by tuning rotor speed. By investigating the upstream influence of the rotor swirl, it was found out that the induced swirl velocity effects upstream were relatively small and they rapidly vanished before reaching one fan diameter upstream of the fan face. The body force model provided insights on aerothermodynamics and aeromechanics of boundary layer ingesting propulsor; these insights could not be obtained by using the uniform pressure jump model.

Propulsor↗
NASA NTRS2016

Thermochemistry of Strong Air Plasmas for Hypervelocity Earth Entry of Asteroids

The specific heat, enthalpy and free energy of nitrogen and oxygen are calculated for pressures between 1 and 100 bars and temperatures between 10,000 and 100,000 degrees Kelvin. The data are presented in a form suitable for input into computational fluid dynamics computer codes. This range of temperature and pressure are typical of the conditions expected in the bow shock layer of a meteor or small asteroid entering Earth's atmosphere. The extent of double and triple ionization of nitrogen and oxygen atoms under these conditions will be estimated.

Ionization Energy Depression↗
NASA NTRS2019

Risk Analysis Associated with Loss of Toxic Gases During Orion Landing and Recovery Operations

Mission, landing and recovery operations for the Orion crew module involve reentry into the Earth's atmosphere and the deployment of three Nomex parachutes to slow the descent before landing along the west coast of the United States. Orion may have residual fuel (hydrazine, N2H4) or coolant (ammonia, NH3) on board which are both highly toxic to crew in the event of exposure. These risks were evaluated using a first principles analysis approach through fluid dynamics modeling. Plume calculations were first performed with the ANSYS Fluent computational fluid dynamics code. Data were then extracted at locations relevant to crew safety such as the snorkel fan inlet and the egress hatch. Mixing calculations were performed to quantify exposure concentrations within the crew bay before and during egress and departure. Finally, results included herein were used to inform the Orion post-landing Concept of Operations (ConOps) so that strategies could be formulated to maintain crew safety in the event of the loss of fuel or coolant.

Orion↗
NASA NTRS2019

Investigation of Transonic Truss-Braced Wing Aircraft Transonic Wing-Strut Interference Effects Using FUN3D

This paper presents a computational study of transonic wing-strut interference effects of Transonic Truss-Braced Wing (TTBW) aircraft using the high-fidelity CFD (Computational Fluid Dynamics) code FUN3D (Fully Unstructured Navier-Stokes Three Dimensional). The study is conducted for the wing-strut and the wing-alone configurations at different Mach numbers and Reynolds numbers. The interference effects are calculated by comparing the wing aerodynamics along the spanwise direction between the wing-strut and the wing-alone configurations. The presence of the strut underneath the wing induces a suction peak on the lower surface of the wing, which causes changes in aerodynamic forces and moments, as well as the aerodynamic center location. The interference effects become more pronounced as the Mach number increases. The Reynolds number has less impact on the interference effects. A transonic wing-strut interference aerodynamic correction method is developed for use in a lower-fidelity tool, VSPAERO [a vortex lattice flow solver], coupled to a finite-element model for rapid flutter analysis.

Xiong, Juntao↗
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 NTRS2020

Computational Investigation of Powered Descent for Human-Scale Mars Landers

NASA human-scale Mars entry, descent, and landing vehicles currently under consideration have system requirements beyond the current state of the art. Specifically, heritage approaches for deceleration using parachutes and initiating engines at subsonic speeds are not adequate to land these large vehicles (50-60 metric ton entry mass) precisely (within 50 meters of a surface target). Therefore, the present approach for landing humans on Mars utilizes entry vehicles with large aerodynamic surface areas that transition to the descent and landing phase by initiating retropropulsion at supersonic speeds. Vehicle performance, including stability and control, depends on accurate characterization of the aerodynamic-propulsive interference effects, especially at engine initiation, when the vehicles are transitioning from an entry attitude to that suitable for descent and landing. This paper summarizes the human Mars entry vehicle configurations, retropropulsion integration, performance assumptions, and expected flight environment at engine initiation. Cases spanning supersonic, transonic, and subsonic conditions during powered descent are analyzed using four computational fluid dynamics codes, all of which have been previously applied to supersonic retropropulsion at varying scales and conditions. The results of the analysis are used to develop an aerodynamic-propulsive interference model that has since been applied in trajectory simulations to augment the fidelity of human-scale Mars entry vehicle performance.

Ashley M Korzun↗
NASA NTRS2020

Tuning Neural Network Models for Improved Prediction of Boundary Layer Transition

Boundary layer transition can strongly impact flight vehicle performance as it influences surface skin friction and aerodynamic heating, making accurate transition prediction a key to designing next generation aircraft. Artificial neural networks (ANNs) have shown promise toward predicting laminar-turbulent transition based on linear stability correlations. The computational efficiency of ANNs and the substantially reduced user involvement in relation to direct computations based on the linear stability theory (LST) makes them an attractive methodology for integrating the LST based correlations in computational fluid dynamics codes. Tollmien-Schlichting (TS) waves correspond to the dominant transition mechanism in 2D or weakly 3D subsonic boundary layers, such as those encountered in general aviation applications. Improvements to neural network model accuracy in predicting the amplification rates of TS instability waves have been investigated by leveraging recent machine learning developments in conjunction with surrogate optimization techniques and via suitable augmentation of the data used to train the networks. The optimized models trained on the modified dataset reduced the average transition location errors on different airfoils at several flow conditions by 51% of the original manually-tuned network’s errors on the same flow cases. The actual transition locations were derived from the Langley Stability and Transition Analysis Code (LASTRAC).

Machine Learning↗
NASA NTRSDate not supplied

An Interactive MATLAB Program for Fitting Transfer Functions to Frequency Responses

A computer program called FRFit (Frequency Response Fitting) is described for fitting single-input single-output transfer function models to empirical frequency response data. The program is interactive in that the user specifies ``elementary factors'' (gain, delay, pure differentiators and integrators, and first- and second-order zeros and poles) by entering numerical values or moving sliders in a graphical user interface. A nonlinear optimization can then be performed to obtain maximum likelihood estimates of transfer function parameters and uncertainties to provide feedback on the modeling and refine estimates. Several examples are discussed, including the identification of aircraft pitch dynamics from simulation data and data reported in the literature, approximating Theodorsen's function of unsteady aerodynamics, and obtaining a reduced-order model of a computational fluid dynamics code describing the unsteady aerodynamics around an aeroelastic wing. The program has some usefulness as a teaching aid, and can be applied to model structure determination, reduced-order modeling, preliminary analysis, and simple system identification problems. The program was written in MATLAB and is planned for public release through the NASA Software Catalog.

System identification↗
NASA NTRSDate not supplied

Development of the 10% NASA CRM High-Lift Common Research Model (CRM-HL)

A 10%-scale Common Research Model-high lift version (CRM-HL) was developed to support extensive experimental investigations in the areas of active flow control, control surface optimization for landing and take-off, acoustic technology development, and computational fluid dynamics code validation to name a few. The CRM-HL was designed to have modular components that will enable the model to be tested in the following three forms: (1) conventional high-lift (CHL) configuration, (2) Active Flow Control (AFC) enabled simple hinged flap high-lift (SHL) configuration, and (3) CHL configuration with noise reduction treatments for the slats and/or the flaps. This paper describes the design, analysis, and fabrication of the CRM-HL for testing in the NASA Langley Research Center (LaRC) 14 by 22-Foot Subsonic Tunnel (14x22) and other wind tunnels as deemed appropriate by the research community such as the Qinetiq 5 Meter tunnel.

Jared S. Fell↗
NASA NTRSDate not supplied

Large Vehicle Lunar Landing Surface Interaction and In-Situ Resource Based Risk Mitigation: Landing & Launch Pads

A key capability required for the exploration of planetary bodies is the ability to land on the surface. Previous work performed by NASA and other institutions has primarily focused on landing small spacecraft on planetary surfaces and the associated small-to-medium thrusters required for the soft landing. In the case of human exploration—particularly the establishment of long duration exploration and habitation outposts—the ability to land large landers, such as the SpaceX Starship, is necessary. These larger landing systems require the use of more powerful engines, with higher engine exhaust temperatures and higher landing loads. Understanding the excavation of material by the engines, as well as the potential for the landing legs to sink into the subsurface, is key in ensuring reliable and safe landings. A further improvement in landing reliability can be achieved by constructing landing / launch pads, especially with in-situ resources. Some material excavation by the plume is inevitable, leaving at least a portion of the surface scoured and uneven under the lander and ejecting regolith particles and rocks at very high velocities. One possible solution would be to robotically build landing / launch pads (ideally autonomously) at the destination using in-situ materials. In this case, the first one or few landers will need to land on unimproved surfaces at higher risk; however, they would bring the required equipment to build the landing pads with mostly local resources, thus increasing the reliability of safe landing for subsequent larger landers. A number of methods to build in-situ landing and launch pads have already been developed. These methods include no, or some, addition of required binder additives to the local regolith material, different processing approaches and result in varying landing pad strengths. A sub-scale rocket engine plume, was used to simulate some of the conditions of a landing on the Moon to assess the effectiveness of various materials for an in-situ built landing pad, The GO2/GCH4rocket engine fired on a 1m2area coupons of representative pad materials. The results will allow continued development towards materials that satisfy the landing pad properties required for the effective risk reduction and increased reliability for landing people and equipment on the lunar surface. This work contained two parts: (1) computer modeling of a large rocket engine plume interacting with regolith on the Moon, using the Granular Gas Flow Solver (GGFS) provided by CFD Research Corporation as well as other computational fluid dynamics codes (CFD) such as Loci/CHEM. (2) Developing landing/launch pad materials that could be used for in-situ construction on the lunar surface in the future, to mitigate the calculated effects of a large vehicle rocket engine landing and launching on the Moon.

Lunar↗
NASA NTRS2020

Space Launch Vehicle Transient Particle Redistribution Modeling and Implications for Optically Sensitive Payloads

The performance of contamination sensitive components—such as optical components—can be degraded by particulate matter depositing on the surfaces. Particles can accumulate during manufacturing, handling and operation. For a space-based system, particles can shed from the fairing and redistribute onto sensitive surfaces during launch. An engineering modeling approach has been developed for modeling particle migration during launch. The approach involves particle detachment from the fairing, particle transport through the venting atmosphere inside the fairing, and attachment to the receiving surface. Particle size and amounts on the fairing surface can be modeled using distributions from standards, such as IEST-STD-CC1246E, as well as from empirical data obtained from tape lifts. Surface interactions are modeled using theoretical as well as empirical data. Commercial computational fluid dynamics codes are used to calculate the gas flow in the fairing during depressurization during launch. This approach not only provides insight into particle redistribution during launch but also can be used to establish fairing cleanliness requirements.

Brieda, Lubos↗
NASA NTRS2022

Simulation and Regression Modeling of Nasa'S X-59 Low-Boom Carpets Across America

NASA’s X-59 aircraft is predicted to produce a significantly quieter cruise sonic boom than traditional N-wave-producing aircraft. A propagation simulation study was undertaken to quantify loudness levels, exposure size, and variability of the X-59’s low-boom carpet using realistic atmospheric profiles across the contiguous United States of America (CONUS). Near-field pressure data of the X-59 in supersonic cruise from NASA’s fully unstructured Navier–Stokes three-dimensional (known as FUN3D) computational fluid dynamics code were propagated using NASA’s PCBoom code, which solves an enhanced Burgers equation along acoustic rays. Atmospheric profiles from the National Oceanic and Atmospheric Administration’s Climate Forecast System Version 2 database were used for propagation at 138 locations across the CONUS. Carpets at each location were generated for aircraft headings in the four cardinal directions. Over one million X-59 carpets were generated in total. The effects of the heading, season, geography, and climate zone on boom levels and exposure size are presented. Multiple linear regression models were developed to estimate carpet width and loudness metrics across the CONUS. These results inform regulators and mission planners on expected variations in boom levels and carpet extent from atmospheric variations. Understanding potential carpet variability is important when planning community noise surveys using the X-59.

X-59↗
NASA NTRSDate not supplied

Simulation and Regression Modeling of X-59 Low-Boom Carpets Across America

The NASA X-59 aircraft is predicted to produce a significantly quieter cruise sonic boom than traditional N-wave-producing aircraft. A propagation simulation study was undertaken to quantify loudness levels, exposure size, and variability of the X-59 low-boom carpet using realistic atmospheric profiles across the contiguous United States of America (CONUS). Near-field pressure data of the X-59 in supersonic cruise from NASA’s fully unstructured Navier–Stokes three-dimensional (known as FUN3D) computational fluid dynamics code were propagated using NASA’s PCBoom code, which solves an enhanced Burgers equation along acoustic rays. Atmospheric profiles from the National Oceanic and Atmospheric Administration’s Climate Forecast System Version 2 database were used for propagation at 138 locations across the CONUS. Carpets at each location were generated for aircraft headings in the four cardinal directions. Over one million X-59 carpets were generated in total. The effects of the heading, season, geography, and climate zone on boom levels and exposure size are presented. Multiple linear regression models were developed to estimate carpet width and loudness metrics across the CONUS. These results inform regulators and mission planners on expected variations in boom levels and carpet extent from atmospheric variations. Understanding potential carpet variability is important when planning community noise surveys using the X-59.

X-59↗
NASA NTRSDate not supplied

Magnetohydrodynamic and Aerodynamic Assessment of Ballistic Entry of a 70deg Spherecone at Mars and Venus

An electrical conductivity database for continuum flow in a CO2 atmosphere over a 70 deg spherecone was created using the Data Parallel Line Relaxation Code computational fluid dynamics software to inform development of future magnetohydrodynamic subsystems at Venus and Mars. Sixteen freestream conditions were considered at Mars with atmospheric relative velocities from 5 to 8 km/s and altitudes between 20 and 80 km. Sixteen freestream conditions were considered at Venus with atmospheric relative velocities from 9 to 12 km/s and altitudes between 85 and 115 km. Results indicate that the total electrical conductivity in the flow volume always increases as velocity increases. At low velocities, the electrical conductivity is higher at high altitudes, while at high velocities, the electrical conductivity is higher at low altitudes. Three of the 80 km altitude computational fluid dynamics solutions show good agreement with Direct Simulation Monte Carlo results. In general, computational fluid dynamics predicts thinner shocks, higher electron number density, and similar vibrational temperatures as direct simulation Monte Carlo. The magnetohydrodynamic force was calculated at both Mars and Venus. Results indicate there may not be sufficient control authority to use magnetohydrodynamics as a trajectory control mechanism at Mars without artificially increasing the electrical conductivity of the flow, but there may be appreciable control authority for a drag-modulated aerocapture at Venus.

Destiny M Fawley↗
NASA NTRS2024

Computational Techniques to Generate Space Launch System Aerodynamic Databases

This document describes the reasoning and trade studies used to evaluate tools for constructing the aerodynamic lineload databases for the liftoff and transition phases of flight for the Space Launch System. Three computational fluid dynamics codes (USM3D, FUN3D, Kestrel) were investigated with various turbulence models, as well as detached eddy simulation variants for the launch vehicle in free air and in proximity to the tower. Decisions were made mostly based on results from brief developmental studies performed in response to specific, unforeseen challenges that were encountered in the analysis of a given configuration. The need to develop databases in a timely manner, as well as accurately capture the expected leeward-wake flowfield characteristics, led to the selection of the Kestrel flow solver with its delayed detached eddy simulation method, the Spalart-Allmaras turbulence model, and the adaptive mesh refinement capability in the off-body Cartesian grid region.

Karen A. Deere↗