ALBERT

Description: Large-eddy simulations are performed using wall-resolved mesh for a Mach 2.29 impinging shock wave/boundary-layer interaction. Flow conditions are based on an experiment and therefore entire span was simulated, including the two sidewalls. Mean flow comparison with the experimental data showed that the predicted interaction length was larger in the simulation. Time-series analysis of a rake of pressure signals immediately downstream of the mean reflected shock position showed a peak in weighted power spectral density occurred about St(sub Lint) = 0.01, owing to a larger interaction length. Budgets of Reynolds-stress transport calculated across the span and along the corner bisector showed high degree of anisotropy. Merging of the secondary flows and separation along the corner gave rise to unstable counter-rotating vortices, which straddle the corner and grow in size. This also leads to a development of new behavior in the viscous sublayer along the corner bisector, where the pressure strain and molecular diffusion mechanisms become prominent.

Description: This article discusses the use of numerical optimization procedures to aid in the calibration of turbulence model coefficients. Such methods would increase the rigor and repeatability of the calibration procedure by requiring clearly defined and objective optimization metrics, and could be used to identify unique combinations of coefficient values for specific flow problems. The approach is applied to the re-calibration of an explicit algebraic Reynolds stress model for the incompressible planar mixing layer using the Nelder-Mead simplex algorithm and a micro-genetic algorithm with minimally imposed constraints. Three composite fitness functions, each based upon the error in the mixing layer growth rate and the normal and shear components of the Reynolds stresses, are investigated. The results demonstrate a significant improvement in the target objectives through the adjustment of three pressure-strain coefficients. Adjustments of additional coefficients provide little further benefit. Issues regarding the effectiveness of the fitness functions and the efficiency of the optimization algorithms are also discussed.

Description: This manual describes the installation and execution of FUN3D (Fully-UNstructured three-dimensional CFD (Computational Fluid Dynamics) code) version 13.5, including optional dependent packages. FUN3D is a suite of computational fluid dynamics simulation and design tools that uses mixed-element unstructured grids in a large number of formats, including structured multiblock and overset grid systems. A discretely-exact adjoint solver enables efficient gradient-based design and grid adaptation to reduce estimated discretization error. FUN3D is available with and without a reacting, real-gas capability. This generic gas option is available only for those persons that qualify for its beta release status.

Description: A computational fluid dynamics code has been developed for large-eddy simulations (LES) of turbulent flow. The code uses high-order of accuracy and high-resolution numerical methods to minimize solution error and maximize the resolution of the turbulent structures. Spatial discretization is performed using explicit central differencing. The central differencing schemes in the code include 2nd- to 12th-order standard central difference methods as well as 7-, 9-, 11- and 13-point dispersion relation preserving schemes. Solution filtering and high-order shock capturing are included for stability. Time discretization is performed using multistage Runge-Kutta methods that are up to 4th order accurate. Several options are available to model turbulence including: Baldwin-Lomax and Spalart-Allmaras Reynolds-averaged Navier-Stokes turbulence models, and Smagorinsky, Dynamic Smagorinsky and Vreman sub-grid scale models for LES. This report presents the theory behind the numerical and physical models used in the code and provides a user's manual to the operation of the code.

Description: Large-eddy simulations are performed using wall-resolved mesh for a Mach 2.29 impinging shock wave/boundary-layer interaction. Flow conditions are based on an experiment and therefore entire span was simulated, including the two sidewalls. Mean flow comparison with the experimental data showed that the interaction was larger in the simulation. Time-series analysis of a rake of pressure probes immediately downstream of the mean reflected shock position showed a peak in weighted power spectral density occurred about $St_{Lint}=0.01$, owing to a larger interaction length. Budgets of Reynolds-stress transport calculated across the span and along the corner bisector showed high degree of anisotropy. Merging of the secondary flows and separation along the corner gives rise to unstablecounter rotating vortices, which straddle the corner and grow in size. This also leads to a development of new behavior in the viscous sublayer along the corner bisector, where the pressure strain andmolecular diffusion mechanisms become prominent.

Description: No abstract available

Description: The Predictive Thermal Control (PTC) technology development project is a multiyear effort initiated in Fiscal Year (FY) 2017, to mature the Technology Readiness Level (TRL) of critical technologies required to enable ultra-thermally-stable telescopes for exoplanet science. A key PTC partner is Harris Corporation (Rochester NY).

Description: Experiments are being conducted in the NASA Ames Hypervelocity Free Flight Aerodynamic Facility to quantify the effects on turbulent convective heat transfer of surface roughness representative of a new class of 3D woven thermal protection system mRough-wall turbulent heat transfer measurements were obtained on ballistic-range models in hypersonic flight in the NASA Ames Hypervelocity Free Flight Aerodynamic Facility. Each model had three different surface textures on segments of the conic frustum: smooth wall, sand roughness, and a pattern roughness, thus providing smooth-wall and sand-roughness reference data for each test. The pattern roughness was representative of a woven thermal protection system material developed by NASA's Heatshield for Extreme Entry Environment Technology project. The tests were conducted at launch speeds of 3.2 km/s in air at 0.15 atm. Roughness Reynolds numbers, k+, ranged for 12 to 70 for the sand roughness, and as high as 200 for the pattern roughness. Boundary-layer parameters required for calculating k+ were evaluated using computational fluid dynamics simulations. The effects of pattern roughness are generally characterized by an equivalent sand roughness determined with a correlation developed from experimental data obtained on specifically-designed roughness patterns that do not necessarily resemble real TPS materials. Two sand roughness correlations were examined: Dirling and van Rij, et al. Both gave good agreement with the measured heat-flux augmentation for the two larger pattern roughness heights tested, but not for the smallest height tested. It has yet to be determined whether this difference is due to limitations in the experimental approach, or due to limits in the correlations used. Future experiments are planned that will include roughness patterns more like those used in developing the equivalent sand roughness correlations.aterials being developed by NASA's Heatshield for Extreme Entry Environment Technology (HEEET) project. Data were simultaneously obtained on sand-grain roughened surfaces and smooth surfaces, which can be compared with previously obtained data. Results are presented in this extended abstract for one roughness pattern. The full paper will include results from three roughness patterns representing virgin HEEET, nominal turbulent ablated HEEET, and twice the roughness of nominal turbulent ablated HEEET. Results will be used to compare with commonly used equivalent sand grain roughness correlations.

Description: Temperature-dependent data of a RUAG six-component block-type balance was analyzed to assess the accuracy of two load prediction methods for temperature-dependent balance data. The supplied data was prepared for the analysis by splitting it into calibration and check load data subsets. The first calibration data subset was obtained at a temperature of 294 Kelvin. The second calibration data subset was obtained at a temperature of 315 Kelvin. A subset of 38 points was extracted from the second data set and used as check loads so that the accuracy of the two load prediction methods could be tested. First, the Iterative Method in combination with an extended independent and dependent variable set was used for the balance load prediction. This approach fits electrical outputs as a function of loads and the temperature and, afterwards, constructs a load iteration scheme from the regression coefficients so that loads can be predicted from outputs and the temperature during a wind tunnel test. The Non-Iterative Method was also used for the load prediction. This alternate method can more easily be implemented in a data system as loads are directly fitted as a function of electrical outputs and the temperature. Analysis results for the axial force are only discussed in the paper as similar results were obtained for the other five load components. Results for both methods clearly show that the cross-product term constructed from either a primary gage load or a primary gage output and the temperature explains the majority of the temperature-dependent part of the predicted balance load. This term models the temperature dependent nature of the gage sensitivity. Therefore, it is recommended to apply primary gage loadings at different temperatures during a balance calibration whenever temperature effects need to be described. These loadings will contain information about the temperature-dependent nature of the gage sensitivities that can be quantified by related cross-product terms in regression models of the data.

Description: Over the last 5 years, the Heatshield for Extreme Entry Environment Technology (HEEET) project has been working to mature a 3-D Woven Thermal Protection System (TPS) to Technical Readiness Level (TRL) 6 to support future NASA missions to destinations such as Venus and Saturn. A key aspect of the project has been the development of the manufacturing and integration processes/procedures necessary to build a heat shield utilizing the HEEET 3D-woven material. This has culminated in the building of a 1-meter diameter Engineering Test Unit (ETU) representative of what would be used for a Saturn probe. The present talk provides an overview of recent testing of NASA's Heatshield for Extreme Entry Environment Technology (HEEET) 3D Woven TPS. Under the current program, the ETU has been subjected to Thermal and Mechanical loads typical of deep space mission to Saturn. Thermal testing of HEEET coupons has performance up to 4,500 watts per centimeter squared at 5 atmospheres stagnation pressure and successful shear performance up to 3000 pascals at 1,650 watts per centimeter squared at 2.6 atmospheres pressure.

Description: Laser Rayleigh scattering was used to investigate clusters in the free-stream flow at Arnold Engineering Development Centers Tunnel 9 (T9). The facility was run at Mach-14, with a pure-N2 flow medium, and at several total pressures and temperatures. Using an excimer laser operating at 248 nm, the Rayleigh instrument imaged scattering from the focused laser beam in the free-stream. As a wind-tunnel flow is accelerated, it cools and approaches the condensation boundary. As a precursor to condensation, small clusters of molecules are first formed, but the individual clusters are too small to be spatially resolved in typical images of the beam. Thus clusters effectively add a spatially smooth background signal to the pure diatomic-molecule Rayleigh signal. The main result of the present work is that clustering was not significant. After correcting for interference by small particles imbedded in the T9 flow, cluster scattering was unobservable or smaller than one standard deviation (1-sigma) of the uncertainties for almost all tunnel runs. The total light scattering level was measured to be 1.05 +/- 0.15 (1-sigma) of the expected diatomic scattering, when averaged over the entire usable data set. This result included flow conditions that were supercooled to temperatures of ~ 20 K, about 25 K below the condensation limit of ~ 45 K. Thus the Mach-14 nozzle flow is essentially cluster-free for many supercooled conditions that might be used to extend the facility operating range to larger Reynolds numbers.

Description: Future urban mobility promises to deliver transformative impact across the value chain. Consequently, suppliers will face disruption in the form of new technologies, stakeholders and market dynamics. How can supply networks be optimized to meet capabilities that are yet unknown? This session will explore what business models and supply chain strategies can best deliver value for urban air transport and will address how to scale these networks at the pace of this rapidly evolving ecosystem.

Description: 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.

Description: During instrument-level or spacecraft-level ground testing, heat pipes may be placed in reflux mode, with condenser above evaporator. A liquid pool will form at the bottom of the heat pipe. If heat is applied to a site below the surface of the liquid pool in a vertical heat pipe, the heat pipe can work properly under reflux mode. A superheat is required for startup. If heat is applied to a site above the liquid pool, the heat pipe is not expected to work unless additional heat is applied to the liquid pool to provide the needed flow circulation. There are many reason to minimize the additional heater power. An experimental investigation was conducted to study the heat pipe behavior under this configuration.

Description: In this report we have catalogued the flow regimes observed in microgravity, summarized correlations for the pressure drop and rate of heat transfer that are commonly used, and discuss the validation of a few correlations from available experimental results. Two-phase flow through some specific components such as bends, tees, filters and pumps are discussed from a physical perspective to guide the designer on how reduced gravity might affect their performance. Phase separation in zero gravity is addressed through the behavior and basic design concepts for devices based on passive centrifugal action, capillary forces, gas extraction through a membrane installed in a channel wall and the use of a syringe with a perforated piston to remove bubbles from small liquid volumes. We address the common instabilities that develop in flow loops owing exclusively to the two-phase nature of the flow, e.g., Ledinegg instability and concentration waves. Finally we briefly review flow metering and gauging; two-phase flow through porous media, where pressure drop and flow regime map correlations in zero-g are a current research topic; and basic operation principles of heat pipes and capillary pumped loops.

Description: Uncertainties in early observations of potentially hazardous asteroids result in preliminary impact corridors that can stretch across large portions of the Earths surface. At this early stage of detection, the corridor width and potential for damage are typically estimated using techniques from nuclear weapons research. These estimates often employ spherical blast assumptions resulting in a constant width impact corridor. In actuality, however, the ground damage footprint of obliquely entering asteroids is generally roughly elliptical or butterfly shaped, with the major axis extending in the cross range direction and the minor axis aligned with ground-track of the meteoroid. Since actual ground footprints for oblique entries may have aspect ratios greater than two or three, the assumption of a circular blast may significantly underestimate the area of the impact swath and the at-risk population. This work develops an engineering model that can be used to quickly estimate the eccentricity of the ground footprint as a function of local impact parameters. This yields vastly improved local estimates of the corridor width and can significantly enhance the accuracy of risk analysis.

Description: ASA & Ames Introduction: Overview of NASA and Ames Research Activities, with a special focus on NASA Aeronautics activities. All materials are overview in nature and have been presented previously in open forums.

Description: NASA & Ames Introduction: Overview of NASA and Ames Research Activities, with a special focus on NASA Aeronautics activities. All materials are overview in nature and have been presented previously in open forums.

Description: Current turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing-fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow-field data in the junction region is needed. In this report, we present an overview of experimental measurements on a wing-fuselage junction model that addresses this need. The experimental measurements were performed in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The model was a full-span wing-fuselage body that was configured with truncated DLR-F6 wings, both with and without leading-edge extensions at the wing root. The model was tested at a fixed chord Reynolds number of 2.4 million, and angles-of-attack ranging from -10 degrees to +10 degrees were considered. Flow-field measurements were performed with a pair of miniature laser Doppler velocimetry (LDV) probes that were housed inside the model and attached to three-axis traverse systems. One LDV probe was used to measure the separated flow field in the trailing-edge junction region. The other LDV probe was alternately used to measure the flow field in the leading-edge region of the wing and to measure the incoming fuselage boundary layer well upstream of the leading edge. Both LDV probes provided measurements from which all three mean velocity components, all six independent components of the Reynolds-stress tensor, and all ten independent components of the velocity triple products were calculated. In addition to the flow-field measurements, static and dynamic pressures were measured at selected locations on the wings and fuselage of the model, infrared imaging was used to characterize boundary-layer transition, oil-flow visualization was used to visualize the separated flow in the leading- and trailing-edge regions of the wing, and unsteady shear stress was measured at limited locations using capacitive shear-stress sensors. Sample results from the measurement techniques employed during the test are presented and discussed.

Description: The InSight Mars Lander successfully landed on the surface on November 26, 2018. This poster will describe the methodologies and margins used in developing the aerothermal environments for design of the thermal protection systems (TPS), as well as a prediction of as-flown environments based on the best estimated trajectory. The InSight mission spacecraft design approach included the effects of radiant heat flux to the aft body from the wake for the first time on a US Mars Mission, due to overwhelming evidence in ground testing for the European ExoMars mission (2009/2010) [1] and 2010 tests in the Electric Arc Shock Tube (EAST) facility [2]. The radiant energy on an aftbody was also recently confirmed via measurement on the Schiaparelli mission [3]. In addition, the InSight mission expected to enter the Mars atmosphere during the dust storm season, so the heatshield TPS was designed to accommodate the extra recession due to the potential dust impact. This poster will compare the predicted aerothermal environments using the reconstructed best estimated trajectory to the design environments. Design Approach: The InSight spacecraft was planned to be a near-design-to-print copy of the Phoenix spacecraft. The determination of the heatshield TPS requirements was approached as if it was a new design due to the new requirement of flying through a dust storm. The baseline for aftbody was build-to-print, and all analyses focused on ensuring adequate margin. This proved to be a challenge because the Phoenix aftbody was designed to withstand only convective heating and the InSight aftbody was evaluated for both convective and radiative heating. Aerothermal environments were predicted using the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) and the Data Parallel Line Relaxation (DPLR) CFD codes, and the Nonequilibrium Radiative Transport and Spectra Program (NEQAIR) utilizing bounding design trajectories derived from Monte Carlo analyses from the Program to Optimize Simulated Trajectories II (POST2). In all cases, super-catalytic flowfields were assigned to ensure the most conservative heating results. Two trajectories were evaluated: 1) the trajectory with the maximum heat flux was utilized to determine the flowfield characteristics and the viability of the selection of TPS materials; and 2) the trajectory with the maximum heat load was used to determine the required thicknesses of the TPS materials. Evaluation of the MEDLI data [4], along with ground test data [5] led to the determination of whether or not the flow would transition from laminar to turbulent on the heatshield, which also determined the TPS sizing location for the heatshield. Aerothermal margins were added for the convective heating and developed for the radiative heating. TPS material sizing was determined with the Reaction Kinetic Ablation Program (REKAP) and the Fully Implicit Ablation and Thermal Analysis program (FIAT) using a three-branched approach to account for aerothermal, material response, and material properties uncertainties. In addition, the heatshield recession was augmented by an analysis of the effect of entry through a potential dusty atmosphere using a methodology developed in References [6] and [7]. These analyses resulted in an increase to the Phoenix heatshield TPS thickness. Reconstruction Efforts: Once the best estimated trajectory is reconstructed by the team, the LAURA/HARA (High-Temperature Aerothermo-dynamic Radiation model) and DPLR/NEQAIR code pairs will be used to predict the as-flown aerothermal conditions. In these runs, fully-catalytic flowfields will be assigned because it is a more physically accurate description of the chemistry in the flow. Once again, determination of the onset of turbulence on the heatshield will be evaluated. The as-flown aerothermal environments will then be compared to the design environments.

Description: This PowerPoint presentation will discuss Aura's current spacecraft and instrument status, highlight any performance trends and impacts to operations, identify any operational changes and express concerns or potential process improvements. Reviewed by Eric Moyer, ESMO Deputy Project Manager.

Description: The rising number of small unmanned aerial vehicles (UAVs) expected in the next decade will enable a new series of commercial, service, and military operations in low altitude airspace as well as above densely populated areas. These operations may include on-demand delivery, medical transportation services, law enforcement operations, traffic surveillance and many more. Such unprecedented scenarios create the need for robust, efficient ways to monitor the UAV state in time to guarantee safety and mitigate contingencies throughout the operations. This work proposes a generalized monitoring and prediction methodology that utilizes realtime measurements of an autonomous UAV following a series of way-points. Two different methods, based on sinusoidal acceleration profiles and high-order splines, are utilized to generate the predicted path. The monitoring approach includes dynamic trajectory re-planning in the event of unexpected detour or hovering of the UAV during flight. It can be further extended to different vehicle types, to quantify uncertainty affecting the state variables, e.g., aerodynamic and other environmental effects, and can also be implemented to prognosticate safety-critical metrics which depend on the estimated flight path and required thrust. The proposed framework is implemented on a simplified, scalable UAV modeling and control system traversing 3D trajectories. Results presented include examples of real-time predictions of the UAV trajectories during flight and a critical analysis of the proposed scenarios under uncertainty constraints.

Description: No abstract available

Description: Abstract and not the Final document is attached. Low Lunar orbit presents a unique thermal environment with high planetary and high solar IR requirements. Orion requires a phase change material heat exchanger (PCM HX) to act as a supplemental heat rejection device (SHReD) during this orbit. As a result, Orion currently uses a PCMHX to meet heat rejection demands in low lunar orbit. This PCM HX weighs 145 lbs, a significant amount of weight on the Crew Module Adaptor. To reduce this weight, a new PCM HX and phase change material is being proposed. This new PCM HX, constructed by Mezzo technologies, was originally designed as a water based PCM HX but is now be repurposed for phase change materials with transition temperatures in Orion's set points and different freeze front propagations. Mezzo's PCM HX utilizes micro tubes which greatly increase the overall heat transfer efficiency allowing for a compact design and significant weight savings. A new phase change material is also being proposed which has a higher latent heat of fusion as well as a higher density. This paper investigates the design, testing, and analysis done on the new Mezzo PCM HX as well as the corresponding phase change material.

Description: Computational ice shapes were generated on the boundary layer ingesting engine nacelle of the D8 Double Bubble aircraft. The computations were generated using LEWICE3D, a well-known CFD icing post processor. A 50-bin global drop diameter discretization was used to capture the collection efficiency due to the direct impingement of water onto the engine nacelle. These discrete results were superposed in a weighted fashion to generate six drop size distributions that span the Appendix C and O regimes. Due to the presence of upstream geometries, i.e. the fuselage nose, the trajectories of the water drops are highly complex. Since the ice shapes are significantly correlated with the collection efficiency, the upstream fuselage nose has a significant impact on the ice accretion on the engine nacelle. These complex trajectories are caused by the ballistic nature of the particles and are thus exacerbated as particle size increases. Shadowzones are generated on the engine nacelle, and due to the curvature of the nose of the aircraft the shadowzone boundary moves from lower inboard to upper outboard as particle size increases. The largest particle impinging one the engine nacelle from the 50-bin discretization was the 47 um drop diameter. As a result, the MVD greater than 40 um Appendix O conditions were characterized by extremely low collection efficiency on the engine nacelle for these direct impingement simulations.

Description: This paper evaluates a thermodynamic ice crystal icing model that has been previously presented to describe the possible mechanisms of icing within the core of a turbofan jet engine. The model functions between two distinct ice accretions based on a surface energy balance: freeze-dominated icing and melt-dominated icing. Freeze-dominated icing occurs when liquid water (from melted ice crystals) freezes and accretes on a surface along with the existing ice of the impinging water and ice mass. This freeze-dominated icing is characterized as having strong adhesion to the surface. The amount of ice accretion is partially dictated by a freeze fraction, which is the fraction of impinging liquid water that freezes. Melt-dominated icing occurs as unmelted ice on a surface accumulates. This melt-dominated icing is characterized by weakly bonded surface adhesion. The amount of ice accumulation is partially dictated by a melt fraction, which is the fraction of impinging ice crystals that melts. Experimentally observed ice growth rates suggest that only a small fraction of the impinging ice remains on the surface, implying a mass loss mechanism such as splash, runback, bounce, or erosion. The fraction of mass loss must be determined in conjunction with the fraction of freezing liquid water or fraction of melting ice on an icing surface for a given ice growth rate. This mass loss parameter, however, along with the freeze fraction and melt fraction, are the only experimental parameters that are currently not measured directly. Using icing growth rates from ice crystal icing experiments, a methodology that has been previously proposed is used to determine these unknown parameters. This work takes ice accretion data from tests conducted by the National Aeronautics and Space Administration (NASA) at the Glenn Research Center in 2018 that examined the fundamental physics of ice crystal icing. This paper continues evaluation of the thermodynamic model from a previous effort, with additions to the model that account for sub-freezing temperatures that have been observed at the leading edge of the airfoil during icing. The predicted temperatures were generally in good agreement with measured temperatures. Other key findings include the total wet-bulb temperature being a good first order indicator of whether icing is freeze-dominated (sub-freezing values) or melt-dominated (above freezing). Maximum sticking efficiency values, the fraction of impinging mass that adheres to a surface, was calculated to be about 0.2, and retained this maximum value for a range of melt ratios (0.3 to 0.65 and possibly higher), which is defined as the ratio of liquid water content to total water content. Higher air velocities reduced the maximum sticking efficiency and shifted the icing regime to higher melt ratio values. Finally, the leading edge ice accretion angle was found to be related to ice growth (lower growth rates for smaller angles) and melt ratio (smaller melt ratios resulted in smaller angles, likely due to erosion effects).

Description: The Honeywell Uncertified Research Engine (HURE), a research version of a turbofan engine that never entered production, was tested in the NASA Propulsion System Laboratory (PSL), an altitude test facility at the NASA Glenn Research Center. The PSL is a facility that is equipped with water spray bars capable of producing an ice cloud consisting of ice particles, having a controlled particle diameter and concentration in the air flow. To develop the test matrix of the HURE, numerical analysis of flow and ice particle thermodynamics was performed on the compression system of the turbofan engine to predict operating conditions that could potentially result in a risk of ice accretion due to ice crystal ingestion. The goal of the test matrix was to have ice accrete in two regions of the compression system: region one, which consists of the fan-stator through the inlet guide vane (IGV), and region two which is the first stator within the high pressure compressor. The predictive analyses were performed with the mean line compressor flow modeling code (COMDES-MELT) which includes an ice particle model. The HURE engine was tested in PSL with the ice cloud over the range of operating conditions of altitude, ambient temperature, simulated flight Mach number, and fan speed with guidance from the analytical predictions. The engine was fitted with video cameras at strategic locations within the engine compression system flow path where ice was predicted to accrete, in order to visually confirm ice accretion when it occurred. In addition, traditional compressor instrumentation such as total pressure and temperature probes, static pressure taps, and metal temperature thermocouples were installed in targeted areas where the risk of ice accretion was expected. The current research focuses on the analysis of the data that was obtained after testing the HURE engine in PSL with ice crystal ingestion. The computational method (COMDES-MELT) was enhanced by computing key parameters through the fan- stator at multiple span wise locations, in order to increase the fidelity with the current mean-line method. The Icing Wedge static wet bulb temperature thresholds were applicable for determining the risk of ice accretion in the fan-stator, which is thought to be an adiabatic region. At some operating conditions near the splitter-lip region, other sources of heat (non-adiabatic walls) were suspected to be the cause of accretion, and the Icing Wedge was not applicable to predict accretion at that location. A simple order-of-magnitude heat transfer model was implemented into the COMDES-MELT code to estimate the wall temperature minimum and maximum thresholds that support ice accretion, as observed by video confirmation. The results from this model spanned the range of wall temperatures measured on a previous engine that experienced ice accretion at certain operating conditions.

Description: Any cluster of parachute systems is subject to effects on performance due to interactions between the parachutes. One such interaction is the twisting of a riser from one parachute around that of another. Due to friction and relative motion between the risers, it is possible for the tension in the riser near the attach point to be different from the tension in the riser towards the suspension lines or canopy. This could result in system failure due to larger than expected loading. The Orion Capsule Parachute Assembly System (CPAS) designed and executed a test to quantify the amplification of the load in a parachute riser due to twist, rocking rate and angle, cluster size, and canopy load. The design of the testing approach, test matrix, and hardware are discussed along with results and findings.

Description: The swirl distortion of a StreamVane (Trademark) was investigated in the NASA Glenn Research Center W8 test facility. The StreamVane (Trademark) was designed and generated by Virginia Tech based on CFD simulations and included a center body at the aerodynamic interface plane. The swirl pattern generated by the distortion was evaluated using a dense grid of 5-hole Pitot probe measurements captured using a rotating array of probes. Good agreement was found between the design intent and the results at 38.5 kg/s mass flow. The StreamVane (Trademark) swirl results were compared to clean facility flow at 5 inlet mass flows and found to be consistent. Additionally, the axial location of the StreamVane (Trademark) relative to the measurement plane was investigated to determine the impact on downstream total pressure loss generated by the vanes. The intent of this work was to assess the viability of using a StreamVane (Trademark) to generate a Type I or Type II distortion into a Boundary Layer Ingesting propulsor to assess its aerodynamic performance and aeromechanic response.

Description: The Honeywell Uncertified Research Engine (HURE), a research version of a turbofan engine that never entered production, was tested in the NASA Propulsion System Laboratory (PSL), an altitude test facility at the NASA Glenn Research Center. The PSL is a facility that is equipped with water spray bars capable of producing an ice cloud consisting of ice particles, having a controlled particle diameter and concentration in the air flow. To develop the test matrix of the HURE, numerical analysis of flow and ice particle thermodynamics was performed on the compression system of the turbofan engine to predict operating conditions that could potentially result in a risk of ice accretion due to ice crystal ingestion. The goal of the test matrix was to have ice accrete in two regions of the compression system: region one, which consists of the fan-stator through the inlet guide vane (IGV), and region two which is the first stator within the high pressure compressor. The predictive analyses were performed with the mean line compressor flow modeling code (COMDES-MELT) which includes an ice particle model. The HURE engine was tested in PSL with the ice cloud over the range of operating conditions of altitude, ambient temperature, simulated flight Mach number, and fan speed with guidance from the analytical predictions. The engine was fitted with video cameras at strategic locations within the engine compression system flow path where ice was predicted to accrete, in order to visually confirm ice accretion when it occurred. In addition, traditional compressor instrumentation such as total pressure and temperature probes, static pressure taps, and metal temperature thermocouples were installed in targeted areas where the risk of ice accretion was expected. The current research focuses on the analysis of the data that was obtained after testing the HURE engine in PSL with ice crystal ingestion. The computational method (COMDES-MELT) was enhanced by computing key parameters through the fan-stator at multiple span wise locations, in order to increase the fidelity with the current mean-line method. The Icing Wedge static wet bulb temperature thresholds were applicable for determining the risk of ice accretion in the fan-stator, which is thought to be an adiabatic region. At some operating conditions near the splitter-lip region, other sources of heat (non-adiabatic walls) were suspected to be the cause of accretion, and the Icing Wedge was not applicable to predict accretion at that location. A simple order-of-magnitude heat transfer model was implemented into the COMDES-MELT code to estimate the wall temperature minimum and maximum thresholds that support ice accretion, as observed by video confirmation. The results from this model spanned the range of wall temperatures measured on a previous engine that experienced ice accretion at certain operating conditions.

Description: This paper presents a new method for transonic pitching airfoils based on a RANS CFD study and the Theodorsen model of an oscillating pitching flat plate. This study quantifies the deviation of the lift coefficient predictions using CFD from that obtained using the Theodorsen model, which is based on the incompressible potential flow assumption. The present method corrects this theoretical model by modulating the Theodorsen functions by coefficient functions that depend on the reduced frequency and the Mach number. It is demonstrated that the modified theoretical model predicts lift coefficient in good agreement with the CFD results in the Mach number range from incompressible (M =0.2) to transonic (M =0.755) flow for a range of reduced frequencies typical of transonic flutter. The simulations are first validated by comparing pitching NACA0012 airfoil results with experimental results at transonic flight conditions, which establishes the requirements for a grid converged unsteady transonic solution. The hysteresis loop, Cl versus , attains a grid independent solution that compares well with experiment. The present correction method will guide the development of a new state space model for the Variable Camber Continuous Trailing Edge Flap (VCCTEF) system and eventually a new transfer function that will be incorporated in a new aeroelastic framework leading to an appropriate transonic flutter model for use in the future aircraft systems in development under the NASA Advanced Air Transportation Technologies (AATT) project.

Description: This work continues the analysis of data obtained during a 2017 NASA DGEN Aeropropulsion Research Turbofan (DART) core/combustor-noise baseline test in the NASA GRC Aero-Acoustic Propulsion Laboratory (AAPL). The DART is a cost-efficient testbed for the study of core-noise physics and mitigation. Acoustic data were simultaneously acquired using the AAPL overhead microphone array in the engine aft-quadrant farfield, a single midfield microphone, and two infinite-tube-probe sensors for unsteady pressures at the core-nozzle exit. The data are here examined on an 1/3-octave basis as a first step in extending and improving core-noise prediction capability.

Description: Radiative heating computations are performed for high speed lunar return experiments conducted in the Electric Arc Shock Tube (EAST) facility at NASA Ames Research Center. The nonequilibrium radiative transport equations are solved via NASA's in-house radiation code NEQAIR using flow field input from US3D flow solver. The post-shock flow properties for the 10 km/s Earth entry conditions are computed using the stagnation line of a blunt-body and a full facility CFD (Computational Fluid Dynamics) simulation of the EAST shock tube. The shocked gas in the blunt-body flow achieves a thermochemical equilibrium away from the shock front whereas EAST flow exhibits a nonequilibrium behavior due to strong viscous dissipation of the shock by boundary layer. The full-tube flow calculations capture the influence of the boundary layer on the shocked gas state and provide a realistic fluid dynamic input for the radiative predictions. The integrated radiance behind the shock is calculated in NEQAIR for wavelength regimes from Vacuum-UltraViolet (VUV) to InfraRed (IR), which are pertinent to the emission characteristics of high enthalpy shock waves in air. These radiance profiles are validated against corresponding EAST shots. The full-tube simulations successfully predict a sharp radiance peak at the shock front which gets smeared in the test data due to the spatial resolution in the measurements. The full facility based radiance behind the shock shows a slightly better match with the test data in the VUV and Red spectral regions, as compared to that from a blunt-body based predictions. The UV radiance is very similar for both geometries and under-predicts the test behavior. The IR test data matches better with the blunt-body based predictions where the full-tube simulations show a significant over-prediction.

Description: NASA is seeking a new baseline aircraft model to assess the state-of-the-art technology for aircraft noise, emissions, and fuel/energy consumption as an update to a 2005 baseline. The process of modeling engine and airframe models as a system has historically required many iterations at NASA between the airframe and engine models. A new internal process presented in this paper contains a method that simultaneously calibrates an airframe and engine model to known data to create an aircraft system model. The work presented in this paper proposes a new framework in creating new aircraft models for future NASA research. This approach is presented as a general outline applicable to any chosen commercial aircraft. As an applied example, the B737 MAX 8 aircraft is chosen as the integrated engine and airframe model subjected to calibration. Initial results show a close match to available data but further refinement in the process is necessary for this ongoing work.

Description: The 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) at NASA Glenn Research Center was built in 1969 in the return leg of the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT). The 8x6 SWT was completed in 1949 and acoustically treated to mitigate community noise issues in 1950. This treatment included the addition of a large muffler downstream of the 8x6 SWT test section and diffuser. The 9x15 LSWT was designed for performance testing of V/STOL aircraft models, but with the addition of the current acoustic treatment in 1986 the tunnel been used principally for acoustic and performance testing of aircraft propulsion systems. The present document describes the status of the acoustic upgrade as of early 2019.

Description: High speed rotorcraft transmissions are subject to load-independent power losses consisting of drag and pumping loss. Tightly conforming shrouds enclosing the transmission gears are often incorporated to reduce the drag component of the total load-independent losses. However, tightly conforming axial shrouding can result in an increase in the pumping loss component. Quantifying the pumping loss of shrouded gear transmissions has been the subject of many studies. This study presents a new approach for estimating pumping loss based on the concept of swept volume borrowed from the positive displacement pump and compressor industry. In this study, pumping loss of shrouded gear transmissions is considered to be related to the swept volume of the gear sets and the downstream flow resistance created by the shroud clearances. The drag loss and pumping loss of a spur gear pair have been determined through testing using the NASA Glenn Research Center Gear Windage Test Facility. The results from this testing have been compared to theoretical results using the formulations presented in this study. Good correlation exist between the test pumping power loss and the predicted pumping power loss for tightly conforming axial shroud configurations.

Description: This presentation serves as an overview of test plans for an upcoming DGEN Aeropropulsion Research Turbofan (DART) test entry at the NASA GRC AeroAcoustic Propulsion Laboratory (AAPL). The test entry includes: (1)a fan intra-stage velocity field survey, which will be compared to a Computational Fluid Dynamics (CFD) survey of DART, (2) an exploratory noise study of DART with several objectives focused on measurement projection to the far-field, source identification improvements and development of a barrier wall for isolation of various sources, (3) advancement of core/combustor noise research on DART using more extensive engine-mounted instrumentation, and (4) high-temperature pressure sensor technology-readiness-level (TRL) advancement.

Description: Early in the Orion CPAS (Capsule Parachute Assembly System) project a main parachute was fabricated with lighter weight broadcloth in the lower part of the parachute skirt in order to look into different options for reducing the mass of the CPAS. At the end of Orion CPAS airdrop testing this parachute was used as a test equipment recovery parachute in order to gather data on the performance of this parachute. The parachute was the single recovery parachute in order to achieve the proper load under the parachute. It was flown on the final CPAS qualification test CQT 4-8 in September 2018.This paper will include imagery analysis, performance analysis based on all the gathered data, a full description of the configuration of the recovery parachute, as well as a comparison between this parachute and other CPAS recovery parachutes and other CPAS Main parachutes.

Description: Numerical investigations of the flowfield inside NASA Ames' Electric Arc Shock Tube have been performed. The focus is to simulate the experiments designed to reproduce shock layer radiation layer relevant to Earth re-entry conditions. This paper assess the current computational capability in simulating time-accurate unsteady nonequilibrium flows in the presence of strong shock waves with state-of-the-art physical models. The technical approach is described with preliminary results presented for one specific flow condition. It was found that the axisymmetric source term generates a numerical instability that appears as shock bending. This instability is time dependent which greatly affects the shock speed. Post-shock conditions are discussed and compared to CEA equilibrium prediction and good agreement was obtained close to the test-section and just behind the shock.

Description: The HEEET project was conceived to develop a heatshield with a high performance ablative thermal protection material that can withstand the extreme entry environment produced as a result of rapid deceleration during high speed entry into Venus, Saturn, Uranus or higher speed entry into Earth's atmosphere. Successful maturation of HEEET supports future New Frontiers and Discovery AO's, as well as Flagship and directed missions in the longer term. In addition, HEEET has the potential to evolve and to support re-entry to Earth, for missions such as Mars Sample Return.The primary goal of the HEEET Project was to develop an ablative TPS heat-shield based on woven TPS technology to Technology Readiness Level (TRL) 6. Key evidence to support the TRL evaluation includes: Demonstration of reproducible manufacturing of a dual layer material over a range of thicknesses and integrated on to a heatshield engineering test unit at a scale that is applicable to near term Discovery as the highest priority and future NF missions as secondary priority set of missions. Demonstration of predictable and stable performance of the dual layer TPS over a range of entry environments that are applicable to near term Discovery and NF missions of interest to SMD.Includes completion of coupon arc jet and laser testing and development of a mid-fidelity thermal response model that correlates with test results. Demonstration of flight heatshield system design for a range of sizes and loads that are relevant to near term Discovery and NF missions of interest to SMD. Includes completion of structural testing to validate analytic thermal/structural models and development of a material property database. Includes structural testing of a ~1m Engineering Test Unit under relevant entry loads.

Description: This paper reports computational analyses and flow characterization studies in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using a wedge model placed in a free jet downstream of new 9-inch diameter conical nozzle in the Ames 60-MW Interaction Heating Facility. Both the nozzle and wedge model were specifically designed for testing in the new Laser-Enhanced Arc-jet Facility. Data were obtained using stagnation calorimeters and wedge models placed downstream of the nozzle exit. Two instrumented wedge calibration plates were used: one water-cooled and the other RCG-coated tile plate. Experimental surveys of arc-jet test flow with pitot and heat flux probes were also performed at three arc-heater conditions, providing assessment of the flow uniformity and valuable data for the flow characterization. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, including the models tested, and comparisons with the experimental measurements. By taking into account nonuniform total enthalpy and mass flux profiles at the nozzle inlet as well as the expansion waves emanating from the nozzle exit and their effects on the model flowfields, these simulations approximately reproduce the probe survey data and predict the wedge model surface pressure and heat flux measurements.

Description: Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional (3D) icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind tunnel testing was carried out for a 13.3-percent-scale semispan wing based upon the Common Research Model airplane configuration. The wind tunnel testing was conducted at the Office National dEtudes et de Recherches Arospatiales (ONERA) F1 pressurized wind tunnel with Reynolds numbers of 1.6 x 10(exp 6) to 11.9 x 10(exp 6 ) and Mach numbers of 0.09 to 0.34. Five different configurations were investigated using fully 3D, high-fidelity artificial ice shapes that maintain nearly all of the 3D ice-accretion features documented in prior icing wind tunnel tests. These large, leading-edge ice shapes were nominally based upon airplane holding in icing conditions scenarios. For three of these configurations, lower fidelity simulations were also built and tested. The results presented in this paper show that while Reynolds and Mach number effects are important for quantifying the clean-wing performance, there is very little to no effect for an iced wing with 3D, high-fidelity artificial ice shapes or 3D smooth ice shapes with grit roughness. These conclusions are consistent with the large volume of past research on iced airfoils. However, some differences were also noted for the associated stalling angle of the iced swept wing and for various lower fidelity versions of the leading-edge ice accretion. More research is planned to further investigate the key features of ice-accretion geometry that must be simulated in lower fidelity versions in order to capture the essential aerodynamics.

Description: In this paper, we introduce a level set method topology optimization method of structures subjected to coupled mechanical and thermal loads. Different examples considering compliance minimization and stress minimization under temperature and volume constraints, and mass minimization under stress and temperature constraints, are presented. The p-norm of the stress field and temperature field is used to approximate the maximum stress and temperature, respectively. The developed method is applied in the design of an L-bracket and a battery package. The results show that designs obtained by ignoring the thermal or structural constraints can result in high values of temperature or stress, respectively.

Description: Urban air taxis, also known as urban air mobility (UAM) vehicles, are anticipated to be an area of significant market growth in the near future. These vehicles are typically vertical take-off and landing (VTOL) designs which are capable of carrying 1 to 30 passengers in an intra-urban environment with flights of less than 50 nautical miles. Development of UAM vehicles and their integration into the airspace will be enabled by advancements in a number of areas including electrified propulsion systems, structures, acoustics, automation, and controls. However, the strong multidisciplinary interactions for these unique vehicles presents a significant new design challenge. This work describes the development of a multidisciplinary analysis and optimization environment which can be used to support the conceptual design of these UAM vehicles, using efficient gradient based optimization with analytic derivatives. The tools included in this multidisciplinary analysis model the aircraft trajectory, vehicle aerodynamics, structures, and electrified propulsion system. The multidisciplinary environment created in this research is unique in that all the physics tools are tightly integrated together, with the trajectory model directly calling the aerodynamics, structures, and propulsion models. This multidisciplinary analysis environment is then demonstrated in the design optimization of a turboelectric tiltwing UAM vehicle concept.

Description: An ice shape database has been created to document ice accretions on a 21-inch chord NACA0012 model and a 72-inch chord NACA 23012 airfoil model resulting from an exposure to a Supercooled Large Drop (SLD) icing cloud with a bimodal drop size distribution. The ice shapes created were documented with photographs, laser scanned surface measurements over a section of the model span, and measurement of the ice mass over the same section of each accretion. The icing conditions used in the test matrix were based upon previously used conditions on the same models but with an alternate approach to evaluation of drop distribution effects. Ice shapes resulting from the bimodal distribution as well as from equivalent monomodal drop size distributions were obtained and compared. Results indicate that the ice shapes resulting from the monomodal and bimodal drop size distributions had similar shapes, but the bimodal distributions had greater mass and volume measurements and icing limits that extended further back on the chord of the model.

Description: As both NASA and the aeronautics industry recognize the need for higher fuel efficiency and lower carbon emissions in both commercial airline and private aviation applications, development of all-electric or hybrid electric aircraft have garnered renewed interest in the aviation community. For the particular example of the hybrid-electric option, the solid oxide fuel cell (SOFC) is an attractive option for the power source, due to its potential to utilize aviation fuels thereby having minimal impact to aviation infrastructure. SOFC stack performance depends upon many factors, one of the most important is the way the oxidant and fuel gases are delivered to the fuel cells. System modeling of various aircraft configurations for FUELEAP (Fostering Ultra-Efficient, Low-Emitting Aviation Power) point to the need to operate SOFC stacks at high current densities. This creates challenges in the thermal profile of the stacks with potential to create large thermal gradients and hot spots. This study investigates two types of commercial solid oxide fuel cell stacks, the cross flow and co-flow gas designs, both convectively cooled with cathode air. High fuel utilization factors were also employed under varying electrical loads expected from the demands of flight. In addition, performance, range of operation and endurance were investigated under conditions of high current loads and thermal cycling. Evaluations include the study of gas kinetic using electrochemical spectroscopy. Testing took place at the facilities of NASA Glenn using a commercial test system (FuelCon AG, Magdeburg Germany). These studies are crucial to the Glenn Research Center's ability to conduct research, evaluation and development of the next-generation SOFC based stacks for cutting-edge energy technologies for aerospace applications. This study supports NASA's Convergent Aeronautics Solutions' (CAS) FUELEAP project.

Description: Understanding the aerodynamic impact of swept-wing ice accretions is a crucial component of the design of modern aircraft. Computer-simulation tools are commonly used to approximate ice shapes, so the necessary level of detail or fidelity of those simulated ice shapes must be understood relative to high-fidelity representations of the ice. Previous tests were performed in the NASA Icing Research Tunnel to acquire high-fidelity ice shapes. From this database, full-span artificial ice shapes were designed and manufactured for both an 8.9%-scale and 13.3%-scale semispan wing model of the CRM65 which has been established as the full-scale baseline for this swept-wing project. These models were tested in the Walter H. Beech wind tunnel at Wichita State University and at the ONERA F1 facility, respectively. The data collected in the Wichita St. University wind tunnel provided a low-Reynolds number baseline study while the pressurized F1 facility produced data over a wide range of Reynolds and Mach numbers with the highest Reynolds number studied being approximately Re=11.9x106. Past work focused on only three different fidelity variations for ice shapes based on multiple icing conditions. This work presents a more detailed investigation into several fidelity representations of a single highly three-dimensional scallop ice accretion. Sensitivity to roughness size and application technique on a low-fidelity smooth ice shape is described. The data indicate that the aerodynamic performance is not especially sensitive to the grit variations. An ice accretion code was also used to generate ice shapes for aerodynamic testing and comparisons. These ice shapes have a general appearance like the low-fidelity smooth ice shapes, but in this case, the computer-generated ice shape is significantly smaller. As such, the impact of that ice shape on the aerodynamic performance of the wing is reduced compared to the smooth ice shape based on the icing experiment for those same conditions. Spanwise discontinuities were also introduced to a low-fidelity ice shape in an attempt to quantify the impact of those variation in the high-fidelity ice shape. While the lift data indicate good agreement between the high-fidelity ice shapes and the low-fidelity ice shapes and the low-fidelity ice shapes with spanwise discontinuities, a closer investigation of the data suggests potential, significant differences in the flowfield. These results were similar at both facilities over the wide range of test conditions utilized.

Description: Recently, heat transfer correlations based on liquid nitrogen (LN2) and liquid hydrogen (LH2) pipe quenching data were developed to improve the predictive accuracy of lumped node codes like SINDA/FLUINT and the Generalized Fluid System Simulation Program (GFSSP). After implementing these correlations into both programs, updated model runs showed strong improvement in LN2 pipe chilldown modeling but only modest improvement in LH2 modeling. Due to large differences in thermal and fluid properties between the two fluids, results indicated a need to develop a separate set of LH2-only correlations to improve the accuracy of the simulations. This paper presents a new set of two-phase convection heat transfer correlations based on LH2 pipe quenching data. A correlation to predict the bulk vapor temperature was developed after analysis showed that high amounts of thermal nonequilibrium of the liquid and vapor phases occurred during film boiling of LH2. Implemented in a numerical model, the new correlations achieve a mean absolute error of 19.5 K in the predicted wall temperature when compared to recent LH2 pipe chilldown data, an improvement of 40% over recent GFSSP predictions. This correlation set can be implemented in simulations of the transient LH2 chilldown process. Such simulations are useful for predicting the chilldown time and boil-off mass of LH2 for applications such as the transfer of LH2 from a ground storage tank to the rocket vehicle propellant tank, or through a rocket engine feedline during engine startup.

Description: Film cooling is used in a wide variety of engineering applications for protection of surfaces from hot or combusting gases. The design of more efficient film cooling geometries/configurations could be facilitated by an ability to accurately model and predict the effectiveness of current designs using computational fluid dynamics (CFD) code predictions. Hence, a benchmark set of flow field property data were obtained for use in assessing current CFD capabilities and for development of better modeling approaches for these turbulent flow fields where accurate calculation of turbulent heat flux is important. Both Particle Image Velocimetry (PIV) and spontaneous rotational Raman scattering (SRS) spectroscopy were used to acquire high quality, spatially-resolved measurements of the mean velocity, turbulence intensity as well as the mean temperature and root mean square (rms) temperatures in a film cooling flow field. In addition to off-body flow field measurements, infrared thermography (IR) and thermocouple measurements on the plate surface enabled estimates of the film effectiveness. Raman spectra in air were obtained across a matrix of axial locations downstream from a 68.07 mm square nozzle blowing heated air over a range of temperatures (up to TR = 2.7) and Mach numbers (up to M0.9), across a 30.48 cm long plate equipped with three patches of 45 small (~1 mm) diameter cooling holes arranged in a staggered configuration. In addition, both centerline streamwise 2-component PIV and cross-stream 3-component Stereo PIV data at 14 axial stations were collected in the same flows. Only a subset of the data collected in the test program is included in this Part I report and are available from the NASA STI office. The final portion of the data will be published in a future report, Part II, along with CFD predictions of the complex cooling film flow.

Description: This summer internship is focused on using CFD and fluid mechanics to optimize the SRL-ADEPT geometry in an attempt to increase drag and area-effectiveness, and reduce flow separation.

Description: ESA recently flew an entry, descent, and landing demonstrator module called Schiaparelli that entered the atmosphere of Mars on the 19th of October, 2016. The instrumentation suite included heatshield and backshell pressure transducers and thermocouples (known as AMELIA) and backshell radiation and direct heatflux-sensing sensors (known as COMARS and ICOTOM). Due to the failed landing of Schiaparelli, only a subset of the flight data was transmitted before and after plasma black-out. The goal of this paper is to present comparisons of the flight data with calculations from NASA simulation tools, DPLR/NEQAIR and LAURA/HARA. DPLR and LAURA are used to calculate the flowfield around the vehicle and surface properties, such as pressure and convective heating. The flowfield data are passed to NEQAIR and HARA to calculate the radiative heat flux. Comparisons will be made to the COMARS total heat flux, radiative heat flux and pressure measurements. Results will also be shown against the reconstructed heat flux which was calculated from an inverse analysis of the AMELIA thermocouple data performed by Astrium. Preliminary calculations are presented in this abstract. The aerodynamics of the vehicle and certain as yet unexplained features of the inverse analysis and forebody data will be investigated.

Description: Overview of Workshop terms and goals for Enabling Autonomous Flight and Operations in the National Airspace.

Description: No abstract available

Description: NASA's Descent System Studies (DSS) Program is studying various concept vehicles to enable landing of heavy payloads on the surface of Mars. While it is desirable to run high-fidelity CFD simulations to accurately assess the aerodynamic and aerothermal effects of various design changes during EDL, it is usually difficult to quickly generate high-quality grids suitable for such analyses. One approach to address this bottleneck in mesh generation is through the use oversetting grids. Although the overset approach is efficient and powerful in solving partial differential equations on complex geometries, new users often find it challenging to apply overset concepts for their simulations. For example, generating hyperbolic grids with sufficient overlap; priority in hole-cutting on multiple overlapping grids; and fixes to assemble overlapping viscous grids at the body surface. The objective of this presentation is to introduce a simple process that combines the advantages of near-body, point-matched, structured grids with oversetting background grids suitable for grid alignment. This approach allows for grids that can be sequenced, reclustering of mesh spacing at the wall, and grid alignment with the bow shock. The current methodology is tested on a Mid-L/D configuration using the overset DPLR code.

Description: This study presents an onboard decision-making architecture for small unmanned aerial systems (sUAS). The decision-maker is part of NASA's SAFE50 project that is working under the UAS Traffic Management (UTM) Technical Capability Level (TCL) 4 to provide autonomous point-to-point UAV flight in BVLOS, high-density urban environments. The decision-maker monitors various metrics to determine the safety and feasibility of the mission and categorizes flight states as Nominal, Off-Nominal, Alternate Land, and Land Now in a finite state machine. Changes in the monitored metrics serve as transitions in the state machine and trigger replanning. Navigation degradation and communication failure are simulated to show the feasibility of the decision-maker framework in appropriately switching the flight state.

Description: Electronics Boxes with high heat dissipations use a thermal interface material to increase heat transfer to the radiator in a vacuum/space environment. There are lots of materials to choose from, but for Spacecraft applications, there are more than high heat transfer metrics which must be met. Contamination (both particle generation and outgassing), ease of cutting, and removal are just as important metrics in material selection. However, vendor data of material thermal conductance is usually based on a 1" X 1" piece of material under high uniform pressures. Large Electronics boxes almost never have optimal pressures, as they are bolted along the perimeter and leave gaps in the center regions. In order to characterize the relative thermal conductance for large Electronics boxes, an 8" X 8" plate was fabricated to simulate an electronics box bottom and bolted around the perimeter to a cold plate. Various thermal interface materials were inserted between the box and cold plate, and overall thermal conductance's were calculated. A table was generated which compares the full gamut of thermal interface materials for large boxes, from a dry joint to a wet joint. Materials were placed in order of high to low conductance's, so an engineer can compare the benefit of each material in a real-world scenario.

Description: The intermediate wake region of a thick flat plate with a circular trailing edge (TE) is investigated with a direct numerical simulation (DNS). The upper and lower separating boundary layers are both turbulent and are statistically identical; the resulting wake is symmetric in the mean. Earlier research dealt with the near/very-near wake of the same plate (x/D 〈 13.0, x is the streamwise distance from the center of the circular TE and D is the plate-thickness/TE-diameter). In the present investigation the emphasis is on the evolution of shed-vortex structure and turbulence intensity distributions with increasing x; the focus is on the region 20.0 〈 x/D 〈 40.0. Profile similarity in wake velocity statistics is explored.

Description: An alternative process technique, namely vacuum-assisted axial injection potting (VaAIP), has been developed to pot the Litz wires in the stator winding of high power density electric motors for the future electrified aircrafts. Initial trials of the process showed significant improvement in potting quality with less voids, thus potential improvement in thermal management of the motors. As an initial effort of pot-ability assessment, microstructures, 2-D and 3-D, of the Litz wires including dimensions and distribution of conductor filament, coating, and open spaces; packing patterns; shape/configuration changes of each bundles or the overall cross-sections per degree of twist were determined and quantified successfully. The microstructure analyses were performed not only for effective potting process development but also for more realistic electro-thermal modeling solutions. This paper will present results of the microstructure analyses, potentials of the VaAIP process from the trials, and future plans for scale-up and implementation of the process into a full-scale prototype stator winding.

Description: Ammonia is used in the Starboard 1 (S1) and Port 1 (P1) External Active Thermal Control System (EATCS) to cool the pressurized modules, and some of the external electrical power distribution hardware. Leaks that develop in these critical cooling systems that deplete in-line tanks can ultimately result in loss of cooling, which can have devastating impacts to the mission, science and crew onboard the ISS. A slow ammonia leak was initially observed from the P1 EATCS in 2011, but later in 2013 the leak rate began to accelerate. The ammonia inventory eventually began to decay exponentially, raising concerns that the inventory could drop to levels where the system would not be operational.The Robotic External Leak Locator (RELL) was built and launched to the ISS to detect and help locate ammonia leaks using the ISS Robotic Arm and remote ground operator control without constant crew involvement. RELL pinpointed the ammonia leak to the two flexible jumper hose assemblies connecting one of two fluid loops in one of the three deployable radiators to the P1 EATCS. The ammonia inside the two hose assemblies and that radiator fluid loop was isolated and vented to space in 2017. This stopped the leak and an Extravehicular Activity was conducted to remove the two hose assemblies so they could be returned to ground for further Test, Teardown and Evaluation (TT&E). The purpose of this presentation is to discuss this leakage scenario and the TT&E efforts.

Description: This presentation summarizes the current plans and efforts at NASA/Goddard to develop new thermal control technology for anticipated future missions. It will also address some of the programmatic developments currently underway at NASA, especially with respect to the NASA Technology Development Program and will highlight some of the latest flight project activities at GSFC. While funding for basic technology development is still scarce, significant efforts are being made in direct support of flight programs. New technology development continues to be driven by the needs of future missions, and applications of these technologies to current Goddard programs will be addressed. Many of these technologies also have broad applicability to DOD, DOE, and commercial programs. Partnerships have been developed with the Air Force, Navy, and various universities to promote technology development. Technology development activities supported by the internal research and development (IRAD) program, the Small Business Innovative Research (SBIR) program, and other technology development programs are reviewed in this presentation. Specific technologies addressed include; micro-channel heat transfer, latest developments of electro-hydro-dynamically pumped systems, Atomic Layer Deposition (ALD), thermal control coatings, and various other research activities.

Description: Normally, in order to characterize multilayer insulation installed onto a test tank, the boil-off of the tank is measured and then heat loads from structural and fluid penetrations are calculated from temperature measurements throughout the system. For the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig testing, it was determined that this approach would have significant uncertainties (over 50%) and that another method was needed to characterize the heat load through the blanket. Heat flux sensors are widely used to measure heat loads and characterize insulation systems at room temperature, however, the heat fluxes measured are usually two orders of magnitude higher than high performance MLI. Three different heat flux sensors were initially checked out on a liquid hydrogen calorimeter. One was chosen for actual implementation and 20 sensors were ordered. Of those sensors, calibration was attempted on 7 of the sensors. The results from testing and calibration are discussed.

Description: The two decades old high order central differencing via entropy splitting and summation-by-parts (SBP) difference boundary closure of Ols- son & Oliger (1994), Gerritsen & Olsson (1996), and Yee et al. (2000) is revisited. The entropy splitting is a form of skew-symmetric splitting of the nonlinear Euler flux derivatives. Central differencing applied to the entropy splitting form of the Euler flux derivatives together with SBP difference operators will, hereafter, be referred to as entropy split schemes. This study is prompted by the recent growing interest in numerical methods for which a discrete entropy conservation law holds, a discrete global entropy conservation can be proved and/or the numerical method possesses a stable entropy in the framework of SBP difference operators and L2-energy norm estimate. The objective of this paper is to recast the entropy split scheme as the re- cent definition of an entropy stable method for central differencing with SBP operators for both periodic and non-periodic boundary conditions for non- linear Euler equations. Standard high order spatial central differencing as well as high order central spatial DRP (dispersion relation preserving) spatial differencing is part of the entropy stable methodology framework. Long time integration of 2D and 3D test cases is included to show the comparison of this efficient entropy stable method with the Tadmor-type of entropy conservative methods. Studies also include the comparison among the three skew-symmetric splittings on their nonlinear stability and accuracy performance without added numerical dissipations for smooth flows. These are, namely, entropy splitting, Ducros et al. splitting and the Kennedy & Grub- ber splitting.

Description: The talk will cover various research and development challenges and opportunities related to enabling autonomous flight and airspace operations. Particularly, it will address the needs and importance for enabling autonomous operations, various technical challenges and opportunities, and minimum viable product strategy.

Description: No abstract available

Description: This paper presents numerical models of boiling in a heated tube using the Generalized Fluid System Simulation Program (GFSSP), a finite-volume-based general-purpose flow network code developed at NASA/Marshall Space Flight Center. The heated tube is discretized into a one-dimensional array of nodes and branches to represent the flow of liquid and vapor in a tube with a prescribed pressure differential. The solid wall is also discretized into solid nodes and conductors to allow for heat transfer between the wall and the fluid. The conservation equations of mass, momentum, and energy of the fluid are solved simultaneously with the energy conservation equation for the solid wall. Two experimental configurations of fluid flowing in a vertical tube have been simulated, one with water and the other with liquid hydrogen. This paper compares experimental data with numerical predictions based on four different published correlations for boiling heat transfer coefficients. Three of these correlations are applicable to the saturated vertical flow conditions of the experiments. One of them is applicable to film boiling and has been used for the liquid hydrogen experiment, which was in film boiling regime. For the case of boiling water, the predictions of wall temperatures using the boiling heat transfer correlations agreed well with the experimental results. However, in the case of boiling hydrogen larger discrepancies were observed between the experimental data and numerical predictions.

Description: No abstract available

Description: These slides present NASA's vision, historical overview , and current Unmanned Systems projects at AFRC.

Description: Additive manufacturing methods for producing single materials are rapidly improving. The resulting material properties and microstructures are becoming more comparable to those of conventionally fabricated materials. However, the need for multi-functional and complex structures and components requires additional innovations in manufacturing such as multi-material and hybrid additive manufacturing approaches. Additive manufacturing machines with multiple print capabilities and combinations of AM, machining, and conventional processing methods will further open up design spaces and possibilities. In this presentation, several examples of the needs and methods for multi-material fabrication will be discussed with a focus on aerospace applications. Direct printing of silver coils in conjunction with fused deposition modeling, machined parts, and, binder jetting is being developed for innovative stator designs. Binder jetting of silicon-based materials with powder bed additions is being developed for heat exchanger applications. Additive manufacturing of bi-material systems is being pursued to fabricate lightweight, integrated, multifunctional structures.

Description: NASA is conducting research under the UAS Integration in the NAS Project to develop standards that will enable mid-size and large unmanned aircraft to fly unrestricted in the National Airspace System. As these efforts move into its second phase, NASA is planning a series of flight tests and demonstrations, integrating industry partners' technologies. These events will not only provide valuable data to inform the RTCA Special Committee 228 DAA and C2 MOPS, but also provide an opportunity for the UAS community to test their technologies in a realistic environment. An overview of NASA UAS-NAS research will be presented touching on human systems integration, modeling and simulation and guidance and control. Plans for Flight Test 6 and Systems Integration Operationalization (SIO) will also be presented. The purpose of this meeting is to share with Kitty Hawk, at a high level, UAS-NAS research and discuss potential future collaboration between NASA and Kitty Hawk.

Description: NASA has been exploring the use of artificial intelligence technologies to improve vehicle and airspace capabilities for over 20 years. This research has ranged from the use of neural nets to allow adaptive control, to autonomous rotorcraft operations, to real-time prognostics, to an eventual goal of autonomous vehicles operating in an autonomous airspace, in the context of smart communities. Examples of past research and future directions in autonomy for aeronautics will be presented.

Description: With the rise in big data and analytics, machine learning is transforming many industries. It is being increasingly employed to solve a wide range of complex problems, producing autonomous systems that support human decision-making. For the aircraft engine industry, machine learning of historical and existing engine data could provide insights that help drive for better engine design. This work explored the application of machine learning to engine preliminary design. Engine core-size prediction was chosen for the first study because of its relative simplicity in terms of number of input variables required (only three). Specifically, machine-learning predictive tools were developed for turbofan engine core-size prediction, using publicly available data of two hundred manufactured engines and engines that were studied previously in NASA aeronautics projects. The prediction results of these models show that, by bringing together big data, robust machine-learning algorithms and automation, a machine learning-based predictive model can be an effective tool for turbofan engine core-size prediction. The promising results of this first study paves the way for further exploration of the use of machine learning for aircraft engine preliminary design.

Description: No abstract available

Description: Throughout the world, especially in dense urban environments, the quality of life is being negatively impacted by ever growing commute time. Travel, beyond commuting, is increasingly driven by door-to-door challenges ? not just gate-to-gate considerations. Air Mobility may be an approach to address these challenges, as it can effectively convert our 2D mobility system to a 3D mobility system, vastly increasing mobility options.

Description: Introduce mobility challenge, how air mobility can address the mobility challenge, then how an Air Mobility Data & Reasoning Fabric can enable the envisioned future air mobility. Poses the question of what is an effective role for NASA in enabling an Air Mobility Data & Reasoning Fabric.

Description: Overview NASA Aeronautics activities, and how NASA Ames Research Center contributes to NASA's Aeronautics activities.

Description: Free-Flight CFD capability has been implemented into the finite-volume solver US3D under the Entry Systems Modeling project. Several simulations of ballistic range experiments have been performed in order to validate the simulation software and methodology. Extension of the software to flight scale trajectories with varying freestream conditions has been carried out. Results show promising ability to predict vehicle behavior as compared to flight. Finally, a multi-body free-flight capability has been developed to generalize the single-body free-flight solver to study multiple bodies in proximal flight.

Description: No abstract available

Description: The durability of gas-turbine engine components can be significantly affected by the ingestion of siliceous particles, which can melt at high temperature and corrode protective coatings that are essential for long life requirements. The silicate debris consists mainly of CaO-MgO-Al2O3-SiO2 (CMAS) and is usually ingested by aircraft engines during and after take-off, sticking to their hot surfaces and resulting in the formation of calcium rare-earth silicate oxyapatites. The thermochemistry of coatings and their reaction products with molten silicate debris are crucial to understand in order to improve the durability of gas-turbine engines. Here we discuss results of high temperature drop solution calorimetry, drop-and-catch calorimetry (DnC) and differential thermal analysis (DTA) techniques for the thermodynamic properties of both thermal barrier coatings (TBCs) and environmental barrier coatings (EBCs) and their reaction with CMAS compositions. The enthalpies of solution of Y2Si2O7, Yb2Si2O7, 31YSZ, and 16RESZ based coatings and the oxyapatite are moderately positive. However, oxyapatite formation is only favorable over coating dissolution in terms of enthalpy for 7YSZ. The enthalpies of mixing between the coatings and the molten silicate are less exothermic for Yb2Si2O7 and CaYb4Si3O13 than for 7YSZ, indicating lower energetic stability of the latter against molten silicate corrosion. We also report for the first time the calorimetric measurements of the enthalpies of formation of rare-earth silicate based EBC coatings and oxyapatites (rare-earth, RE = Y, Yb, Gd, Dy, Er, Nd and Sm).

Description: Despite the introduction of flight, duty, and rest time regulations to reduce the risk of sleepiness, airline pilots often encounter elevated sleepiness during flight. To combat this sleepiness, in some instances, pilots can take a short nap on the flight deck (controlled rest) to improve their alertness. Little is known, however, as to when and how often this countermeasure is used operationally. Methods: Forty-four pilots from a European carrier wore actiwatches and filled in an electronic sleep and work diary for approximately 2 weeks resulting in data from 239 flights. Self-reported in-flight rest periods were used to set rest intervals and sleep was estimated within these intervals using Philips Actiware 6.0.9. Wake threshold selection was set to medium; sleep threshold detection algorithm was set to 10 immobile minutes at sleep onset and sleep end. Timing of sleep periods was analyzed relative to home base time. Results: Preliminary analyses showed that controlled rest was taken on 46% (n=110) of flights. On 23 flights (10%) pilots reported taking two controlled rest periods. Sleep, as estimated by actigraphy, was achieved during 80% (n=106) of controlled rest periods. The mean sleep duration was 32 ( 12) minutes estimated within successful controlled rest periods. Approximately two-thirds (67.5%, n=81) of all rest periods were initiated during home base time night (0000h-0800h). On 11% (n=26) of flights, pilots also reported taking bunk rest (longer rest period in a designated sleeping facility).Conclusion:This study shows that controlled rest is commonly used as a countermeasure to sleepiness on the flight deck. Further analysis is required to determine what other factors contribute to the decision to take controlled rest, and how effective it is in reducing sleepiness on the flight deck.

Description: No abstract available

Description: This paper explores the novel Strayton engine concept. This engine combines the cycles of a Brayton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation). The purpose of this paper is to detail the system level modeling techniques used to generate the simulation, perform a cycle analysis of the combined cycle engine, identify key technologies and challenges associated with the concept, and compare potential performance gains with existing gas turbine engines and internal combustion engines. Topics such as controls, blade cooling effects, engine weight, and heat transfer using heat pipe are also explored. Results from this work show potential architectures that could provide the required heat transfer rates, potential control strategies, and performance benefits, including efficiency gains between 10% and 3% on engines ranging from 200HP to 670HP with the combined cycle engine.

Description: NASAs Flight Imagery Launch Monitoring Real-time System (FILMRS) cameras were originally developed for the Space Launch System (SLS) Core Stage. These Commercial Off the Shelf (COTS) cameras have been redesigned and reduced by an order of magnitude in size for the Exploration Upper Stage (EUS). The change in thermal environment has led to the application of various passive thermal control methods and the addition of a heater option. This paper will give a summary of the design and development test effort associated with adapting the COTS camera for the demands of the space environment and associated thermal mitigations applied as the project prepares to complete the design. The application of this camera for other space systems is discussed.

Description: Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient (C(sub )) input. The results indicated that a collective C(sub ) of about 1% could increase the side force by 3050%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The C(sub ) is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.

Description: Plant Water Management is a technology demonstration of recent advances in micro-g capillary fluidics research applied to plant growth systems. It has applications in long-term food production systems for missions to the Moon and Mars, as well as the immediate need for ISS food supplements to the crew diet. PWM will demonstrate the low-gravity role of surface tension, wetting, and system geometry to effectively replace the role of gravity in certain terrestrial plant growth systems.

Description: Numerical investigations of the ow field inside NASA Ames' Electric Arc Shock Tube have been performed. The focus is to simulate the experiments designed to reproduce shock layer radiation layer relevant to Earth re-entry conditions. This paper assess the current computational capability in simulating unsteady nonequilibrium flows in the presence of strong shock waves with state-of-the-art physical models. The technical approach is described with preliminary results presented for one specific ow condition. The numerical problems encountered during the computation of these flows are detailed, along with the methods used to resolve them. Post-shock conditions are discussed and compared to CEA equilibrium prediction.

Description: In order to improve the cryogenic propellant management technologies for a liquid hydrogen rocket with high specific impulse, JAXA, the University of Tokyo, and the NASA Glenn Research Center have jointly organized a multi-agency model validation collaboration project. As part of this project, JAXA's boiling simulation was validated with NASA's experimental data on vertical pipeline chill-down. Simulation results were in good agreement with the experimental data obtained using an improved boiling model to reproduce the spray flow. This activity achieved liquid hydrogen turbo-pump simulation at JAXA for grasping the boiling flow phenomenon from engine cut-off to re-ignition. This joint research resulted in an international cooperative relationship for discussing the cryogenic propellant management technologies necessary to develop next-generation liquid rockets.

Description: "Heat pipes are being used on many spacecraft to acquire heat dissipated by the payload and transport the heat to a remote radiator. In instrument-level or spacecraft-level ground testing, many heat pipes are placed in a gravity-driven reflux mode where the condenser is well above the evaporator, resulting in the formation of a liquid pool at the bottom of the heat pipe. If a head load is applied to a site that is in contact with the liquid pool, the generated vapor will flow upward to the condenser and the condensate will fall back to the evaporator due the influence of gravity. Hence, the heat pipe can operate steadily under reflux mode because the heated site always has sufficient liquid supply to sustain the fluid flow. In contrast, when a heat load is applied to a site remote from the liquid pool, the heat pipe will be unable to transfer heat through liquid evaporation unless the heated site has a chance to be in contact with liquid. This can be accomplished by applying an additional heat load to the liquid pool to establish a reflux flow so that the remote site can capture the falling condensate. An experimental investigation was conducted to study the effect of gravity on the thermal performance of a heat pipe under reflux mode with multiple heat loads. An aluminum ammonia heat pipe with internal axial grooves was placed in a vertical position. Cooling was provided to the top of the heat pipe, and heat was applied to three sites below the condenser with various heat distributions. One of the heated sites was above the liquid pool, and two were in direct contact with the liquid pool. Test results showed that when a heat load was applied to either one or both of the lower sites, the heat pipe could run steadily under reflux mode. After a reflux flow had been established, a heat load could be applied to the upper site. If the upper site could capture sufficient liquid falling from the condenser to handle its heat load solely by liquid evaporation, the heat pipe could reach steady operation. Otherwise, the temperature of the upper site would oscillate due to its intermittent contact with the falling liquid. For a given heat load to the upper site, the amplitude of temperature oscillation decreased with an increasing heat load to the lower sites because there was more falling condensate available for the upper site to capture. Moreover, the temperature oscillation disappeared completely when the total heat loads to lower sites exceeded a threshold power, and the threshold power increased with an increasing heat load to the upper site."

Description: NEQAIR v15.0 provides the first steps to improved coupling between NEQAIR and the DPLR CFD code, which will be fully realized in v15.1. The plan is to release NEQAIR v15.1 and DPLR 4.05 at the same time. The improvements implemented in NEQAIR v15.0 have focused on improving stability, solution robustness, usability and providing different options for running the code. It is also the first version of the code to have a new input file and line of sight format since 2009. Backward compatibility with previous formats of the input files (neqair.inp and LOS.dat) has also been provided. NEQAIR v15.0 supersedes the prerelease of this version, as well as NEQAIR v14.0, v13.2, v13.1 and the suite of NEQAIR2009 versions. These updates have predominantly been performed by Brett Cruden and Aaron Brandis from AMA Inc at NASA Ames Research Center between 2016 and 2018. NEQAIR v15.0 is a standalone software tool for line-by-line spectral computation of radiative intensities and/or radiative heat flux, with one-dimensional transport of radiation. In order to accomplish this, NEQAIR v15.0, as in previous versions, requires the specification of distances (in cm), temperatures (in K) and number densities (in parts/cc) of constituent species along lines of sight. Therefore, it is assumed that flow quantities have been extracted from flow fields computed using other tools, such as CFD codes like DPLR or LAURA, and that lines of sight have been constructed and written out in the format required by NEQAIR v15.0. There are two principal modes for running NEQAIR v15.0. In the first mode NEQAIR v15.0 is used as a tool for creating synthetic spectra of any desired resolution (including convolution with a specified instrument/slit function). The first mode is typically exercised in simulating/interpreting spectroscopic measurements of different sources (e.g. shock tube data, plasma torches, etc.). In the second mode, NEQAIR v15.0 is used as a radiative heat flux prediction tool for flight projects. Correspondingly, NEQAIR has also been used to simulate the radiance measured on previous flight missions. This report summarizes the database updates, corrections that have been made to the code, changes to input files, parallelization, the current usage recommendations, including test cases, and an indication of the performance enhancements achieved.

Description: No abstract available

Description: NASAs Commercial Supersonic Technology (CST) project has formulated a technical challenge to design a quiet propulsion system for a low boom supersonic aircraft that meets Federal Aviation Authoritys airport noise regulations with sufficient margin. Several proposed configurations take advantage of shielding from the wing or other air-frame components. Development of carefully validated computational tools are necessary for critically evaluating installation concepts that are currently being proposed to meet the technical challenge. Semi-empirical models that predict the noise reduction potential of arbitrary shielding surfaces are yet to mature. Another key challenge is the systematic assessment of additional noise from the interaction between high speed jet turbulence and a surface in its vicinity. As a first step towards predicting noise reduction due to radical installation concepts from first principles, we simulate the noise generated by a high speed turbulent round jet near a simple planar surface. Detailed comparisons are made with a dedicated experiment conducted at NASAs Glenn Research Center. Sensitivity of far-field noise predictions to grid resolution is systematically documented. A permeable Ffowcs Williams Hawkings (FWH) surface enclosing both the jet and the shielding surface is used to predict far-field noise from the simulated flowfield. Details of the structured overset grids, numerical discretization, and turbulence model are provided. Near-field comparisons to PIV data and far-field comparisons to microphone array measurements are discussed. Excellent agreement for an initial validation study on an isolated free round jet was obtained and the findings were utilized in the jet surface interaction study. The split between shielded and reflected side of the microphone array was captured with good agreement, as well as the peak in the noise spectra due to scattering of turbulent energy into sound by the trailing edge of the surface.

Description: A validated computational fluid-structure interaction method for simulating the complex interaction between the large deformation of very thin, highly deformable structures and compressible flows is extended to consider large-scale problems in supersonic flows using parallel computing. The coupled fluid-structure interaction system is solved in a partitioned, or weakly-coupled, manner. The foundations of the applied fluid-structure interaction method are a higher-order, block-structured Cartesian, sharp immersed boundary method for the compressible Navier-Stokes equations and a computational structural dynamics solver employing a geometrically nonlinear 3-node shell element based on the mixed interpolation of tensorial components formulation. The method is applied to large deformation fluid-structure interaction validation cases before being applied to the inflation of a supersonic parachute in the upper Martian atmosphere where the goal is to demonstrate the capabilities of the solver when considering large-scale problems in supersonic flows.

Description: No abstract available

Description: No abstract available

Description: Noise has been identified as a major challenge to community acceptance of Urban Air Mobility Systems. The purpose of this paper is to assist designers, developers, and implementers in understanding the various factorsfrom the noise source itself to the conditions of the communitythat influence how the community is likely to respond to the introduction of a new noise source. Particular consideration is given to the role of non-acoustical factors and suggestions are offered as to how future research could help advance the understanding of community acceptance of Urban Air Mobility and other emergent vehicle systems.

Description: What is needed, on a regional level, to prepare for future urban air mobility operations in terms of airspace management, infrastructure, community, and aircraft.

Description: No abstract available

Description: The Mars Science Laboratory (MSL) was protected during entry into the Martian atmosphere by a thermal protection system that used NASAs Phenolic Impregnated Carbon Ablator (PICA). The heat shield of the probe was instrumented with the Mars Entry Descent and Landing Instrument (MEDLI) suite of sensors. MEDLI Integrated Sensor Plugs (MISP) included thermocouples that measured in-depth temperatures at various locations on the heatshield. The flight data has been used as a benchmark for validating ablation codes within NASA. This work seeks to refine the estimate of the material properties for the MSL heat shield and the aerothermal environment during Mars entry using estimation methods in DAKOTA on the temperature data obtained from MEDLI.

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Description: No abstract available

Description: This paper describes the use of detailed multidisciplinary fluid/thermal/ structural/neutronic simulations to predict performance of the nuclear fuel elements of a Nuclear Thermal Propulsion rocket reactor. To achieve maximum performance, a rocket reactor's fuel must operate near thermal hydraulic, structural and neutronic limits where multidisciplinary interactions are important. Yet physical testing is expensive, time- consuming and risky. Lower-fidelity correlations (heat transfer) and simulations have always existed for design, and one role of detailed numerical analysis is to confirm correlation validity and accuracy. For complex and subtle issues, detailed numerical simulations may prove their value. The paper gives examples of both of these situations. Limitations of the methods and potential extensions will be explored.