ALBERT

Description: A fundamental exploratory experiment is conducted assessing the performance of a one-sided ejector with the eventual goal of noise reduction for jet engines. The hardware is comprised of an 8:1 rectangular nozzle together with an ejector box whose lower surface is flush with the lower lip of the nozzle. Secondary flow is allowed through a gap between the upper lip of the nozzle and a flap that constitutes the upper surface of the ejector. Wall static pressures and Pitot probe surveys are conducted to evaluate the performance of the ejector with variation of geometric parameters. It is found that addition of vortex generating tabs at the upper lip of the nozzle significantly increases secondary flow entrainment. The entrainment is further enhanced by a divergence of the ejector upper surface. Limited noise measurements are done. The baseline ejector (without tabs) often encounters flow resonance with accompanying tones. The tabs have the additional benefit of eliminating those tones in all cases. However, for the tabbed case, addition of the ejector produces insignificant further noise reduction. This is due to the fact that the flow remains unmixed on the lower half of the ejector. The focus of ongoing and future efforts is to achieve sufficient mixing of the flow so that the exhaust velocities are uniformly low, while keeping the ejector hardware short and lightweight.

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: 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: Outline - Introduction: X-57 CFD task overview; Motivation. Part I, Computational simulations without propulsion: Establishing CFD (Computational Fluid Dynamics) Best Practices - Grid generation - Mesh refinement study - Numerical methods - Wind tunnel validation study; Power-Off Aerodynamic Database Results. Part II, Computational simulations with propulsion: Cruise Power-On Database; High-Lift Power-On Database. Summary.

Description: A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile.

Description: A simulator to artificially generate turbofan broadband signatures using the ANCF (Advanced Noise Control Fan) test article is presented. [Development of a Broadband Acoustic Emulator to Mature Propulsion Noise Reduction (CFANS-BB: Configurable Fan Artificial Noise Source- Broadband)]

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: Reynolds-Averaged Navier-Stokes simulations have been performed on a three-stream inverted velocity profile nozzle with and without various configurations of chevrons attached. The nozzle was mounted on a planform to imitate an engine mounted above a wing, shielding ground observers from engine noise. Several chevron designs intended to aggressively mix the jet and move noise sources upstream for shielding were examined to investigate there effects on noise and thrust. Numerical results for the baseline nozzle and one chevron configuration were compared with far-field noise and particle image velocimetry data obtained in NASA Glenn Research Center's Aero-Acoustic Propulsion Laboratory. A configuration in which chevrons alternate penetration into the primary stream and tertiary fan stream was explored using the Modern Design of Experiments approach. Short, high-penetration chevrons demonstrated a significant noise reduction for a relatively small thrust penalty.

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: 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: 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: This paper describes the design of a turboshaft engine for a tiltwing air taxi application. In this case, the tiltwing air taxi is intended to fly a 400 nm mission with up to fifteen passengers. Engine requirements for the concept engine are taken from aircraft system studies where thrust is produced by four propellers driven by electric motors and powered by a single gas turbine engine. The purpose of this paper is to perform a cycle design optimization that minimizes fuel consumption and weight while respecting current technology limitations to meet mission requirements. To achieve results, the engine overall pressure ratio and maximum temperature at the exit of the combustor are set as the design parameters. Several sensitivity studies are also performed to visualize optimization trends. Results of the optimization study show solutions are heavily dependent on engine cooling flow requirements and exact mission requirements. This engine is intended for use in large system optimization research.

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: Time accurate simulation of non-equilibrium flows inside shock tube facilities presents several challenges from both physical and mathematical aspects. Furthermore, the large computational cost makes it impractical to support a real-time experimental test campaign. In this work, we explore other methods for modeling the shock tube problem with the main focus on the post-shock region and the absolute radiation emanating from it. The proposed alternative approach is several orders of magnitude less computationally expensive while still accurate enough with regards to the quantities of interest. Excellent agreement is found with the established stagnation-line approach. Comparison with time-accurate simulations shows good agreement close to the peak values and disagreement of the temperatures relaxation and radiance profiles toward equilibrium.

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: Magnetic gearing is being investigated at NASA as a replacement to conventional mechanical gearing in aerospace applications. Some potential benefits of magnetic gears over mechanical gearing are torque transmission without mechanical contact, decreased transmission noise, and no required lubrication. However, in order to be a viable alternative for aerospace applications, magnetic gearing must be shown to provide high enough specific torque (torque per unit mass). NASA's second magnetic gearing prototype (PT-2) was able to achieve promising specific torque on par with low torque mechanical gearboxes. This work will briefly review the electromagnetic and structural design of PT-2, provide detailed information on fabrication and assembly, examine build errors, walk through rebuild efforts to improve operation, and conclude with remarks on build difficulties and opportunities for improvement in future prototypes.

Description: NASA Acoustic Stirling IRAD (Internal Research and Development) Thermal Recovery Energy Efficient System (TREES) Energy Conversion and Management in Aircraft. Presentation on energy conversion on aircraft. Thermal energy recovery changes aircraft thermal management from being a necessary burden on aircraft performance to a desirable asset. It improves the engine performance by recycling waste heat and ultimately rejecting all collected aircraft heat out through the engine nozzle.

Description: Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design in support of the National Aeronautics and Space Administration's (NASA) Aeronautics Mission Research Directorate's (ARMD) Revolutionary Vertical Lift Technology (RVLT) Project. These models were developed using open data sources, assuming current and advanced technology levels, and range from 650 to 7,500 shaft output horsepower (485 to 5,600 kilowatts). Documenting the methodology, assumptions, and resulting performance realizes important benefits for NASA and the aviation community. NASA concept vehicle efforts using these propulsion models can more readily shared among the government, industry and university community as common baselines to support current and future work. Assessing the benefits of advanced technologies and new configurations can be facilitated using these models, which helps guide technology investment. As the various modeling conceptual vehicle and mission analysis environments advance, these models can be used directly for broader systems analysis studies, including optimization within the propulsion model itself. To perform this effort, the turboshaft engine is briefly discussed, highlighting the specific components and their expected performance characteristics over the power range and technology levels considered. Engine configurations will also be discussed as they will vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be reported, noting trends that are important for system studies. The effect of advanced propulsion technologies on RVLT-concept vehicles are also reported. Finally, potential future propulsion modeling work will be proposed.

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: A model-scale exhaust system was tested to validate low-noise concepts and noise prediction methods. The tests involved far-field acoustics, translating phased array, and particle image velocimetry; this report covers the far-field acoustic measurements. Data were acquired for a series of nozzles with different chevron designs, both uninstalled and installed on a representative aircraft planform. The impact of the various chevron treatments on the far-field noise was documented, along with the impact of the pylon and planform. For the baseline nozzle, installation produced a 2 EPNdB (Effective Perceived Noise in deciBels) reduction, as assumed in system studies. Chevrons were used to shift noise sources upstream to maximize the installation benefits and to reduce unshielded sources downstream. These resulted in reductions of 4-5 EPNdB...

Description: Turboshaft engine performance and weight models were developed to support conceptual propulsion and vehicle mission design and performance under the Revolutionary Vertical Lift Technology (RVLT) Project. These models were developed using open data sources, assuming current and advanced technology levels, and range from 650 to 7,500 shaft output horsepower (485 to 5,600 kW). Documenting the methodology, assumptions, and resulting performance realizes important benefits NASA and the aviation community. NASA concept vehicle efforts using these propulsion models can be more readily shared among the government, industry and university community as common baselines to support current and future work. Assessing the benefits of advanced technologies and new configurations can be facilitated using these models, which helps guide technology investment. As the various modeling conceptual vehicle and mission analysis environments advanced, these models can be used directly for broader systems analysis studies, including optimization within the propulsion model itself. To perform this effort, the turboshaft engine is briefly discussed, highlighting the specific components and their expected performance characteristics over the power range and technology levels considered. Engine configurations will also be discussed as they will vary based on power output and assumed technology level. Engine performance, such as airflow, power output and weight will be reported, noting trends that are important for system studies. The effect of advanced propulsion technologies on RVLT concept vehicles are also reported. Finally, potential future propulsion modeling work will be proposed.

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: This paper presents the design, development, operation, and test capabilities of a proposed superconducting coil testbed to measure alternating current (AC) losses at the NASA Glenn Research Center. Superconducting AC losses are important in the design of electric stators and rotors, power transmission lines, transformers, fault current limiters, magnets, and superconducting energy storage (not batteries). The new liquid-hydrogen-based rig will allow superconducting testing across a wide range of test parameters, including injected current up to 400 A, frequency (0 to 400 Hz), magnetic field (0 to 0.6 T), phase angle between induced voltage and injected current (180 to 180), coil coolant temperature (18 to 28 K), and AC power loss (5 to 30 W). While the target application of interest is 20 K superconducting MgB2 (the only superconductor that can presently be made with low losses) stator coils for future electric machines, the rig can accommodate test articles (TAs) with straight wire, tape, cables, coils of any shape, any allowable combination of superconducting wire and fluid (e.g., yttrium barium copper oxide (YBCO) coils and liquid nitrogen), and AC or direct current (DC) testing. The new spin rig builds upon the existing Air Force spin rig through a more flexible mode of fluid control, a wider gap space (up to 10.2 cm) for TAs, and the ability to accommodate TAs over a wider range of operating temperatures (18 to 95 K) using liquid hydrogen, gaseous helium, or liquid nitrogen as the working fluid, thus supporting direct cooled machines below 77 K.

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: The pressure gain combustion (PGC) community is currently investigating rotating detonation engine (RDE) configurations where the flow direction is predominantly radial while the detonation travels circumferentially. These configurations are sometimes referred to as disk rotating detonation engines (DRDE) due to their nominal appearance as two disks in parallel with a gap between them. Having radial flow between disks, as opposed to the conventional RDE with axial flow in an annulus, may have profound effects on both the flow field and the performance. It may also yield extraordinarily compact devices which are well suited to particular propulsion and power applications. This presentation describes a preliminary effort to model the DRDE using a modified computational fluid dynamics (CFD) code originally written for analyzing ordinary RDE's. The quasi-two-dimensional code modifications are described, and some simple test flows are analyzed to insure that the modifications are functioning as envisioned. The code is then used to examine several DRDE scenarios such as radially inward and radially outward devices to see if stable operation is possible and if so, to assess the performance in terms of pressure gain. It is found that several flow scenarios are not only stable, but show superior performance to the ordinary RDE.

Description: No abstract available

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: Magnetic gears are currently being developed for use in a variety of industries such as wind and automotive, because of their higher reliability and lower maintenance cost than their mechanical counterparts. The bulk of magnetic gear development to date has focused on maximizing the technology's volumetric torque density. In contrast, the primary performance metrics for an aircraft's gear box are its mass and efficiency. To that end this paper presents a study of the achievable electromagnetic specific torque and efficiency of concentric magnetic gears. NASA's second magnetic gear prototype is used as the baseline for this study. Achievable electromagnetic specific torque and efficiency trends are presented with respect to higher level design variables such as gear ratio and radius.

Description: Magnetic gears are an attractive alternative to mechanical gears for electrified aircraft drive systems due to their ability to transmit torque without mechanical tooth contact. Consequently, magnetic gears enable electrified aircraft to take advantage of the benefits of gearing without introducing most of the contact-related reliability concerns associated with mechanical gearing. Magnetic gears however, have not been shown to match the specific torque (torque/mass) and efficiency of their mechanical counterparts in an aerospace application to date. In this paper, the design of a concentric magnetic gear for a personal air transport NASA reference vehicle is presented to demonstrate the feasibility of a magnetic gear for aerospace applications.

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: 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: 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: An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves Blasius-like mean velocity profiles but a highly disturbed laminar state with large turbulence intensities in a range of subsonic Mach numbers.

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: Electrified aircraft propulsion seeks to address ambitious goals in the commercial airline industry, including significant decreases in fuel burn, emissions, noise, and takeoff field length. In order to move these electrified propulsion concepts forward, analysis tools are needed that can model propulsion systems containing both gas turbine and power system components. This work presents the definition of an electric port, a set of electrical power systems tools, and simulation examples for the Numerical Propulsion System Simulation (NPSS) software. NPSS is the industry standard modeling and simulation package for aircraft propulsion systems, and the ability to design, size, integrate, and analyze electric power systems will enable industry efforts towards the development of electrified aircraft propulsion.

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: An overview is given of an effort that focused on using CFD analysis to complement design and configuration definition of Lean-Direct Injection (LDI) combustion concepts for NASA's Commercial Supersonic Transport (CST) program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for second and third generation LDI configurations at CST cruise conditions. All computations were performed with a consistent approach of mesh-generation, spray modeling, ignition and kinetics modeling. Emissions (EINOx) characteristics were predicted for CST cruise conditions, and compared with emissions data from experimental measurements to evaluate the fidelity of the CFD modeling approach to predict emissions changes in response to changes in supersonic cycle conditions.

Description: Some of the challenges associated with developing electric aircraft propulsion systems include developing powertrain components that are both efficient and light-weight. In particular, electric motors must simultaneously achieve high efficiency by minimizing electrical and mechanical losses while also achieving high specific power by increasing the torque and/or speed. Normally increasing torque or speed will increase electrical and mechanical losses. The High Efficiency Megawatt Machine (HEMM) minimizes electrical losses by incorporating a superconductor to enable increased current on the rotor. And the rotor spins in a vacuum to minimize thermal and mechanical losses. Some organizations have been developing superconducting rotors for similar reasons using either cryogenic fluid transfer systems, fully immersed cryogenic cooling, and in a few cases utilized built-in cryogenic cooling on the rotor using a Brayton or Stirling system but the implementation was too large or inefficient for effective motor integration. Instead, a new approach for cryogenically cooling the superconducting rotor coil with an embedded rotating cryocooler is presented that fits completely within the rotating shaft.

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: Time accurate simulation of non-equilibrium flows inside shock tube facilities presents several challenges from both physical and mathematical aspects. Furthermore, the drastic computational cost makes it non-practical to support real-time experimental test campaign. In this work, we explore other methods for modeling the shock tube prob- lem with the main focus on the post-shock region and the absolute radiation emanating from it. The proposed alternative approach is several orders of magnitude less computa- tionally expansive while still accurate enough with regards to the quantities of interest. Excellent agreement is found with the well-established stagnation-line approach. Comparison with the time-accurate simulation shows good agreement close to the peak values and disagreement of the temperatures relaxation and radiance profiles toward equilibrium, due to shock speed unsteadiness.

Description: No abstract available

Description: The X-57 60kW Permanent Magnet Synchronous Motor for cruise applications was modeled utilizing a two-dimensional electromagnetics simulation software called Finite Element Method Magnets (FEMM, D. Meeker). Through FEMM, the simulated induction and torque characteristics of the X-57 PMSM were obtained. These parameters and other values were compared to actual static laboratory measurements. A three-dimensional electromagnetic model of the X-57 cruise motor was created utilizing OperaFEA (Dassault Systemes SE, Velizy-Villacoublay, France). Torque, RPM, power, resistance, and inductance characteristics were examined along with establishing work to begin examining heat flow and heat dissipation for efficiency purposes.

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: The adoption of SiC devices in high power applications enables higher switching speed, which requires lower circuit parasitic inductance to reduce the voltage overshoot. This paper presents the design of a busbar for a 500 kVA three-level active natural clamped converter. The layout of the busbar is discussed in detail based on the analysis of the multiple commutation loops, magnetic cancelling effect, and DC-link capacitor placement. The loop inductance of the designed busbar is verified with simulation, impedance measurements and converter experiment. The results can match with each other and the inductances of small and large loop are 6.5 nH and 17.5 nH respectively, which is significantly lower than the busbars of NPC type converters in other references.

Description: NASA is broadly engaged in Electrified Aircraft Propulsion (EAP) efforts across air vehicle sizes and electric aircraft propulsion approaches. EAP enables a wide range of propulsion airframe integration options as well as the use of rechargeable energy storage in an aircraft. This paper is limited to a discussion of boundary layer ingestion (BLI) systems which are located on the fuselage of the aircraft and use electrical drive systems. We term that combination an "electrical propulsive fuselage". The benefits, challenges, and design parameters of an electrically driven fuselage BLI system are considered. Five existing types of fuselage BLI implementation approaches which can be implemented using either electrical or mechanical drive systems are reviewed. An overview of boundary layer types, fan response to boundary layer, and electrical system for aircraft propulsion is presented. An idea distributed electric propulsive fuselage is proposed.

Description: An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASAs N+3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flame tube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a medium-power cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

Description: Magnetic gears are currently being explored to replace mechanical gears in various industries such as wind and automotive due to their higher reliability and lower maintenance requirements. In these applications volume minimization has been the goal of magnetic gear development. In contrast, the primary performance metrics for electrified aircraft drives are mass and efficiency. This paper presents the first ever study of design tradeoffs between electromagnetic mass and efficiency of concentric magnetic gears and the feasibility of achieving the low mass and high efficiency required for electrified aircraft applications. Higher level design variables are considered, including gear ratio, number of magnetic pole pairs, and number of magnets per pole pair.

Description: This presentation describes experimental and computational approaches to measuring pressure gain in the various devices currently under investigation wherein the working fluid undergoes a pressure gain combustion (PGC) process. Pressure gain is essentially a measure of the fluid availability for work or thrust production. The devices covered are Resonant Pulse Combustors, Internal Combustion Wave Rotors, Pulse Detonation Engines, and Rotating Detonation Engines. The approaches to pressure gain measurement differ in each device. However, all of the approaches attempt to address the fundamental challenges of PGC system measurement: the extremely harsh environment which makes instrumentation difficult, and the temporal and spatial non-uniformity associated with the exhausting flow which makes assigning a single value to the total pressure difficult. As part of the two-day 2019 International Constant Volume and Detonative Combustion Workshop, held in conjunction with the 2019 AIAA Propulsion and Energy Forum, this presentation is intended to foster discussion and eventual consensus on acceptable measurement methods.

Description: Transition from fossil fuels to synthetic drop-in fuels without the need to change existing combustors is the current research topic. The combustor performances such as cold-day ignition limits, lean blow-out (LBO) limits and altitude relight limits are the main focus points. The objective of this work is to evaluate the effect of different fuel candidates on the operability of gas turbines by comparing a conventional petroleum-based fuel with one other alternative fuel candidate. Time filtered Navier-Stokes simulations (TFNS) and K-LES are performed to examine the performance of these fuels at the stable conditions close to blow-out in a referee combustor rig.

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: This paper continues a parametric study in which we consider the effect of air swirler configuration on the flame structure and combustor performance using a circular 7-point Lean Direct Injector Array for gas turbine applications. The injector array consists of a center swirler element surrounded by six swirler elements. Parameters considered in this study include swirler angle (60 or 52), handedness (co-swirling or counter-swirling) and center swirler offset. The primary focus considers flame stability, comparing four key air swirler configurations: for 1) fuel-lean flames; 2) high cold flow air reference velocity flames. We determined that the baseline swirler configuration had the best lean stability and could sustain the highest reference velocity. For this baseline configuration, we also compare the lean-blowout limits of four aircraft gas turbine reference fuels. With regard to lean blow-out, we determined that C4 could sustain the leanest flame, followed closely by A2. A1 was a poor performer.

Description: In rotorcraft, one of the main sources of mechanical failure is the gearbox, because of the many wear and failure modes associated with tooth contact in traditional mechanical gear boxes. Magnetic gears transmit torque without mechanical tooth contact between gear bodies and therefore they have none of the tooth contact related failure modes associated with mechanical gearing. As a result, magnetic gears have the potential to enable more reliable rotorcraft gearboxes. However, magnetic gears have not been demonstrated to match the performance of mechanical gearboxes at a high enough technology readiness level (TRL) to be used on an aircraft to date. To that end, NASA's Revolutionary Vertical Lift Technologies project has made an investment in developing magnetic gearboxes specifically for electrified vertical lift vehicles (EVTOL). In this presentation, the results of that investment to date will be discussed.

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: 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: Presentation to the International Forum on Aviation (IFAR) at the Electric Hybrid Propulsion Workshop #2 in Budapest, Hungary. This presentation is to provide an overview of NASA's investments in electrified propulsion as a starting point for the workshop, which will concentrate on the safety of electrified airplanes and potential for international collaboration.

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.

Description: No abstract available

Description: No abstract available

Description: This is for an invited lecture at Cleveland State University for a combustion course. The presentation gives an overview of some of the optical diagnostic techniques the Combustion Branch uses to characterize research fuel injection concepts to reduce emissions. Examples are provided that come from previously presented conference papers

Description: No abstract available

Description: Porous Microstructure Analysis (PuMA) software is used to perform simulations of molecular beam scattering experiments of hyperthermal atomic oxygen striking FiberFormr, a carbon preform material used commonly as a precursor in thermal protection systems (TPS). The purpose of this study is to investigate the reactive interaction of fibrous carbon with atomic oxygen in a complex microstructure, which is the primary source of carbon removal at lower temperatures. The detailed micro-structure of FiberFormr obtained from X-ray micro-tomography is used in the PuMA simulations to capture the complexity of the porous and fibrous characteristic of FiberFormr. A finite-rate surface chemistry model recently constructed from the molecular beam scattering experiments on vitreous carbon is applied to each fiber of the FiberFormr material. This model consists of detailed surface reaction mechanisms such as adsorption, desorption, and several types of Langmuir-Hinshelwood (LH) reactions to characterize the oxygen-carbon interactions at the surface. Comparison between the experimental and PuMA time-of-flight (TOF) distributions of both O and CO show good agreement. It is also found that a significantly higher amount of CO is generated when the beam interacted with FiberFormr, when compared with vitreous carbon. This is postulated to be primarily a result of multiple collisions of oxygen with the fibers, resulting in an higher effective rate of CO production. Multiple collisions with the different fibers, resulting from the porous nature of FiberFormr is also found to thermalize the O atoms, in addition to the adsorption/desorption process. The effect of micro-structure is concluded to be crucial in determining the final composition and energy distributions of the products. Thus, an effective model for the oxygen interaction with FiberFormr, fully accounting for the detailed micro-structure, for use in Computational Fluid Dynamics (CFD) and material response codes, is presented.

Description: MIT, Aurora Flight Sciences, and USC have collaborated to assess the feasibility of electric, hybridelectric, and turbo-electric propulsion for ultra-efficient commercial transportation. The work has drawn on the team expertise in disciplines related to aircraft design, propulsion-airframe integration, electric machines and systems, engineering system design, and optimization. A parametric trade space analysis has been carried out to assess vehicle performance across a range of transport missions and propulsion architectures to establish how electrified propulsion systems scale. An optimization approach to vehicle conceptual design modeling was taken to enable rapid multidisciplinary design space exploration and sensitivity analysis. The results of the analysis indicate vehicle aero-propulsive integration benefits enabled by electrification are required to offset the increased weight and loss associated with the electric system and achieve enhanced performance; the report describes the conceptual configurations than can offer such enhancements. The main contribution of the present work is the definition of electric vehicle design attributes for potential efficiency improvements at different scales. Based on these results, key areas for future research are identified, and extensions to the trade space analysis suitable for higher fidelity electrified commercial aircraft design and analysis have been developed.

Description: The high power density of emerging electronic devices is driving the transition from remote cooling, which relies on conduction and spreading, to embedded cooling, which extracts dissipated heat on-site. Two-phase microgap coolers employ the forced flow of dielectric fluids undergoing phase change in a heated channel within or between devices. Such coolers must work reliably in all orientations for a variety of applications (e.g., vehicle-based equipment), as well as in microgravity and high-g for aerospace applications, but the lack of acceptable models and correlations for orientation- and gravity-independent operation has limited their use. Reliable criteria for achieving orientation- and gravity-independent flow boiling would enable emerging systems to exploit this thermal management technique and streamline the technology development process. As a first step toward understanding the effect of gravity in two-phase microgap flow and transport, in an earlier effort, the authors studied the effects of evaporator orientation, mass flux, and heat flux on flow boiling of HFE7100 in a 1.01 mm tall by 13.0 mm wide by 12.7 mm long microgap channel. Orientation-independence, defined as achieving similar critical heat fluxes, heat transfer coefficients, and flow regimes across orientations, was achieved for mass fluxes of 400 kg/sq.m-s and greater (corresponding to a Froude number of about 0.8). In the present effort, the authors have studied the effects of gravity, mass flux, and subcooling on flow boiling of HFE7100 in a 0.17 mm tall by 13.0 mm wide by 12.7 mm long microgap channel. The Flow Boiling in Microgap Coolers payload experienced about three minutes of weightlessness and shorter periods of high-g (up to about 5 g) during two recent flights aboard the Blue Origin New Shepard reusable launch vehicle. The results from the flight experiments will be presented and compared with published criteria for achieving gravity-independence.

Description: An infrared (IR) camera provides a way of examining temperature trends associated with simulated microgravity flame spread in the Narrow Channel Apparatus (NCA). The IR camera measures the surface temperature of solid poly methyl methacrylate (PMMA) fuel. These tests examine the forward conduction of heat ahead of the flame front in the non-thermally thin fuel.The NCA is a combustion wind tunnel that simulates a microgravity flame spread environment by employing a narrow gap between the fuel and ceiling of the device, limiting the effects of buoyancy. Test conditions of a 5 mm gap, mean opposed flow velocity of 15 cm/s, and fuel thickness of 3 mm are used.PMMA is selected as the fuel due to repeatability of test results, ease of computational modeling, and known combustion mechanics. Using specific lens and bandpass filter combinations the PMMA can be imaged as effectively opaque. The spectral emissivity for PMMA was calculated and incorporated into the calibration of the camera.Surface temperatures from the IR camera are compared to results from thermocouples embedded in the surface of the fuel. The IR camera results show that nontrivial forward conduction occurs during tests, and therefore must be included in computational models of the process.

Description: Hybrid electric propulsion architectures provide the infrastructure to enable additional benefits to the propulsion system that are otherwise unrealizable with the sole use of the current, state-of-the-art, gas-driven, turbine engines. The presence of electric machines (EMs) coupled to the shaft(s) of the turbine engine provide the ability to actively alter the operation of the engine to the benefit of the propulsion system and the aircraft it propels. This is the goal of the Turbine Electrified Energy Management (TEEM) concept, which at its broadest level addresses the management of energy across the electrified propulsion system. Prior work has demonstrated the use of this concept to alter steady-state operation and improve transient operability of a hybrid-electric propulsion system. The main benefits previously illustrated include the elimination of stability bleeds and expansion of the turbomachinery design space in order to enable more efficient designs. This paper focuses on the development of control strategies to implement the TEEM concept, and it explores several possible architecture variants for applying this concept. Comparison studies are conducted between a purely gas-driven turbofan (baseline engine configuration) and TEEM augmented variants of the baseline engine. The variants are distinguished by the shaft(s) that possess an EM. The configurations consider EMs on both shafts, an EM on the high pressure spool (HPS) only, and an EM on the low pressure spool (LPS) only. These configurations are referred to as the dual-spool configuration, the HPS configuration, and LPS configuration, respectively. The studies expose several options in configuring and controlling the system, including the use of a single EM coupled to a single shaft of a two-spool engine to positively impact the operability of both shafts. The studies also demonstrate the use of independently designed controllers for the electric machine(s) that allow for a decoupled control design process.

Description: NASA's broad investments in Electrified Aircraft Propulsion (EAP) are reviewed in this paper. NASA investments are guided by an assessment of potential market impacts, technical key performance parameters, and technology readiness attained through a combination of studies, enabling fundamental research, and flight research. NASA has determined that the impact of EAP varies by market and NASA is considering three markets: national/international, on-demand mobility, and short haul regional air transport. Flight research is underway to demonstrate integrated solutions and inform standards and certification processes. This paper focuses on the vehicle related activities, however there are related NASA activities in air space management and vehicle autonomy activities as well as a breakthrough technology project called the Convergent Aeronautics Solutions Project. A key finding is that sufficient technical advances in key areas have been made which indicate EAP is a viable technology for aircraft. Significant progress has been made to reduce EAP adoption barriers and further work is needed to transition the technology to a commercial product and improve the technology so it is applicable to large transonic aircraft. This paper will review the activities of the Hybrid Gas Electric Subproject of the Advanced Air Transport Technology Project, the Revolutionary Vertical Lift Technology Project, and the X-57 Flight Demonstration Project, and discuss the potential EAP benefits for commercial and military applications.

Description: An overview is given of an effort that focused on using CFD analysis to complement design and configuration definition of Lean-Direct Injection (LDI) combustion concepts for NASA's Commercial Supersonic Transport (CST) program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for second and third generation LDI configurations at CST cruise conditions. All computations were performed with a consistent approach of mesh-generation, spray modeling, ignition and kinetics modeling. Emissions (EINOx) characteristics were predicted for CST cruise conditions, and compared with emissions data from experimental measurements to evaluate the fidelity of the CFD modeling approach to predict emissions changes in response to changes in supersonic cycle conditions.

Description: Some of the challenges associated with developing electric aircraft propulsion systems include developing powertrain components that are both efficient and light-weight. In particular, electric motors must simultaneously achieve high efficiency by minimizing electrical and mechanical losses while also achieving high specific power by increasing the torque and/or speed. Normally increasing torque or speed will increase electrical and mechanical losses. The High Efficiency Megawatt Machine (HEMM) minimizes electrical losses by incorporating a superconductor to enable increased current on the rotor. And the rotor spins in a vacuum to minimize thermal and mechanical losses. Some organizations have been developing superconducting rotors for similar reasons using either cryogenic fluid transfer systems, fully immersed cryogenic cooling, and in a few cases utilized built-in cryogenic cooling on the rotor using a Brayton or Stirling system but the implementation was too large or inefficient for effective motor integration. Instead, a new approach for cryogenically cooling the superconducting rotor coil with an embedded rotating cryocooler is presented that fits completely within the rotating shaft.

Description: An experiment is conducted with hot-wire anemometry to document the exit boundary layer characteristics of two nozzle configurations at jet Mach numbers up to 0.82. Far-field noise and jet plume experimental data from these two configurations have been used in Large Eddy Simulations (LES) of jets by colleagues at other Institutions. The current experiment provides the boundary layer data which have been identified as being critical for validation of the simulations since the initial conditions can significantly affect subsequent jet evolution and its radiated noise. The data exhibit fully turbulent boundary layers for the case with a pipe attached upstream of the nozzle. The case without the pipe involves Blasius-like mean velocity profiles but a highly disturbed laminar state with large turbulence intensities in a range of subsonic Mach numbers.

Description: Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Description: Electrified aircraft propulsion (EAP) systems hold potential for the reduction of aircraft fuel burn, emissions, and noise. Currently, NASA and other organizations are actively working to identify and mature technologies necessary to bring EAP designs to reality. This paper specifically focuses on the envisioned control technology challenges associated with EAP designs that include gas turbine technology. Topics discussed include analytical tools for the dynamic modeling and analysis of EAP systems, and control design strategies at the propulsion and component levels. This includes integrated supervisory control facilitating the coordinated operation of turbine and electrical components, control strategies that seek to minimize fuel consumption and lessen the challenges associated with thermal management, and dynamic control to ensure engine operability during system transients. These dynamic control strategies include innovative control approaches that either extract or supply power to engine shafts dependent upon operating phase, which may improve performance and reduced gas turbine engine weight. Finally, a discussion of control architecture design considerations to help alleviate the propulsion/aircraft integration and certification challenges associated with EAP systems is provided.

Description: An efficient strategy for propagating sonic boom signatures from a near-field Computational Fluid Dynamics (CFD) solution to the mid-field is presented. The method is based on a high-order accurate finite-difference discretization of the 3D Euler equations on a specially designed curvilinear grid and a single sweep space marching solution algorithm. The new approach leads to more than a factor of two reduction in overall computational resources compared to the current method used to propagate near-field sonic booms to the ground. Accuracy and efficiency of the near-field to mid-field process is demonstrated using a selection of test cases from the AIAA Sonic Boom Prediction Workshops. Azimuthal dependence of nonlinear wave propagation from the near-field to mid-field is analyzed along with its effects on the ground level noise.

Description: The accuracy of the scale-resolving simulations for practical geometries strongly depends on the inflow boundary conditions. Imposing experimentally observed turbulent inflow profiles for the numerical simulations is a major challenge. Existing methods available in the literature assume self-similar behavior, which is not true for most of the experiments. In the present work, we formulate the turbulent inflow profile generation technique as an optimization problem. An adjoint technique is exploited to evaluate the sensitivities of multiple input parameters for the present problem. The present formulation is then tested to generate a laminar boundary layer profile, turbulent boundary layer profile, and turbulent jet profile.

Description: Magnetic gearing is being investigated at NASA as a replacement to conventional mechanical gearing in aerospace applications. Some key benefits of magnetic gears over mechanical gearing are torque transmission without mechanical contact, decreased transmission noise, less frequent maintenance, and lack of lubrication. In order to take advantage of these benefits in aerospace applications, magnetic gearing must be shown to provide high enough specific torque (torque per unit mass). Prototype 2 (PT-2), developed to maximize specific torque, and fabricated at NASA Glenn Research Center, has shown promising specific torque comparable to low torque mechanical gears. This work will briefly review the electromagnetic and structural design of PT-2, provide detailed information on fabrication and assembly, examine build errors, walk through rebuild efforts to improve operation, and conclude with remarks on build difficulties and opportunities for improvement in future prototypes.

Description: Detailed spectrally and spatially resolved radiance has been measured in the Electric Arc Shock Tube at NASA Ames Research Center for conditions relevant to Titan entry, with varying atmospheric composition, free-stream density (equivalently, altitude) and shock velocity. The test campaign measured radiation at velocities from 4.7 km/s to 8 km/s and free-stream pressures of 0.1, 0.28 and 0.47 Torr with a variety of compositions. Radiances measured in this work are substantially larger compared to that reported both in past EAST test campaigns and in other shock tube facilities. Depending on the metric used for comparison, the discrepancy can be as high as an order of magnitude. Due to the difference with previously reported data, a substantial effort was undertaken to provide confidence in the new results. The present work provides a new benchmark set of data to replace those published in previous studies. The effect of gas impurities identified in previous shock tube studies was also examined by testing in pure N2 and deliberate addition of air to the CH4/N2 mixtures. Furthermore, a test campaign in pure N2 was also conducted with the aim of providing data for improving fundamental understanding of high enthalpy flows containing N2, such as high-speed entries into Earth or Titan. These experiments cover conditions from approximately 6 km/s to 11 km/s at an initial pressure of 0.2 Torr. It is the intention of this paper to motivate code comparisons benchmarked against this data set.

Description: Modifications to key coefficients in a k E based explicit algebraic stress model (EASM) are examined with the objective of improving the prediction of turbulent jet flows. The pressure strain coefficient, C2 and the turbulent diffusion coefficients, k and E were investigated. For a series of benchmark subsonic jets at heated and unheated conditions, lowering C2 from the default value of 0.36 to 0.10 resulted in a significant improvement in the jet mixing, when compared to experimental data. Changing k and E from default values of 1.00 and 1.4489, respectively, to 0.50 and 0.7244, respectively, improved the initial mixing rate, while reducing the farfield mixing rate and the peak turbulent kinetic energy along the centerline. A high-speed mixing layer was also investigated for performance of baseline and modified EASM coefficients, with similar results as for the jet cases. A flat plate boundary layer was briefly examined to determine the effects of changing the coefficients on the turbulent skin friction coefficient. The change to the pressure strain coefficient, C2 = 0.10 is recommended for future EASM calculation of jets flow; however, it is also recommended that the diffusion coefficients remain at their default values.

Description: Two full seven-equation turbulence models available in the FUN3D code are evaluated for their ability to improve the computation of challenging mixing flows encountered in aerospace propulsion. These models are the SSG/LRR and Wilcox full second-moment Reynolds stress models. They solve equations for the six components of the Reynolds stress and a seventh equation for the turbulent length scale. Two standard eddy viscosity models are also evaluated for comparison, the Spalart-Allmaras (SA) one-equation model and the Menter Shear Stress Transport (SST-V) two-equation turbulence model. Flow through an axisymmetric reference nozzle is examined at three flow conditions: subsonic unheated, subsonic heated, and near sonic unheated. Centerline profiles of velocity and turbulent kinetic energy and radial profiles of velocity, turbulent kinetic energy and turbulent stresses are examined. Results showed that the SA model did well at predicting the jet potential core length, but over-mixed the downstream flow, whereas the SST-V model over-predicted the potential core length. The Wilcox-model significantly over-predicted the potential core length and under-predicted the mixing and was not well-suited for the jet flows evaluated, however the SSG/LRR Reynolds stress model did well at predicting the mixing rate and mean velocity for all cases examined.

Description: An experimental study is conducted in an effort to advance the understanding of flow physics associated with a boundary layer ingesting, distributed propulsion system. The influence of incoming boundary layer thickness and flow distortions are studied on the flow downstream and the overall performance of the system. The propulsion model, fabricated using additive manufacturing and integrated with electrical fans, is mounted on a flat plate and tested at subsonic speeds. Detailed characterization of the incoming boundary layer and subsequent assessment of the downstream flow field is performed using hotwire anemometry. Modification of the incoming boundary layer is achieved by placing tripping devices, such as rods and vortex generating ramps, near the leading edge of the flat plate. The overall performance of the system for different incoming flow conditions is analyzed by comparing magnitudes of exhaust velocities as well as estimated propulsive power to the corresponding baseline values. For a constant input power to the fans, smaller upstream flow distortions and moderately thickened boundary layers result in marginal changes in the flow field downstream. On the other hand, notable reductions in downstream flow velocities and propulsive power are observed in the case of a significantly thicker and/or distorted incoming boundary layer. It is hoped that this study will serve as a database for this technologically relevant flow field that has not been explored adequately before.

Description: An instability is described which arises in computational fluid dynamic (CFD) simulations of semi-idealized rotating detonation engines (RDE) configured with a throat at the exit. Its existence is verified by examining output from two independently developed CFD codes simulating the same configuration and producing solutions that agree well. The instability is shown to be thermo-acoustic in that a spatial integral of the product of pressure and heat release fluctuations develops a regular oscillation which grows in time. The instability can become severe enough to cause detonation failure. Its onset is shown to be closely linked to the size of the exit throat and the size of the inlet restriction; both parameters that strongly influence RDE performance. It is shown that the instability places a cap on ideal RDEperformance, but that an optimized exhaust throat and inlet restriction combination still yields substantial pressure gain. Other parametric sensitivities are also examined in terms ofinstability growth. These include axial length, inlet manifold pressure, and air-fuel ratio.

Description: A fundamental exploratory experiment is conducted assessing the performance of a one-sided ejector with the eventual goal of noise reduction for jet engines. The hardware is comprised of an 8:1 rectangular nozzle together with an ejector box whose lower surface is flush with the lower lip of the nozzle. Secondary flow is allowed through a gap between the upper lip of the nozzle and a flap that constitutes the upper surface of the ejector. Wall static pressures and Pitot probe surveys are conducted to evaluate the performance of the ejector with variation of geometric parameters. It is found that addition of vortex generating tabs at the upper lip of the nozzle significantly increases secondary flow entrainment. The entrainment is further enhanced by a divergence of the ejector up-per surface. Limited noise measurements are done. The baseline ejector (without tabs) often encounters flow resonance with accompanying tones. The tabs have the additional benefit of eliminating those tones in all cases. However, for the tabbed case, addition of the ejector produces insignificant further noise reduction. This is due to the fact that the flow remains unmixed on the lower half of the ejector. The focus of ongoing and future efforts is to achieve sufficient mixing of the flow so that the exhaust velocities are uniformly low, while keeping the ejector hardware short and light-weight.

Description: This paper describes the design of a turboshaft engine for a tiltwing air taxi application. In this case, the tiltwing air taxi is intended to fly a 400-nautical mile mission with up to fifteen passengers. Engine requirements for the concept engine are taken from aircraft system studies where thrust is produced by four propellers driven by electric motors and powered by a single gas turbine engine. The purpose of this paper is to perform a cycle design optimization that minimizes fuel consumption and weight while respecting current technology limitations to meet mission requirements. To achieve results, the engine overall pressure ratio and maximum temperature at the exit of the combustor are set as the design parameters. Several sensitivity studies are also performed to visualize optimization trends. Results of the optimization study show solutions are heavily dependent on engine cooling flow requirements and exact mission requirements. This engine is intended for use in large system optimization research.