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NASA NTRS2009

Large Eddy Simulation of Transonic Flow Field in NASA Rotor 37

The current paper reports on numerical investigations on the flow characteristics in a transonic axial compressor, NASA Rotor 37. The flow field was used previously as a CFD blind test case conducted by American Society of Mechanical Engineers in 1994. Since the CFD blind-test exercise, many numerical studies on the flow field in the NASA Rotor 37 have been reported. Although steady improvements have been reported in both numerical procedure and turbulence closure, it is believed that all the important aspects of the flow field have not been fully explained with numerical studies based on the Reynolds Averaged Navier-Stokes (RANS) solution. Experimental data show large dip in total pressure distribution near the hub at downstream of the rotor at 100% rotor speed. Most original numerical solutions from the blind test exercise did not predict this total pressure deficit correctly. This total pressure deficit at the rotor exit was attributed to a hub corner flow separation by the author. Several subsequent numerical studies with different turbulence closure model also calculated this dip in total pressure rise. Also, several studies attributed this total pressure deficit to a small leakage flow coming from the hub in the test article. As the experimental study cannot be repeated, either explanation cannot be validated. The primary purpose of the current investigation is to investigate the transonic flow field with both RANS and a Large Eddy Simulation (LES). The RANS approach gives similar results presented at the original blind test exercise. Although the RANS calculates higher overall total pressure rise, the total pressure deficit near the hub is calculated correctly. The numerical solution shows that the total pressure deficit is due to a hub corner flow separation. The calculated pressure rise from the LES agrees better with the measured total pressure rise especially near the casing area where the passage shock interacts with the tip clearance vortex and flow becomes unsteady due to this interaction. The LES simulation also calculates the total pressure rise deficit near the hub and it agrees well with the measured data.

Hah, Chunill↗

NASA NTRS1960

Experimental and Theoretical Study of a Rectangular Wing in a Vortical Wake at Low Speed

A systematic study has been made, experimentally and theoretically, of the effects of a vortical wake on the aerodynamic characteristics of a rectangular wing at subsonic speed. The vortex generator and wing were mounted on a reflection plane to avoid body-wing interference. Vortex position, relative to the wing, was varied both in the spanwise direction and normal to the wing. Angle of attack of the wing was varied from -40 to +60. Both chordwise and spanwise pressure distributions were obtained with the wing in uniform and vortical flow fields. Stream surveys were made to determine the flow characteristics in the vortical wake. The vortex-induced lift was calculated by several theoretical methods including strip theory, reverse-flow theory, and reverse-flow theory including a finite vortex core. In addition, the Prandtl lifting-line theory and the Weissinger theory were used to calculate the spanwise distribution of vortex-induced loads. With reverse-flow theory, predictions of the interference lift were generally good, and with Weissinger's theory the agreement between the theoretical spanwise variation of induced load and the experimental variation was good. Results of the stream survey show that the vortex generated by a lifting surface of rectangular plan form tends to trail back streamwise from the tip and does not approach the theoretical location, or centroid of circulation, given by theory. This discrepancy introduced errors in the prediction of vortex interference, especially when the vortex core passed immediately outboard of the wing tip. The wake produced by the vortex generator in these tests was not fully rolled up into a circular vortex, and so lacked symmetry in the vertical direction of the transverse plane. It was found that the direction of circulation affected the induced loads on the wing either when the wing was at angle of attack or when the vortex was some distance away from the plane of the wing.

Smith, Willard G.↗

DOE OSTI2016

Advances in Decommissioning Using Laser Cutting - 16035

This paper describes advancements made in laser cutting in the UK Collaborative Project 'LaserSnake2'. In particular it will focus on the use of special laser cutting heads and optics to enhance cutting performance, whilst at the same time, reducing the effects of the residual laser beam passing through the cut. The benefits of using lasers for applications in decommissioning include the high speeds available, the tolerance of the process, the lightness of the cutting head, the lack of a reaction force with the part being cut and the ease of automation of the laser cutting operation. Laser cutting, however, is a thermal process and a potential detriment, is that the residual laser beam passing through the kerf, might damage or indeed set fire to, something positioned behind the part being cut. This paper describes the use of a diffractive optical element in the laser beam forming optics, designed to extend the depth of focus of the system, without increasing the physical focal length of the optic. In this way, for cutting thick materials, the goal is to achieve the cutting performance of a long focal length lens, with the beam divergence of a short focal length lens. A design of diffractive optical element is presented, which when used with a laser power of 5 kW, in conjunction with a set of conventional beam forming optics, including a 250 mm focal length focusing lens, is able to cut 40 mm thick C-Mn steel plate at 50 mm/min, while varying the nozzle tip to plate stand-off distance by 100 mm. For the same laser power in the focusing beam, this performance can be equaled by removing the diffractive optic and 250 mm lens and replacing these with just a 500 mm lens. However, when using the diffractive optical element, the power density in the residual beam has been reduced by a factor four when compared to that from the optical arrangement using the 500 mm lens. As well as the cutting performance, the effect of the two different power density residual beams on concrete will be presented. (authors)

12 MANAGEMENT OF RADIOACTIVE AND NON-RADIOACTIVE W↗

NASA NTRS1992

RTJ-303: Variable geometry, oblique wing supersonic aircraft

This document is a preliminary design of a High Speed Civil Transport (HSCT) named the RTJ-303. It is a 300 passenger, Mach 1.6 transport with a range of 5000 nautical miles. It features four mixed-flow turbofan engines, variable geometry oblique wing, with conventional tail-aft control surfaces. The preliminary cost analysis for a production of 300 aircraft shows that flyaway cost would be 183 million dollars (1992) per aircraft. The aircraft uses standard jet fuel and requires no special materials to handle aerodynamic heating in flight because the stagnation temperatures are approximately 130 degrees Fahrenheit in the supersonic cruise condition. It should be stressed that this aircraft could be built with today's technology and does not rely on vague and uncertain assumptions of technology advances. Included in this report are sections discussing the details of the preliminary design sequence including the mission to be performed, operational and performance constraints, the aircraft configuration and the tradeoffs of the final choice, wing design, a detailed fuselage design, empennage design, sizing of tail geometry, and selection of control surfaces, a discussion on propulsion system/inlet choice and their position on the aircraft, landing gear design including a look at tire selection, tip-over criterion, pavement loading, and retraction kinematics, structures design including load determination, and materials selection, aircraft performance, a look at stability and handling qualities, systems layout including location of key components, operations requirements maintenance characteristics, a preliminary cost analysis, and conclusions made regarding the design, and recommendations for further study.

Antaran, Albert↗

NASA NTRS1987

Low-cost FM oscillator for capacitance type of blade tip clearance measurement system

The frequency-modulated (FM) oscillator described is part of a blade tip clearance measurement system that meets the needs of a wide class of fans, compressors, and turbines. As a result of advancements in the technology of ultra-high-frequency operational amplifiers, the FM oscillator requires only a single low-cost integrated circuit. Its carrier frequency is 42.8 MHz when it is used with an integrated probe and connecting cable assembly consisting of a 0.81 cm diameter engine-mounted capacitance probe and a 61 cm long hermetically sealed coaxial cable. A complete circuit analysis is given, including amplifier negative resistance characteristics. An error analysis of environmentally induced effects is also derived, and an error-correcting technique is proposed. The oscillator can be calibrated in the static mode and has a negative peak frequency deviation of 400 kHz for a rotor blade thickness of 1.2 mm. High-temperature performance tests of the probe and 13 cm of the adjacent cable show good accuracy up to 600 C, the maximum permissible seal temperature. The major source of error is the residual FM oscillator noise, which produces a clearance error of + or - 10 microns at a clearance of 0.5 mm. The oscillator electronics accommodates the high rotor speeds associated with small engines, the signals from which may have frequency components as high as 1 MHz.

Barranger, John P.↗

DOE OSTI2014

SU-D-BRF-06: A Brachytherapy Simulator with Realistic Haptic Force Feedback and Real-Time Ultrasounds Image Simulation for Training and Teaching

Purpose: Surgical procedures require dexterity, expertise and repetition to reach optimal patient outcomes. However, efficient training opportunities are usually limited. This work presents a simulator system with realistic haptic force-feedback and full, real-time ultrasounds image simulation. Methods: The simulator is composed of a custom-made Linear-DELTA force-feedback robotic platform. The needle tip is mounted on a force gauge at the end effector of the robot, which responds to needle insertion by providing reaction forces. 3D geometry of the tissue is using a tetrahedral finite element mesh (FEM) mimicking tissue properties. As the needle is inserted/retracted, tissue deformation is computed using a mass-tensor nonlinear visco-elastic FEM. The real-time deformation is fed to the L-DELTA to take into account the force imparted to the needle, providing feedback to the end-user when crossing tissue boundaries or needle bending. Real-time 2D US image is also generated synchronously showing anatomy, needle insertion and tissue deformation. The simulator is running on an Intel I7 6- core CPU at 3.26 MHz. 3D tissue rendering and ultrasound display are performed on a Windows 7 computer; the FEM computation and L-DELTA control are executed on a similar PC using the Neutrino real-time OS. Both machines communicate through an Ethernet link. Results: The system runs at 500 Hz for a 8333-tetrahedron tissue mesh and a 100-node angular spring needle model. This frame rate ensures a relatively smooth displacement of the needle when pushed or retracted (±20 N in all directions at speeds of up to 2 m/s). Unlike commercially-available haptic platforms, the oblong workspace of the L-DELTA robot complies with that required for brachytherapy needle displacements of 0.1m by 0.1m by 0.25m. Conclusion: We have demonstrated a real-life, realistic brachytherapy simulator developed for prostate implants (LDR/HDR). The platform could be adapted to other sites or training for other types of needle-based procedures.

62 RADIOLOGY AND NUCLEAR MEDICINE↗

DOE OSTI2019

Predictive LES Modeling and Validation of High-Pressure Turbulent Flames and Flashback in Hydrogen-enriched Gas Turbines

The overall objective of this project is to make advances in large-eddy simulation modeling of swirl-flame flashback at elevated pressure, where the reactants can be premixed or stratified, and composed of methane enriched with hydrogen. The modeling effort was conducted with a companion experimental validation effort that focused on high-speed measurements of flashback in a model swirl-flame combustor. The modeling effort focused on the development of manifold-based combustion model for flashback simulations. The main complexity in this configuration comes from the flames’ propagation through a stratified fuel-air mixture along with heat loss near the walls. For this purpose, a non-adiabatic flamelet model was developed and tested using canonical flow configurations. Further, extensive validation in premixed configurations is used to establish the model’s accuracy in capturing heat loss and flame propagation. Finally, the model was tested using the UT flame flashback experiment, and was shown to accurately capture the flame flashback process. The end products from this work are a) a full validated flameout-based approach for modeling flame flashback; b) a complete implementation of the model in the open source OpenFOAM framework, which has been made available to both industry and academia; c) an extensive suite of validation studies performed using the OpenFOAM solver. In addition, new techniques for flame modeling, subfilter modeling of scalars, and simulation of low Mach number flows were developed and tested. In the experimental validation effort, flashback was studied in a premixed and stratified swirl flame by using high-speed imaging diagnostics. Fuels tested included CH 4 enriched with H 2 , with H 2 enrichment up to 84% by volume. These flashback experiments are conducted at pressures ranging from 1 to 5 atm. A swirler-based fuel-injection system was used to create the fuel-air stratification in the radial direction. For the high-pressure measurements, an optically accessible elevated pressure chamber was used. The spatial distribution of the equivalence ratio under non-reacting conditions was characterized using planar laser-induced fluorescence with acetone as the fuel tracer. It was observed that fuel-air pockets were distributed across the mixing tube width, although in an average sense, the fuel-air mixture was radially stratified. The global behavior of upstream flame propagation was investigated for different levels of hydrogen-enrichment. For stratified hydrogen-rich flashback, the propagation path of the flame changed from the inner wall to outer wall, which was induced by the faster chemistry of stoichiometric mixtures that are frequently present near the outer wall. This behavior of hydrogen-rich flashback persists even at elevated pressures up to 5 atm, although the propagation of the flame occurs as a wide flame tongue as opposed to the acute-tipped flame structures present in the atmospheric cases. The end product of the experimental effort is greatly improved knowledge of flashback physics and a unique database that can be used for model validation.

08 HYDROGEN↗

DOE OSTI2024

In-vivo Raman microspectroscopy reveals differential nitrate concentration in different developmental zones in Arabidopsis roots

Abstract Background Nitrate (NO 3 − ) is one of the two major forms of inorganic nitrogen absorbed by plant roots, and the tissue nitrate concentration in roots is considered important for optimizing developmental programs. Technologies to quantify the expression levels of nitrate transporters and assimilating enzymes at the cellular level have improved drastically in the past decade. However, a technological gap remains for detecting nitrate at a high spatial resolution. Using extraction-based methods, it is challenging to reliably estimate nitrate concentration from a small volume of cells (i.e., with high spatial resolution), since targeting a small or specific group of cells is physically difficult. Alternatively, nitrate detection with microelectrodes offers subcellular resolution with high cell specificity, but this method has some limitations on cell accessibility and detection speed. Finally, optical nitrate biosensors have very good ( in-vivo ) sensitivity (below 1 mM) and cellular-level spatial resolution, but require plant transformation, limiting their applicability. In this work, we apply Raman microspectroscopy for high-dynamic range in-vivo mapping of nitrate in different developmental zones of Arabidopsis thaliana roots in-situ . Results As a proof of concept, we have used Raman microspectroscopy for in-vivo mapping of nitrate content in roots of Arabidopsis seedlings grown on agar media with different nitrate concentrations. Our results revealed that the root nitrate concentration increases gradually from the meristematic zone (~ 250 µm from the root cap) to the maturation zone (~ 3 mm from the root cap) in roots grown under typical growth conditions used for Arabidopsis, a trend that has not been previously reported. This trend was observed for plants grown in agar media with different nitrate concentrations (0.5–10 mM). These results were validated through destructive measurement of nitrate concentration. Conclusions We present a methodology based on Raman microspectroscopy for in-vivo label-free mapping of nitrate within small root tissue volumes in Arabidopsis. Measurements are done in-situ without additional sample preparation. Our measurements revealed nitrate concentration changes from lower to higher concentration from tip to mature root tissue. Accumulation of nitrate in the maturation zone tissue shows a saturation behavior. The presented Raman-based approach allows for in-situ non-destructive measurements of Raman-active compounds.

Fernández González, Alma↗

DOE OSTI2022

Development of Explainable Data-Driven Turbulence Models with Application to Liquid Fuel Nuclear Reactors

Liquid fuel nuclear reactors offer innovative possibilities in terms of nuclear reactor designs and passive safety systems. Molten Salts Reactors (MSRs) with a fast spectrum are a particular type of these reactors using liquid fuel. MSFRs often involve large open cavities in their core in which the liquid fuel circulates at a high speed to transport the heat generated by the nuclear reactions into the heat exchangers. This high-speed flow yields a turbulent field with large Reynolds numbers in the reactor core. Since the nuclear power, the neutron precursor’s transport and the thermal exchanges are strongly coupled in the MSFR’s core cavity, having accurate turbulent models for the liquid fuel flow is necessary to avoid introducing significant errors in the numerical simulations of these reactors. Nonetheless, high-accuracy simulations of the turbulent flow field in the reactor cavity of these reactors are usually prohibitively expensive in terms of computational resources, especially when performing multiphysics numerical calculations. Therefore, in this work, we propose a novel method using a modified genetic algorithm to optimize the calculation of the Reynolds Shear Stress Tensor (RST) used for turbulence modeling. The proposed optimization methodology is particularly suitable for advanced liquid fuel reactors such as the MSFRs since it allows the development of high-accuracy but still low-computational-cost turbulence models for the liquid fuel. We demonstrate the applicability of this approach by developing high accuracy Reynolds-Averaged Navier–Stokes (RANS) models (averaged flow error less than 5%) for a low and a large aspect ratio in a Backward-Facing Step (BFS) section particularly challenging for RANS models. The newly developed turbulence models better capture the flow field after the boundary layer tipping, over the extent of the recirculation bubble, and near the boundary layer reattachment region in both BFS configurations. The main reason for these improvements is that the developed models better capture the flow field turbulent anisotropy in the bulk region of the BFS. Then, we illustrate the interest in using this turbulence modeling approach for the case of an MSFR by quantifying the impact of the turbulence modeling on the reactor key parameters.

42 ENGINEERING↗

NASA NTRS1952

Altitude Investigation of Gas Temperature Distribution at Turbine of Three Similar Axial-Flow Turbojet Engines

An investigation of the effect of inlet pressure, corrected engine speed, and turbine temperature level on turbine-inlet gas temperature distributions was conducted on a J40-WE-6, interim J40-WE-6, and prototype J40-WE-8 turbojet engine in the altitude wind tunnel at the NAC.4 Lewis laboratory. The engines were investigated over a range of simulated pressure altitudes from 15,000 to 55,000 feet, flight Mach numbers from 0.12 to 0.64, and corrected engine speeds from 7198 to 8026 rpm, The gas temperature distribution at the turbine of the three engines over the range of operating conditions investigated was considered satisfactory from the standpoint of desired temperature distribution with one exception - the distribution for the J40-WE-6 engine indicated a trend with decreasing engine-inlet pressure for the temperature to exceed the desired in the region of the blade hub. Installation of a compressor-outlet mixer vane assembly remedied this undesirable temperature distribution, The experimental data have shown that turbine-inlet temperature distributions are influenced in the expected manner by changes in compressor-outlet pressure or mass-flow distribution and by changes in combustor hole-area distribution. The similarity between turbine-inlet and turbine-outlet temperature distribution indicated only a small shift in temperature distribution imposed by the turbine rotors. The attainable jet thrusts of the three engines were influenced in different degrees and directions by changes in temperature distributions with change in engine-inlet pressure. Inability to match the desired temperature distribution resulted, for the J40-WE-6 engine, in an 11-percent thrust loss based on an average turbine-inlet temperature of 1500 F at an engine-inlet pressure of 500 pounds per square foot absolute. Departure from the desired temperature distribution in the Slade tip region results, for the prototype J40-WE-8 engine, in an attainable thrust increase of 3 to 4 percent as compared with that obtained if tip-region temperature limitations were observed.

Prince, W.R.↗

NASA NTRS1994

Special relativity effects for space-based coherent lidar experiments

There is a great need to develop a system that can measure accurately atmospheric wind profiles because an accurate data of wind profiles in the atmosphere constitutes single most input for reliable simulations of global climate numerical methods. Also such data helps us understand atmospheric circulation and climate dynamics better. Because of this need for accurate wind measurements, a space-based Laser Atmospheric Winds Sounder (LAWS) is being designed at MSFC to measure wind profiles in the lower atmosphere of the earth with an accuracy of 1 m/s at lower altitudes to 5m/s at higher altitudes. This system uses an orbiting spacecraft with a pulsed laser source and measures the Doppler shift between the transmitted and received frequencies to estimate the atmospheric wind velocities. If a significant return from the ground (sea) is possible, the spacecraft speed and height are estimated from it and these results and the Doppler shift are then used to estimate the wind velocities in the atmosphere. It is expected that at the proposed wavelengths, there will be enough backscatter from the aerosols but there may no be significant return from the ground. So a coherent (heterodyne) detection system is being proposed for signal processing because it can provide high signal to noise ratio and sensitivity and thus make the best use of low ground return. However, for a heterodyne detection scheme to provide the best results, it is important that the receiving aperture be aligned properly for the proposed wind sounder, this amounts to only a few microradians tolerance in alignment. It is suspected that the satellite motion relative to the ground may introduce errors in the order of a few microradians because of special relativity. Hence, the problem of laser scattering off a moving fixed target when the source and receiver are moving, which was not treated in the past in the literature, was analyzed in the following, using relativistic electrodynamics and applied to the case of the space-based coherent lidar, assuming flat ground. Here an interest in developing analytical expression for the location of the receiving point for the return with respect to the satellite, receiving angle and Doppler shift in frequency and amount of tip, all as measured in the satellite moving coordinate system and the diffuse scattering angle at the ground which does not require any compensation. All the three cases of retro-reflection, specular reflection and diffuse scattering by the ground should be treated though retro-reflection and diffuse scattering are more important.

Raogudimetla, V. S.↗

DOE OSTI2021

Life cycle of streaks in the buffer layer of wall-bounded turbulence

In this work, streaks in the buffer layer of wall-bounded turbulence are tracked in time to study their life cycle. Spatially and temporally resolved direct numerical simulation data are used to analyze the strong wall-parallel movements conditioned to low-speed streamwise flow. The analysis of the streaks shows that there is a clear distinction between wall-attached and detached streaks, and that the wall-attached streaks can be further categorized into streaks that are contained in the buffer layer and the ones that reach the outer region. The results reveal that streaks are born in the buffer layer, coalescing with each other to create larger streaks that are still attached to the wall. Once the streak becomes large enough, it starts to meander due to the large streamwise-to-wall-normal aspect ratio, and consequently the elongation in the streamwise direction, which makes it more difficult for the streak to be oriented strictly in the streamwise direction. While the continuous interaction of the streaks allows the superstructure to span extremely long temporal and length scales, individual streak components are relatively small and short-lived. Tall-attached streaks eventually split into wall-attached and wall-detached components. These wall-detached streaks have a strong wall-normal velocity away from the wall, similar to ejections or bursts observed in the literature. Conditionally averaging the flow fields to these split events show that the detached streak has not only a larger wall-normal velocity compared to the wall-attached counterpart, it also has a larger (less negative) streamwise velocity, similar to the velocity field at the tip of a vortex cluster.

42 ENGINEERING↗

NASA NTRS1949

The Static-Pressure Error of a Wing Airspeed Installation of the McDonnell XF-88 Airplane in Dives to Transonic Speeds

Measurements were made, in dives to transonic speeds, of the static-pressure position error at a distance of one chord ahead of the McDonnell XF-88 airplane. The airplane incorporates a wing which is swept back 35 deg along the 0.22 chord line and utilizes a 65-series airfoil with a 9-percent-thick section perpendicular to the 0.25-chord line. The section in the stream direction is approximately 8-percent thick. Data up to a Mach number of about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Data at Mach numbers above about 0.97 were obtained within an airplane normal-force-coefficient range from about 0.05 to about 0.68. Results of the measurements indicate that the static-pressure error, within the accuracy of measurement, is negligible from a Mach number of 0.65 to a Mach number of about 0.97. With a further increase in Mach number, the static-pressure error increases rapidly; at the highest Mach number attained in these tests (about M = 1.038), the error increases to about 8 percent of the impact pressure. Above a Mach number of about 0.975, the recorded Mach number remains substantially constant with increasing true Mach number; the installation is of no value between a Mach number of about 0.975 and at least 1.038, as the true Mach number cannot be obtained from the recorded Mach number in this range. Previously published data have shown that at 0.96 chord ahead of the wing tip of the straight-wing X-l airplanes, a rapid rise of position error started at a Mach number of about 0.8. In the case of the XF-88 airplane, this rise of position error was delayed, presumably by the sweep of the wing, to a Mach number of about 0.97.

Goodman, Harold R.↗

NASA NTRS2017

Impact of Wake Dispersion on Axial Compressor Performance

Rotor wake dispersion in a low-speed, one and half stage axial compressor is investigated in detail with a Large Eddy Simulation (LES). The primary focus is to quantify the total pressure recovery due to wake stretching and the total pressure loss from the rotor wake interaction with the stator blade boundary layer. The relative magnitude of the aerodynamic loss due to these two effects is examined at several radial locations. The spacing between the rotor and the stator was varied from 29% to 112% of the rotor axial chord at the mid span to investigate the effects of rotor wake decay before entering the stator passage. The current analysis indicates that the efficiency through the compressor stage is increased about 0.5% when the spacing between the rotor and the stator is decreased from 112% to 29% of the rotor axial chord at mid-span. 22% of the efficiency gain from the narrower axial gap is due to the wake recovery and 63% is due to the stronger unsteady pressure field at the exit of the rotor due to stage interaction. Total pressure loss/recovery across the stator varies significantly in the radial direction for the current compressor, which has a much lower aspect ratio. The total pressure recovery due to wake stretching is larger than the total pressure loss due to the unsteady boundary layer development on the stator blade from 20% to 35% of the span from the hub for 29% spacing and from 35% to 55% of the span for 112% spacing. Above 50% of the span, rotor tip clearance flow affects wake dispersion and the overall wake recovery is less than expected.

Axial Compressor↗

NASA NTRS2012

Implementation of a Transition Model in a NASA Code and Validation Using Heat Transfer Data on a Turbine Blade

The purpose of this report is to summarize and document the work done to enable a NASA CFD code to model laminar-turbulent transition process on an isolated turbine blade. The ultimate purpose of the present work is to down-select a transition model that would allow the flow simulation of a variable speed power turbine to be accurately performed. The flow modeling in its final form will account for the blade row interactions and their effects on transition which would lead to accurate accounting for losses. The present work only concerns itself with steady flows of variable inlet turbulence. The low Reynolds number k- model of Wilcox and a modified version of the same model will be used for modeling of transition on experimentally measured blade pressure and heat transfer. It will be shown that the k- model and its modified variant fail to simulate the transition with any degree of accuracy. A case is thus made for the adoption of more accurate transition models. Three-equation models based on the work of Mayle on Laminar Kinetic Energy were explored. The three-equation model of Walters and Leylek was thought to be in a relatively mature state of development and was implemented in the Glenn-HT code. Two-dimensional heat transfer predictions of flat plate flow and two-dimensional and three-dimensional heat transfer predictions on a turbine blade were performed and reported herein. Surface heat transfer rate serves as sensitive indicator of transition. With the newly implemented model, it was shown that the simulation of transition process is much improved over the baseline k- model for the single Reynolds number and pressure ratio attempted; while agreement with heat transfer data became more satisfactory. Armed with the new transition model, total-pressure losses of computed three-dimensional flow of E3 tip section cascade were compared to the experimental data for a range of incidence angles. The results obtained, form a partial loss bucket for the chosen blade. In time the loss bucket will be populated with losses at additional incidences. Results obtained thus far will be discussed herein.

Ameri, Ali A.↗

DOE OSTI2017

Mapping of exhaust gas recirculation and combustion-residual backflow in the intake ports of a heavy-duty diesel engine using a multiplexed multi-species absorption spectroscopy sensor [Mapping of EGR and combustion residual backflow in the intake ports of a heavy duty diesel engine using a multiplexed multi-species absorption spectroscopy sensor]

Here, the combustion-residual backflow into the intake ports of a commercial diesel engine (Cummins ISX series) was spatiotemporally mapped using a multiplexed multi-species absorption spectroscopy sensor system; the resulting cycle- and cylinder-resolved measurements are applicable for assessing cylinder charge uniformity, control strategies, and computational fluid dynamics tools. On-engine measurements were made using four compact (3/8 in Outside Diameter) stainless steel probes which enabled simultaneous multi-point measurements, required minimal engine hardware modification, and featured a novel tip design for measurement of gas flows parallel to the probe axis. Three sensor probes were used to perform simultaneous backflow measurements in intake runners corresponding to three of the six engine cylinders, and a fourth probe was installed in the intake manifold plenum for tracking dynamics introduced by an external exhaust gas recirculation mixer. Near-crank-angle resolved measurements (5 kHz, that is, 1.2 crank angle resolution at 1000 RPM) were performed during steady-state engine operation at various levels of external exhaust gas recirculation to measure the gas properties and penetration distance of the backflow into the intake runners on a cylinder- and cycle-basis. Validation of computational fluid dynamics model results is also presented to demonstrate the utility of such measurements in advancing engine research.

33 ADVANCED PROPULSION SYSTEMS↗

DOE OSTI2016

Eulerian frequency analysis of structural vibrations from high-speed video

An approach for the analysis of the frequency content of structural vibrations from high-speed video recordings is proposed. The techniques and tools proposed rely on an Eulerian approach, that is, using the time history of pixels independently to analyse structural motion, as opposed to Lagrangian approaches, where the motion of the structure is tracked in time. The starting point is an existing Eulerian motion magnification method, which consists in decomposing the video frames into a set of spatial scales through a so-called Laplacian pyramid [1]. Each scale — or level — can be amplified independently to reconstruct a magnified motion of the observed structure. The approach proposed here provides two analysis tools or pre-amplification steps. The first tool provides a representation of the global frequency content of a video per pyramid level. This may be further enhanced by applying an angular filter in the spatial frequency domain to each frame of the video before the Laplacian pyramid decomposition, which allows for the identification of the frequency content of the structural vibrations in a particular direction of space. This proposed tool complements the existing Eulerian magnification method by amplifying selectively the levels containing relevant motion information with respect to their frequency content. This magnifies the displacement while limiting the noise contribution. The second tool is a holographic representation of the frequency content of a vibrating structure, yielding a map of the predominant frequency components across the structure. In contrast to the global frequency content representation of the video, this tool provides a local analysis of the periodic gray scale intensity changes of the frame in order to identify the vibrating parts of the structure and their main frequencies. Validation cases are provided and the advantages and limits of the approaches are discussed. The first validation case consists of the frequency content retrieval of the tip of a shaker, excited at selected fixed frequencies. The goal of this setup is to retrieve the frequencies at which the tip is excited. The second validation case consists of two thin metal beams connected to a randomly excited bar. It is shown that the holographic representation visually highlights the predominant frequency content of each pixel and locates the global frequencies of the motion, thus retrieving the natural frequencies for each beam.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

NASA NTRS2022

Computational Analysis of the X-57 Maxwell Airplane at Unpowered Conditions (Preliminary Fuselage)

The X-57 Maxwell is an all-electric airplane that implements a distributed electric propulsion system to demonstrate that high-efficiency electric propulsion can be integrated with aerodynamics to increase the performance of an airplane. To this end, distributed electric fans were installed on the wing to provide increased flow over the wing at the low takeoff and landing speeds of the X-57. The low-speed lift augmentation allows for a reduction in wing area for cruise optimization. The X-57 wing area was reduced to 42 percent of the wing area of the baseline aircraft, a Tecnam P2006T. With this reduced wing area and the electric propulsion system, it is estimated that the X-57 will cruise on less than one-third the total energy compared to the baseline aircraft. To meet the cruise performance goal at a Mach number of 0.233 at an altitude of 8000 feet, the X-57 has a cruise lift coefficient of 0.7516 and needs to have a cruise drag coefficient of 0.05423 or less. The USM3D computational solver was used to investigate the X-57 performance, without the distributed electric propulsion high-lift system operating. The unpowered X-57 performance is of interest to quantify if the X-57 can meet the cruise drag performance goal, and to document the lift performance of the very small wing at takeoff and landing conditions. The primary configurations investigated in this paper include the cruise configuration with no flap deflection, a takeoff configuration with a 10◦ flap deflection, and a landing configuration with a 30◦ flap deflection. The conditions for the cruise configuration were a flight unit Reynolds number of 1.32E+06 per foot, an altitude of 8000 feet, a Mach number of 0.233, and angles of attack from −2° to 24° . At the cruise lift coefficient of 0.7516, the computed drag coefficient is 0.05275. This computed drag is less than the drag coefficient of 0.05423 that is required to meet the X-57 airplane performance goal. However, the computational airplane is a completely smooth geometry and does not account for protuberance drag, nor the drag from steps and gaps in the actual X-57 airplane. Therefore, based upon the CFD drag calculation there is a 10-percent margin to account for some of the differences between the as-built metal fuselage and empennage construction, and the smooth computational geometry. The computed cruise drag also does not account for an induced drag reduction due to the wing-tip propellers and a drag reduction due to laminar flow achieved on the wing. The computed lift to drag ratio is 14.14 at the cruise lift coefficient of 0.7516, and the maximum computed lift to drag ratio is 15.8. The maximum lift coefficient for the cruise configuration was 2.13 at an angle of attack of 15°. The conditions for the takeoff configuration with a 10° flap deflection were a flight unit Reynolds number of 0.986E+06 per foot, an altitude of 2500 feet, a Mach number of 0.149, and angles of attack from −2° to 22°. The maximum lift coefficient for the takeoff configuration was 2.21 at an angle of attack of 16°. The conditions for the landing configuration with a 30° flap deflection were a flight unit Reynolds number of 0.922E+06 per foot, an altitude of 2500 feet, a Mach number of 0.139, and angles of attack from −2° to 24°. The maximum lift coefficient for the landing configuration was 2.58 and occurred at two angles of attack, 10° and 14°. Based on the unpowered maximum lift coefficient of 2.58 for the 30° flap deflection, along with computations of the distributed electric propulsion lift augmentation (not shown in this paper), the X-57 Maxwell is estimated to meet its powered landing goal of a maximum lift coefficient of 4.0.

X-57 Maxwell Airplane↗