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NASA NTRS2021

Oblique High Mach Number Heliospheric Shocks: The Role of α Particles

Spacecraft observations of heliospheric shocks often find oscillations in the magnetic field and density both, upstream and downstream. The downstream magnetic oscillations of oblique collisionless shocks were detected by Wind with 10.9 samples/s and DSCOVR spacecraft with high temporal resolution of 50 samples/s. The density oscillations associated with the shocks are also evident in proton and α particle density by Wind (with much lower temporal resolution). Recently, we have investigated low Mach number low-β oblique shock oscillations using satellite data and 2.5D hybrid modeling with electrons modeled as fluid and ions modeled as particles and found that α particles—an important component of heliospheric plasma—may affect considerably the downstream oscillations and the shock structure. The objective of the present study is to investigate the effects of α particles on high Mach number heliospheric shocks dynamics, oscillations, nonstationarity, and shock-front rippling. We extend the study to high Mach number shocks (M > 3), investigate several α particle typical densities, and compare the results for the various shock parameters. We model the effects of α particles on the shock ramp, wake, and downstream oscillation structure and the kinetic properties of proton and α particle velocity distributions at various locations downstream of the shocks. Using the 2.5D hybrid model we found that the modeled high Mach number quasi-perpendicular shock magnetic and density structures are significantly affected by α particles with typical solar wind relative abundances, suggesting that the observed high Mach number shocks are similarly affected by α particles.

Leon Ofman↗

NASA NTRSDate not supplied

A Comparison of Rotor Disk Modeling and Blade-Resolved CFD Simulations for NASA's Tiltwing Air Taxi

A multi-fidelity computational fluid dynamics analysis is carried out for NASA’s tiltwing air taxi concept operating in airplane and helicopter mode. High-fidelity simulations are computationally expensive due to individual rotor blade modeling in a time-dependent computational domain with rotating grids. The mid-fidelity rotor disk option, in its source term implementation, is explored as a more affordable alternative. Computations are performed with NASA’s OVERFLOW flow solver loosely-coupled with the comprehensive code CAMRAD II for appropriate rotor trim. Detailed comparisons are shown for the trim solution, airloads, wake geometry, and rotor performance. While the rotor disk model is able to capture the flow field with satisfactory agreement in airplane mode, it faces difficulties in helicopter mode due to the three-dimensional effects of the wake. Although this study is limited to a specific vehicle geometry, it is expected that the results are somewhat generalizable to the analysis of multi-rotor configurations.

ARMD↗

NASA NTRS1990

Rapid mix concepts for low emission combustors in gas turbine engines

NASA LeRC has identified the Rich burn/Quick mix/Lean burn (RQL) combustor as a potential gas turbine combustor concept to reduce NOx emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NOx levels, NASA LeRC soon will test a flametube version of an RQL combustor. The critical technology needed for the RQL combustor is a method of quickly mixing combustion air with rich burn gases. Two concepts were proposed to enhance jet mixing in a circular cross-section: the Asymmetric Jet Penetration (AJP) concept; and the Lobed Mixer (LM) concept. In Phase 1, two preliminary configurations of the AJP concept were compared with a conventional 12-jet radial-inflow slot design. The configurations were screened using an advanced 3-D Computational Fluid Dynamics (CFD) code named REFLEQS. Both non-reacting and reacting analyses were performed. For an objective comparison, the conventional design was optimized by parametric variation of the jet-to-mainstream momentum flux (J) ratio. The optimum J was then employed in the AJP simulations. Results showed that the three-jet AJP configuration was superior in overall mixedness compared to the conventional design. However, in regards to NOx emissions, the AJP configuration was inferior. The higher emission level for AJP was caused by a single hot spot located in the wake of the central jet as it entered the combustor. Ways of maintaining good mixedness while eliminating the hot spot were identified for Phase 2 study. Overall, Phase 1 showed the viability of using CFD analyses to evaluate quick-mix concepts. A high probability exists that advancing mixing concepts will reduce NOx emissions in RQL combustors, and should be explored in Phase 2, by parallel numerical and experimental work.

Talpallikar, Milind V.↗

NASA NTRS1998

Grid Generation Techniques Utilizing the Volume Grid Manipulator

This paper presents grid generation techniques available in the Volume Grid Manipulation (VGM) code. The VGM code is designed to manipulate existing line, surface and volume grids to improve the quality of the data. It embodies an easy to read rich language of commands that enables such alterations as topology changes, grid adaption and smoothing. Additionally, the VGM code can be used to construct simplified straight lines, splines, and conic sections which are common curves used in the generation and manipulation of points, lines, surfaces and volumes (i.e., grid data). These simple geometric curves are essential in the construction of domain discretizations for computational fluid dynamic simulations. By comparison to previously established methods of generating these curves interactively, the VGM code provides control of slope continuity and grid point-to-point stretchings as well as quick changes in the controlling parameters. The VGM code offers the capability to couple the generation of these geometries with an extensive manipulation methodology in a scripting language. The scripting language allows parametric studies of a vehicle geometry to be efficiently performed to evaluate favorable trends in the design process. As examples of the powerful capabilities of the VGM code, a wake flow field domain will be appended to an existing X33 Venturestar volume grid; negative volumes resulting from grid expansions to enable flow field capture on a simple geometry, will be corrected; and geometrical changes to a vehicle component of the X33 Venturestar will be shown.

Alter, Stephen J.↗

NASA NTRSDate not supplied

Assessment of Turbulence Models for a Single-Injector Cooling Flow

This work assesses the ability of Reynolds-averaged Navier-Stokes turbulence models to predict a fundamental thermal cooling flow configuration. This geometry consists of a plate with a single cooling hole positioned downstream of a heated exhaust nozzle. The linear Shear Stress Transport (SST), non-linear Quadratic Constitutive Relation (QCR), and differential SSG/LRR Reynolds Stress Model (RSM) turbulence models were investigated for flow conditions with a blowing ratio of 1.0 and 2.0. Despite significant differences in turbulence model complexity, the results miss important details observed in the data and large eddy simulation (LES). Injection penetration is over-predicted. Mixing and centerline wake recovery is too slow. Temperature predictions are worse than velocity, suggesting that perhaps the simple gradient diffusion approximation used to model the turbulent heat flux may be to blame. These observations are put into context using other studies to identify shortcomings in the modeling of both the turbulent stress and turbulent heat flux.

Film Cooling↗

NASA NTRSDate not supplied

Assessment of Turbulence Models for a Single-Injector Cooling Flow

Film Cooling↗

NASA NTRS1994

The Roles of Large- and Small-Scale Structures in Turbulent Free Shear Layers

The occurrence of large-scale coherent structures in turbulent free shear flows (especially the planar mixing layer) has been recognized for some time. Indeed, the observation of such structures in mixing layers did much to promote interest in the study of coherent structures in turbulence. It has been widely assumed that the large-scale structures in these flows are responsible for the entrainment of free-stream fluid and the overall growth of the layer, while the small-scale structures provide mixing and dissipation. A model of scalar mixing based on these ideas was proposed for these flows. However, recent experimental and computational evidence suggests that the dominance of the large-scale structures in turbulent mixing layers is not universal. In addition, there is a substantial variation among experiments in several statistical measures of self-similar mixing layers, for example growth rate and velocity variances. To investigate the importance of large-scale structures, several free shear flows (mixing layers and wakes) have been simulated via direct numerical simulation. The simulations are designed to mimic experimental mixing layers in which the splitter plate boundary layers are turbulent. Different levels of two-dimensional forcing are included resulting in large-scale structures of differing strength and importance. These simulations are used to investigate the role of large-scale coherent structures in free shear layers and the effect of these structures on relevant turbulence statistics and scalar mixing. It is found that the statistics and structures in several experiments involving turbulent mixing layers are in better agreement with simulations that do not exhibit dominant large-scale structures than those in which the common mixing layer structures do dominate. It is also found that the level of forcing can have a profound effect on the qualitative and quantitative features of these shear layer, even when they are nominally self-similar.

Moser, Robert D.↗

NASA NTRS2015

Ocular Outcomes Comparison Between 14- and 70-day Head-down Tilt Bed Rest

INTRODUCTION: Ophthalmological changes, including optic disc edema with optic nerve sheath distension, posterior globe flattening with hyperopic shift, choroidal folds and cotton wool spots have been detected in some astronauts involved in long-duration spaceflights. (sup 1) It is hypothesized that elevated intracranial pressure resulting from microgravity-induced cephalad fluid shifts may be responsible for most of these findings. Head-down tilt bed rest (HTDBR) is a ground-based microgravity analog which also produces cephalad fluid shifts. It is conceivable that prolonged HDTBR exposure may induce ocular changes similar to those experienced in microgravity. Therefore, the purpose of the present study was to compare structural and functional ocular outcomes between 14- and 70-day HDTBR in healthy human subjects. It is hypothesized that 70-d HDTBR induced ocular changes of greater magnitude as compared to 14-d HDTBR. METHODS: Two HDTBR studies were conducted at the NASA Flight Analogs Research Unit, located at The University of Texas Medical Branch at Galveston, TX. Identical NASA standard screening procedures and BR conditions (e.g., strict sleep-wake cycle, standardized diet, continuous video monitoring) were implemented in both studies. Participants spent 14 and/or 70 consecutive days in a 6deg HDT position and did not engage in exercise. Subjects received weekly ocular examinations before, during, and after HDTBR. Ocular testing included: distance and near best-corrected visual acuity (BCVA), cycloplegic refraction, intraocular pressure (IOP) measurement, color vision, red dot test, modified Amsler grid test, confrontational visual field, color fundus photography and Spectral-domain OCT scans of the macula and the optic disc. Pre/post HDTBR differences between the two studies will be evaluated for BCVA, spherical equivalent, IOP, retinal nerve fiber layer (RNFL) thickness and macular OCT parameters. RESULTS: 16 (12 males and 4 females) and 6 (5 males and 1 female) subjects participated in the 14 and 70-day HDTBR studies, respectively. One subject participated in both studies. The demographic and ophthalmological characteristics of the 14-day HDTBR study have been presented elsewhere.2 In the 70- day study, subjects age averaged 39.5 +/- 7.8 years; pre- and post-HDTBR Goldmann IOP were, on average, 15.4 and 14.6 mmHg in the right eye, and 15.3 and 14.4 mmHg in the left eye, respectively; preand post-HDTBR Spectralis OCT (Heidelberg Engineering, GmbH, Heidelberg, Germany) average RNFL thickness were 103.8 and 106 μm in the right eye, and 103 and 104.3 μm in the left eye, respectively. In both studies, color vision, red dot test, modified Amsler grid test and confrontational visual field were within normal limits at all time points; no changes were detected on stereoscopic color optic disc photography. For the above outcomes, preliminary results of the pre/post HDTBR differences between the two studies will be presented at the 2015 Human Research Program Investigators' Workshop. CONCLUSIONS: Although 14-day HDTBR did not seem to induce clinically relevant ocular changes, (sup 2) a systematic evaluation and comparison with the 70-day HDTBR study results will elucidate whether the magnitude of ocular structural and functional changes is affected by the time spent in the recumbent position. Further research will also address the effects of integrated resistance and aerobic training during HDTBR on the ocular outcomes. REFERENCES: 1. Mader TH et al. Optic disc edema, globe flattening, choroidal folds, and hyperopic shifts observed in astronauts after long-duration space flight. Ophthalmology. 2011;118(10):2058-69. 2. Taibbi G et al. Ocular outcomes evaluation in a 14-day head-down bed rest study. Aviat Space Environ Med. 2014;85(10):983-92.

Cromwell, R.L.↗

NASA NTRS2001

Velocity Measurement in a Dual-Mode Supersonic Combustor using Particle Image Velocimetry

Temporally and spatially-resolved, two-component measurements of velocity in a supersonic hydrogen-air combustor are reported. The combustor had a single unswept ramp fuel injector and operated with an inlet Mach number of 2 and a flow total temperature approaching 1200 K. The experiment simulated the mixing and combustion processes of a dual-mode scramjet operating at a flight Mach number near 5. The velocity measurements were obtained by seeding the fuel with alumina particles and performing Particle Image Velocimetry on the mixing and combustion wake of the ramp injector. To assess the effects of combustion on the fuel air-mixing process, the distribution of time-averaged velocity and relative turbulence intensity was determined for the cases of fuel-air mixing and fuel-air reacting. Relative to the mixing case, the near field core velocity of the reacting fuel jet had a slower streamwise decay. In the far field, downstream of 4 to 6 ramp heights from the ramp base, the heat release of combustion resulted in decreased flow velocity and increased turbulence levels. The reacting measurements were also compared with a computational fluid dynamics solution of the flow field. Numerically predicted velocity magnitudes were higher than that measured and the jet penetration was lower.

Goyne, C. P.↗

NASA NTRSDate not supplied

Tiltwing Transition Flight Analysis Using High-Fidelity CFD

Transition from hover to forward flight and vice versa represents the most critical flight phase of tiltwing aircraft. Despite its importance to ensure a safe operation, the aerodynamics of this maneuver are not sufficiently understood. This paper focuses on the study of transition flight for NASA’s six-passenger tiltwing air taxi by means of high-fidelity computational fluid dynamics simulations. On the basis of a static trim solution, four points within the transition corridor are analyzed: transition mode at wing tilt angles of60and44, and airplane mode at airspeeds of110 ktand155 kt. We investigate the balance of forces and moments for rotor-borne and wing-borne regimes, and how rotor-on-rotor and rotor-on-wing interactions affect performance. The simulations indicate that during the early stages of transition, the vortices remain in close proximity to the proprotors, inducing large fluctuations on the order of the mean blade loading. Additionally, the blowing and swirling effects of the proprotor wakes delay flow separation over a portion of the wing. The mid-transition conditions appear to be critical, with extensive regions of separated flow over the wing. In airplane mode, proprotor-wing slipstream effects can be exploited to enhance lift generation. The results of this paper contribute to a deeper understanding of the complex aerodynamic interactions during transition flight to enable a safer and more efficient operation of tiltwing aircraft.

ARMD↗

NASA NTRS1991

Temporally and spatially resolved flow in a two-stage axial compressor. II - Computational assessment

Fluid dynamics of turbomachines are complicated due to aerodynamic interactions between rotors and stators. It is necessary to understand the aerodynamics associated with these interactions in order to design turbomachines that are both light and compact as well as reliable and efficient. The current study uses an unsteady, thin-layer Navier-Stokes zonal approach to investigate the unsteady aerodyamics of a multistage compressor. Relative motion between rotors and stators is made possible by use of systems of patched and overlaid grids. Results have been computed for a 2 1/2-stage compressor configuration. The numerical data compares well with experimental data for surface pressures and wake data. In addition, the effect of grid refinement on the solution is studied.

Gundy-Burlet, K. L.↗

NASA NTRS1990

Temporally and spatially resolved flow in a two-stage axial compressor. Part 2: Computational assessment

Fluid dynamics of turbomachines are complicated due to aerodynamic interactions between rotors and stators. It is necessary to understand the aerodynamics associated with these interactions in order to design turbomachines that are both light and compact as well as reliable and efficient. The current study uses an unsteady, thin-layer Navier-Stokes zonal approach to investigate the unsteady aerodynamics of a multi-stage compressor. Relative motion between rotors and stators is made possible by use of systems of patched and overlaid grids. Results have been computed for a 2 1/2-stage compressor configuration. The numerical data compares well with experimental data for surface pressures and wake data. In addition, the effect of grid refinement on the solution is studied.

Gundy-Burlet, Karen L.↗

NASA NTRS1997

Laser Anemometer Measurements of the Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller

A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code. The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.

Skoch, G. J.↗

NASA NTRS2001

An Experimental Study of the Ground Transportation System (GTS) Model in the NASA Ames 7- by 10-Ft Wind Tunnel

The 1/8-scale Ground Transportation System (GTS) model was studied experimentally in the NASA Ames 7- by 10-Ft Wind Tunnel. Designed for validation of computational fluid dynamics (CFD), the GTS model has a simplified geometry with a cab-over-engine design and no tractor-trailer gap. As a further simplification, all measurements of the GTS model were made without wheels. Aerodynamic boattail plates were also tested on the rear of the trailer to provide a simple geometry modification for computation. The experimental measurements include body-axis drag, surface pressures, surface hot-film anemometry, oil-film interferometry, and 3-D particle image velocimetry (PIV). The wind-averaged drag coefficient with and without boattail plates was 0.225 and 0.277, respectively. PIV measurements behind the model reveal a significant reduction in the wake size due to the flow turning provided by the boattail plates. Hot-film measurements on the side of the cab indicate laminar separation with turbulent reattachment within 0.08 trailer width for zero and +/- 10 degrees yaw. Oil film interferometry provided quantitative measurements of skin friction and qualitative oil flow images. A complete set of the experimental data and the surface definition of the model are included on a CD-ROM for further analysis and comparison.

Storms, Bruce L.↗

NASA NTRS2007

Loss-of-Control-Inhibitor Systems for Aircraft

Systems to provide improved tactile feedback to aircraft pilots are being developed to help the pilots maintain harmony between their control actions and the positions of aircraft control surfaces, thereby helping to prevent loss of control. A system of this type, denoted a loss-of-control-inhibitor system (LOCIS) can be implemented as a relatively simple addition to almost any pre-existing flight-control system. The LOCIS concept offers at least a partial solution to the problem of (1) keeping a pilot aware of the state of the control system and the aircraft and (2) maintaining sufficient control under conditions that, as described below, have been known to lead to loss of control. Current commercial aircraft exhibit uneven responses of primary flight-control surfaces to aggressive pilot control commands, leading to deterioration of pilots ability to control their aircraft. In severe cases, this phenomenon can result in loss of control and consequent loss of aircraft. For an older aircraft equipped with a purely mechanical control system, the loss of harmony between a pilot s command action and the control- surface response can be attributed to compliance in the control system (caused, for example, by stretching of control cables, flexing of push rods, or servo-valve distortion). In a newer aircraft equipped with a fly-by-wire control system, the major contributions to loss of harmony between the pilot and the control surfaces are delays attributable to computer cycle time, control shaping, filtering, aliasing, servo-valve distortion, and actuator rate limiting. In addition, a fly-by-wire control system provides no tactile feedback that would enable the pilot to sense such features of the control state as surface flutter, surface jam, position limiting, actuator rate limiting, and control limiting imposed by the aircraft operational envelope. Hence, for example, when a pilot is involved in aggressive closed-loop maneuvering, as when encountering a wake-vortex upset on final landing approach, the control-surface delay can lead to loss of control. Aggressive piloting can be triggered and exacerbated by control-system anomalies, which the pilot cannot diagnose because of the lack of symptoms caused by the absence of feedback through the controls. The purpose served by a LOCIS is to counteract these adverse effects by providing real-time feedback that notifies the pilot that the aircraft is tending to lag the pilot s commands. A LOCIS (see figure) includes cockpit control input-position sensors, control-surface output-position sensors, variable dampers (for example, shock absorbers containing magneto-rheological fluids such that the damping forces can be varied within times of the order of milliseconds by varying applied magnetic fields) attached to the cockpit control levers, electromagnet coils to apply the magnetic fields, and feedback control circuits to drive the electromagnet coils. The feedback control gains are chosen so that the current applied to each electromagnet coil results in a damping force that increases in a suitable nonlinear manner (e.g., exponentially) with the difference between the actual and commanded positions of the affected control surface. The increasing damping force both alerts the pilot to the onset of a potentially dangerous situation and resists the pilot s effort to command a control surface to change position at an excessive rate

AHarrah, Ralph C.↗

NASA NTRS1993

The WINCOF-I code: Detailed description

The performance of an axial-flow fan-compressor unit is basically unsteady when there is ingestion of water along with the gas phase. The gas phase is a mixture of air and water vapor in the case of a bypass fan engine that provides thrust power to an aircraft. The liquid water may be in the form of droplets and film at entry to the fan. The unsteadiness is then associated with the relative motion between the gas phase and water, at entry and within the machine, while the water undergoes impact on material surfaces, centrifuging, heat and mass transfer processes, and reingestion in blade wakes, following peal off from blade surfaces. The unsteadiness may be caused by changes in atmospheric conditions and at entry into and exit from rain storms while the aircraft is in flight. In a multi-stage machine, with an uneven distribution of blade tip clearance, the combined effect of various processes in the presence of steady or time-dependent ingestion is such as to make the performance of a fan and a compressor unit time-dependent from the start of ingestion up to a short time following termination of ingestion. The original WINCOF code was developed without accounting for the relative motion between gas and liquid phases in the ingested fluid. A modification of the WINCOF code was developed and named WINCOF-1. The WINCOF-1 code can provide the transient performance of a fan-compressor unit under a variety of input conditions.

Murthy, S. N. B.↗

NASA NTRS2016

Ice Accretions and Full-Scale Iced Aerodynamic Performance Data for a Two-Dimensional NACA 23012 Airfoil

This report documents the data collected during the large wind tunnel campaigns conducted as part of the SUNSET project (StUdies oN Scaling EffecTs due to ice) also known as the Ice-Accretion Aerodynamics Simulation study: a joint effort by NASA, the Office National d'Etudes et Recherches Aérospatiales (ONERA), and the University of Illinois. These data form a benchmark database of full-scale ice accretions and corresponding ice-contaminated aerodynamic performance data for a two-dimensional (2D) NACA 23012 airfoil. The wider research effort also included an analysis of ice-contaminated aerodynamics that categorized ice accretions by aerodynamic effects and an investigation of subscale, low- Reynolds-number ice-contaminated aerodynamics for the NACA 23012 airfoil. The low-Reynolds-number investigation included an analysis of the geometric fidelity needed to reliably assess aerodynamic effects of airfoil icing using artificial ice shapes. Included herein are records of the ice accreted during campaigns in NASA Glenn Research Center's Icing Research Tunnel (IRT). Two different 2D NACA 23012 airfoil models were used during these campaigns; an 18-in. (45.7-cm) chord (subscale) model and a 72-in. (182.9-cm) chord (full-scale) model. The aircraft icing conditions used during these campaigns were selected from the Federal Aviation Administration's (FAA's) Code of Federal Regulations (CFR) Part 25 Appendix C icing envelopes. The records include the test conditions, photographs of the ice accreted, tracings of the ice, and ice depth measurements. Model coordinates and pressure tap locations are also presented. Also included herein are the data recorded during a wind tunnel campaign conducted in the F1 Subsonic Pressurized Wind Tunnel of ONERA. The F1 tunnel is a pressured, high- Reynolds-number facility that could accommodate the full-scale (72-in. (182.9-cm) chord) 2D NACA 23012 model. Molds were made of the ice accreted during selected test runs of the full-scale model in the IRT. From these molds, castings were made that closely replicated the features of the accreted ice. The castings were then mounted on the full-scale model in the F1 tunnel, and aerodynamic performance measurements were made using model surface pressure taps, the facility force balance system, and a large wake rake designed specifically for these tests. Tests were run over a range of Reynolds and Mach numbers. For each run, the model was rotated over a range of angles-of-attack that included airfoil stall. The benchmark data collected during these campaigns were, and continue to be, used for various purposes. The full-scale data form a unique, ice-accretion and associated aerodynamic performance dataset that can be used as a reference when addressing concerns regarding the use of subscale ice-accretion data to assess full-scale icing effects. Further, the data may be used in the development or enhancement of both ice-accretion prediction codes and computational fluid dynamic codes when applied to study the effects of icing. Finally, as was done in the wider study, the data may be used to help determine the level of geometric fidelity needed for artificial ice used to assess aerodynamic degradation due to aircraft icing. The structured, multifaceted approach used in this research effort provides a unique perspective on the aerodynamic effects of aircraft icing. The data presented in this report are available in electronic form upon formal approval by proper NASA and ONERA authorities.

Addy, Harold E., Jr.↗

NASA NTRSDate not supplied

High-Fidelity Simulations of a Quadrotor Vehicle for Urban Air Mobility

High-fidelity computational fluid dynamics simulations have been carried out to analyze NASA’s quadrotor air taxi concept for urban air mobility. High-order accurate schemes, dual-time stepping, and the delayed detached-eddy simulation model have been employed. The flow solver has been loosely coupled with the rotorcraft comprehensive analysis code. The vehicle simulated is a six-passenger quadrotor for air taxi operations. The quadrotor has been simulated in hover and cruise, with flight conditions representative of an air taxi mission. Initially, the quadcopter vehicle is analyzed using low-fidelity tools. Then, a complete analysis using high-fidelity simulations of the quadcopter is performed in conjunction with collective pitch control. Airloads, wake geometry, and performance are studied and compared. A comparison between isolated rotors and complete vehicle simulations is presented. NASA’s quadrotor is one of NASA’s concept vehicles designed in support of aircraft development for vertical take-off and landing air taxi operations.

ARMD↗