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All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Effect of Inflow Boundary Conditions on Hovering Tilt-Rotor Flows

A study on the effect of inflow turbulence boundary conditions on the local flow on and around a body in flight has been carried out. The study has been carried out using OVERFLOW2 flow solver using the default Spalart-Allmaras turbulence model in OVERFLOW. Many OVERFLOW turbulent flow simulations have been reported using the SA-fv3 model1 over the years. The present study demonstrates that the turbulence levels imposed as a boundary condition (b.c.) in the far-field as implemented in OVERFLOW are not correct. In fact, very low level of turbulence at the far-field boundary as implemented in OVERFLOW results in the SA-fv3 model predicting transition-like profiles on a given body. By choosing sufficiently high levels of the Reynolds number of turbulence, Rt, as a boundary condition, this anomalous behavior of SA-fv3 model is eliminated. Since numerous papers using OVERFLOW have been presented in the literature including that by the author2 using low level of inflow turbulence, it will be beneficial to the CFD community at large and in particular to the OVERFLOW community to understand this effect of the inflow b.c. in the SA-fv3 model. Various results reported over the years in domains such as high-lift applications, drag prediction applications and rotorcraft flow applications can be revisited using the new turbulence inflow boundary condition prescription as suggested in this study. It has been demonstrated in the literature that the effect of inflow turbulence level on the downstream development of flow is too important to be ignored. Demonstrative results are shown and compared with experiment for the case of hovering XV-15 rotor flow at a tip Reynolds number of 4.9 x 10(exp 6) and Mach number of 0.69 in Figs. 1 and 2 below. Fig. 1(a,b) shows skin friction predictions at radial station of r/R = 0.28 corresponding to 5 different inflow boundary conditions obtained with the SA-fv3 model. As the inflow turbulence is increased monotonically, from a value of Rt = 0.1 to 100.0 as shown in Fig. 1, a fully developed turbulent skin friction profile is realized progressively as shown in Fig. 1(b). Similarly skin friction results shown in Fig. 2(a,b) corresponding to radial stations of r/R = 0.72 reinforce the postulate made here that as the inflow turbulence level is increased gradually, fully turbulent on-the-blade simulation is realized uniformly. In the full paper extensive results on skin friction, pressure and velocity profiles as well as the Figure of Merit will be presented and compared with experiment that will illustrate the substantial effect of the inflow boundary conditions on the development of hovering rotor flow downstream.

Kaul, Upender K.↗
NASA NTRS2012

Cascading Tesla Oscillating Flow Diode for Stirling Engine Gas Bearings

Replacing the mechanical check-valve in a Stirling engine with a micromachined, non-moving-part flow diode eliminates moving parts and reduces the risk of microparticle clogging. At very small scales, helium gas has sufficient mass momentum that it can act as a flow controller in a similar way as a transistor can redirect electrical signals with a smaller bias signal. The innovation here forces helium gas to flow in predominantly one direction by offering a clear, straight-path microchannel in one direction of flow, but then through a sophisticated geometry, the reversed flow is forced through a tortuous path. This redirection is achieved by using microfluid channel flow to force the much larger main flow into this tortuous path. While microdiodes have been developed in the past, this innovation cascades Tesla diodes to create a much higher pressure in the gas bearing supply plenum. In addition, the special shape of the leaves captures loose particles that would otherwise clog the microchannel of the gas bearing pads.

Dyson, Rodger↗
NASA NTRS2012

Analysis of Ares Crew Launch Vehicle Transonic Alternating Flow Phenomenon

A transonic wind tunnel test of the Ares I-X Rigid Buffet Model (RBM) identified a Mach number regime where unusually large buffet loads are present. A subsequent investigation identified the cause of these loads to be an alternating flow phenomenon at the Crew Module-Service Module junction. The conical design of the Ares I-X Crew Module and the cylindrical design of the Service Module exposes the vehicle to unsteady pressure loads due to the sudden transition between a subsonic separated and a supersonic attached flow about the cone-cylinder junction as the local flow randomly fluctuates back and forth between the two flow states. These fluctuations produce a square-wave like pattern in the pressure time histories resulting in large amplitude, impulsive buffet loads. Subsequent testing of the Ares I RBM found much lower buffet loads since the evolved Ares I design includes an ogive fairing that covers the Crew Module-Service Module junction, thereby making the vehicle less susceptible to the onset of alternating flow. An analysis of the alternating flow separation and attachment phenomenon indicates that the phenomenon is most severe at low angles of attack and exacerbated by the presence of vehicle protuberances. A launch vehicle may experience either a single or, at most, a few impulsive loads since it is constantly accelerating during ascent rather than dwelling at constant flow conditions in a wind tunnel. A comparison of a windtunnel- test-data-derived impulsive load to flight-test-data-derived load indicates a significant over-prediction in the magnitude and duration of the buffet load. I. Introduction One

Sekula, Martin K.↗
NASA NTRS2015

Resource Prospector Propulsion System Cold Flow Testing

Resource Prospector (RP) is a NASA mission being led by NASA Ames Research Center with current plans to deliver a scientific payload package aboard a rover to the lunar surface. As part of an early risk reduction activity, Marshall Space Flight Center (MSFC) and Johnson Space Flight Center (JSC) have jointly developed a government-version concept of a lunar lander for the mission. The spacecraft consists of two parts, the lander and the rover which carries the scientific instruments. The lander holds the rover during launch, cruise, and landing on the surface. Following terminal descent and landing the lander portion of the spacecraft become dormant after the rover embarks on the science mission. The lander will be equipped with a propulsion system for lunar descent and landing, as well as trajectory correction and attitude control maneuvers during transit to the moon. Hypergolic propellants monomethyl hydrazine and nitrogen tetroxide will be used to fuel sixteen 70-lbf descent thrusters and twelve 5-lbf attitude control thrusters. A total of four metal-diaphragm tanks, two per propellant, will be used along with a high-pressure composite-overwrapped pressure vessel for the helium pressurant gas. Many of the major propulsion system components are heritage missile hardware obtained by NASA from the Air Force. In parallel with the flight system design activities, a simulated propulsion system based on flight drawings was built for conducting a series of water flow tests to characterize the transient fluid flow of the propulsion system feed lines and to verify the critical operation modes such as system priming, waterhammer, and crucial mission duty cycles. The primary objective of the cold flow testing was to simulate the RP propulsion system fluid flow operation through water flow testing and to obtain data for anchoring analytical models. The models will be used to predict the transient and steady state flow behaviors in the actual flight operations. All design and build efforts, including the analytical modeling, have been performed. The cold flow testing of the propulsion system was set up and conducted at a NASA MSFC test facility. All testing was completed in the summer of 2014, and this paper documents the results of that testing and the associated fluid system modeling efforts.

Williams, Hunter↗
NASA NTRS1951

Viscous Effects on Impact Probes in a Subsonic Rarified Gas Flow

The problem of time response of pressure gaging systems used with low density flows has been discussed in Ref ]4, and the effect of outgassing on pressure magnitudes was indicated in the same reference. Briefly, gases or vapors adhering to the internal walls of the pressure gaging system behave like gas sources and produce a pressure rise in the gage system which has no relation to the external flow. The effect can also occur in the reverse direction, with "in-gassing" or the action of an effective sink in the gage system as gases entering through the probe orifice are adsorbed to the walls. For given surface conditions, the magnitude of the pressure error to be expected due to out-gassing depends on the dimensions of the probe system. In the present tests, it was desirable to use the smallest possible probe to yield the lowest possible Reynolds number. The lower limit on size was fixed by outgassing effects, evaluated by the following procedure. 2.0 PROCEDURE The pumping system was adjusted to give a pressure, measured at the reservoir, of 9 microns Hg with no flow into the wind tunnel. The upstream metering valve was then opened and the air flow rate adjusted to give a pressure of 100 microns in the reservoir under steady flow conditions, The pressure read by one of the impact tubes, inserted into the flow, was measured. The upstream metering valve was then closed rapid)y, and the reservoir pressure, p0, and the impact probe pressure, Pj, were measured simultaneously at definite time intervals. Another probe was inserted into the flow and the procedure repeated, until the pressure-time data had been obtained for each probe investigated. The results for a series of tests Involving probe Nos.l, 2 and 3 are shown on HYD 2616 300 RESULTS From HYD 2616, It is clear that the smallest probe (No.3) requires the longest time to reach equilibrium. After a sufficient time has elapsed (about 180 seconds), this probe read the same, within the accuracy of measurement, as the other two, When a similar experiment was performed utilizing a probe which was one-half the size of probe No.2, it indicated a pressure, after 180 seconds, which was almost 10 microns Hg higher than the other probes. Accordingly, only probe Nos. 1, 2, 3 and 4 were employed in the experiments, and a time of at least 180 seconds was allowed to elapse between a change of setting and the reading of the instruments0

Kane, E. D.↗
NASA NTRS2016

A Dual-Plane PIV Study of Turbulent Heat Transfer Flows

Thin film cooling is a widely used technique in turbomachinery and rocket propulsion applications, where cool injection air protects a surface from hot combustion gases. The injected air typically has a different velocity and temperature from the free stream combustion flow, yielding a flow field with high turbulence and large temperature differences. These thin film cooling flows provide a good test case for evaluating computational model prediction capabilities. The goal of this work is to provide a database of flow field measurements for validating computational flow prediction models applied to turbulent heat transfer flows. In this work we describe the application of a Dual-Plane Particle Image Velocimetry (PIV) technique in a thin film cooling wind tunnel facility where the injection air stream velocity and temperatures are varied in order to provide benchmark turbulent heat transfer flow field measurements. The Dual-Plane PIV data collected include all three components of velocity and all three components of vorticity, spanning the width of the tunnel at multiple axial measurement planes.

Turbulent heat transfer↗
NASA NTRS2016

Particle Streak Velocimetry of Supersonic Nozzle Flows

A novel velocimetry technique to probe the exhaust flow of a laboratory scale combustor is being developed. The technique combines the advantages of standard particle velocimetry techniques and the ultra-fast imaging capabilities of a streak camera to probe high speed flows near continuously with improved spatial and velocity resolution. This "Particle Streak Velocimetry" technique tracks laser illuminated seed particles at up to 236 picosecond temporal resolution allowing time-resolved measurement of one-dimensional flows exceeding 2000 m/s as are found in rocket nozzles and many other applications. Developmental tests with cold nitrogen have been performed to validate and troubleshoot the technique with supersonic flows of much lower velocity and without background noise due to combusting flow. Flow velocities on the order of 500 m/s have been probed with titanium dioxide particles and a continuous-wave laser diode. Single frame images containing multiple streaks are analyzed to find the average slope of all incident particles corresponding to the centerline axial flow velocity. Long term objectives for these tests are correlation of specific impulse to theoretical combustion predictions and direct comparisons between candidate green fuels and the industry standard, monomethylhydrazine, each tested under identical conditions.

Willits, J. D.↗
NASA NTRS2004

Gas-Liquid Two-Phase Flow Through Packed Beds in Microgravity

Experimental data on flow pattern transitions, pressure drop, and flow characteristics for cocurrent gas–liquid flow through packed columns in microgravity is analyzed. The pulse flow regime exists over a wider range of gas and liquid flow rates under microgravity conditions. Furthermore, maps used to predict transition boundaries in normal gravity do not apply in microgravity. The flow regime data are compared to the widely used Talmor map and a new transition criterion between bubble and pulse flow in microgravity is proposed. The pressure-drop data clearly show that interfacial effects can increase the pressure drop by as much as 300% compared to those predicted by the single-phase Ergun equation. A two-phase friction factor is correlated to the superficial gas and liquid Reynolds numbers and the Suratman number. New data are also presented on the influence of gravity on the pulse amplitude and frequency.

Brian J. Motil↗
NASA NTRS2017

Models of Sector Flows Under Local, Regional and Airport Weather Constraints

Recently, the ATM community has made important progress in collaborative trajectory management through the introduction of a new FAA traffic management initiative called a Collaborative Trajectory Options Program (CTOP). FAA can use CTOPs to manage air traffic under multiple constraints (manifested as flow constrained areas or FCAs) in the system, and it allows flight operators to indicate their preferences for routing and delay options. CTOPs also permits better management of the overall trajectory of flights by considering both routing and departure delay options simultaneously. However, adoption of CTOPs in airspace has been hampered by many factors that include challenges in how to identify constrained areas and how to set rates for the FCAs. Decision support tools providing assistance would be particularly helpful in effective use of CTOPs. Such DSTs tools would need models of demand and capacity in the presence of multiple constraints. This study examines different approaches to using historical data to create and validate models of maximum flows in sectors and other airspace regions in the presence of multiple constraints. A challenge in creating an empirical model of flows under multiple constraints is a lack of sufficient historical data that captures diverse situations involving combinations of multiple constraints especially those with severe weather. The approach taken here to deal with this is two-fold. First, we create a generalized sector model encompassing multiple sectors rather than individual sectors in order to increase the amount of data used for creating the model by an order of magnitude. Secondly, we decompose the problem so that the amount of data needed is reduced. This involves creating a baseline demand model plus a separate weather constrained flow reduction model and then composing these into a single integrated model. A nominal demand model is a flow model (gdem) in the presence of clear local weather. This defines the flow as a function of weather constraints in neighboring regions, airport constraints and weather in locations that can cause re-routes to the location of interest. A weather constrained flow reduction model (fwx-red) is a model of reduction in baseline counts as a function of local weather. Because the number of independent variables associated with each of the two decomposed models is smaller than that with a single model, need for amount of data is reduced. Finally, a composite model that combines these two can be represented as fwx-red (gdem(e), l) where e represents non-local constraints and l represents local weather. The approaches studied to developing these models are divided into three categories: (1) Point estimation models (2) Empirical models (3) Theoretical models. Errors in predictions of these different types of models have been estimated. In situations when there is abundant data, point estimation models tend to be very accurate. In contrast, empirical models do better than theoretical models when there is some data available. The biggest benefit of theoretical models is their general applicability in wider range situations once the degree of accuracy of these has been established.

Kulkarni, Deepak↗
NASA NTRS2015

Visualization and Quantification of Rotor Tip Vortices in Helicopter Flows

Helicopter aeromechanics encompasses a highly vortical flow field. The vortices generated at each blade tip contain unsteady, complex, three-dimensional structures, which interact with each other, other blades, the fuselage and various components of the helicopter. It is crucial to understand vortex kinematics and their subsequent dynamic evolution. Much research has been devoted to the understanding of helicopter vortex dynamics, including a number of experimental studies.1-6 In May 2010 Particle Image Velocimetry (PIV) measurements of a full-scale UH-60A rotor were acquired in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel.1 These measurements were taken at a plane just downstream of the advancing blade in the vicinity of the blade tip—the so-called PIV plane. The resulting PIV data were then processed using an ensemble-average approach to create graphical representations of the vortical wake velocity and vorticity fields, which, in turn, have enhanced the understanding of rotorcraft vortical wake flow field physics and have provided a more detailed validation of vortical wake computer simulations.7 A common approach used to analyze flow field features is to compute and plot color contour maps of various scalar quantities such as pressure, velocity magnitude and vorticity magnitude. For example, the color map of the vorticity magnitude is typically used to determine vortical flow structure. With this approach the vortex core may appear larger or smaller, depending on the contour levels that are selected. Thus, the resulting visualization is sensitive to user-specified contour levels. For vortex core radius measurements, it is more accurate to calculate the vortex core radius using the cross-flow velocity profile across the vortex core. The task of extracting the cross-flow velocity profile can be time consuming with existing tools since the user needs to manually select the core center then specify sampling points along the profile axis. The task becomes even more challenging when the associated grid system uses AMR (Adaptive Mesh Refinement) where the profile axis could span multiple grid blocks. There are a number of existing techniques for profiling of vortex core attributes;8-9 however, these techniques are not fully automatic in that the user still needs to select the vortex core center to compute the cross-flow velocity profile. The present study introduces a new color map scheme that is based on the vortex core radius, which is fully automatic and does not require user intervention. Analysis and visualization of blade tip vortices on the PIV plane using the proposed new color map scheme are described in Section II. The new approach is evaluated using two case studies, which are described in Section III. The paper ends with a summary in Section IV.

Quantification↗
NASA NTRS2019

Microgravity Liquid-Gas Two-Phase Flow: Review of Pressure Drop and Heat Transfer Correlations and Guidelines for Equipment Operability

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.

liquid-gas↗
NASA NTRS2019

Gravity Effects in Two-Phase Microgap Flow

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.

Robinson, Franklin L.↗
NASA NTRS2019

100-kHz Rate Rayleigh Imaging for Combustion and Flow Diagnostics

Two-dimensional (2D) Rayleigh scattering (RS) imaging at an ultrahigh repetition rate of 100 kHz is demonstrated in non-reacting and reacting flows employing a high-energy burst-mode laser system. Image sequences of flow mixture fraction were directly derived from high-speed RS images. Additionally, a 2D instantaneous flow velocity field at 100 kHz was obtained through optical-flow-based analysis of the RS images. The technique was also applied to study turbulent flames having a near-constant Rayleigh cross section. The demonstrated high-speed RS technique in conjunction with optical-flow-based analysis provides non-intrusive, simultaneous measurements of the flow mixing and velocity field, extending the measurement capability of the RS technique to high-speed non-reacting and reacting flows.

Jiang, Naibo↗
NASA NTRS2019

Surface Flow Visualization of the High-Lift Common Research Model

A 10% scale version of the High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The CRM-HL experiment included various configurations such as conventional and simple-hinged flaps, with and without engine nacelle/pylon, with and without nacelle chine, different Active Flow Control (AFC) methods (sweeping jets, alternating pulsed jets, and preconditioned boundary layer blowing), and their various parameters. This particular study is focused on the surface flow visualization of the conventional CRM-HL model at landing configuration. The conventional CRM-HL model with the single-slotted Fowler flap system serves as a baseline for the AFC-enabled simplified high-lift configuration as well as a high-lift technology development platform due to its publicly open geometry. Surface flow visualizations were performed using fluorescent minitufts, which were found to be nonintrusive to the aerodynamic performance. Tuft flow visualizations are supplemented with the relevant pressure and force measurements in order to understand the flow characteristics developed on the conventional CRM- HL model. In addition, three dimensional, unsteady, compressible Computational Fluid Dynamic (CFD) simulations were performed for selective cases. The surface streamlines and transverse velocity fluctuations obtained by the CFD simulations are qualitatively compared to the tuft direction and tuft unsteadiness, respectively. Force measurements of the CRM-HL model show performance degradation at higher angles of attack. Surface flow visualizations revealed the performance loss due to the nacelle/pylon wake that grows with angle of attack and eventually promotes flow separation over the inboard wing. This performance loss was successfully recovered by placing a chine on the engine nacelle.

Koklu, Mehti↗
NASA NTRSDate not supplied

Flow Physics at a Large Rotor Tip Gap in a Multi-Stage Axial Compressor

The flow physics in a large rotor tip gap in a 1.5-stage axial compressor is investigated in the current study. The flow structure in the rotor tip region is complex with several dominant vortical structures of opposite rotation, resulting in inhomogeneous and highly anisotropic turbulence. Earlier measurements show that eddy viscosity is negative over large parts of the tip region and eddy viscosity varies among stress/strain components. The present study aims to understand how the complex nature of rotor tip leakage flow affects compressor performance when the tip gap size is greater than 4–5% of the rotor span, which is typical of advanced small core engines. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques are applied to study flow physics in a large rotor tip gap (5.5% of rotor span) in a 1.5-stage axial compressor. Calculated flow fields from the two different approaches are compared with available measurements and examined in detail. LES calculates the pressure rise in the present compressor fairly well, while URANS with a standard two-equation turbulence closure underpredicts the pressure rise by 15-20% of the measured values. The current study shows that URANS with the current turbulence closure produces much higher all-positive eddy viscosity in the tip-gap region compared to measurements and LES. The distribution of eddy viscosity in the URANS simulation is also wrong. Consequently, the flow in the tip region is highly damped with significantly larger blockage generation, which results in the tip leakage vortex (TLV) staying closer to the blade suction side compared to the measurement. When the TLV stays closer to the blade, both flow turning and the pressure rise across the compressor are reduced compared to the measurements. It appears that this effect is amplified by a large rotor tip gap.

tip gap↗
NASA NTRS2019

Novel concept for detection of a fluid flow fault in a pumped fluid heat rejection system

A pumped fluid heat rejection system (HRS) requires continuous flow of the working fluid to ensure that the components controlled by the HRS stay within their allowable temperature limits. An interruption of flow could result in violations of hardware qualification limits and mission failure. Some of these violations can happen within a few hours. Hence, quick detection of a flow fault to invoke mitigation measures is very critical. In typical pumped fluid HRS, dual or triple pumps are employed to switch the backup units in case the primary unit were to fail. The key metric for the selection of a fault detection system is that it should detect the fault much before the fault’s impact on the thermal health of the HRS controlled components is realized. A trade study was conducted to select such a system for the Europa Clipper Mission to Europa, a moon of Jupiter that is planned for a launch in 2023. Out of the several concepts studied, the most attractive one was a novel and simple concept that uses a low power film heater attached to a section of the HRS tubing. While the fluid flows at its nominal rate, the high thermal coupling of the flowing fluid leads to the tube being close to the fluid’s temperature. However, when the flow stops, the heater warms the small thermal mass of the tubing to a high temperature in a short span of time (~15 minutes). This large temperature rise would then imply that the flow must have stopped. This paper will describe the various concepts considered, the chosen concept, its implementation, and the results of developments tests to validate its performance.

Schmidt, Tyler↗
NASA NTRSDate not supplied

Investigation of the 30P30N Slat Flow Field with Passive Control Devices Using Particle Image Velocimetry

The leading-edge slat of a high-lift wing is one of the main noise contributors during approach and landing. This paper describes an experimental investigation of the velocity field associated with multiple passive noise treatments, including slat-cusp extensions, a cove filler, and a gap filler, on a two-dimensional multi-element high-lift 30P30N airfoil. Previous work documented comparisons of both surface and far field pressure fluctuations in the presence of these devices with those for a baseline case at different flow conditions. However, important information related to the velocity fields was missing from the previous measurements, hindering our ability to elucidate the changes in the flow physics associated with the noise treatments. Therefore, two-component Particle Image Velocity has been used to investigate the influence of passive noise treatments on the flow fields. All measurements are taken at an effective, free-air angle of attack of 5.5 degrees and a stowed-chord-based Reynolds number of 1.71 × 10e6. The measurements show that the slat extensions shorten the slat-cove shear layer trajectory, resulting in reduced growth of disturbances within the slat-cove shear layer. This leads to a shift of tonal peaks to higher frequencies and a reduction in the tonal amplitudes. The gap filler blocks the flow path through the gap, causing the reattachment location to shift to the main wing leading edge lower surface. Consequently, the feedback loop associated with the flow-acoustic interaction in the baseline case is eliminated, and the turbulent kinetic energy in the slat-cove shear layer is significantly reduced. However, extensive flow separation is observed on the suction side of the gap filler, which does not eliminate the noise reduction benefit due to the gap filler, but does degrade the aerodynamic performance of the high-lift configuration. Finally, the overall flow field in the presence of the cove filler is similar to that in the baseline case at the design angle of attack, but the slat-cove shear layer is eliminated leading to a suppression of the cavity tones associated with the shear layer. This change accounts for the reduction in slat noise as measured in previous work.

Aeroacoustics↗
NASA NTRSDate not supplied

Flow Characterization of the NASA Langley Unitary Plan WindTunnel, Test Section 2: Computational Results

Flow in the empty Unitary Planform Wind Tunnel at the NASA Langley Research Center issimulated with computational fluid dynamics methods. The objectives are to assess CFD’s truepredictive capability and to generate flow-field maps upstream of the tunnel’s test section foruse in associated vehicle tests. Multiple CFD solvers, grid adaption methods, and turbulencemodels are used by five teams doing the simulations. The simulation domain is as large aspractical: from the start of the settling chamber, just downstream of the last set of turningvanes, to well downstream of the test section. Simulations are done at three Mach numbers andseveral Reynolds numbers that cover most of the tunnel’s operating range. The sensitivity ofthe CFD solutions to simulation process details, including turbulence model, spatial resolutionand settling-chamber inflow velocity profile, is characterized. Significant sensitivities in somedetails of the predicted test section flow are described. CFD predicts that details of flow inthe settling chamber can affect the flow in the test section. Turning of the flow through thetunnel contraction and past nozzle-block hardware generates streamwise vortices, some ofwhich survive into the core of the test section, principally at low Mach number. Preliminarycomparisons to experimental measurements are given. The major flow characteristics andthe dynamical vortices predicted by CFD exist in the experiment. Several details in CFD andexperiment are observed to differ, including the effects of the vortices and features that appearto be Mach waves in the test section. Sensitivity to grid resolution and turbulence modelingis noted in the vortices; no potential cause for the differences in Mach waves has yet beenidentified in the CFD sensitivity studies.

wind tunnel↗