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At least 91 records · Page 5

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NASA NTRS1985

Discussion of flight experiments with an entry research vehicle

The focus of interest is the maneuvering flight of advanced entry vehicles operating at altitudes above 50 km and at velocities of 5 to 8 km/s. Information resulting in more accurate aerodynamic analysis is sought and measurement techniques that appear to be applicable are identified. Measurements discussed include: shock layer or boundary layer profiles of velocity, temperature, species mass fractions, and other gas properties associated with aerodynamic heating; surface energy transfer process; nonequilibrium flow processes and pressure distribution; separated, vortic leeside flow of nonequilibrium fluid; boundary layer transition on highly swept configurations; and shock and surface slip and gas/surface interaction. Further study should focus on evolving measurement techniques, installation requirements, and on identification of the portions of flights where successful results seem probable.

Potter, J. L.↗

NASA NTRS2005

Scanning Backscatter Lidar Observations for Characterizing 4-D Cloud and Aerosol Fields to Improve Radiative Transfer Parameterizations

Clouds have a powerful influence on atmospheric radiative transfer and hence are crucial to understanding and interpreting the exchange of radiation between the Earth's surface, the atmosphere, and space. Because clouds are highly variable in space, time and physical makeup, it is important to be able to observe them in three dimensions (3-D) with sufficient resolution that the data can be used to generate and validate parameterizations of cloud fields at the resolution scale of global climate models (GCMs). Simulation of photon transport in three dimensionally inhomogeneous cloud fields show that spatial inhomogeneities tend to decrease cloud reflection and absorption and increase direct and diffuse transmission, Therefore it is an important task to characterize cloud spatial structures in three dimensions on the scale of GCM grid elements. In order to validate cloud parameterizations that represent the ensemble, or mean and variance of cloud properties within a GCM grid element, measurements of the parameters must be obtained on a much finer scale so that the statistics on those measurements are truly representative. High spatial sampling resolution is required, on the order of 1 km or less. Since the radiation fields respond almost instantaneously to changes in the cloud field, and clouds changes occur on scales of seconds and less when viewed on scales of approximately 100m, the temporal resolution of cloud properties should be measured and characterized on second time scales. GCM time steps are typically on the order of an hour, but in order to obtain sufficient statistical representations of cloud properties in the parameterizations that are used as model inputs, averaged values of cloud properties should be calculated on time scales on the order of 10-100 s. The Holographic Airborne Rotating Lidar Instrument Experiment (HARLIE) provides exceptional temporal (100 ms) and spatial (30 m) resolution measurements of aerosol and cloud backscatter in three dimensions. HARLIE was used in a ground-based configuration in several recent field campaigns. Principal data products include aerosol backscatter profiles, boundary layer heights, entrainment zone thickness, cloud fraction as a function of altitude and horizontal wind vector profiles based on correlating the motions of clouds and aerosol structures across portions of the scan. Comparisons will be made between various cloud detecting instruments to develop a baseline performance metric.

Schwemmer, Geary K.↗

NASA NTRS1992

Flight test results of riblets at supersonic speeds

A flight experiment to test and evaluate the skin friction drag characteristics of a riblet surface in turbulent flow at supersonic speeds was conducted at NASA Dryden. Riblets of groove sizes 0.0030 and 0.0013 in. were mounted on the F-104G flight test fixture. The test surfaces were surveyed with boundary layer rakes and pressure orifices to examine the boundary layer profiles and pressure distributions of the flow. Skin friction reductions caused by the riblet surface were reported based on measured differences of momentum thickness between the smooth and riblet surfaces obtained from the boundary layer data. Flight test results for the 0.0030 in. riblet show skin friction reductions of 4 to 8 % for Mach numbers ranging from 1.2 to 1.6 and Reynolds numbers ranging from 2 to 3.4 million per unit foot. The results from the 0.0013 in. riblets show skin friction reductions of 4 to 15 % for Mach 1.2 to 1.4 and Reynolds numbers ranging from 3.6 to 6 million per unit foot.

Zuniga, Fanny A.↗

NASA NTRS1947

Boundary-layer and stalling characteristics of two symmetrical NACA low-drag airfoil sections

Two symmetrical airfoils, an NACA 633-018 and an NACA 631-012, were investigated for the purpose of determining their stalling and boundary-layer characteristics with a view toward the eventual application of this information to the problem of boundary-layer control. Force measurements, pressure distributions, tuft studies, and boundary-layer-profile measurements were made at a value of 5,800,000 Reynolds number. It was found that the 18-percent-thick airfoil stalled progressively from the trailing edge because of separation of the turbulent boundary layer. In contrast, the12-percent-thick airfoil stalled abruptly from a separation of flow near the leading edge before the turbulent boundary layer became subject to separation. From this it was concluded that if high values of lift are to be obtained with thin, high-critical-speed sections by means of boundary-layer control, the work must be directed toward delaying the separation of flow near the leading edge. It was found that the presence of a nose flap on the 12-percent-thick section caused the airfoil to stall in a manner similar to that of the 18-percent-thick section.

Mccullough, George B↗

NASA NTRS1999

Crossflow Stability and Transition Experiments in Swept-Wing Flow

An experimental examination of crossflow instability and transition on a 45deg swept wing was conducted in the Arizona State University Unsteady Wind Tunnel. The stationary-vortex pattern and transition location are visualized by using both sublimating chemical and liquid-crystal coatings. Extensive hot-wire measurements were obtained at several measurement stations across a single vortex track. The mean and travelling wave disturbances were measured simultaneously. Stationary crossflow disturbance profiles were determined by subtracting either a reference or a span-averaged velocity profile from the mean velocity data. Mean, stationary crossflow, and traveling wave velocity data were presented as local boundary layer profiles and contour plots across a single stationary crossflow vortex track. Disturbance mode profiles and growth rates were determined. The experimental data are compared with predictions from linear stability theory.

Dagenhart, J. Ray↗

NASA NTRS1986

Temperature and velocity profiles in sooting free boundary layer flames

Temperature and velocity profiles are presented for cyclohexane, n-heptane, and iso-octane free, laminar, boundary layer, sooting, diffusion flames. Temperatures are measured with 3 mil Pt/Pt-13 percent Rh thermocouples. Corrected gas temperatures are derived by performing an energy balance of convection to and radiation from the thermocouple bead incorporating the variation of air conductivity and platinum emissivity with temperature. Velocities are measured using laser doppler velocimetry techniques. Profiles are compared with previously reported analytic temperature and velocity fields. Comparison of theoretical and experimental temperature profiles suggests improvement in the analytical treatment is needed, which accounts more accurately for the local soot radiation. The velocity profiles are in good agreement, with the departure of the theory from observation partially due to the small fluctuations inherent in these free flows.

Ang, J. A.↗

NASA NTRS2010

Receptivity of Hypersonic Boundary Layers over Straight and Flared Cones

The effects of adverse pressure gradients on the receptivity and stability of hypersonic boundary layers were numerically investigated. Simulations were performed for boundary layer flows over a straight cone and two flared cones. The steady and the unsteady flow fields were obtained by solving the two-dimensional Navier-Stokes equations in axi-symmetric coordinates using the 5th order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The mean boundary layer profiles were analyzed using local stability and non-local parabolized stability equations (PSE) methods. After the most amplified disturbances were identified, two-dimensional plane acoustic waves were introduced at the outer boundary of the computational domain and time accurate simulations were performed. The adverse pressure gradient was found to affect the boundary layer stability in two important ways. Firstly, the frequency of the most amplified second-mode disturbance was increased relative to the zero pressure gradient case. Secondly, the amplification of first- and second-mode disturbances was increased. Although an adverse pressure gradient enhances instability wave growth rates, small nose-tip bluntness was found to delay transition due to the low receptivity coefficient and the resulting weak initial amplitude of the instability waves. The computed and measured amplitude-frequency spectrums in all three cases agree very well in terms of frequency and the shape except for the amplitude.

Balakumar, Ponnampalam↗

NASA NTRS2004

Effect of Blowing on Boundary Layer of Scarf Inlet

When aircraft operate in stationary or low speed conditions, airflow into the engine accelerates around the inlet lip and pockets of turbulence that cause noise and vibration can be ingested. This problem has been encountered with engines equipped with the scarf inlet, both in full scale and in model tests, where the noise produced during the static test makes it difficult to assess the noise reduction performance of the scarf inlet. NASA Langley researchers have implemented boundary layer control in an attempt to reduce the influence of the flow nonuniformity in a 12-in. diameter model of a high bypass fan engine mounted in an anechoic chamber. Static pressures and boundary layer profiles were measured in the inlet and far field acoustic measurements were made to assess the effectiveness of the blowing treatment. The blowing system was found to lack the authority to overcome the inlet distortions. Methods to improve the implementation of boundary layer control to reduce inlet distortion are discussed.

Gerhold, Carl H.↗

NASA NTRS1992

Crossflow Stability and Transition Experiments in a Swept-Wing Flow

An experimental examination of crossflow instability and transition on a 45 degree swept wing is conducted in the Arizona State University Unsteady Wind Tunnel. The stationary-vortex pattern and transition location are visualized using both sublimating-chemical and liquid-crystal coatings. Extensive hot-wire measurements are conducted at several measurement stations across a single vortex track. The mean and travelling-wave disturbances are measured simultaneously. Stationary-crossflow disturbance profiles are determined by subtracting either a reference or a span-averaged velocity profile from the mean-velocity data. Mean, stationary-crossflow, and travelling-wave velocity data are presented as local boundary-layer profiles and as contour plots across a single stationary-crossflow vortex track. Disturbance-mode profiles and growth rates are determined. The experimental data are compared to predictions from linear stability theory.

Dagenhart, John Ray↗

NASA NTRS1981

Some observations of surface pressures and the near wake of a blunt trailing edge airfoil

Experiments with a truncated and untruncated airfoils of profiles NACA 640A10, were carried out in subsonic wind tunnels in a velocity range of 19m/s to 54m/s corresponding to Reynolds numbers of 200,000 to 468,000 based on the chord. Airfoil spanned the test section to achieve two dimensionality of the model. Velocity measurements, pressure measurements, and vortex shedding in the wake were measured using a hotwire and pressure transducers. The measured chordwise static pressure distribution on the smooth trailing edge airfoil along the midspan plane, agreed with the theoretical results calculated on the basis of the potential flow for that airfoil. Boundary layer profiles measured in the midspan plane, behind the maximum thickness of the airfoil show no separation of the flow. Spanwise distribution of the measured static pressure on the upper surface of the airfoil shows uniformity for both configurations with and without the boundary layer trip. This uniformity of pressure distribution and separation indicates that the flow on the airfoil was uniform and two dimensional in character.

Digumarthi, R. V.↗

NASA NTRS1972

Turbulent boundary layer in an adverse pressure gradient without effect of wall curvature

The hypersonic compressible turbulent boundary layer in an adverse pressure gradient along a cylindrical axisymmetric body was studied. The tests were conducted in a Mach 6 contoured axisymmetric nozzle. An external compression cowl was used to produce the gradual adverse pressure gradient and a maximum pressure rise of 7 times the freestream static pressure was achieved in a test region of 23 cm. Boundary layer profiles of static pressure, total pressure, and total temperature, as well as wall transient heat transfer rates were measured. Comparisons of the velocity total temperature profiles to linear and quadratic relations were made. Measured heat transfer data were in good agreement with the prediction from the flat-plate reference enthalpy method. Integral parameters were also in good agreement with results of numerical solutions for compressible turbulent boundary layer equations.

Zakkay, V.↗

NASA NTRS1991

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 5: Propagation of propeller tone noise through a fuselage boundary layer

An analysis of tone noise propagation through a boundary layer and fuselage scattering effects was derived. This analysis is a three dimensional and the complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The outer wave field is constructed analytically from an incident wave appropriate to the source and a scattered wave in the standard Hankel function form. For the incident wave, an existing function - domain propeller noise radiation theory is used. In the boundary layer region, the wave equation is solved by numerical methods. The theoretical analysis is embodied in a computer program which allows the calculation of correction factors for the fuselage scattering and boundary layer refraction effects. The effects are dependent on boundary layer profile, flight speed, and frequency. Corrections can be derived for any point on the fuselage, including those on the opposite side from the source. The theory was verified using limited cases and by comparing calculations with available measurements from JetStar tests of model prop-fans. For the JetStar model scale, the boundary layer refraction effects produce moderate fuselage pressure reinforcements aft of and near the plane of rotation and significant attenuation forward of the plane of rotation at high flight speeds. At lower flight speeds, the calculated boundary layer effects result in moderate amplification over the fuselage area of interest. Apparent amplification forward of the plane of rotation is a result of effective changes in the source directivity due to boundary layer refraction effects. Full scale effects are calculated to be moderate, providing fuselage pressure amplification of about 5 dB at the peak noise location. Evaluation using available noise measurements was made under high-speed, high-altitude flight conditions. Comparisons of calculations made of free field noise, using a current frequency-domain propeller noise prediction method, and fuselage effects using this new procedure show good agreement with fuselage measurements over a wide range of flight speeds and frequencies. Correction factors for the JetStar measurements made on the fuselage are provided in an Appendix.

Magliozzi, B.↗

NASA NTRS1985

A novel centrifugal diffuser test device

A large scale (2 in.-wide test vanes at 50-in. inlet diameter), low speed (100-200 ft/sec) steady flow radial cascade wind tunnel for diffuser studies was designed, built and tested. The apparatus was shown to provide flow angles (from radial) in the range of 58-72 deg with suitable spanwise profiles. The novel flow angle control mechanism was shown to work, but measured flow angles were somewhat smaller than expected. The mechanism for controlling case-wall boundary layer profile did not behave as predicted. Attempts to predict the generated flow both analytically and with computational codes are compared with initial measurements. The low speed apparatus will be used to obtain detailed data for diffuser design and analysis code verification, and to provide experience toward the design of a high speed device.

Shreeve, R. P.↗

NASA NTRS1987

The measurement of boundary layers on a compressor blade in cascade. III - Pressure surface boundary layers and the near wake

A one-component laser Doppler velocimeter (LDV) has been used to measure the two-dimensional periodic flow field about a double circular arc, compressor blade in cascade. Eleven boundary layer profiles were taken on both the pressure and suction surfaces of the blade, and two profiles were taken in the near wake. In this part of the study, the detailed LDV studies are described. The measurements indicate that the onset of transition occurs near 60 percent chord. The lack of a logarithmic region in the data measured at 97.9 percent chord indicates that transition is not complete. The thin laminar boundary layers near the leading edge led to some measurement problems, characterized by large turbulence intensities, in using the LDV.

Deutsch, Steven↗

NASA NTRS1979

A field study of air flow and turbulent features of advection fog

The setup and initial operation of a set of specialized meteorological data collection hardware are described. To study the life cycle of advection fogs at a lake test site, turbulence levels in the fog are identified, and correlated with the temperature gradients and mean wind profiles. A meteorological tower was instrumented to allow multiple-level measurements of wind and temperature on a continuous basis. Additional instrumentation was: (1)hydrothermograph, (2)microbarograph, (3)transmissometers, and (4)a boundary layer profiler. Two types of fogs were identified, and important differences in the turbulence scales were noted.

Connell, J. D.↗

NASA NTRS2016

Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D

This paper presents results from an exploratory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected and the resulting steady-state analyses using NASA's FUN3D CFD software.

Chwalowski, Pawel↗

NASA NTRSDate not supplied

Design Exploration of a Transonic Cruise Slotted Airfoil

A knowledge-based aerodynamic design method, CDISC, has been leveraged for the computational design exploration of the cruise slotted airfoil as a drag-saving technology for single aisle, transonic transport aircraft. Aerodynamic predictions were generated using the NASA USM3D Reynolds-averaged Navier-Stokes flow solver, and laminar flow assessments were conducted using the NASA BLSTA3D boundary layer profile solver paired with the LASTRAC stability analysis and transition prediction software. The CDISC design method was used to parametrically vary several slotted-airfoil design variables to understand their impact on aerodynamic performance. In addition to establishing best practices for slotted airfoil design, it was observed that the historically noted skin-friction penalty for cruise slotted airfoils was attributed to the low-Reynolds-number boundary layer of the flap. At cruise, this skin-friction penalty negated any potential decrease in pressure drag enabled by the cruise slotted airfoil architecture. This observation led to the development of an Aft-Laminar Multi-element Airfoil concept that uses airfoil shaping to promote natural laminar flow on the flap. At cruise conditions, the concept is predicted to offer a 2.7% reduction in sectional drag relative to a fully turbulent supercritical airfoil. Off-design analyses showed that laminar flow could be maintained with near-cruise variations in angle of attack and Mach number, resulting in sustained aerodynamic efficiency improvements and a delay in the drag rise Mach number. A low-speed, high-lift analysis at takeoff conditions predicted a maximum lift coefficient of 2.4 could be achieved by use of a variable-camber leading edge on the main element and simple deflection of the flap. These computationally-predicted benefits have motivated future research toward an aft-laminar cruise slotted wing design as an incremental step toward the application of natural laminar flow technology on transonic commercial transports.

Brett R Hiller↗

NASA NTRS1997

Reentry-F Flowfield Solutions at 80,000 ft.

Three equilibrium-air numerical solutions are presented for the Reentry-F flight-test vehicle at Mach 20, 80,000 Ft. conditions, including turbulent flow predictions. The three solutions are from a thin-layer Navier-Stokes code, coupled thin-layer and parabolized Navier-Stokes codes, and an approximate viscous shock-layer code. Boundary-layer and shock-layer profiles are presented and compared between the three solutions, revealing close agreement between the three solution methods. Notable exceptions to the close agreement, with 7-10 percent discrepancies, occur in the density profiles at the boundary-layer edge, in the boundary-layer velocity profiles, and in the shock-layer profiles in regions influenced by the nose bluntness.

Wood, William A.↗