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Kegerise, Michael A.

Publications and source records attributed to Kegerise, Michael A..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

An Experimental Investigation of a Wing-Fuselage Junction Model in the NASA Langley 14- by 22-Foot Subsonic Tunnel

Current turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing-fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow-field data in the junction region is needed. In this report, we present an overview of experimental measurements on a wing-fuselage junction model that addresses this need. The experimental measurements were performed in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The model was a full-span wing-fuselage body that was configured with truncated DLR-F6 wings, both with and without leading-edge extensions at the wing root. The model was tested at a fixed chord Reynolds number of 2.4 million, and angles-of-attack ranging from -10 degrees to +10 degrees were considered. Flow-field measurements were performed with a pair of miniature laser Doppler velocimetry (LDV) probes that were housed inside the model and attached to three-axis traverse systems. One LDV probe was used to measure the separated flow field in the trailing-edge junction region. The other LDV probe was alternately used to measure the flow field in the leading-edge region of the wing and to measure the incoming fuselage boundary layer well upstream of the leading edge. Both LDV probes provided measurements from which all three mean velocity components, all six independent components of the Reynolds-stress tensor, and all ten independent components of the velocity triple products were calculated. In addition to the flow-field measurements, static and dynamic pressures were measured at selected locations on the wings and fuselage of the model, infrared imaging was used to characterize boundary-layer transition, oil-flow visualization was used to visualize the separated flow in the leading- and trailing-edge regions of the wing, and unsteady shear stress was measured at limited locations using capacitive shear-stress sensors. Sample results from the measurement techniques employed during the test are presented and discussed.

Kegerise, Michael A.↗
NASA NTRS2018

Transition Induced by Tandem Rectangular Roughness Elements on a Supersonic Flat Plate

The flow behind two rectangular roughness elements with a height approximately 38-41 percent of the boundary layer thickness was examined with a hot-wire probe. The rectangular roughness elements are oriented so that one element was at a plus 45-degree angle relative to the leading edge of the plate. A second roughness element was placed 7.16 millimeters downstream of the first one with either the same orientation relative to the leading edge of the plate, or an opposing orientation of minus 45 degrees from the leading edge. Mean mass-flux and total-temperature profiles of the flow field downstream of the tandem roughness elements were examined for mean-flow distortion. Using streak strength as a measure of mean-flow distortion, the tandem roughness elements had approximately the same amount of distortion, regardless of their relative orientation. Mass-flux fluctuation profiles show that the dominant mode downstream of the tandem roughness elements with the same orientation was similar to that of a single roughness element and centered at a frequency of approximately 55 kilohertz (kHz). The dominant instability downstream of the tandem roughness elements with opposing orientation was centered at a frequency of 65 kHz and grew more slowly than the instabilities behind the single roughness element.

Chou, Amanda↗
NASA NTRS2017

Receptivity and Forced Response to Acoustic Disturbances in High-Speed Boundary Layers

Supersonic boundary-layer receptivity to freestream acoustic disturbances is investigated by solving the Navier–Stokes equations for Mach 3.5 flow over a 7 deg half-angle cone. The freestream disturbances are generated from a wavy wall placed at the nozzle wall. The freestream acoustic disturbances radiated by the wavy wall are obtained by solving the linearized Euler equations. The results show that no noticeable instability modes are generated when the acoustic disturbances impinge the cone obliquely. The results show that the perturbations generated inside the boundary layer by the acoustic disturbances are the response of the boundary layer to the external forcing. The amplitude of the forced disturbances inside the boundary layer are about 2.5 times larger than the incoming field for zero azimuthal wave number, and they are about 1.5 times for large azimuthal wave numbers.

Balakumar, P.↗
NASA NTRS2017

Transition Induced by a Streamwise Array of Roughness Elements on a Supersonic Flat Plate

Roughness is unavoidable on practical high-speed vehicles, so it is critical to determine its impact on boundary layer transition. The flow field downstream of a streamwise array of cylindrical roughness elements is probed with hot-wire anemometry in this experiment. Mean flow distortion is examined in several measurement planes in the wake of the cylindrical roughness using the streak strength profiles and contour plots of the mass flux and total temperature. The roughness element heights and spacings were varied and their instability modes were examined. Cylindrical roughness elements approximately 140 micron tall produce an odd instability mode that grows weakly with downstream distance in the measurement range of this experiment. Cylindrical roughness elements approximately 280 micron tall produce an even instability mode that grows, becomes nonlinear, and then breaks down. Transition onset remains constant relative to the most downstream roughness in the streamwise array when the 280 micron roughness elements are spaced 2 diameters apart. Transition onset occurs at an earlier upstream location relative to the most downstream roughness in the streamwise array when the roughness elements are spaced 4 diameters appear to recover before the next downstream roughness element, so the location of transition shifts with the location of the most downstream roughness element in the array. When the rough- apart. The wake behind roughness elements spaced 2 diameters apart do not ness elements are spaced 4 diameters apart, the flow behind the first roughness element has enough space to recover before feeding into the second roughness element, and thus, moves transition forward.

Chou, Amanda↗
NASA NTRS2016

Wind Tunnel Test of a Risk-Reduction Wing/Fuselage Model to Examine Juncture-Flow Phenomena

A wing/fuselage wind-tunnel model was tested in the Langley 14- by 22-foot Subsonic Wind Tunnel in preparation for a highly-instrumented Juncture Flow Experiment to be conducted in the same facility. This test, which was sponsored by the NASA Transformational Tool and Technologies Project, is part of a comprehensive set of experimental and computational research activities to develop revolutionary, physics-based aeronautics analysis and design capability. The objectives of this particular test were to examine the surface and off-body flow on a generic wing/body combination to: 1) choose a final wing for a future, highly instrumented model, 2) use the results to facilitate unsteady pressure sensor placement on the model, 3) determine the area to be surveyed with an embedded laser-doppler velocimetry (LDV) system, 4) investigate the primary juncture corner- flow separation region using particle image velocimetry (PIV) to see if the particle seeding is adequately entrained and to examine the structure in the separated region, and 5) to determine the similarity of observed flow features with those predicted by computational fluid dynamics (CFD). This report documents the results of the above experiment that specifically address the first three goals. Multiple wing configurations were tested at a chord Reynolds number of 2.4 million. Flow patterns on the surface of the wings and in the region of the wing/fuselage juncture were examined using oil- flow visualization and infrared thermography. A limited number of unsteady pressure sensors on the fuselage around the wing leading and trailing edges were used to identify any dynamic effects of the horseshoe vortex on the flow field. The area of separated flow in the wing/fuselage juncture near the wing trailing edge was observed for all wing configurations at various angles of attack. All of the test objectives were met. The staff of the 14- by 22-foot Subsonic Wind Tunnel provided outstanding support and delivered exceptional value to the experiment, which exceeded expectations. The results of this test will directly inform the planning for the first of a series of instrumented-model tests at the same Reynolds number. These tests will be performed on a slightly larger-scale model with the selected wing, and will include off-body measurements with LDV and PIV, steady and unsteady pressure measurements, and the flow-visualization techniques that are discussed in this report.

Kegerise, Michael A.↗
NASA NTRS2016

Receptivity and Forced Response to Acoustic Disturbances in High-Speed Boundary Layers

Supersonic boundary-layer receptivity to freestream acoustic disturbances is investigated by solving the Navier-Stokes equations for Mach 3.5 flow over a sharp flat plate and a 7-deg half-angle cone. The freestream disturbances are generated from a wavy wall placed at the nozzle wall. The freestream acoustic disturbances radiated by the wavy wall are obtained by solving the linearized Euler equations. The results for the flat plate show that instability modes are generated at all the incident angles ranging from zero to highly oblique. However, the receptivity coefficient decreases by about 20 times when the incident angle increases from zero to a highly oblique angle of 68 degrees. The results for the cone show that no instability modes are generated when the acoustic disturbances impinge the cone obliquely. The results show that the perturbations generated inside the boundary layer by the acoustic disturbances are the response of the boundary layer to the external forcing. The amplitude of the forced disturbances inside the boundary layer are about 2.5 times larger than the incoming field for zero azimuthal wavenumber and they are about 1.5 times for large azimuthal wavenumbers.

Balakumar, P.↗
NASA NTRS2016

The NASA Juncture Flow Experiment: Goals, Progress, and Preliminary Testing (Invited)

NASA has been working toward designing and conducting a juncture flow experiment on a wing-body aircraft configuration. The experiment is planned to provide validation-quality data for CFD that focuses on the onset and progression of a separation bubble near the wing-body juncture trailing edge region. This paper describes the goals and purpose of the experiment. Although currently considered unreliable, preliminary CFD analyses of several different configurations are shown. These configurations have been subsequently tested in a series of "risk-reduction" wind tunnel tests, in order to help down-select to a final configuration that will attain the desired flow behavior. The risk-reduction testing at the higher Reynolds number has not yet been completed (at the time of this writing), but some results from one of the low-Reynolds-number experiments are shown.

Rumsey, Christopher L.↗
NASA NTRS2016

Measurements in a Transitioning Cone Boundary Layer at Freestream Mach 3.5

An experimental study was conducted in the Supersonic Low-Disturbance Tunnel to investigate naturally-occurring instabilities in a supersonic boundary layer on a 7 deg half- angle cone. All tests were conducted with a nominal freestream Mach number of M(sub infinity) = 3:5, total temperature of T(sub 0) = 299:8 K, and unit Reynolds numbers of Re(sub infinity) x 10(exp -6) = 9:89, 13.85, 21.77, and 25.73 m(exp -1). Instability measurements were acquired under noisy- ow and quiet- ow conditions. Measurements were made to document the freestream and the boundary-layer edge environment, to document the cone baseline flow, and to establish the stability characteristics of the transitioning flow. Pitot pressure and hot-wire boundary- layer measurements were obtained using a model-integrated traverse system. All hot- wire results were single-point measurements and were acquired with a sensor calibrated to mass ux. For the noisy-flow conditions, excellent agreement for the growth rates and mode shapes was achieved between the measured results and linear stability theory (LST). The corresponding N factor at transition from LST is N 3:9. The stability measurements for the quiet-flow conditions were limited to the aft end of the cone. The most unstable first-mode instabilities as predicted by LST were successfully measured, but this unstable first mode was not the dominant instability measured in the boundary layer. Instead, the dominant instabilities were found to be the less-amplified, low-frequency disturbances predicted by linear stability theory, and these instabilities grew according to linear theory. These low-frequency unstable disturbances were initiated by freestream acoustic disturbances through a receptivity process that is believed to occur near the branch I locations of the cone. Under quiet-flow conditions, the boundary layer remained laminar up to the last measurement station for the largest Re1, implying a transition N factor of N greater than 8:5.

King, Rudolph A.↗
NASA NTRS2014

An Experimental Study of Roughness-Induced Instabilities in a Supersonic Boundary Layer

Progress on an experimental study of laminar-to-turbulent transition induced by an isolated roughness element in a supersonic laminar boundary layer is reported in this paper. Here, the primary focus is on the effects of roughness planform shape on the instability and transition characteristics. Four different roughness planform shapes were considered (a diamond, a circle, a right triangle, and a 45 degree fence) and the height and width of each one was held fixed so that a consistent frontal area was presented to the oncoming boundary layer. The nominal roughness Reynolds number was 462 and the ratio of the roughness height to the boundary layer thickness was 0.48. Detailed flow- field surveys in the wake of each geometry were performed via hot-wire anemometry. High- and low-speed streaks were observed in the wake of each roughness geometry, and the modified mean flow associated with these streak structures was found to support a single dominant convective instability mode. For the symmetric planform shapes - the diamond and circular planforms - the instability characteristics (mode shapes, growth rates, and frequencies) were found to be similar. For the asymmetric planform shapes - the right-triangle and 45 degree fence planforms - the mode shapes were asymmetrically distributed about the roughness-wake centerline. The instability growth rates for the asymmetric planforms were lower than those for the symmetric planforms and therefore, transition onset was delayed relative to the symmetric planforms.

Kegerise, Michael A.↗
NASA NTRS2014

Characterization of a Laser-Generated Perturbation in High-Speed Flow for Receptivity Studies

A better understanding of receptivity can contribute to the development of an amplitude-based method of transition prediction. This type of prediction model would incorporate more physics than the semi-empirical methods, which are widely used. The experimental study of receptivity requires a characterization of the external disturbances and a study of their effect on the boundary layer instabilities. Characterization measurements for a laser-generated perturbation were made in two different wind tunnels. These measurements were made with hot-wire probes, optical techniques, and pressure transducer probes. Existing methods all have their limitations, so better measurements will require the development of new instrumentation. Nevertheless, the freestream laser-generated perturbation has been shown to be about 6 mm in diameter at a static density of about 0.045 kg/cubic m. The amplitude of the perturbation is large, which may be unsuitable for the study of linear growth.

Chou, Amanda↗
NASA NTRS2013

Flow Disturbance Measurements in the National Transonic Facility

Recent flow measurements have been acquired in the National Transonic Facility to assess the test-section unsteady flow environment. The primary purpose of the test is to determine the feasibility of the facility to conduct laminar-flow-control testing and boundary-layer transition-sensitive testing at flight-relevant operating conditions throughout the transonic Mach number range. The facility can operate in two modes, warm and cryogenic test conditions for testing full and semispan-scaled models. Data were acquired for Mach and unit Reynolds numbers ranging from 0.2 less than or equal to M less than or equal to 0.95 and 3.3 × 10(exp 6) less than Re/m less than 220×10(exp 6) collectively at air and cryogenic conditions. Measurements were made in the test section using a survey rake that was populated with 19 probes. Roll polar data at selected conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. This paper focuses primarily on the unsteady pressure and hot-wire results. Based on the current measurements and previous data, an assessment was made that the facility may be a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

King, Rudolph A.↗
NASA NTRS2012

An Experimental and Numerical Study of Roughness-Induced Instabilities in a Mach 3.5 Boundary Layer

Progress on a joint experimental and numerical study of laminar-to-turbulent transition induced by an isolated roughness element in a high-speed laminar boundary layer is reported in this paper. The numerical analysis suggests that transition is driven by the instability of high- and low-speed streaks embedded in the wake of the isolated roughness element. In addition, spatial stability analysis revealed that the wake flow supports multiple modes (even and odd) of convective instabilities that experience strong enough growth to cause transition. The experimental measurements, which included hot-wire and pitot-probe surveys, confirmed the existence of embedded high- and low-speed streaks in the roughness wake. Furthermore, the measurements indicate the presence of both even and odd modes of instability, although their relative magnitude depends on the specifics of the roughness geometry and flow conditions (e.g., the value of Re(sub kk) or k/delta. For the two test cases considered in the measurements (Re(sub kk) values of 462 and 319), the even mode and the odd mode were respectively dominant and appear to play a primary role in the transition process.

Kegerise, Michael A.↗
NASA NTRS2012

Flow Disturbance Characterization Measurements in the National Transonic Facility

Recent flow measurements have been acquired in the National Transonic Facility (NTF) to assess the unsteady flow environment in the test section. The primary purpose of the test is to determine the feasibility of the NTF to conduct laminar-flow-control testing and boundary-layer transition sensitive testing. The NTF can operate in two modes, warm (air) and cold/cryogenic (nitrogen) test conditions for testing full and semispan scaled models. The warm-air mode enables low to moderately high Reynolds numbers through the use of high tunnel pressure, and the nitrogen mode enables high Reynolds numbers up to flight conditions, depending on aircraft type and size, utilizing high tunnel pressure and cryogenic temperatures. NASA's Environmentally Responsible Aviation (ERA) project is interested in demonstrating different laminar-flow technologies at flight-relevant operating conditions throughout the transonic Mach number range and the NTF is well suited for the initial ground-based demonstrations. Roll polar data at selected test conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired from the rake probes included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. . Based on the current measurements and previous data, an assessment was made that the NTF is a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

King, Rudolph A.↗
NASA NTRS2011

Receptivity of Hypersonic Boundary Layers to Acoustic and Vortical Disturbances

Boundary layer receptivity to two-dimensional acoustic disturbances at different incidence angles and to vortical disturbances is investigated by solving the Navier-Stokes equations for Mach 6 flow over a 7deg half-angle sharp-tipped wedge and a cone. Higher order spatial and temporal schemes are employed to obtain the solution. The results show that the instability waves are generated in the leading edge region and that the boundary layer is much more receptive to slow acoustic waves as compared to the fast waves. It is found that the receptivity of the boundary layer on the windward side (with respect to the acoustic forcing) decreases when the incidence angle is increased from 0 to 30 degrees. However, the receptivity coefficient for the leeward side is found to vary relatively weakly with the incidence angle. The maximum receptivity is obtained when the wave incident angle is about 20 degrees. Vortical disturbances also generate unstable second modes, however the receptivity coefficients are smaller than that for the acoustic waves. Vortical disturbances first generate the fast acoustic modes and they switch to the slow mode near the continuous spectrum.

Balakamar, P.↗
NASA NTRS2011

Off-Body Boundary-Layer Measurement Techniques Development for Supersonic Low-Disturbance Flows

Investigations were performed to develop accurate boundary-layer measurement techniques in a Mach 3.5 laminar boundary layer on a 7 half-angle cone at 0 angle of attack. A discussion of the measurement challenges is presented as well as how each was addressed. A computational study was performed to minimize the probe aerodynamic interference effects resulting in improved pitot and hot-wire probe designs. Probe calibration and positioning processes were also developed with the goal of reducing the measurement uncertainties from 10% levels to less than 5% levels. Efforts were made to define the experimental boundary conditions for the cone flow so comparisons could be made with a set of companion computational simulations. The development status of the mean and dynamic boundary-layer flow measurements for a nominally sharp cone in a low-disturbance supersonic flow is presented.

Owens, Lewis R.↗
NASA NTRS2010

High-Speed Boundary-Layer Transition Induced by an Isolated Roughness Element

Progress on an experimental effort to quantify the instability mechanisms associated with roughness-induced transition in a high-speed boundary layer is reported in this paper. To simulate the low-disturbance environment encountered during high-altitude flight, the experimental study was performed in the NASA-Langley Mach 3.5 Supersonic Low-Disturbance Tunnel. A flat plate trip sizing study was performed first to identify the roughness height required to force transition. That study, which included transition onset measurements under both quiet and noisy freestream conditions, confirmed the sensitivity of roughness-induced transition to freestream disturbance levels. Surveys of the laminar boundary layer on a 7deg half-angle sharp-tipped cone were performed via hot-wire anemometry and pitot-pressure measurements. The measured mean mass-flux and Mach-number profiles agreed very well with computed mean-flow profiles. Finally, surveys of the boundary layer developing downstream of an isolated roughness element on the cone were performed. The measurements revealed an instability in the far wake of the roughness element that grows exponentially and has peak frequencies in the 150 to 250 kHz range.

Kegerise, Michael 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↗