Experiments on single oblique laminar-instability waves in a boundary layer: Introduction, growth, and transition

The laminar-turbulent transition in an incompressible flat-plate boundary layer was studied experimentally by using a spanwise array of computer-controlled surface heating elements to generate small disturbances. Oblique Tollmien-Schlichting waves were successfully introduced, and their downstream development into the intermittent region was studied using flush-mounted hot-film wall-shear sensors and dye flow visualization. Comparative studies of the development of single oblique waves were made for various wave angles, frequencies, and amplitudes. As these single oblique waves grew and began to break down, higher harmonics and subharmonics appeared in the wall shear. The amplitude of the subharmonic component decreased rapidly with increasing oblique-wave angle, so that a 10 degrees oblique wave had a subharmonic amplitude an order of magnitude below that for a two-dimensional (2-D) wave. Thus, the nonlinear mechanism that produces the subharmonic is affected by the symmetry of the primary wave. Intermittency measurements, carried out farther downstream, show that a 2-D wave is most effective in moving the transition point upstream, for a given power input

Recent insights into instability and transition to turbulence in open-flow systems

Roads to turbulence in open-flow shear layers are interpreted as sequences of often competing instabilities. These correspond to primary and higher order restructurings of vorticity distributions which culminate in convected spatial disorder (with some spatial coherence on the scale of the shear layer) traditionally called turbulence. Attempts are made to interpret these phenomena in terms of concepts of convective and global instabilities on one hand, and of chaos and strange attractors on the other. The first is fruitful, and together with a review of mechanisms of receptivity provides a unifying approach to understanding and estimating transition to turbulence. In contrast, current evidence indicates that concepts of chaos are unlikely to help in predicting transition in open-flow systems. Furthermore, a distinction should apparently be made between temporal chaos and the convected spatial disorder of turbulence past Reynolds numbers where boundary layers and separated shear layers are formed

Institute for Computational Mechanics in Propulsion (ICOMP)

The Institute for Computational Mechanics in Propulsion (ICOMP) is a combined activity of Case Western Reserve University, Ohio Aerospace Institute (OAI) and NASA Lewis. The purpose of ICOMP is to develop techniques to improve problem solving capabilities in all aspects of computational mechanics related to propulsion. The activities at ICOMP during 1991 are described

Experimental investigation of induced supersonic boundary layer transition

Turbulence onset within an initially laminar flow is one of the most common phenomenon in Fluid Mechanics, yet is an open field of research. This is due to the many and diverse causes that can trigger turbulence, which often add to each other, change their effect upon the flow velocity, and are difficult to single out in real-world situations. This is why laminar-toturbulent transition experiments have been and still are a fundamental tool for the researcher. This thesis work has addressed the study of turbulence onset in supersonic flows from an experimental point of view. Two test campaigns have been carried out each in a dedicated supersonic wind tunnel. The first aimed at tracking turbulence onset triggered by a single tetrahedral roughness element. It has been carried out in a multiple-Mach-number supersonic wind tunnel whose qualification campaign has been completed within this thesis work. Two roughness heights were tested, both for two Mach numbers, 1.6 and 2.3. The second test campaign has investigated the effects of an oblique shock wave impinging onto a Mach-2 transitional boundary layer. The transitional state of a boundary layer is that state during which the boundary layer passes from fully laminar to fully turbulent. This passage can be either induced or natural, and it was represented in this thesis work by the trigger wake and by the boundary layer flow downstream of the release point of an electric spark, respectively. Surface sensors, as thin-films and piezoelectric pressure transducers, were used to measure steady and unsteady highfrequency flow evolutions. Different wall temperatures were set for the thin-films insert as to allow the experimental estimation of the recovery temperature. Convective heat-flux trends have been extracted from the steady measurements, which, together with the recovery temperature trends, allowed the calculation of the Stanton number trends. All these quantities let to conclude on the general state of the boundary layer investigated in the first test campaign. The post-processing of the unsteady measurements yielded temperature and pressure fluctuations spectra and RMS streamwise evolutions, along with spectral time evolution at a given position. For the first test campaign, they allowed the characterization of the unsteadiness produced by the roughness within the supersonic boundary layer at different downstream locations. In so doing, they helped conclude on the state of the boundary layer, thus on the effectiveness of the roughness in triggering transition to turbulence. For the second test campaign, they allowed to single out the unsteady effects of the shock impinging downstream of the single roughness and downstream of the electric spark release point. In this way, differences in the shock effect between the roughness configuration and the clean-plate configuration have been highlighted, and the effects of different spark release frequencies compared.El desarrollo de la turbulencia en el interior de un flujo inicialmente laminar, a pesar de ser uno de los fenómenos más comunes en la mecánica de fluidos, continúa siendo un campo abierto de investigación. Esto es debido a las muchas y diversas causas que condicionan la transición de flujo laminar a régimen turbulento, a menudo actuando de modo combinado, cuyo efecto cambia con la velocidad del flujo y las cuales son difíciles de aislar en situaciones reales. Este es el motivo por el cual los experimentos que estudian la transición de régimen laminar a turbulento han sido y continúan siendo una herramienta fundamental para el investigador. Esta tesis doctoral ha abordado el estudio del comienzo de la turbulencia en flujos supersónicos desde un punto de vista experimental. Dos series de experimentos fueron realizados, cada uno en un túnel de viento supersónico específico. La primera serie tuvo como objetivo el seguimiento del inicio de la turbulencia causado por un único elemento de rugosidad de forma hexaédrica. Este tipo de experimentos fue realizado en un túnel de viento supersónico capaz de operar en un cierto rango de números de Mach, cuya caracterización fue completada en paralelo a esta investigación. Los experimentos fueron realizados a dos números de Mach, 1.6 y 2.3, y dos niveles de rugosidad diferentes variando la temperatura de pared. El objetivo de la segunda serie de medidas consistió en investigar los efectos del impacto de una onda de choque oblicua en una capa límite transitoria, en un flujo a Mach 2. El estado transitorio de una capa límite es aquel durante el cual la capa límite pasa de enteramente laminar a enteramente turbulenta. Esta transformación puede ser tanto natural como inducida. Ambos escenarios han sido reproducidos en esta tesis. La transformación natural ha sido representada mediante la presencia de un elemento de rugosidad, mientras que la transformación inducida se corresponde con el desarrollo de la capa límite aguas abajo del punto de liberación de chispas eléctricas. Para medir la evolución de flujos estacionarios y no estacionarios de alta frecuencia se emplearon sensores superficiales tales como thin-films y transductores piezoeléctricos de presión. Para determinar experimentalmente la temperatura de recuperación (recovery temperature) se aplicaron distintas temperaturas de pared a la pieza contenedora de los thin-films. De las medidas estacionarias se extrajeron las diferentes distribuciones del flujo de calor por convección, las cuales, junto con la distribución de la temperatura de recuperación, permitieron el cálculo de la distribución espacial del número de Stanton. Estos resultados permitieron llegar a una conclusión sobre el estado general de la capa límite investigada en la primera serie de experimentos. Las medidas no estacionarias proporcionaron espectros de fluctuación de temperatura y presión, y evoluciones longitudinales de residuos cuadráticos medios (MSR), así como espectros de evolución temporal en una posición dada. Estos datos permitieron caracterizar, en la primera serie de experimentos, la inestabilidad producida por la rugosidad en la capa límite supersónica en distintas posiciones aguas abajo. Ayudaron, por tanto, a determinar el estado de la capa límite y, co n ello, concluir en la efectividad de la rugosidad para provocar la transición a régimen turbulento. En la segunda serie de experimentos se pudieron señalar los efectos no estacionarios de la onda de choque incidiendo aguas debajo del elemento de rugosidad, y los efectos no estacionarios de la onda incidente aguas abajo del punto de liberación de chispas eléctricas. De este modo, se han identificado las diferencias entre el efecto de una onda de choque en presencia de un elemento de rugosidad y en el caso de la configuración limpia, y se han comparado los efectos del uso de diferentes frecuencias de descargas eléctrica

Director's Discretionary Fund Report for Fiscal Year 1997

This technical memorandum contains brief technical papers describing research and technology development programs sponsored by the Ames Research Center Director's Discretionary Fund during fiscal year 1997 (October 1996 through September 1997). Appendices provide administrative information for each of the sponsored research programs

On the question of instabilities upstream of cylindrical bodies

In an attempt to understand the unsteady vortical phenomena in perturbed stagnation regions of cylindrical bodies, a critical review of the theoretical and experimental evidence was made. Current theory is revealed to be incomplete, incorrect, or inapplicable to the phenomena observed experimentally. The formalistic approach via the principle of exchange of instabilities should most likely be replaced by a forced-disturbance approach. Also, many false conclusions were reached by ignoring that treatment of the base and perturbed flows in Hiemenz coordinate eta is asymptotic in nature. Almost surely the techniques of matched asymptotic expansions are expected to be used to capture correctly the diffusive and vorticity amplifying processes of the disturbances regarding the mean-flow boundary layer and outer potential field as eta and y/diameter approach infinity. The serious uncertainties in the experiments are discussed in detail

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

Aeronautical Engineering: A continuing bibliography with indexes (supplement 177)

This bibliography lists 469 reports, articles and other documents introduced into the NASA scientific and technical information system in July 1984

Measurements of entropy-layer instabilities over cone-ogive-cylinders at Mach 6

Predicting the onset of boundary layer transition is critical in hypersonic flight. To improve transition prediction methods, it is necessary to understand the underlying instability mechanisms that cause transition. Entropy-layer instabilities are of particular interest in the design of blunt reentry vehicles and other blunt supersonic and hypersonic vehicles. Entropy-layer instabilities from outside the boundary layer may enter the boundary layer and have a significant effect on transition. There is little experimental data for entropy-layer instabilities. ^ Experimental measurements of what appear to be entropy-layer instabilities have been made in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) using surface pressure transducers and hot-wire anemometry. A long cone-ogive-cylinder model with interchangeable cone-ogive noses was used to generate the shock curvature that resulted in an entropy layer conducive to instability growth. The nosetip angles of the cone-ogive range from 25 to 40 degrees, with a majority of the measurements taken with the sharp 30 to 35-degree nosetips. ^ Surface measurements of the entropy-layer instabilities using the 30 to 35-degree configurations show disturbances between 15 and 50 kHz. As the nosetip angle increases, the frequency of the instability decreases slightly. Results also show that the instability magnitude as measured on the model surface increases with downstream distance, then decreases, before starting to increase again. The decrease is likely due to stabilization that occurs during the entropy-layer swallowing process. ^ Off-surface measurements using hot wires have also been made for each of the cone-ogive-cylinder configurations. These measurements show the location, frequency, and relative magnitude of the entropy-layer instability. As the instability progresses downstream, it grows inside the entropy layer, then at a certain distance downstream, the instability approaches the model surface and enters the boundary layer. Results show a smooth variation of the location of this instability descent with nosetip angle. As the angle increases, the instability approaches the model further upstream. ^ Cross-correlations between the surface transducer and hot-wire anemometry measurements confirm that the same instability is being measured at both locations. Cross-correlations between axially-displaced surface sensors were used to calculate an instability convection velocity that is approximately equal to the numerically-calculated flow velocity. And cross-correlations between azimuthally-displaced sensors show that the instability is primarily axisymmetric. The model angle of attack for all measurements was nominally zero. However, the actual angle of attack may vary by up to 0.1 degrees. The experimental results were also compared with mean-flow computations for several of the model configurations