Experimental Study of Bypass Transition In a Boundary Layer

Experimental Thermal and Fluid Science, 2003

This work describes a hot-wire anemometer system and an experimental procedure designed for mapping of the flow structure in disturbed boundary layers. It also briefly presents results from periodically disturbed wing boundary layers. The hotwire system consists of a number of individual anemometer units, which can be combined to operate multi-sensor probes for simultaneous measurement of velocities in many points. In the present paper, however, only a single probe was used mounted in a dedicated traverse system in order to demonstrate how ensemble averaging can provide knowledge about the growth and decay of organised flow structures, as long as they are periodic in both time and space. The information from this flow structure can then be used in the design of multi-array hot-wire probes for further studies of the non-deterministic stages of the boundary layer transition.

50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2012

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. In the current test, we acquired data for Mach and unit Reynolds numbers ranging from 0.2 ≤ M ≤ 0.95 and 3.3 × 10 6 < Re/m < 220 × 10 6 (1 × 10 6 < Re/ft < 67 × 10 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 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. The results presented in this report focus primarily on the unsteady pressure and hot-wire results in the form of P s /q, m /(m), etc. 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.

Journal of Applied Fluid Mechanics

An experimental investigation was carried out to study the turbulent flow over a flat plate in a subsonic wind tunnel. The enhanced level of turbulence was generated by five wicker grids with square meshes, and different parameters (diameter of the grid rod d = 0.3 to 3 mm and the grid mesh size M = 1 to 30 mm). The velocity of the flow was measured by means of a 1D hot-wire probe, suitable for measurements in a boundary layer. The main aim of the investigation was to explore the influence of the free stream turbulence length scale on the onset of laminar-turbulent bypass transition in a boundary layer on a flat plate. For this purpose, several transition correlations were presented, including intensity and length scales of turbulence, both at the leading edge of a plate and at the onset of transition. The paper ends with an attempt to create a correlation, which takes into account a simultaneous impact of turbulence intensity and turbulence scale on the boundary layer transition. To assess the isotropy of turbulence, the skewness factor of the flow velocity distribution was determined. Also several longitudinal scales of turbulence were determined and compared (integral scale, dissipation scale, Taylor microscale and Kolmogorov scale) for different grids and different velocities of the mean flow U = 4, 6, 10, 15, 20 m/s.

(NFAC) 40 foot by 80 foot wind tunnel at NASA Ames Research Center in the summer of 2011. The development of two measurement techniques is discussed in this work, both with the objective of making measurements on AMELIA for CFD validation. First, the work on the application of the Fringe-Imaging Skin Friction (FISF) technique to AMELIA is discussed. The FISF technique measures the skin friction magnitude and direction by applying oil droplets on a surface, exposing them to flow, measuring their thickness, and correlating their thickness to the local skin friction. The technique has the unique ability to obtain global skin friction measurements. A two foot, nickel plated, blended wing section test article has been manufactured specifically for FISF. The model is illuminated with mercury vapor lamps and imaged with a Canon 50D with a 546 nm bandpass filter. Various tests are applied to the wing in order to further characterize uncertainties related with the FISF technique. Human repeatability has uncertainties of ±2.3% of fringe spacing and ±2.0° in skin friction vector direction, while image post processing yields ±25% variation in skin friction coefficient. A method for measuring photogrammetry uncertainty is developed. The effect of filter variation and test repeatability was found to be negligible. A validation against a Preston tube was found to have 1.8% accuracy. Second, the validation of a micro flow measurement device is investigated. Anemometers have always had limited capability in making near wall measurements, driving the design of new devices capable of measurements with increased wall proximity. Utilizing a thermocouple boundary layer rake, wall measurements within 0.0025 inches of the surface have been made. A Cross Correlation Rake (CCR) has the advantage of not requiring calibration but obtaining the same proximity and resolution as the thermocouple boundary layer rake. The flow device utilizes time of flight measurements computed via cross correlation to calculate wall velocity profiles. The CCR was designed to be applied to AMELIA to measure flow velocities above a flap in a transonic flow regime. The validation of the CCR was unsuccessful. Due to the fragile construction of the CCR, only one data point at 0.10589 inches from the surface was available for validation. The subsonic wind tunnel's variable frequency drive generated noise which could not be filtered or shielded, requiring the use of a flow bench for validation testing. Since velocity measurements could not be made in the flow bench, a v comparison of a fast and slow velocity was made. The CCR was not able to detect the difference between the two flow velocities. Currently, the CCR cannot be applied on AMELIA due to the unsuccessfully validation of the device.

Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy, 2007

The present paper reports the results of a detailed experimental study on a large-scale flat plate with prescribed adverse pressure gradient producing a turbulent boundary layer separation.

International journal of engineering research and technology, 2013

Fluid mechanics is considered as one of the core subjects in mechanical engineering. The concepts of this subject are easy to understand when they are shown practically. One of such important concept in fluid mechanics is boundary layer. It has variety of applications in many fields. In practice to study this phenomenon wind tunnel is used as one can have a good control over the test section environment with this instrument. Wind tunnel is not only useful in getting the concept of boundary layer but also to have study of boundary layer over different geometries for different velocities. In this project, we have developed a specialized fixture for wind tunnel which can take readings over the different points on flat plate and thus will give the nature of boundary layer.

Journal of Applied Mechanics and Technical Physics, 1988

2007

The Pegasus wing-glove flight experiment 1-4 was designed to provide crossflow transition data at high Mach numbers, specifically to help validate stability based predictions for transition onset in a flight environment. This paper provides an analysis of the flight experiment, with emphasis on computational results for crossflow disturbances and the correlation of disturbance growth factors with in-flight transition locations via the method. Implications of the flight data for attachment line stability are also examined. Analysis of the thermocouple data reveals that transition (from turbulent to laminar flow) was first detected during the ascending flight of the rocket when the free stream Mach number exceeded about 4. Therefore, computations have been performed for flight Mach numbers of 4.13, 4.35, 4.56 and 4.99. Due to continually decreasing unit Reynolds number at higher altitudes, the entire wing-glove boundary layer became laminar at the highest flight Mach number computed above. In contrast, the boundary layer flow over the inboard tile region remained transitional up to and somewhat beyond the time of laminarization over the instrumented glove region. Linear stability predictions confirmed that the tile boundary layer is indeed more unstable to crossflow disturbances than the much colder stainless steel glove boundary layer. The transition locations based on thermocouple data from both the glove and the tile regions are found to correlate with stationary-crossflow N-factors within the range of 7 to 12.4 and with traveling mode N-factors between 7.6 and 14.1. Data from the thermocouples and hot film sensors indicates that transition from turbulent to laminar flow (i.e., laminarization) at a fixed point over the glove is generally completed within a flight time interval of 3 seconds. However, the times at which transition begins and ends as inferred from the hot film sensors are found to differ by about 2 seconds from the corresponding estimates based on the thermocouple data. N e