Investigation Of Boundary Layer Transition Research Articles

Experimental investigation of boundary layer transition by wall-mounted elements

Experimental investigation of boundary layer transition by wall-mounted elements

The Experiments and Stability Analysis of Hypersonic Boundary Layer Transition on a Flat Plate

Experimental and linear stability theory (LST) investigation of boundary layer transition on a flat plate was conducted with a flow of Mach number 5. The temperature distributions and second-mode disturbances on the flat plate surface at different unit Reynolds number (Reunit) values were captured by infrared thermography and PCB technology, respectively, which revealed the transition location of the flat-plate boundary layer. The PCB sensors successfully captured the second-mode disturbances within the boundary layer initially at a frequency of about 100 kHz, with a gradually expanding frequency range as the distance travelled downstream increased. The evolution characteristics of the second-mode instabilities were also investigated by LST and obtained for the second mode, ranging from 100 to 250 kHz. The amplitude amplification factor (N-factor) of the second-mode instabilities was calculated by the eN method. The N-factor of the transition location in the wind tunnel experiment predicted by LST is about 0.98 and 1.25 for Reunit = 6.38 × 106 and 8.20 × 106, respectively.

Experimental investigation of boundary layer transition on a swept flat plate under variable Reynolds number

Experimental investigation of boundary layer transition on a swept flat plate under variable Reynolds number

Experimental investigation of boundary layer transition over a delta wing at Mach number 6

An experimental study on the boundary layer transition over a delta wing was carried out at Mach number 6 in a quiet wind tunnel. The Nano-tracer-based Planar Laser Scattering (NPLS) and Temperature-Sensitive Paints (TSP) techniques were used to measure the fine flow field structure and the wall Stanton number of the delta wing. The influence of factors such as the angle of attack and the Reynolds number was studied. The following results were obtained. The boundary layer transition between the leading edge and the centerline was dominated by the crossflow instability. At the location of the initial appearance of the traveling crossflow waves, the Stanton number began to rise. The Stanton number reached a maximum when the crossflow waves were broken up to turbulence. Increasing the angle of attack increased the spanwise pressure gradient at the windward side of the delta wing, thereby increasing the crossflow instability and advancing the boundary layer transition front. However, increasing the angle of attack caused the transition front to move backward at the leeward side. In addition, the sensitivity of the boundary layer transition to the Reynolds number varied with the angle of attack and the region.

Experimental investigation of boundary layer transition in flow past a bluff body

We explore the phenomenon of drag crisis observed for the flow over bluff bodies at high Reynolds numbers. The drag coefficient reduces significantly beyond a certain Re due to the transition of the boundary layer from laminar to turbulent state. Flow past a smooth sphere and a circular cylinder is experimentally investigated for 1.0 × 105 ≤ Re ≤ 5.0 × 105 via unsteady force, surface–pressure and 2–D Particle Image Velocimetry(PIV) measurements. In case of a smooth sphere, the drag crisis is observed for Re > 3.3 × 105. The unsteady force measurements reveal that the fluctuations in the force coefficients initially increase with Re in the high subcritical regime and then experience a steep fall in the critical regime. It is found from the PIV measurements that the normal Reynolds stresses in the separated shear layer from the sphere are one order lower in magnitude for the supercritical regime in comparison to the subcritical regime. In the case of flow past a smooth circular cylinder, a two–stage drag crisis is captured using surface–pressure measurements where the boundary layer over one side of the cylinder undergoes transition around Re = 3.9 × 105 and that over the second side transitions around Re = 4.8 × 105. The transition is accompanied with increased fluctuations in the surface–pressure coefficients near the shoulders of the cylinder.

Experimental investigation of boundary layer transition on rotating cones in axial flow in 0° and 35° angle of attack

In this paper, experimental results of using hot wire anemometer and smoke visualization are presented. The results obtained on the hot wire anemometer for critical Reynolds number and transitional Reynolds number are compared with previous results. Excellent agreement is found for the transitional Reynolds number. The results for the transitional Reynolds number are also compared with previous linear stability results. The results of the smoke visualization clearly show the crossflow vortices which arise in the transition process from a laminar to a turbulent flow. A non-zero angle of attack is also considered. We compare our results with linear stability theory put forward by Garrett and Peake (Eur J Mech B Fluid 26:344–353, 2007). Also the visualization and hot wire anemometer results have been compared graphically. The goal of this paper is to check reliability of using hot wire anemometer and smoke visualization in transition problems and check reliability of linear stability theory for this case and compare our results with some trusty experimental works.

Behaviour of a natural laminar flow aerofoil in flight through atmospheric turbulence

AbstractAtmospheric turbulence is encountered frequently in flight and it creates oncoming flow disturbances of varying direction and magnitude for aircraft passing through turbulent zones. The unique measurement set-up on a motorised glider enables the investigation of the flow processes acting on a laminar wing section in flight through atmospheric turbulence. The expected quasi-steady aerofoil characteristics are deduced from an investigation of boundary-layer transition under calm flight conditions. Spanwise frequency–wavenumber spectra and comparisons with linear stability theory (LST) yield insight into the linear and weakly nonlinear stages of transition. Simultaneous measurement of the oncoming flow, characteristic flow quantities on the wing section and the motion of the aerofoil enables correlations between these quantities and provides insight into the unsteady flight physics. Emphasis is placed on the response of laminar–turbulent transition to moderate free-stream turbulence on both sides of the wing section. On the lower side of the aerofoil significant and rapid upstream fluctuations of transition are observed, which correspond closely to variations in the pressure distribution. Wavelet analysis is applied to gain insight into the composition of these streamwise excursions of the transition front in the time-pseudo-frequency domain. It is shown that they are driven by rapid transient base-flow changes and that transition is initiated by a short growth stage of Tollmien–Schlichting (TS) waves.

An experimental investigation of boundary layer transition on rotating cone in axial flow

An experimental investigation of boundary layer transition on rotating cone in axial flow

The Use of Hot Films for Investigation of Boundary-Layer Transition

The constant temperature hot-film technique has been used to investigate the transition region on a modified flat plate and a large two-dimensional wing. Time histories and r.m.s. values of the output signals together with the power spectral density (PSD) clearly indicate the build up of energy associated with the development of Tollmien–Schlichting waves prior to the onset of transition. A large increase in the low-frequency energy content of the spectrum occurs at the onset of transition. Measurements of skin friction through the transition region have been obtained based on the mean heat loss of the hot film.

Fully elliptic incompressible flow calculations on regular grid

References *Mack, L. M., Stability Theory and the Problem of Supersonic Boundary-Layer Transition, AIAA Journal, Vol. 13, No. 3, 1975, pp. 278-298. Mack, L. M., Layer Linear Stability Theory, AGARD Kept. 709, June 1984. Potter, J. L., Review of the Influence of Cooled Walls on Boundary Layer Transition, AIAA Journal, Vol. 18, No. 1980, pp. 1010-1012. Sanator, R. J., DeCarlo, J. P., and Torrillo, D. T., Boundary Layer Transition Data for a Cold-Wall Slender Cone, AIAA Journal, Vol. 3, No. 4, 1965, pp. 758-760. Stetson, K. F., and Rushton, G. H., Shock Tunnel Investigation of Boundary Layer Transition at M = 5.5, AIAA Journal, Vol. 5, No. 5, 1967, pp. 899-906. Mateer, G. G., Effects of Wall Cooling and Angle of Attack on Boundary-Layer Transition on Sharp Cones at Moo = 7.4, NASA TD D-6908, Aug. 1972. Stetson, K. F., Kimmel, R. L., Thompson, E. R., Donaldson, J. C., and Siler, L. G., A Comparison of Planar and Conical Boundary Layer Stability and Transition at a Mach Number of 8, AIAA Paper 91-1639, June 1991. Kendall, J. M., Wind Tunnel Experiments Relating to Supersonic and Hypersonic Boundary Layer Transition, AIAA Journal, Vol. 13, No. 3, 1975, pp. 290-299. Demetriades, A., New Experiments on Hypersonic Boundary Layer Stability Including Wall Temperature Effects, Proceedings of the Heat Transfer and Fluid Mechanics Institute, edited by C. T. Crowe and W. L. Grosshandler, Stanford Univ. Press, Stanford, CA, 1978, pp. 39-54.

Closure to “Discussion of ‘Investigations of Boundary Layer Transition in an Adverse Pressure Gradient’” (1989, ASME J. Turbomach., 111, p. 374)

Closure to “Discussion of ‘Investigations of Boundary Layer Transition in an Adverse Pressure Gradient’” (1989, ASME J. Turbomach., 111, p. 374)

Investigations of Boundary Layer Transition in an Adverse Pressure Gradient

Boundary layer transition was measured on a flat plate for four different turbulence levels. A range of adverse pressure gradients was imposed for one of these. The zero pressure gradient results were in agreement with accepted data for transition inception, length, and turbulent spot formation rate. They were also well represented by Narasimha’s universal intermittency distribution. A surprisingly strong similarity was also exhibited by intermittency distributions under adverse pressure gradients. Dimensionless velocity profiles were reasonable for the zero pressure gradient cases but difficulties with skin-friction prediction were experienced under adverse pressure gradient conditions. For this moderate turbulence level the transition inception Reynolds number remained reasonably constant with pressure gradient. Transition lengths, however, were greatly reduced by the imposition of even a weak adverse pressure gradient. This was associated with a strong increase in turbulent spot formation rate.

Experimental investigation of boundary-layer transition at supersonic velocities

It is usually assumed that for geometrically similar bodies the state of the boundary layer and the location of the laminar-turbulent transition region are determined by the principal similarity parameters-the Mach M and Reynolds Re numbers. However, it has recently been found that at supersonic velocities the parameter (⋎/U), calculated from the free-stream velocity and viscosity and having the dimension of length, has a considerable influence on the transition.

Flight Measurements of Aerodynamic Heating and Boundary Layer Transition on the Viking 10 Nose Cone

JET aircraft and missiles are now flying at velocities high enough in the supersonic regime to experience an excessive rise in surface temperature due to the conversion of some of the kinetic energy of the vehicle into heat by compression and friction in the boundary layer. To prevent this thermal barrier from setting an upper limit on plane and missile performance, it is necessary to be able to predict this temperature rise for a particular design configuration in order that structural failure, melting, fuel evaporation, and other thermal barrier effects can be prevented in the final design. At present, it is very difficult to predict the temperature rise on a particular plane or missile. Because of the urgency of the problem, much recent theoretical work has been done on aerodynamic heating, but relatively little flight data— particularly at high Mach numbers—have been available for correlation with theoretical prediction. The experiment performed in Viking 10 on May 7, 1954, was designed to add to the small store of flight measurements of aero-heating and free-flight boundary layer transition. In order to predict this in-flight temperature rise, it is necessary to know the temperature recovery factor, the local heat-transfer coefficient, and the conditions for boundary layer transition. Except for the work of Fischer and Norris (l) on V-2 rockets No. 19 and No. 27, Sternberg (2) on V-2 No. 61, and NACA (3), very few free-flight measurements of heat transfer coefficients under transient flight conditions have been made. Most of the experimental investigations of boundary layer transition have been conducted in wind tunnels, (e.g., Czarnecki and Sinclair, NACA RM L53I18a) where there is reason to believe that local shocks, tunnel turbulence level, and angularity of the tunnel airstream may affect the supersonic transition data. The object of the present study was to correlate the state of the boundary layer determined from flight data with the Van Driest condition for complete stability of the laminar boundary layer (4). Local supersonic convective heat-transfer coefficients, derived from resistance-thermometer measurements at twentytwo points on the inner surface of the Viking No. 10 nose cone, were correlated nondimensionally with current theoretical solutions and found to agree within the accuracy of measurement. The ratio of measured wall temperature to local stream temperature was plotted vs. local Mach number for turbulent, transition, and laminar flow, and the Van Driest