Investigations on Hypersonic Laminar to Turbulent Boundary Layer Transition in a Shock Tunnel

Abstract

The phenomenon of laminar to turbulent boundary layer transition has baffled researchers in high speed fluid dynamics and aerodynamics for over half a century. The reason for departure of ordered laminar boundary layer to chaotic transitional and turbulent boundary layer is still not understood properly. With so many subtle factors affecting the transition onset the problem of estimating the transition onset location on an aerodynamic surface is still wide open. The boundary layer transition onset causes an increase in heat transfer over the surface, an increase in skin friction which leads to an increase in drag, an increase in mixing capability of the boundary layer. A transitional and turbulent boundary layer is better equipped to negotiate adverse pressure gradients, encountered in high speed vehicles, than laminar boundary layer. There are both pros and cons to having a transition onset in a given boundary layer, which is an inevitable phenomena given the operational conditions of high speed vehicles. The right strategy is to cleverly manipulate the phenomena of transition onset to suit the design requirements. In order to do so, a clear understanding of the phenomena is required and necessary. Plethora of work has been conducted to understand the boundary layer transition in high speed flows which has resulted in development of several key methods and correlations to estimate transition onset in simple geometries like flat plate and axisymmetric cone. The research gap in this field lies in the characterization of transitional boundary layers in hypersonic flow regime. Furthermore, techniques pertaining to noise environment characterization of shock tunnels are not well explored due to small test duration. The intermittent nature of hypersonic boundary layer is less explored and the characterization of turbulent spots needs more statistically significant data. The measurement of second mode instabilities and their association with local boundary layer properties also needs further probing. Although correlations are available for estimating transition onset location due to presence of a roughness element, their applicability when dealing with hypersonic boundary layers is still debatable as most of these correlations were developed for space shuttle orbiter and reentry vehicles. Furthermore, characterization of the wake generated by these roughness elements is seldom reported in high speed flows. In this regard, the present work, conducted in a shock tunnel facility HST4, involves design and development of a Mach 6 contoured nozzle and its characterization in terms of freestream noise spectrum. The existing HST4 was modified in the present work for a higher test time of 4.5 ms from an earlier test duration of 2.0 ms. This involved the extension of both shock tube and the test section to accommodate the increased mass of the driver and driven (test) gas. The freestream noise environment in the shock tunnel was measured by simultaneously incorporating both experimental and numerical techniques. The effect of shock ahead of the probe was removed by the use of high order numerical computations of Navier-Stokes equations, for which a higher order Navier Stokes solver for compressible flows under finite difference framework was developed in the present work. The freestream noise for HST4 in the present work was found to be around 4.32% for the freestream unit Reynolds number of 4.5 × 106/m. Once the freestream noise was calibrated and compared with the old and contemporary facilities across the world, experiments were conducted to investigate boundary layer transition onset on smooth surfaces (with Rek < 1000.0) as well as surfaces with an isolated roughness element. Some initial work was also conducted to investigate transition onset in three dimensional boundary layer developed over an elliptic cone. The transition onset on a smooth flat plate with sharp leading edge and a smooth axisymmetric cone with a sharp tip was investigated in the present work. The intermittent nature of the transitional boundary layer was probed by measurement of heat transfer on the plate and cone surfaces. Turbulent spots which are inherent to these transitional boundary layers were characterized by measuring their leading edge and trailing edge velocities. It was found that the turbulent spots developed in the transitional boundary layer developed over the cone had the leading edge velocities of 0.92Ue while the trailing edge had the velocities of 0.57Ue associated with them thereby indicating that the turbulence was spread from the turbulent spot to the surrounding laminar boundary layer. In the case of flat plate the turbulent spots had a leading edge velocity of 0.90Ue while the trailing edge velocity was found to be 0.42Ue. Spot generation rate in such transitional boundary layers was also measured and it was found that the spot generation rate varied linearly with edge unit Reynolds number. The spot generation rate in the present work varied between 500000−4500000 spots/meter/second for both the test models. The intermittency of the turbulent spots was experimentally measured and compared with the universal intermittency curves. The second mode instabilities, which are one of the prominent features in hypersonic transitional boundary layers, were measured by using high frequency response pressure transducers in the frequency bandwidth of 11 − 700 kHz. It was found that the second mode instabilities in the present work has a wavelength that was 2.5 times the local boundary layer thickness. This in turn lead to a decrease in dominant frequency of the second mode as the instabilities are convected downstream along the boundary layer. The present dataset on natural transition was also compared with the correlations and available data from facilities and flight tests conducted across the world. The boundary layer transition onset due to the presence of an isolated roughness element in the form of a cubic protuberance and a three dimensional cavity was explored in the present work. Both protuberance and cavity led to critical as well as effective transition onset in the boundary layer developed over an axisymmetric right angled cone. The data obtained from the present work also indicated the inability of shuttle correlations to predict transition onset due the presence of an isolated protuberance on the test surface. A correlation for estimating critical transition onset on an axisymmetric cone due to the presence of an isolated protuberance element was found and expressed by the following equation Reθ Me

k δ ! = Ck = 313 (1) . The pressure fluctuations in the wake of both kinds of roughness element were measured. A single frequency oscillation with a narrow bandwidth centered around 23 kHz corresponding to hair pin vortices in the wake of roughness element was found. Present work also revealed that while the protuberance suppressed the second mode instabilities in the boundary layer, the cavity aided in development of high frequency instabilities akin to second mode. Finally in order to investigate transition onset in three dimensional boundary layers, experiments were conducted to probe boundary layer transition onset on an elliptic cone for three distinct freestream unit Reynolds number in the range of 3−6 million/m. It was found that the cross-flow instabilities dominated the transition phenomena and led to an early transition onset when compared to the axisymmetric cone.