There is, I think, something ethereal about i —the square root of minus one. I remember first hearing about it at school. It seemed an odd beast at that time—an intruder hovering on the edge of reality.

Usually familiarity dulls this sense of the bizarre, but in the case of i it was the reverse: over the years the sense of its surreal nature intensified. It seemed that it was impossible to write mathematics that described the real world in …

I and i

Review of combustion stabilization for hypersonic airbreathing propulsion

: The effects of nosetip bluntness on cone frustrum boundary layer transition were investigated by conducting hypersonic wind tunnel experiments. Frustrum transition Reynolds numbers were found to have large variations within the entropy layer swallowing region. Investigation results indicated that several nosetip blunting effects were competing for the dominant role in the frustrum transition process and it was speculated that the large variations in transition Reynolds number were associated with the shifting of dominant roles. Based upon these studies, dominant nosetip bluntness effects are tentatively identified with three regions of the entropy layer. Transition movement and patterns on the frustrum of a cone at small angles of attack were investigated at a Mach number of 5.9, using an 8-deg half angle cone with both sharp and blunt nosetips. Small angles of attack produced large asymmetries in the frustrum transition pattern for both sharp and blunt tipped configurations. For the configurations with nosetip bluntness most of the frustrum transition asymmetry occurred between meridian angles of 60 and 120 deg. Data obtained with simulated laminar ablated nosetips were generally representative of what would be expected with a larger spherically blunt nosetip. Keywords: Boundary layer transition; Angle of attack.

/pdf/hypersonic-laminar-boundary-layer-transition-part-1-nosetip-4b8gjsvvyz.pdf

Hypersonic Laminar Boundary Layer Transition. Part 1. Nosetip Bluntness Effects on Cone Frustum Transition. Part 2. Mach 6 Experiments of Transition on a Cone at Angle of Attack

Shock-vortex shear-layer interaction in the transonic flow around a supercritical airfoil at high Reynolds number in buffet conditions

Experimental techniques for imaging laminar-turbulent transition of boundary layers using IR thermography are presented for both flight and wind tunnel test environments. A brief overview of other transition detection techniques is discussed as motivation. A direct comparison is made between IR thermography and naphthalene flow visualization. A technique for obtaining quantitative transition location is presented.

https://file.scirp.org/pdf/OPJ_2013071515221947.pdf

Laminar-Turbulent Boundary Layer Transition Imaging Using IR Thermography

A 30° swept-wing model with a movable, leading-edge extending to 15% chord is used in low-disturbance wind-tunnel tests to study the effect of two-dimensional, step excrescences in a three-dimensional boundary layer. Forwardand aft-facing steps are modulated during the tests. Pressure measurements are compared with computational results, infrared thermography is used to globally detect boundary-layer transition, and hotwire measurements provide details of the boundary-layer profiles in the vicinity of the steps. An analysis of the results is provided including data from tests in a flight environment and from experimental studies of an unswept model of equivalent 2-D pressure gradient.

Effects of Step Excrescences on a Swept Wing in a Low-Disturbance Wind Tunnel

An in-flight test article mounted on a Cessna O-2A Skymaster studying the effect of forwardand aft-facing steps on swept-wing boundary-layer transition is described. An accurate instrumentation suite, consisting of a five-hole probe, static pressure ports, and an infrared camera is presented. Pressure coefficient profiles compare favorably with a computational validation. Infrared thermography is used as a global transition detection technique. Step height and transition location based Reynolds numbers are computed and compared with unswept transition studies. Based on the current data, the addition of the crossflow instability is believed to reduce the critical step height Reynolds number for both forwardand aft-facing steps.

Effects of Step Excrescences on Swept-Wing Transition

The interaction between an existing stationary crossflow instability and a two-dimensional step excrescence placed into a swept boundary layer on a lifting surface has been investigated computationally. Backward-facing steps were found to not amplify the stationary modes. The existence of an additional local traveling instability was found in the recirculation region following the step. Further time-resolved calculations were recommended to confirm this mode as the cause of breakdown. Forward-facing steps were found to affect the transition location only once a critical step height was exceeded. This height was associated with sudden amplification of stationary crossflow modes and transition moving forward to the step. A physics-based correlation based on linear stability theory and associated with the height of the core of the stationary crossflow vortex was proposed for that critical height, which agreed well with results for the geometry and Reynolds numbers tested experimentally. Critical forward-faci...

Computational Investigation of Step Excrescence Sensitivity in a Swept-Wing Boundary Layer

Micrometer-sized spanwise-periodic discrete roughness elements (DREs) were applied to and tested on a 30 deg swept-wing model in order to study their effects on boundary-layer transition in flight ...

https://arc.aiaa.org/doi/pdf/10.2514/1.J056897

Experiments on Discrete Roughness Element Technology for Swept-Wing Laminar Flow Control

The Subsonic Aircraft Roughness Glove Experiment (SARGE) is a hybrid natural laminar flow and passive laminar flow control flight test that will be carried out under the auspices of the NASA Environmentally Responsible Aviation initiative. The primary goals of the SARGE experiment are to 1) achieve natural laminar flow to 0.60 chord on the suction side at up to 22 million chord Reynolds number and 2) at conditions of at least 22 million chord Reynolds number, demonstrate the effectiveness of a passive array of discrete roughness elements in extending laminar flow beyond the natural transition location. The test will be conducted on a test article having 35° leading-edge sweep at 0.75 Mach and lift coefficient of at least 0.5. In cooperation with NASA Dryden Flight Research Center, Texas A&M has completed the optimized aerodynamic design of the wing glove, as well as flight test and instrumentation planning.

Matthew W. Tufts

Papers

Effects of Step Excrescences on a Swept Wing in a Low-Disturbance Wind Tunnel

Effects of Step Excrescences on Swept-Wing Transition

Computational Investigation of Step Excrescence Sensitivity in a Swept-Wing Boundary Layer

Experiments on Discrete Roughness Element Technology for Swept-Wing Laminar Flow Control

A Transonic Laminar-Flow Wing Glove Flight Experiment: Overview and Design Optimization