Hypersonic boundary layer transition control using non-uniform surface temperature distributions
File(s)
Author(s)
Ozawa, Kazuki
Type
Thesis or dissertation
Abstract
Hypersonic vehicle performance is strongly constrained by laminar–turbulent transition, which produces drastic increases in skin-friction drag and surface heat flux. This thesis presents the first experimental demonstration that passive spanwise non-uniform wall-temperature variations can generate streamwise velocity streaks and measurably modify instability amplification in high-speed boundary layers. To achieve this, a novel, experimentally realisable actuation architecture was developed and tested. Two practical actuation concepts were developed: a multi-material conduction approach and a variable-thickness conduction approach. Both are designed using a three-dimensional thermal model.
A hybrid copper–ceramic flat-plate model is tested at Mach 2.75 to validate the generation of controlled spanwise temperature gradients produced passively by supersonic cooling. For an initial model temperature of 423 K, a maximum spanwise temperature difference of approximately 50K is achieved, corresponding to Θ = ΔT/T∞ ≈ 0.4, and Laser Doppler anemometry indicates a streak amplitude of approximately 3% in the highest-ΔT condition. Infrared thermography confirms the imposed surface-temperature field and the repeatability of the thermal actuation.
The concept is then assessed on a Mach 5 configuration, where the spanwise temperature nonuniformity is imposed prior to the run and provides ΔT ≈ 90K across cases (e.g. Θ ≈ 0.75 for a representative condition). High-frequency wall-pressure measurements show attenuation of the dominant instability band centred near 170 kHz, with additional reductions in peaks in the 200 − 300 kHz range relative to a uniform-temperature baseline, and infrared thermography is used to assess the corresponding surface heat-transfer response. The results demonstrate that wall-temperature patterning can be implemented in a physically realisable manner and can measurably attenuate dominant instability bands through thermally induced streak generation.
A hybrid copper–ceramic flat-plate model is tested at Mach 2.75 to validate the generation of controlled spanwise temperature gradients produced passively by supersonic cooling. For an initial model temperature of 423 K, a maximum spanwise temperature difference of approximately 50K is achieved, corresponding to Θ = ΔT/T∞ ≈ 0.4, and Laser Doppler anemometry indicates a streak amplitude of approximately 3% in the highest-ΔT condition. Infrared thermography confirms the imposed surface-temperature field and the repeatability of the thermal actuation.
The concept is then assessed on a Mach 5 configuration, where the spanwise temperature nonuniformity is imposed prior to the run and provides ΔT ≈ 90K across cases (e.g. Θ ≈ 0.75 for a representative condition). High-frequency wall-pressure measurements show attenuation of the dominant instability band centred near 170 kHz, with additional reductions in peaks in the 200 − 300 kHz range relative to a uniform-temperature baseline, and infrared thermography is used to assess the corresponding surface heat-transfer response. The results demonstrate that wall-temperature patterning can be implemented in a physically realisable manner and can measurably attenuate dominant instability bands through thermally induced streak generation.
Version
Open Access
Date Issued
2026-03-02
Date Awarded
2026-09-01
Copyright Statement
Attribution-NonCommercial 4.0 International Licence (CC BY-NC)
License URL
Advisor
Bruce, Paul
Sponsor
The Nakajima Foundation
Imperial College London
Publisher Department
Department of Aeronautics
Publisher Institution
Imperial College London
Qualification Level
Doctoral
Qualification Name
Doctor of Philosophy (PhD)