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Effects of wall cooling on hypersonic shock wave turbulent boundary layer interaction

Published online by Cambridge University Press:  24 July 2025

A.Z. Tang Open the ORCID record for A.Z. Tang [Opens in a new window] ,
B.H. Xu ,
C.X. Li  and
D.J. Ren
A.Z. Tang*
Affiliation:
Tsinghua University, Beijing, China
B.H. Xu
Affiliation:
Tsinghua University, Beijing, China
C.X. Li
Affiliation:
Tsinghua University, Beijing, China
D.J. Ren
Affiliation:
Tsinghua University, Beijing, China
*
Corresponding author: Zhenyuan Tang; Email: tangzy21@mails.tsinghua.edu.cn
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Abstract

Wall cooling is a promising method in controlling compressible flows, including hypersonic shock wave turbulent boundary layer interaction (STBLI). Based on the verified DNS method, a 30-degree compression ramp is used to generate STBLI for Ma of 5 and wall to recovery temperature ratio ranging from 0.2 to 1.0. The results indicate that the separation zone decreases for cold wall conditions and quantitatively validate the wall-temperature-corrected interaction scaling theory in recent literature. The heat transfer results show that the wall cooling greatly influences the heat flux distribution and peak values in STBLI. The two-stage heat flux increase disappears for the cold wall, which corresponds to the reduced separation bubble. The local decrease of the recovery temperature is observed after the shock, which causes the negative heat flux minimum for near ‘adiabatic’ wall conditions and can be attributed to the acceleration of the near-wall supersonic fluid in the turning process. On the whole, the decrease of the wall temperature leads to the 24.3% decrease of the peak heat flux enhancement, and the underlying mechanism is the decrease of the near-wall turbulent aerodynamic heat dissipation enhancement for the wall cooling.

Keywords

aerodynamic heatdirect numerical simulationhypersonicshock wave turbulent boundary layer interactionwall cooling

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Type
Research Article
Information
The Aeronautical Journal , Volume 129 , Issue 1341 , November 2025 , pp. 3184 - 3201
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Copyright
© The Author(s), 2025. Published by Cambridge University Press on behalf of Royal Aeronautical Society

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