Measurement technique and application of boundary layer transition in shock tunnel

LI Qiang; JIANG Tao; CHEN Suyu; CHANG Yu; ZHAO Lei; ZHANG Kouli

doi:10.7527/S1000-6893.2019.22740

ACTA AERONAUTICAET ASTRONAUTICA SINICA >

2019 , Vol. 40 >Issue 8: 122740 - 122740

DOI: https://doi.org/10.7527/S1000-6893.2019.22740

Fluid Mechanics and Flight Mechanics

Measurement technique and application of boundary layer transition in shock tunnel

LI Qiang ,
JIANG Tao ,
CHEN Suyu ,
CHANG Yu ,
ZHAO Lei ,
ZHANG Kouli

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Hypervelocity Aerodynamics Institute, China Aerodynamics Research and Development Center, Mianyang 621000, China

Received date: 2018-10-19

Revised date: 2018-11-11

Online published: 2019-02-15

Supported by

National Key Research and Development Program of China(2016YFA0401201)

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Abstract

Hypersonic boundary layer transition has an important influence on friction drag and heat transfer. The development of hypersonic vehicles is expected to accurately predict and control boundary layer transition. As the main ground equipment for hypersonic aerodynamic thermal environment test, the shock tunnel is an important equipment for studying hypersonic boundary layer transition. However, the original measurement technology of the tunnel is designed for engineering experiments, and it needs to be adapted and upgraded according to the characteristics of hypersonic boundary layer transition in shock tunnel. Therefore, according to the changes in physical parameters such as heat flux pressure density during the hypersonic boundary layer transition process, we develop the thin film heat flux sensor technique, the temperature-sensitive paint technique, the high frequency fluctuation pressure measurement, and the high definition schlieren visualization technique. The boundary layer transition experiment is carried out in China Aerodynamics Research and Development Center Ø2 m shock tunnel (FD-14A). The model is a 7° half-angle cone model with 0.05 mm nosetip bluntness, 10 nominal Mach number, the unit Reynolds number of 1.2×10⁷/m. Simultaneous measurements are carried out by using all the above techniques, obtaining good experimental results of boundary layer transition condition, fluctuation pressure spectrum characteristics, clear schlieren photos of the second mode wave and turbulent spots. The results obtained by these different measurement techniques can cross reference each other, and the theoretical analyses of linear stability are consistent with the experimental results.

Key words： shock tunnel; hypersonic boundary layer; transition measurement technology; high frequency fluctuation pressure; high definition schlieren visualization; second mode wave; turbulent spots

Cite this article

LI Qiang , JIANG Tao , CHEN Suyu , CHANG Yu , ZHAO Lei , ZHANG Kouli . Measurement technique and application of boundary layer transition in shock tunnel[J]. ACTA AERONAUTICAET ASTRONAUTICA SINICA, 2019 , 40(8) : 122740 -122740 . DOI: 10.7527/S1000-6893.2019.22740

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