Analysis of stagger effects on Busemann supersonic biplane airfoil in shock tube tests by point diffraction interferometer method

• Published in: Aerospace science and technology 2022-12, Vol.131, p.107957, Article 107957
• Main Authors: Nguyen, T.D., Taguchi, M., Tsuji, K., Kashitani, M., Tanno, H., Kusunose, K.
• Format: Article
• Language: English
• Subjects: Busemann supersonic biplane; Diaphragmless shock tube system; Point diffraction interferometer; Staggered biplane
• ISSN: 1270-9638; 1626-3219
• DOI: 10.1016/j.ast.2022.107957

The Busemann biplane concept is proposed for low-boom low-drag supersonic aircraft. Studies on the supersonic biplane have been focused on supersonic flow, whereas the understanding of transonic aerodynamic characteristics is essential for aircraft design. Significantly, the complicated flow field with a shock and boundary layer interaction between Busemann biplane elements in transonic flow remains unclear. In the present study, the transonic aerodynamic characteristics of the two-dimensional Busemann biplane are investigated by shock tube tests and numerical simulations, focusing on the flow field between the biplane elements. In addition, the stagger effects, which change the relative position of the lower and upper elements in the freestream direction to avoid the choked flow between the biplane elements, are also analyzed. A shock tube system creates a test flow at Mach numbers of 0.6 to 0.8, and the Reynolds number is set at 2.85 × 105. The Point Diffraction Interferometer (PDI) method is applied to determine the density distribution around the model and the pressure coefficient on the surfaces of the biplane elements. Symmetrical density distribution is confirmed in the Busemann biplane (the baseline model, or the biplane with no stagger). A normal shock wave is generated between the biplane elements. When the Mach number increases, the normal shock wave moves downstream. The normal shock position is quantitatively stable with small fluctuation amplitudes. The fringes are strongly curved near the surfaces of the biplane elements due to the flow separation near the airfoil surfaces. When the Mach number increases, the separation area increases. When the Mach number increases, the peak value of the pressure coefficient on the surfaces of the biplane elements increases. On the other hand, the normal shock wave between the biplane elements is eliminated in the staggered models. The changes in density distribution between the biplane elements are minor compared to the case of the baseline model, which indicates that the stagger reduces the choked flow between the biplane elements. The peak value of the pressure coefficient on the wing surface increases when the stagger value increases. The stagger effects on reducing the total drag coefficient of the biplane models are confirmed.