Mach and Reynolds number effects on transonic buffet on the XRF-1 transport aircraft wing at flight Reynolds number

Abstract

AbstractThis work provides an overview of aerodynamic data acquired in the European Transonic Windtunnel using an XRF-1 transport aircraft configuration both at cruise conditions and at the edges of the flight envelope. The goals and design of the wind tunnel test were described, highlighting the use of the cryogenic wind tunnel’s capability to isolate the effects of $$M_{\infty }$$ M ∞ , $${\text{Re}}_{\infty }$$ Re ∞ and the dynamic pressure q/E. The resulting dataset includes an aerodynamic baseline characterization of the full span model with vertical and horizontal tailplanes and without engine nacelles. The effects of different inflow conditions were studied using data from continuous polars, evaluating the changes in aeroelastic deformation which are proportional to q/E and the influence of $$M_{\infty }$$ M ∞ and $${\text{Re}}_{\infty }$$ Re ∞ on the shock position. Off-design data was analyzed at the lowest and highest measured Mach numbers of 0.84 and 0.90, respectively. Wing lower surface flow and underside shock motion was analyzed at negative angles of attack using $${{c}_{\text{p}}}$$ c p distribution and unsteady pressure transducer fluctuation data, identifying significant upstream displacement of the shock close to the leading edge. Wing upper-side flow and the shock motion near buffet onset and beyond was analyzed using unsteady pressure data from point transducers and unsteady pressure-sensitive paint (PSP) measurements. Buffet occurs at lower angles of attack at high Mach number, and without clearly defined lift break. Spectral contents at the acquired data points in the buffet range suggest broadband fluctuations at Strouhal numbers between 0.2 and 0.6, which is consistent with recent literature. The spanwise shock propagation velocities were determined independently via analysis of unsteady PSP and pressure transducers to be in the range between $$u_{\text{s}} / u_{\infty } = 0.24$$ u s / u ∞ = 0.24 and 0.32, which is similarly in line with published datasets using other swept wing aircraft models.