## 2 Model and Experimental/Numerical Methods  
### Experimental Model and Method  
The experimental model was designed by integrating a forebody with a 2D inlet, which was equipped with two pieces of sweepback sideboards, as shown in Figure 1. The mode consisted of four main parts: I. the body; II. the pre-compression bump; III. the inlet cowl lip; and IV. the sideboards. An on-off throttle system was installed at the end of the model so that the back pressure could be imitated in the experimental process. The total length (l) and width (w) of the model were 350 mm and 100 mm, respectively. The inlet width was 32 mm, and the compression angle of the leading edge of the lip was 1\({}^{\circ}\) (angle with \(x\) axis). The height of the inlet entrance (h\({}_{1}\)) was 18.5 mm and the height of the inlet throat (h\({}_{2}\)) was 9.9 mm. More detailed dimension parameters are shown in Figure 1. Because of the small size of the model, laminar flow might extend to an excessive range on the compression surface of the forebody, which is very different from the case in full-scale aircrafts; the proportion of the laminar flow area on the compression surface is never too large. To avoid the influence of laminar flow, a transition zone was installed in front of the compression surface of the forebody. Figure 1: _Cont._  
Considering the support and motion requirements, the model was installed upside down in the wind tunnel; thus, the original Schlieren was also upside down. For the convenience of analysis, the Schlieren images shown in this paper are all rotated by 180\({}^{\circ}\). Therefore, in all the Schlieren images of this paper, the flow direction is from left to right. The sideboards could be replaced conveniently. Two pairs of sideboards named SB-1 and SB-2 were employed to conduct this experiment, as shown in Figure 2. Then, we defined the area ratio of the inner section, which was perpendicular to the flow direction and went through the point indicated by the red arrow (shown in Figure 1) into the inlet throat, as the nominal internal contraction ratio. Thus, the nominal internal contraction ratio of Model/SB-1 and Model/SB-2 was 1.485 and 1.557, respectively. Fourteen conventional pressure-measuring points and three high-frequency pressure-measuring points were arranged in the model, as shown in Table 1. The measure range of conventional pressure measuring points was 0-103 kPa, the sampling frequency was 100 Hz, and the measurement accuracy was \(\pm\)0.05%FS. The high-frequency pressure points adopted the NS-2 pressure sensor, the working voltage was 5 V, the measure range was 0-200 kPa, and the measurement accuracy was \(\pm\)0.25% FS. Combined with the DH5926 dynamic signal acquisition and processing system, the sampling frequency could reach 20 kHz. A Butterworth filter was used to filter and process the high-frequency signal, where the cutoff frequency was equal to the sampling frequency divided by 2.56. Figure 1: Experiment model: aircraft with integrated forebody and a 2D air inlet—sideboards 1 (SB-1). (**a**) Three-dimensional digital model of the experiment model—sideboards 1 (SB-1). (**b**) Photograph of the experiment model installed in wind tunnel. Figure 2: Installation photos of sideboards. (**a**) SB-1—nominal internal contraction ratio 1.485. (**b**) SB-2—nominal internal contraction ratio 1.557. Figure 3 shows the wind tunnel used for this experiment. A high-pressure gas source combined with downstream vacuum suction was the operation mechanism for the wind tunnel. Four components made up the wind tunnel. They were a thermal storage heating system, a settling chamber/attenuation chamber, a nozzle (\(\varphi\)220 mm), and a vacuum test chamber. A variable attack/sideslip angle mechanism with five degrees of motion was installed in the vacuum test chamber so that the state of the model could be adjusted continuously during the experiment.