Developing and applying Mach 4.5 nozzle in hypersonic wind tunnel

Mach 4.5 tests in a conventional trans-supersonic wind tunnel are often accompanied by the air liquefaction phenomenon, resulting in the low reliability of test data. The Mach 4.5 nozzle developed in a hypersonic wind tunnel is able to heat airflow and provide more accurate test data. At present, China does not have the capability to test the Mach 4.5 nozzle in the 0.5-meter hypersonic wind tunnel. This gap may be filled by developing the Mach 4.5 nozzle in the hypersonic wind tunnel. The axisymmetric nozzle profile was calculated by the inviscid flow calculation method, and the boundary layer was modified by the Sivells-Payne method. Then, the numerical simulation was carried out, and the simulation results prove that the nozzle profile thus calculated meets the design requirements of the Mach number. For its structural design, a three-section design method is adopted to ensure the continuity and smoothness of the inner surface so as to better calibrate the flow field. Standard model tests were also carried out. The test results show that the velocity field of the Mach 4.5 nozzle we developed meets technical requirements. The standard model test data provide data reliable support for the development of aircraft.

Experimental investigation of shock–shock interactions with variable inflow Mach number

Experiments on shock–shock interactions were conducted in a transonic–supersonic wind tunnel with variable free-stream Mach number functionality. Transition between the regular interaction (RI) and the Mach interaction (MI) was induced by variation of the free-steam Mach number for a fixed interaction geometry, as opposed to most previous studies where the shock generator angles are varied at constant Mach number. In this paper, we present a systematic flow-based post-processing methodology of schlieren data that enables an accurate tracking of the evolving shock system including the precise and reproducible detection of RI$$\rightleftarrows $$ ⇄ MI transition. In line with previous experimental studies dealing with noisy free-stream environments, transition hysteresis was not observed. However, we show that establishing accurate values of the flow deflections besides the Mach number is crucial to achieve experimental agreement with the von Neumann criterion, since measured flow deflections deviated significantly, up to $$1.2^{\circ }$$ 1 . 2 ∘ , from nominal wedge angles. We also report a study conducted with a focusing schlieren system with variable focal plane that supported the image processing by providing insights into the three-dimensional side-wall effects integrated in the schlieren images.

Supersonic Windtunnel Noise Attenuation

The aerodynamic generated noise in the supersonic wind tunnel during operation at Ryerson University has exceeded the threshold of hearing damage. An acoustic silencer was to be designed and added to the wind tunnel to reduce the noise level. The main sources of noise generated from the wind tunnel with the silencer were identified to be located at the convergent divergent nozzle and the turbulent region downstream of the shock wave at the diffuser with the maximum acoustic power level of the entire wind tunnel at 161.09 dB. The designed silencer provided an overall sound pressure level reduction of 21.41 db which was considered as acceptable. Refinement to the mesh size and changes to the geometry of the mixing chamber was suggested for a more accurate result in noise output as well as flow conditions would match up to the physical flow. Additional acoustic treatment should be applied to the wind tunnel to further reduce sound pressure level since the noise level still exceeded the threshold of hearing loss.

Supersonic Windtunnel Noise Attenuation

The aerodynamic generated noise in the supersonic wind tunnel during operation at Ryerson University has exceeded the threshold of hearing damage. An acoustic silencer was to be designed and added to the wind tunnel to reduce the noise level. The main sources of noise generated from the wind tunnel with the silencer were identified to be located at the convergent divergent nozzle and the turbulent region downstream of the shock wave at the diffuser with the maximum acoustic power level of the entire wind tunnel at 161.09 dB. The designed silencer provided an overall sound pressure level reduction of 21.41 db which was considered as acceptable. Refinement to the mesh size and changes to the geometry of the mixing chamber was suggested for a more accurate result in noise output as well as flow conditions would match up to the physical flow. Additional acoustic treatment should be applied to the wind tunnel to further reduce sound pressure level since the noise level still exceeded the threshold of hearing loss.

INTEGRATED APPROACH TO GAS-DYNAMIC DESIGNING OF SUPERSONIC AIR INTAKES OF AIRCRAFT

The analytical and experimental studies of the aircraft’s supersonic air intakes have been carried out. An integrated approach to the gas-dynamic designing of aircraft’s supersonic air intakes that eliminates the scale effect problem of a wind tunnel with a small-sized testing area is proposed. The designing approach accelerates the development process and reduces the resource intensity due to the rational distribution of tasks between numerical and physical experiments. The results of the unique tests of the scaled ramjet’s air intake physical model in the supersonic wind tunnel are presented.

Wavelet-based Optical Flow Analysis (wOFA) for Background Oriented Schlieren (BOS) Image Processing

A wavelet-based optical flow analysis (wOFA) method for processing background oriented schlieren (BOS) images is presented and demonstrated on synthetic and experimental data. Optical flow is inherently well-suited to BOS, since the background pattern and lighting conditions are specified and controlled by the user, and can be chosen to play to the strengths of optical flow processing. Analysis of the synthetic BOS data show that a 2D sinusoidal background produces the highest reconstruction accuracy for both wOFA and iterative least squares (ILS) algorithms. wOFA outperforms ILS in terms of overall accuracy for displacement fields with sufficiently high spatial frequency content. In addition, wOFA provides higher spatial resolution, about an order of magnitude in terms of the total number of pixels in the final BOS image. Finally, wOFA is demonstrated on two sets of experimental data, a heat gun plume experiment with nearly ideal imaging characteristics, and experiments in a supersonic wind tunnel flow with more realistic restrictions on the acquisition of images. BOS images computed with wOFA are shown to have higher spatial resolution and sensitivity than ILS, without introducing additional noise. Therefore, wOFA of BOS images are able to reveal flow features not detected by ILS analysis.