شاهد ملکی پور

The performance of a Mach 6 Hypersonic nozzle designed to be installed in a reflective shock tunnel has been experimentally investigated in this paper. The purpose of nozzle performance is to create a uniform flow at the outlet and a suitable test time considering the starting flow inside the nozzle. This nozzle is designed using modern optimization methods based on a classic converging-diverging nozzle. Also, the design and construction of a total pressure rake is presented to evaluate the flow quality in the test section. The complexity of investigating hypersonic flows is the sensitivity of the flow to the existence of various types of disturbances and also fluctuations caused by acoustic waves. Therefore, in hypersonic flow, considerations related to the design of the test section as well as the geometry of the pressure rake are very important. In this paper, the dynamics of the reflected shock wave in the shock tube and the starting waves in the test section are investigated. The pressure upstream of the nozzle is in good agreement with the design values with an error of 4%. The Mach number distribution is presented by measuring the pressure distribution at the end of the nozzle. Finally, the Mach number distribution at the nozzle outlet has been compared with the numerical results which shows an error about 3%. The uniformity of the flow at the end of the nozzle, which has been observed numerically and experimentally, shows the effectiveness of the optimal design method of the ultrasonic nozzle.

In this paper, a systematic approach is considered to Design an optimal hypersonic Nozzle of a shock tunnel. After assigning the requirements and accomplishment of conceptual and preliminary design phases, a modern optimization strategy based on genetic algorithm and a CFD solver has been used to fine tune the nozzle convergent divergent contour. In this way, parameterization of the overall nozzle contour was done with a few control points and a Bezier curve. This arrangement showed a good flexibility to generate appropriate curves for nozzle shape. Design objectives were evaluated with a N-S viscous solver with a two equation turbulence model. Three objective functions were scalerized in a term with summation of weighted parameters: minimum total pressure loss, Mach number uniform distribution along test section and minimum axial flow deviation. A number of geometrical and physical constraints such as nozzle length, throat area, inlet and outlet diameters and inlet boundary conditions were also considered and finally, an optimized nozzle contour showed a significant improvement of about 3% in quality of the Mach 6 flow in the test section.

Most space missions fall in the hypersonic reentry regime. There are several open questions and unknown phenomena in aerothermodynamics around vehicles at flow regimes exceeding Mach 4 due to its inherent complexities. Strong entropy gradients, thick shock layers, dissociated high temperature flows, chemically reacting boundary layers, and viscous interactions are some of the unique flow features dominating the hypersonic flight regimes. It is possible to carry out a detailed analysis on spacecraft aerodynamics leading to an optimized design. Considering the complexity of the problem, desired accuracy of the analysis, and restrictions of numerical methods such as CFD, the experimental approach is preferred in most real-world projects. Due to the increasing importance of Hypersonic Aerodynamicsin modern aerospace studies, this paper reviews the recent advances, technical achievements as well asindustrial and scientific trends subsequent topresenting a brief introduction to the field.