The present research investigates the effect of the location and the width of single shallow circumferential groove casing treatment on the flow field and the stability improvement of NASA Rotor 37 utilizing the help of computational fluid dynamics. At first, steady state simulation of Rotor37 was presented for smooth casing (without groove). Then, forty five various grooved casing were simulated and compared with the smooth casing. The results indicated that narrow grooves had slight effect on the adiabatic efficiency but as the width of the groove was increased, a decline in efficiency was observed. The investigation on the stall margin revealed that narrow grooves next to the leading edge could improve the stall margin by a reduction in the size of vortex breakdown zone. Medium-width grooves displayed an effective role in delaying the separation-produced by shock wave and boundary layer interaction-on the blade suction side near the casing. This type of grooves could improve the stall margin more than narrow grooves when located on the top of separation zone near the blade suction side. Wide grooves had negative effect on the stall margin and caused a significant drop in the efficiency and the total pressure ratio of the compressor.

In this paper, a code is developed in C++ programming language aiming to select and introduce a well-chosen propulsion system for appropriate operation in an aircraft. By using a suitable turbocharged engine, the inlet pressure of the engine manifold will increase to a level equal to the pressure at sea level. Therefore, the aircraft engine will not notice the drop of pressure caused by the increase in altitude. Consequently, the engine power will not be reduced. On the other hand, at high altitude, using only one turbocharger is not adequate to supply engine inlet pressure and flow rate equivalent to sea level conditions, requiring the use of more turbochargers. The code developed in this study will be able to introduce the appropriate turbocharged engine based on the flight altitude and the required power engine of an air vehicle. The altitude defined in this code ranges from 5 to 30 kilometers, which leads to a selection of one to three turbochargers plus a number of intercoolers, according to user input parameters. The objective function of this optimization problem is defined as a function of turbochargers efficiency. However, this objective function can be modified according to the user requests and requirements.

In this paper, aerodynamic optimization of the tangential stacking line of the NASA Rotor 37 as a transonic axial flow compressor rotor is carried out using computational fluid dynamics and genetic algorithm. To cover a wide range of curves with a minimum number of design parameters, a B-spline curve with three control points at 33, 66 and 100% of the blade span is used to define the blade stacking lines. Firstly, by rotating the tangential position of the control points, different rotors have been created and are simulated using the Navier-Stokes governing equations. Then, using genetic algorithm operators, based on the adiabatic efficiency as an objective function, new blades are created and numerically simulated. This process is repeated to achieve maximum adiabatic efficiency. The comparison of the optimum blade and the original blade indicates that optimal tangential stacking line causes the shock wave to move downstream and reduce the secondary flow which has led to an improvement of about 1.7% of the adiabatic efficiency.

The conventional method for cavity analysis is solving two-dimensional equations. The two-dimensional implicit and densitybased Reynolds averaged Navier-Stokes equations and the two-equation standard k-ε turbulence model have been employed to numerically simulate the cold flow field in a single-cavity flame-holding configuration of a supersonic combustor. The cross section of the combustor is assumed to be rectangular. The supersonic inlet is supposed for the steady and unsteady flow conditions along with normal directions to the inlet. For the validation purpose, the numerical results are compared with those of the experimental data available in the current literature. It is quite well-known that the cavity in supersonic combustors helps to separate the fuel from the wall configuration while improving the mixing process in supersonic flows. However, the selection of the most efficient depth for the cavity is crucial in obtaining optimum conditions. In the present research work, the role of the cavity length-to-height (L/D) ratio, the channel height-to-cavity height (H/D) ratio and Mach number are studied numerically. The obtained results indicate that the wall static pressure profiles of validation case predicted by the numerical approaches are well in agreement with those of the experimental data. Also, H/D = 2 was found to be the best choice for the combustion chamber height relative to the cavity depth. The designed geometry is modeled in a commercial software using two-dimensional density-based energy equations while the turbulent characteristics are modeled using standard k-ε turbulence model.