In this article, vibration and Supersonic flutter analyses are studied for trapezoidal sandwich panels. Functionally graded trapezoidal panel as well as reinforced sandwich panel by graphene nano platelets are considered. It is assumed that the graphene platelet (GPL) nanofillers are distributed in the matrix either uniformly or non-uniformly in the direction of thickness. UD, FG-X, FG-V, FG-O and FG-A are the distribution patterns of GPLs. Based on the Kant higher-order theories, the dynamic equations of sandwich panels reinforced with graphene nanoplates are obtained using extended Hamilton’ s principle. Dynamic pressure is estimated according to the quasi-stable theory of Supersonic piston. Then, using a transformation of coordinates, the governing equations and boundary conditions are converted from the original coordinates into new computational ones. Finally, the differential squares method (DQM) to obtain the natural frequencies, the shape of the modes, and the critical aerodynamic pressure is used. The effect of different porosity distribution, porosity coefficients, distribution of graphene nanoplates, weight fraction, geometry of graphene nanofillers and geometric dimensions on natural frequencies and system instability behavior are studied.

This study investigates Supersonic Flow characteristics over circular and elliptic cones at various angles of attack. Simulations were conducted on the cones with the same base area and length-to-diameter ratio. The elliptic cones considered had axis ratios of 1. 5 and 3. The angle of attack varied from 0o to 50o, with two different Mach numbers (1. 97 and 2. 94) employed for the analysis. The numerical results were compared with the experimental and theoretical findings from existing literature. The results revealed that increasing the ellipticity ratio of the cones resulted in higher lift generation. The pressure distributions on the windward and leeward sides of the cones were also examined. The results demonstrated that elliptic cones outperformed circular cones in terms of lift production, and this advantage increased with higher ellipticity ratios. Specifically, when the ellipticity ratio was increased from 1 to 3, the maximum increase in lift coefficient was 96% and 100% at Mach numbers 2. 94 and 1. 97, respectively. Additionally, by changing the ellipticity ratio from 1 to 1. 5, the maximum gain in the lift-drag ratio was 16% and 22% at Mach numbers 1. 97 and 2. 94, respectively. Notably, an elliptic cone with an ellipticity ratio of 3 achieved a remarkable 46% gain in lift-to-drag ratio compared to a circular cone. However, as the angle of attack increased, a primary bow shock formed on the windward side of the cone, with an embedded shock appearing on the leeward side.

In this work, a Supersonic turbulent Flow over a long axisymmetric body was investigated, both experimentally and computationally. The experimental study consisted of a series of wind tunnel tests for the Flow over an ogive-cylinder body at a Mach number of 1.6 and at a Reynolds number of 8´106, at angles of attack between -2 and 6 degrees. It included the surface static pressure and the boundary layer profile measurements. Further, the Flow around the model was visualized using a Schlieren technique. All tests were conducted in the transonic wind tunnel of the Qadr Research Center (QRC). Also, the same Flow at zero angle of attack was computationally simulated using a multi-block grid (with patched method around the block interfaces) to solve the thin layer Navier-Stokes (TLNS) equations. The numerical scheme used was implicit Beam and Warming central differencing, while a Baldwin-Lomax turbulence model was used to close the Reynolds Averaged Navier-Stokes (RANS) equations. The static surface pressure results show that the circumferential pressure at different nose sections varies significantly with angle of attack (in contrast to the circumferential pressure signatures along the cylindrical part of the body), while the total pressure measurements in the boundary layer vary significantly both radically and longitudinally. Two belts with various leading edge angles were installed at different locations along the cylindrical portion of the model. The computational results obtained were compared with some experimental ones (found by these authors), showing considerably close agreements.

This article reports the active control of base Flows using the experimental procedure. Active control of base pressure helps in reducing the base drag in aerodynamic devices having suddenly expanded Flows. Active control in the form of microjets having 0. 5 mm radius placed at forty-five degrees apart is employed to control the base pressure. The Mach numbers of the present analysis are 1. 7, 2. 3, and 2. 7. The length to diameter (L/D) ratio is varied from 10 to 1 and the nozzle pressure ratio (NPR) being changed from 1 to 10 in steps of 1 for base pressure measurements. The area ratio for the entire analysis is fixed at 2. 56. Wall pressure distribution along the enlarged duct is also recorded. No change in base pressure increase/decrease is thoroughly analysed as well. From the experimental investigation, it is found that control plays an important in modifying the base pressure without disturbing the wall pressure distribution. The base pressure variation is entirely different at L/D = 1 compared to a higher L/D ratio due to change in reattachment length and the requirement of the duct length at higher inertia levels. The quality of the Flow in the duct in the presence and absence of control remained the same.

An experimental investigation is conducted to calculate the shock standoff (SSO) distance in front of an acute-angled wedge. For this experimentation, simple water Flows channel analysis is carried out. The Flow velocity is varied from 13. 2 cm/s to 25. 5 cm/s increasing in steps of 1 cm/s. A velocity of 13. 2 cm/s corresponds to Froude number 1. 13 and velocity of 25. 5 cm/s to Froude number 1. 41. The Froude number ranged from 1. 13 to 1. 41 in steps of 0. 04. The study is conducted on 5 mm thick acrylic sheets and of wedge angles 50° , 60° , and 75° to obtain a relation for calculating the SSO distance concerning the Froude number. It is found that the pressure uphill strongly depends upon the Fr and wedge angle. The SSO distance determined experimentally and using the proposed correlation are found to be in good agreement.