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Experiments were carried out to assess the effect of rear face geometrical modifications of a wall mounted rectangular cavity on pressure fluctuations at supersonic speed. Initially, tests were conducted on a base cavity of length to depth ratio of 3.6, in a 50 × 100 mm at a Mach number of 2, and a Reynolds number of 2 × 106 based on cavity length. Geometrical modifications to the rear face of the base cavity in the form of single ramp, double ramp and partial ramp circle were done to study their acoustic emission characteristics. Oil flow images revealed the complex nature of the flow inside the cavities. Time resolved flow visualization revealed attenuation of receptive interaction at the cavity leading edge, and minimized mass flux into the cavity by adopting these modifications. Unsteady measurements indicated reduction in the peak tones, broadband levels, and coherence. Among the modified geometries, double ramp configuration significantly reduced the overall sound pressure levels and the fluctuating pressures of the order of 10 dB and 50% respectively.

Exterior panels forming the skin of flight vehicles traveling through the atmosphere at high supersonic speed are often susceptible to the occurrence of limit-cycle type self excited vibrations called flutter. The development of methods for panel flutter control is desirable and it is the objective of this work by using piezoelectric materials. The nonlinear equations of motion for panel flutter using Von-Karman's large deflection plate theory and the quasi-steady aerodynamic theory are modified to consider the composite plate with the piezoelectric layer. The composite plate consists of two piezoelectric layers bonded perfectly to the upper and lower surface of the original panel layer. When a voltage is applied to the piezoelectric layers, an induced strain in the plate material takes place and consequently tensile in-plane load is obtained. These tensile loads shift the onset of flutter point to about twice the value of the critical dynamic pressure for the plate alone. The values of control voltage and the effect of piezoelectric material properties are investigated. The method gives good results and succeeds in suppressing the panel flutter at its operating condition.

Exterior panels forming the exterior skin of flight vehicles traveling through the atmosphere at high supersonic speed are often susceptible to the occurrence of limit cycle type self- excited vibrations called flutter. The development of methods for the prevention and control of flutter, without adversely affecting the weight and cost, is desirable and it is the objective of this work by using damping materials. The nonlinear equations of motion for panel flutter using Von-Karman's farge deflection plate theory are modified to include damping material layers. At first, a viscoelastic damping layer is considered to be perfectly bonded to the surface of the plate to investigate the extension damping treatment. Secondly, the viscoelastic damping layer over the plate is constrained by a third layer called constrained layer to investigate the shear damping treatment. The two types of damping treatment proved to be able to prevent panel flutter successfully at prescribed operating conditions. But for the same panel weight, the constrained damping layer technique is more efficient than the free damping layer technique.

Exterior panels forming the exterior skin of flight vehicles traveling through the atmosphere at high supersonic speeds are often susceptible to the occurrence of limit-cycle type self-excited vibrations called flutter. Panel flutter is resulting from the dynamic instability of the aerodynamic, inertia, and elastic forces of the system. The equations of motion for panel flutter are derived using Von Karman's large deflection plate theory and quasi-steady aerodynamic theory. The equations are given in terms of the displacement and are presented in non-dimensional form. Galerkins method is used to transform the system of nonlinear partial differential equations into a system of nonlinear ordinary differential equations in time variables. The obtained equations are then solved by the method of harmonic balance and Newton-Raplison algorithm. The results are compared with both theoretical and experimental results given in publications. Then, a parametric study is performed to study the effects of aerodynamic loading, aerodynamic damping, structural damping, in-plane applied loads, thermal stresses and cavity pressure.

Results of experimental studies of the round hydrogen microjet interaction with a coaxial air for supersonic speed jets efflux are presented in this work. It is shown that interaction of round and coaxial air jets at their supersonic efflux leads to appearance of a train small/large-scale and, on the contrary, large/small-scale supersonic cells on a resultant jet with growth and reduction of speed of its efflux, respectively. It is found that the result of the round and coaxial air jets interaction at their supersonic efflux does not depend on geometrical parameters of a round and coaxial exit nozzle.

Results of experimental studies of the round air microjet interaction with a coaxial hydrogen jet at its combustion for supersonic speed jets efflux are presented in this work. It is revealed that combustion of the coaxial hydrogen jet with growth of its speed efflux is accompanied by all scenarios, observed at study of the round and plane hydrogen microjets diffusion combustion. However, “bottleneck flame region” undergoes considerable geometrical deformations because of specifics of a flame of a coaxial jet. It is shown that “bottleneck flame region” is transformed from Y-shaped to spherical shape in the activity of growth of a coaxial jet speed efflux. It is found that a round air microjet interaction with a coaxial hydrogen jet at its combustion is accompanied by several new phenomena: existence of cone-shaped area a coaxial jet combustion near a nozzle exit; existence of small-scale supersonic cells on a resultant flame; absence of the hydrogen combustion efflux from combustion region of a coaxial jet near nozzle exit; flame-out from combustion region of a coaxial jet near nozzle exit that leads to hydrogen ignition downstream, its intensive combustion and sharp acoustic noise occurrence; existence of a turbulent flame, to its separation from a nozzle exit and transition to supersonic combustion of a resultant jet.

A body with blunt nose moving at supersonic speed generates a strong bow shock ahead of its nose. The bow shock creates high pressure on the blunt nose surface that increases aerodynamic drag on the body. The drag forcecan be substantially reduced if bow shock is replaced with oblique shockby attaching a spike in front of the blunt nose. Initially,in order to observe the effect of bluntness, various configurations are analyzed. It is observed that the drag force increases nonlinearly with bluntness. Next, a validation study is carried out to benchmark simulation methodology by reproducing published work through CFD. The purpose of present study is to optimize the shape and geometry of spike in order to reduce the drag force. A nose with small bluntness has no significant drag reduction with spike. Therefore a bluntness ratio of 0.7 was selected to see a significant effect of spike on drag reduction. A disk spike is selected from reference [1] that is validated with CFD. The effect of varying spike length on drag reduction is studied. A spike lengthis finalized that produces lower drag force at zero angle of attack over a supersonic Mach regime. Later on, the disk spike has also been analyzed with angle of attacks.

Experiments were carried out to bring out the effects of the single vertical tail (SVT) and the twin vertical tail (TVT) of a typical winged reentry vehicle on aerodynamic coefficients at supersoni...

Experimental and computational studies have been made to obtain the effect of spike shape and length on the flow over a blunt body at supersonic Mach number of 2at zero angles of attack. The spike used had a sharp tip, spherically blunt tip, and a flat aerospike tip. It is observed that use of a sharp spike leads to the reduction in drag by about 30%, whereas blunt or aerospike having the same length leads to about 45% reduction. Increasing the length of sharp spike reduces the drag. In general, the agreement between the computed and experimental result is found to be good. Flow unsteadiness has been captured through experiment and computation. An increase in unsteadiness is observed to be of the order of 20 dB due to the adoption of a spike, and the influence is mainly in the region close to spike and the main body. It has been observed that with an increase in spike length the pressure fluctuations increase and with the change in shape to blunt and aerospike the fluctuations decrease.

The uniform acceleration of a body in the axisymmetric channel has been numerically simulated. The dependencies of the flow-rate characteristics of the “channel-body” system have been obtained for various cases of geometrical blockage of the channel by the flying body. The geometrical way of passage has been proposed – by means of diffraction of the leading shock wave in the short section of sporadic channel extension.

Using a conical nozzle as an example, the gas-dynamic and thrust characteristics of high-altitude nozzles are studied in a wide range of atmospheric pressures. A geometric model of the nozzle block and the computational domain is developed in a two-dimensional axisymmetric formulation. A calculation was carried out with setting the parameters of the standard atmosphere in the altitude range from 0 to 80 km. In one of the parts of this work, the most common shock-wave structures (triple configurations of shock waves) that arise during flights at supersonic and high supersonic speeds are considered, during jet outflows from propulsion systems. We also studied the influence of the extreme values of the discontinuities of the flow field parameters on the transmitted electromagnetic signals.