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The purpose of this paper is to conduct an experimental investigation on the shock cell structure of jets emanating from a four-lobed corrugated nozzle using Schlieren imaging technique.

The Schlieren images were captured for seven different nozzle pressure ratios (NPR = 2, 3, 4, 5, 6, 7 and 8) and compared with the shock cell structure of a round nozzle with an identical exit area. The variation in the length of the shock cell, width of boundary interaction between adjacent shock cells, maximum width of first shock cell, Mach disk position and diameter for different NPR was measured from the Schlieren images and analysed.

A three-layer shock net observed in the jet emanating from the four-lobed corrugated nozzle is a novel concept in the field of under-expanded jet flows. A shock net represents interconnected layers of shock cells developed because of the interaction between the core and peripheral shock waves in a jet emanating from a corrugated lobed nozzle. Also, the pattern of shock net is different while taking Schlieren images across the groove and lobe sections. Thus, the shock net emerging from a corrugated lobed nozzle varies azimuthally and primarily depends on the nozzle exit cross section. The length of the shock cell, width of boundary interaction between adjacent shock cells, maximum width of first cell, Mach disk position and diameter were found to exhibit increasing trend with NPR.

A novel concept of interconnected layers of shock waves defined as “shock net” developed from a single jet emanating from a four-lobed corrugated nozzle was observed.

The purpose of this paper is to carry out an extensive numerical study in order to understand the flow structures and the resulting noise generated by a supersonic impinging jet on a flat plate. One of the parameters varied in this study is the distance between the jet exit plane and the flat plate.

Because of the unsteady nature of the problem a time-dependent computation is carried out using the detached eddy simulation turbulence model. The OVERFLOW 2 CFD code was used with a highly resolved grid and small time steps.

The authors found that as the separation distance increases, the dominant frequencies in the noise spectrum decrease. In addition, the relative strength of the various frequencies to each other changes with changing distance, indicating the changing modes of the jet. The CFD results indicate a strong interaction between the acoustic waves emanating from the impingement plate and the jet plume. This feedback mechanism is responsible for destabilizing the jet shear layer leading to the jet changing modes. The computed near field spectra, convection velocities of the jet vortical structures and mean jet centerline velocity profile are in good agreement with experimental measurements. The results also show very high sound pressure levels all over the impingement plate but especially near the impingement point. These levels, if sustained, are detrimental to both human operators as well as the surrounding structures.

Given the large-scale nature of the computations carried out, it is very costly to run the computations long enough to collect a good, statistically steady time sample to achieve a low frequency bandwidth resolution. Such a long time sample could actually improve the results in terms of frequency resolution and obtained an even better agreement with experiments. Off course there is always the issue of grid resolution as well, but given the good agreement with experiments that the authors obtained, the authors are confident in their results.

The practical implications of the results the authors obtained are significant in that, the authors now know that hybrid RANS-large eddy simulation methods can be used for this complex, unsteady engineering problems. In addition, the results also show the high noise level both on the impingement surface and in the surroundings of the jet. This could have a negative impact on the structural integrity of the flat surface.

Noisy environments are never desirable anywhere especially in places where human operations take place. Therefore, given the high noise levels obtained in the simulations and confirmed by experiments, any human presence around the jet will be harmful to hearing and precautions need to be taken.

This is a physics-based study; i.e. understanding the physical phenomena involved in supersonic jet impingement. Of particular interest is the interaction of the jet shear layer with the acoustic waves emanating from the impingement area. This feedback loop is found to be responsible for intensifying the instability of the jet shear layer.

This paper presents an experimental investigation to study the effectiveness of micro jets under the influence of Over, Under, and Correct expansion to control the base pressure in suddenly expanded axi‐symmetric ducts.

Four micro jets of 1 mm orifice diameter located at 90° intervals along a pitch circle diameter of 1.3 times the nozzle exit diameter in the base region was employed as active controls. The tests were conducted for Mach numbers 1.5, 1.3, 1.6, 1.8, 2.0, 2.5, and 3.0. The jets were expanded suddenly into an axi‐symmetric tube with cross‐sectional area 2.56 times that of nozzle exit area. The L/D ratio of the sudden expansion tube was varied from 10 to 1.

From the present studies, it was found that the maximum increase in base pressure is 152 percent for Mach number 2.58. It is found that the micro jets do not adversely influence the wall pressure distribution.

As a result of developments in space flights and missile technology, the base flows at high Reynolds numbers continues to be an important area of research. Our understanding of many features of base flows remains poor, due to inadequate knowledge of turbulence, particularly in the presence of strong pressure gradient.

The paper shows that micro jets can serve as an effective controller raising the base suction to almost zero level for some combination for parameters. The nozzle pressure ratio has a definite role to play in fixing the base pressure with and without control. There is no adverse effect of the micro jets on the flow field in the duct.

Under‐ and over‐expanded jet flows are calculated for the Euler equations. The solution procedure is based on the two‐stage Runge‐Kutta time‐stepping scheme. The studies of the flow field structure in diffusers, free jets and impinging jets have been investigated for a range of jet‐to‐stream total pressure ratios and for different exit Mach numbers. The resulting flows show a complex shock‐shear expansion structure with Mach discs. Numerical results are compared with available experimental data and with previously published data. An oscillating phenomenon was observed in the case of free jets operating at sonic exit Mach number and in another case of impingement of under‐expanded jets on a flat plate.

This paper aims to present the results of an experimental investigation carried out to control the base pressure in a suddenly expanded axi‐symmetric passage.

Four micro‐jets of 1 mm orifice diameter located at 90° interval along a pitch circle diameter of 1.3 times the nozzle exit diameter in the base region was employed as active controls. The test Mach numbers were 1.25, 1.3, 1.48, 1.6, 1.8, 2.0, 2.5 and 3.0. The jets were expanded suddenly into an axi‐symmetric tube with cross‐sectional area 4.84 times that of nozzle exit area. The length‐to‐diameter ratio of the sudden expansion tube was varied from 10 to 1. Nozzles generating the above jet Mach numbers were operated with nozzle pressure ratio in the range 3‐11. As high as 40 per cent increase in base pressure was achieved. In addition to base pressure, the wall pressure in the duct was also measured.

It is found that the wall pressure is not adversely influenced by the micro jets.

The paper provides information on internal supersonic flow.

Ideal and practical performance of ram‐jet units in steady flight in the stratosphere at Mach numbers from 1·5 to 4 is examined. The effects of combustion, temperature, altitude, intake and exhaust nozzle design are considered.

This study aims to present the numerical study on supersonic jet mixing characteristics of the co-flow jet by varying lip thickness (LT). The LT chosen for the study is 2 mm, 7.75 mm and 15 mm.

The primary nozzle is designed for delivering Mach 2.0 jet, whereas the secondary nozzle is designed for delivering Mach 1.6 jet. The Nozzle pressure ratio chosen for the study is 3 and 5. To study the mixing characteristics of the co-flow jet, total pressure and Mach number measurements were taken along and normal to the jet axis. To validate the numerical results, the numerical total pressure values were also compared with the experimental result and it is proven to have a good agreement.

The results exhibit that, the 2 mm lip is shear dominant. The 7.75 mm and 15 mm lip is wake dominant. The jet interaction along the jet axis was also studied using the contours of total pressure, Mach number, turbulent kinetic energy and density gradient. The radial Mach number contours at the various axial location of the jet was also studied.

The effect of varying LT in exhaust nozzle plays a vital role in supersonic turbofan aircraft.

Supersonic co-flowing jet mixing effectiveness by varying the LT between the primary supersonic nozzle and the secondary supersonic nozzle has not been analyzed in the past.

The purpose of this paper is to present a procedure to design an experimental setup meant to validate an innovative approach for simulating, via computational fluid dynamics, a high-pressure gas release from a rupture (e.g. on an offshore oil and gas platform). The design is based on a series of scaling exercises, some of which are anything but trivial.

The experimental setup is composed of a wind tunnel, the instrumented scaled (1:10) mock-up of an offshore platform and a gas release system. A correct scaling approach is necessary to define the reference speed in the wind tunnel and the conditions of the gas release to maintain similarity with respect to the real-size phenomena. The scaling of the wind velocity and the scaling of the gas release were inspired by the approach proposed by Hall et al. (1997): a dimensionless group was chosen to link release parameters, wind velocity and geometric scaling factor.

The theoretical scaling approaches for each different part of the setup were applied to the design of the experiment and some criticalities were identified, such as the existence of a set of case studies with some release parameters laying outside the applicability range of the developed scaling methodology, which will be further discussed.

The resulting procedure is one of a kind because it involves a multi-scaling approach because of the different aspects of the design. Literature supports for the different scaling theories but, to the best of the authors’ knowledge, fails to provide an integrated approach that considers the combined effects of scaling.

The purpose of this paper is to design a double parabolic nozzle and to compare the performance with conventional nozzle designs.

The throat diameter and divergent length for Conical, Bell and Double Parabolic nozzles were kept same for the sake of comparison. The double parabolic nozzle has been designed in such a way that the maximum slope of the divergent curve is taken as one-third of the Prandtl Meyer (PM) angle. The studies were carried out at Nozzle Pressure Ratio (NPR) of 5 and also at design conditions (NPR = 3.7). Experimental measurements were carried out for all the three nozzle configurations and the performance parameters compared. Numerical simulations were also carried out in a two-dimensional computational domain incorporating density-based solver with RANS equations and SST k-ω turbulence model.

The numerical predictions were found to be in reasonable agreement with the measured experimental values. An enhancement in thrust was observed for double parabolic nozzle when compared with that of conical and bell nozzles.

Even though the present numerical simulations were capable of predicting shock cell parameters reasonably well, shock oscillations were not captured.

The double parabolic nozzle design has enormous practical importance as a small increase in thrust can result in a significant gain in pay load.

The thrust developed by the double parabolic nozzle is seen to be on the higher side than that of conventional nozzles with better fuel economy.

The overall performance of the double parabolic nozzle is better than conical and bell nozzles for the same throat diameter and length.

This study aims to investigate effects of sonic jet injection into supersonic cross-flow (JISC) numerically in different dynamic pressure ratio values and free stream Mach numbers.

Large Eddy simulation (LES) with dynamic Smagorinsky model is used as the turbulence approach. The numerical results are compared with the experimental data, and the comparison shows acceptable validation.

According to the results, the dynamic pressure ratio has critical effects on the zone related to barrel shock. Despite free stream Mach number, increasing dynamic pressure ratio leads to expansion of barrel shock zone. Consequently, expanded barrel shock zone would bring about more obstruction effect. In addition, the height of counter-rotating vortex pair increases, and the high-pressure area before jet and low-pressure area after jet will rise. The results show that the position of barrel shock is deviated by increasing free stream Mach number, and the Bow shock zone becomes stronger and close to barrel shock. Moreover, high pressure zone, which is located before the jet, decreases by high free stream Mach number.

In this study, LES with a dynamic Smagorinsky model is used as the turbulence approach. Effects of sonic JISC are investigated numerically in different dynamic pressure ratio values and free stream Mach numbers.

As summary, the following are the contribution of this paper in the field of JISC subjects: several case studies of jet condition have been performed. In all the cases, the flow at the nozzle exit is sonic, and the free stream static pressure is constant. To generate proper grid, a cut cell method is used for domain modelling. Boundary condition effect on the wall pressure distribution around the jet and velocity profiles, especially S shape profiles, is investigated. The results show that the relation between representing the location of Mach disk centre and at transonic regime is a function of second-order polynomial, whereas at supersonic regime, the relationship is modelled as a first-order polynomial. In addition, the numerical results are compared with the experimental data demonstrating acceptable validation.