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Article
Publication date: 3 October 2016

Mojtaba Tahani, Mohammad Hojaji and Seyed Vahid Mahmoodi Jezeh

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.

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Practical implications

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.

Originality/value

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.

Details

Aircraft Engineering and Aerospace Technology, vol. 88 no. 6
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 11 July 2018

Xin Liu, Yuming Xing and Liang Zhao

The purpose of this study is to investigate structure parameters that influence the mixing process of droplets-gas in underwater depth-adjustable launcher cooling chamber and help…

Abstract

Purpose

The purpose of this study is to investigate structure parameters that influence the mixing process of droplets-gas in underwater depth-adjustable launcher cooling chamber and help engineers who design the launcher to distinguish the most important factor that impacts mixing performance in the cooling chamber.

Design/methodology/approach

Euler–Lagrangian droplet tracking method was used to simulate droplets-gas mixing process in the cooling chamber. The SST k-w model was adopted to simulate turbulence. Droplet breakup was described by KHRT hybrid model using modified contains which are more fit to the supersonic main flow condition.

Findings

The results show the counter-rotating vortex pairs which caused by injected liquid accelerate the mixing process. High-pressure supersonic freestream makes the liquid jet break into more small droplets due to the high momentum of the main stream. Axial injection angle has the greatest influence on Sauter mean diameter (SMD). Penetration height, SMD and total pressure loss slightly change in different tangential injection conditions. However, mixedness decreases with reduction of tangential injection angle due to a more limited space for spray developing. Enlarging orifice diameter raises penetration and mixedness greatly, while SMD and total pressure loss increase slightly.

Originality/value

The findings of this study confirm the key structure parameter to improve mixing performance in the cooling chamber. Engineers who design the underwater depth-adjustable launcher can refer the findings in this study to make control of launching power more accurate.

Details

Engineering Computations, vol. 35 no. 5
Type: Research Article
ISSN: 0264-4401

Keywords

Article
Publication date: 1 May 1996

P.Y. Tzeng and J.H. Sheu

This paper describes a study concerning the numerical simulation of asonic helium jet through a transverse nozzle in a flat plate exhaustingnormally into a supersonic air flow

Abstract

This paper describes a study concerning the numerical simulation of a sonic helium jet through a transverse nozzle in a flat plate exhausting normally into a supersonic air flow. Three‐dimensional Reynolds‐averaged Navier—Stokes equations coupled with the modified Baldwin‐Lomax algebraic turbulence model and relevant species equations are solved by using a finite‐volume upwind scheme. In this approach, Roe’s flux function, explicit multi‐stage integration and multi‐block procedure are applied to achieve the steady state solution efficiently. The Roe’s flux function is modified to suit the simulation of helium‐air mixing. The comparison between two‐dimensional calculated results with experimental data of surface pressure shows good agreement. The results of three‐dimensional computations for square, circular and rectangular jets are presented, and the essential flow features including induced shocks, upstream separations, and downstream primary and secondary vortices are adequately simulated.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 6 no. 5
Type: Research Article
ISSN: 0961-5539

Keywords

Article
Publication date: 30 October 2020

AmirMahdi Tahsini

The purpose of this paper is to analyze the effect of pressure fluctuations on the combustion efficiency of the hydrogen fuel injected into the supersonic oxidizing cross flow

Abstract

Purpose

The purpose of this paper is to analyze the effect of pressure fluctuations on the combustion efficiency of the hydrogen fuel injected into the supersonic oxidizing cross flow. The pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave.

Design/methodology/approach

The pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave. The unsteady turbulent reacting flow solver is developed to simulate the supersonic flow field in the combustion chamber with detail chemical kinetics, to predict the time-variation of the combustion efficiency due to the imposed pressure fluctuations.

Findings

The results show that the response of the reacting flow field depends on both the frequency of fluctuations and the existence of the inlet shock wave. In addition, the inlet standing shock wave has some attenuating role, but the reacting flow shows an amplifying role on imposed oscillations which is also augmented by imposing anti-phase fluctuations on both inlet and fuel flow streams.

Originality/value

This study is performed to analyze the instabilities in the supersonic combustion which has not been considered before in this manner.

Details

Aircraft Engineering and Aerospace Technology, vol. 93 no. 1
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 15 July 2022

Harish Soundararajan and Sridhar B.T.N.

This study aims to numerically study the three-dimensional (3D) flow field characteristics in a conical convergent divergent (CD) nozzle with an internal strut system to describe…

Abstract

Purpose

This study aims to numerically study the three-dimensional (3D) flow field characteristics in a conical convergent divergent (CD) nozzle with an internal strut system to describe the effect of struts on producing a side force for thrust vectoring applications.

Design/methodology/approach

Struts are solid bodies. When inserted into the supersonic region of the axisymmetric CD nozzle, it induces a shock wave that causes an asymmetric pressure distribution predominantly over the internal surface of the diverging wall of the C-D nozzle, creating a net side force similar to the secondary injection thrust vectoring control method. Numerical simulations were performed by solving Unsteady Reynolds Averaged Navier–Stokes equations with re-normalized group k–ϵ turbulence model. Cylindrical struts of various heights positioned at different locations in the divergent section of the nozzle were investigated at a nozzle pressure of 6.61.

Findings

Thrust vectoring angle of approximately 3.8 degrees was obtained using a single cylindrical strut with a dimensionless thrust (%) and total pressure loss of less than 2.36% and 2.67, respectively. It was shown that the thrust deflection direction could also be changed by changing the strut insertion location. A strut located at half of the diverging length produced a higher deflection per unit total pressure loss.

Practical implications

Using a lightweight and high-temperature resistant material, such as a strut, strut insertion-based thrust vectoring control might provide an alternative thrust vectoring method in applications where a longer period of control is needed with a reduced overall system weight.

Originality/value

This study describes the 3D flow field characteristics which result in side force generation by a supersonic nozzle with an internal strut.

Details

Aircraft Engineering and Aerospace Technology, vol. 95 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 20 July 2023

Mehdi Mohamadi and AmirMahdi Tahsini

The purpose of this study is to investigate the combustion of the n-Heptane droplets in the supersonic combustor with a cavity-based fuel injection configuration. The focus is on…

Abstract

Purpose

The purpose of this study is to investigate the combustion of the n-Heptane droplets in the supersonic combustor with a cavity-based fuel injection configuration. The focus is on the impacts of the droplet size on combustion efficiency.

Design/methodology/approach

The finite volume solver is developed to simulate the two-phase reacting turbulent compressible flow using a single step reaction mechanism as finite rate chemistry. Three different fuel injection settings are studied for the considered physical geometry and flow conditions: the gas fuel injection, small droplet liquid fuel injection and big droplet fuel. The fuel is injected as a slot wall jet from the bottom of the cavity.

Findings

The results show that using the small droplet size, the complete fuel consumption and combustion efficiency can be achieved but using the big droplet sizes, most fuel exit the combustor in the liquid phase and gasified unburned fuel. It is also demonstrated that the cavity's temperature distribution of the liquid fuel case is different from the gas fuel, and two flame branches are observed there due to the droplet evaporation and combustion in the cavity.

Originality/value

To the best of the authors’ knowledge, this study is performed for the first time on the combustion of the n-Heptane fuel droplets in scramjet configuration, which is promising propulsion system for the future economic flights.

Details

Aircraft Engineering and Aerospace Technology, vol. 95 no. 10
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 25 September 2021

Sathish Kumar K, Naren Shankar R, Anusindhiya K and Senthil Kumar B.R.

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…

Abstract

Purpose

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.

Design/methodology/approach

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.

Findings

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.

Practical implications

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

Originality/value

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.

Details

Aircraft Engineering and Aerospace Technology, vol. 94 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 9 January 2024

Kathiravan Balusamy, Vinothraj A. and Suresh V.

The purpose of this study is to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body…

Abstract

Purpose

The purpose of this study is to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body. Specifically, the study aims to analyze the impact of varying the length of the cylindrical rod in the aerospike (ranging from 0.5 to 2.0 times the diameter of the blunt body) and the diameter of the hemispherical disk (ranging from 0.25 to 0.75 times the blunt body diameter). CFD simulations were conducted at a supersonic Mach number of 2 and a Reynolds number of 2.79 × 106.

Design/methodology/approach

ICEM CFD and ANSYS CFX solver were used to generate the three-dimensional flow along with its structures. The flow structure and drag coefficient were computed using Reynolds-averaged Navier–Stokes equation model. The drag reduction mechanism was also explained using the idea of dividing streamline and density contour. The performance of the aero spike length and the effect of aero disk size on the drag are investigated.

Findings

The separating shock is located in front of the blunt body, forming an effective conical shape that reduces the pressure drag acting on the blunt body. It was observed that extending the length of the spike beyond a specific critical point did not impact the flow field characteristics and had no further influence on the enhanced performance. The optimal combination of disk and spike length was determined, resulting in a substantial reduction in drag through the introduction of the aerospike and disk.

Research limitations/implications

To predict the accurate results of drag and to reduce the simulation time, a hexa grid with finer mesh structure was adopted in the simulation.

Practical implications

The blunt nose structures are primarily employed in the design of rockets, missiles, and re-entry capsules to withstand higher aerodynamic loads and aerodynamic heating.

Originality/value

For the optimized size of the aero spike, aero disk is also optimized to use the benefits of both.

Details

Aircraft Engineering and Aerospace Technology, vol. 96 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 23 August 2021

Jefte da Silva Guimarães, Valéria Serrano Faillace Oliveira Leite, Marco Antonio Sala Minucci and Dermeval Carinhana

The purpose of this paper is to demonstrate the aerodynamic behavior of a supersonic combustion test bench (SCTB) components, as the transition piece and the combustor of a…

Abstract

Purpose

The purpose of this paper is to demonstrate the aerodynamic behavior of a supersonic combustion test bench (SCTB) components, as the transition piece and the combustor of a scramjet (supersonic combustion ramjet), manufactured by 3D printing or additive manufacturing (AM).

Design/methodology/approach

For the dimensional and structural analysis of the manufactured models, a portable 3D scanner was used to generate the mesh of its dimensions, and to compare them before and after the experiments, a roughness measuring system was also used to verify the roughness inside the models before and after the tests, as roughness is an important parameter because it directly affects the boundary layer. For the visualization of the flow, the non-intrusive schlieren optical technique was used.

Findings

The experiments were carried out on the SCBT for Mach 2 flows, using the manufactured prototypes and showed that there was no structural and dimensional change of the model after the test batteries. It was found that the roughness presented by the material did not affect the quality of the flow generated. This shows that the investigated material can also be applied in experiments with supersonic flow.

Originality/value

This paper presents that it is possible to use in ground test facilities, for the studies of supersonic flow (in cold condition), pieces and models manufactured by 3D printing without affecting the quality of the flow generated during the experiments. This study presents a new perspective to approach AM applied in the studies of supersonic flows.

Details

Rapid Prototyping Journal, vol. 27 no. 8
Type: Research Article
ISSN: 1355-2546

Keywords

Article
Publication date: 1 August 1946

D.M. Davies

THE earlier classical treatises on aerodynamics concerned themselves with the properties of incompressible fluids. The theory developed on this basis gave an excellent theoretical…

Abstract

THE earlier classical treatises on aerodynamics concerned themselves with the properties of incompressible fluids. The theory developed on this basis gave an excellent theoretical background to the aeronautical engineer and made possible a scientific approach to the problems of aircraft flight. With the steady increase of aircraft speed, however, it soon became evident that the theory would have to be extended to take compressibility into account. One important result, brought out by Glauert's analysis, was the modification of the flow pattern with increasing Mach number. A more striking divergence of compressible from incompressible flow, first encountered at near sonic speeds, is the occurrence of shock waves. A shock wave, in the specialized aeronautical sense, is a pressure impulse travelling through the flow causing a sudden transition from supersonic to subsonic speeds (normal to the wave front) with an attendant increase in pressure and temperature. A brief statement of this sort, however, is of little or no value in giving an idea of the physical nature of the phenomenon. A considerable amount of attention is now focused on the repercussions of shock waves on aeroplane design. It is far easier to understand these design trends if one has a good grasp of the fundamentals underlying the problem. This article sets out to give a brief survey of these fundamentals. It is not easy also to give a complete physical picture of a shock wave but at least a discussion of their formation, propagation, etc. goes a long way towards clarifying one's ideas.

Details

Aircraft Engineering and Aerospace Technology, vol. 18 no. 8
Type: Research Article
ISSN: 0002-2667

1 – 10 of 752