Effect of finite width cavity in axisymmetric supersonic flow field

In this research, the effect of finite width cavities in supersonic flow field is experimentally investigated. The test facility consists of a supersonic nozzle, which provides flow Mach number of 1.9. A circular cross-sectional test section is fastened at the exit of the nozzle. Cavities are incorporated in the test section at a distance of 20 mm from the inlet. Cavities of constant length and width, and varying depth are used for the study. Entrainment of flow in the main stream is observed immediate downstream of the cavity aft edge due to three dimensional effect. As the depth of the cavity is increased, residence time of the fluid as well as the mixing characteristics are enhanced and stagnation pressure loss also increased. Twin cavities are also arranged symmetrically, which further leads to an improvement in mixing with marginal rise in stagnation pressure loss.

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Experimental study on the characteristics of axisymmetric cavity actuated supersonic flow

Proceedings of the Institution of Mechanical Engineers Part G Journal of Aerospace Engineering ◽

10.1177/0954410016667149 ◽

2016 ◽

Vol 231 (14) ◽

pp. 2570-2577

Author(s):

S Jeyakumar ◽

Shan M Assis ◽

K Jayaraman

Keyword(s):

Supersonic Flow ◽

Pressure Loss ◽

Static Pressure ◽

Effective Means ◽

Supersonic Nozzle ◽

Stagnation Pressure ◽

Complex Flow ◽

Wall Angle ◽

Blow Down ◽

Axisymmetric Cavity

The effective means of air fuel mixing and flame holding can be achieved by incorporating cavity in supersonic combustor. Understanding the complex flow field of cavity flow is essential for the design of supersonic combustor. An attempt is made to understand the characteristics of supersonic flow past axisymmetric cavity, and a series of nonreacting experiments are carried out in a blow-down type supersonic flow facility. The facility consists of a supersonic nozzle, issues a flow Mach number of 1.80 into a circular cross sectional supersonic combustor in which axisymmetric cavity is placed. Cavity of two consecutive aft wall angles is the key parameter for the study. The performance of the cavity is investigated based on the static pressure measurement, momentum flux distribution at the exit plane of the combustor, and the stagnation pressure loss of the flow. Wall static pressure distribution revealed that pressure increases with decrease in the secondary aft wall angle below 45° due to stronger recompression of shear layers. Moreover, decreasing primary aft wall angle provides a uniform mixing profile along with decrease in stagnation pressure loss across the combustor.

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Effect of Axisymmetric Aft Wall Angle Cavity in Supersonic Flow Field

International Journal of Turbo and Jet Engines ◽

10.1515/tjj-2016-0027 ◽

2018 ◽

Vol 35 (1) ◽

pp. 29-34 ◽

Cited By ~ 6

Author(s):

S. Jeyakumar ◽

Shan M. Assis ◽

K. Jayaraman

Keyword(s):

Supersonic Flow ◽

Flow Field ◽

Supersonic Nozzle ◽

Stagnation Pressure ◽

Wall Angle ◽

Blow Down ◽

Cross Sectional ◽

Axisymmetric Cavity ◽

Scramjet Combustors ◽

Stable Flow

AbstractCavity plays a significant role in scramjet combustors to enhance mixing and flame holding of supersonic streams. In this study, the characteristics of axisymmetric cavity with varying aft wall angles in a non-reacting supersonic flow field are experimentally investigated. The experiments are conducted in a blow-down type supersonic flow facility. The facility consists of a supersonic nozzle followed by a circular cross sectional duct. The axisymmetric cavity is incorporated inside the duct. Cavity aft wall is inclined with two consecutive angles. The performance of the aft wall cavities are compared with rectangular cavity. Decreasing aft wall angle reduces the cavity drag due to the stable flow field which is vital for flame holding in supersonic combustor. Uniform mixing and gradual decrease in stagnation pressure loss can be achieved by decreasing the cavity aft wall angle.

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Effect of Finite Diaphragm Rupture Process on Microshock Tube Flows

Journal of Fluids Engineering ◽

10.1115/1.4024196 ◽

2013 ◽

Vol 135 (8) ◽

Author(s):

Arun K. R ◽

H. D. Kim ◽

T. Setoguchi

Keyword(s):

Supersonic Flow ◽

Flow Field ◽

Shock Front ◽

Supersonic Nozzle ◽

Propagation Speed ◽

Propagation Distance ◽

Rupture Process ◽

Shock Propagation ◽

Diaphragm Rupture ◽

Contact Distance

The study of flow physics in microshock tubes is of growing importance with the recent development of microscale technology. The flow characteristics in a microshock tube is considerably different from that of the conventional macroshock tube due to the boundary layer effects and high Knudsen number effects. In the present study an axisymmetric computational fluid dynamics (CFD) method was employed to simulate the microshock tube flow field with Maxwell's slip velocity and temperature jump boundary conditions, to accommodate the rarefaction effects. The effects of finite diaphragm rupture process and partial diaphragm rupture on the flow field and the wave propagations were investigated, in detail. The results show that the shock propagation distance attenuates rapidly for a microshock tube compared to a macroshock tube. For microshock tubes, the contact surface comes closer to the shock front compared to the analytical macroshock tube case. Due to the finite diaphragm rupture process the moving shock front will be generated after a certain distance ahead of the diaphragm and get attenuated rapidly as it propagates compared to the sudden rupture case. The shock-contact distance reduces considerably for the finite diaphragm rupture case compared to the sudden diaphragm rupture process. A partially burst diaphragm within a microshock tube initiates a supersonic flow in the vicinity of the diaphragm similar to that of a supersonic nozzle flow. The supersonic flow expansion leads to the formation of oblique shock cells ahead of the diaphragm and significantly attenuates the moving shock propagation speed.

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Effect of Cavities on Mixing of Coaxial High-Speed Jets

International Journal of Engineering and Advanced Technology - Regular Issue ◽

10.35940/ijeat.a1076.1291s419 ◽

2019 ◽

Vol 9 (1S4) ◽

pp. 1-4

Keyword(s):

Pressure Loss ◽

Total Pressure ◽

High Speed ◽

Flux Distribution ◽

Supersonic Nozzle ◽

Atmospheric Temperature ◽

Stagnation Pressure ◽

Primary Flow ◽

Mixing Performance ◽

Cavity Configuration

Cavity performance in coaxial supersonic jets is investigated experimentally in a supersonic flow facility. The coaxial jets issued from the supersonic nozzle enter into a supersonic combustor in which the cavities are incorporated. The primary jet is maintained at a Mach of 1.32, while secondary jet is designed for Mach of 1.00, 1.11 and 1.45. The open type cavities are axisymmetric. The primary flow is maintained at a temperature of 1050K and the secondary flow is at atmospheric temperature. Static and stagnation pressures are measured by using a conventional pitot probe to analyze the quantitative mixing performance and total pressure loss. Uniform momentum flux distribution is observed in cavity configurations compared with nocavity. A more uniform mixing, as well as minimum stagnation pressure loss, is observed for cavity configuration - 4, L/D = 1.53, than other cavity configurations

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The Study of Design and Installation Process for Supersonic Nozzle

Advanced Materials Research ◽

10.4028/www.scientific.net/amr.569.500 ◽

2012 ◽

Vol 569 ◽

pp. 500-503

Author(s):

Lian Sheng Wu ◽

Guang Li Li ◽

Qi Fu

Keyword(s):

Numerical Simulation ◽

Wind Tunnel ◽

Flow Field ◽

Test Section ◽

Supersonic Nozzle ◽

Simulation Method ◽

Supersonic Wind Tunnel ◽

Numerical Simulation Method ◽

Field Quality ◽

The Impact

A practical optimal design method of supersonic nozzle is proposed for a supersonic wind tunnel’s design. Design a set of nozzle wall lines with the same nozzle length and different Mach numbers 1.5, 2.0, 2.5. Use numerical simulation method for the verify and analysis of the designed nozzle. Mainly study the impact of the installation gradient between nozzle and test section on flow field quality. This wind tunnel is the subsonic, transonic and supersonic wind tunnel and its test section cross is 0.2 m × 0.2 m .The impact on flow field quality of the test section was studied quantitatively by using the numerical simulation method. The installation gradient index was given. It has some practical value to the construction of supersonic wind tunnel. At present, this study has been applied in construction of the wind tunnel. The gradient of the test section import shall not be greater than 0.5 mm.

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Stagnation pressure effect on the supersonic flow parameters with application for air in nozzles

The Aeronautical Journal ◽

10.1017/aer.2015.13 ◽

2016 ◽

Vol 120 (1224) ◽

pp. 313-354 ◽

Cited By ~ 1

Author(s):

Merouane Salhi ◽

Toufik Zebbiche ◽

Abderrahmane Mehalem

Keyword(s):

Supersonic Flow ◽

Supersonic Nozzle ◽

Molecular Size ◽

State Equation ◽

Stagnation Pressure ◽

Design Parameters ◽

Perfect Gas ◽

Thrust Coefficient ◽

Real Gas ◽

Flow Parameters

ABSTRACTWhen the stagnation pressure of a perfect gas increases, the specific heat and their ratio do not remain constant anymore and start to vary with this pressure. The gas doesn't stay perfect. Its state equation changes and it becomes a real gas. In this case, the effects of molecular size and intermolecular attraction forces intervene to correct the state equation. The aim of this work is to determine the effect of stagnation pressure on the thermodynamic, physical and geometrical supersonic flow parameters in order to find a general form for real gas. With the assumptions that Berthelot's state equation accounts for molecular size and intermolecular force effects, expressions are developed for analysing the supersonic flow for thermally and calorically imperfect gas lower than the dissociation molecules threshold. The design parameters of the supersonic nozzle-like thrust coefficient depend directly on the stagnation parameters of the combustion chamber. The application made for air. A computation of error was made in this case to give a limit of the perfect gas model compared to the real gas model.

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A Numerical Study on the Characteristics of Air–Fuel Mixing Using a Fluidic Oscillator in Supersonic Flow Fields

Energies ◽

10.3390/en12244758 ◽

2019 ◽

Vol 12 (24) ◽

pp. 4758 ◽

Cited By ~ 1

Author(s):

Eun Cheol Lee ◽

Seung-Won Cha ◽

Hee-Soo Kwon ◽

Tae-Seong Roh ◽

Hyoung Jin Lee

Keyword(s):

Supersonic Flow ◽

Flow Field ◽

Mass Flow Rate ◽

Flow Rate ◽

Pressure Loss ◽

Total Pressure ◽

Total Pressure Loss ◽

Fuel Distribution ◽

Fluidic Oscillator ◽

Fuel Mixing

In this study, numerical simulations were conducted to confirm the possibility of improved mixing performance by using a fluidic oscillator as a fuel injector. Three-dimensional URANS non-reacting simulations were conducted to examine air–fuel mixing in a supersonic flow field of Mach 3.38. The numerical methods were validated through simulations of the oscillating flow generated from the fluidic oscillator. The results show that the mass flow rate and momentum are reduced at the outlet because the total pressure loss increases inside the fluidic oscillator, which means that higher pressure needs to be applied to supply the same mass flow rate. The simulation showed that the flow structure varies over time as the injected flow is swept laterally. With lateral injection, the fuel distribution is long and narrow, and asymmetric vortexes are generated. However, with central injection, the fuel distribution is relatively similar to the case of using a simple injector. Compared to the simple injector, the penetration length, flammable area, and mixing efficiency were improved. However, the total pressure loss in the flow field increases as well. The results showed that the supersonic fluidic oscillator could be fully utilized as a means to enhance the mixing effect, however a method to reduce the total pressure loss is necessary for practical application.

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Improvement of the Performance of a Supersonic Nozzle by Riblets

Journal of Fluids Engineering ◽

10.1115/1.1286991 ◽

2000 ◽

Vol 122 (3) ◽

pp. 585-591 ◽

Cited By ~ 4

Author(s):

Kazumi Tsunoda ◽

Tomohiko Suzuki ◽

Toshiaki Asai

Keyword(s):

Mach Number ◽

Pressure Loss ◽

Static Pressure ◽

Supersonic Nozzle ◽

Pressure Rise ◽

Internal Flow ◽

Stagnation Pressure ◽

Separation Region ◽

Loss Reduction ◽

Riblet Surface

This paper describes an experimental study of supersonic internal flow over a riblet surface mounted on a channel wall to reduce pressure loss and improve the performance of a supersonic nozzle. The magnitude of the static pressure in the pressure-rise region observed in channels with riblet surface became lower than that for a smooth surface, and the significance of its difference was indicated by uncertainty analysis estimated at 95 percent coverage. The Mach number distributions obtained by traversing a Pitot-tube showed that the separation point moved downstream and the size of the separation region became small when using riblets. Furthermore, it was found that the stagnation pressure loss reduction was as large as 56 percent in the uniform supersonic flow field at a Mach number of 2.0, and 29 percent in the separation region. [S0098-2202(00)00103-6]

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Design, Flow Field and Heat Transfer Characterization of the Conjugate Aero-Thermal Test Facility at NETL

Volume 7B: Heat Transfer ◽

10.1115/gt2020-15644 ◽

2020 ◽

Author(s):

S. Ramesh ◽

E. Robey ◽

S. A. Lawson ◽

D. Straub ◽

J. Black

Keyword(s):

Heat Transfer ◽

Flow Field ◽

Test Section ◽

Film Cooling ◽

Conjugate Heat Transfer ◽

Design Flow ◽

Test Facility ◽

Engine Operation ◽

Ir Camera ◽

Hot Gas

Abstract A new aerothermal test facility was constructed for the purpose of studying film cooling performance in an environment that accurately simulates conjugate heat transfer characteristics that exist in engine operation. This paper details the design of the facility and the plan for conducting steady-state film cooling experiments to improve the understanding of conjugate heat transfer scaling from laboratory to engine conditions. The test facility consists of two separate flow channels (hot gas/coolant) and each gas path has a flow conditioning section, a convergent nozzle and a test section/channel with viewports. Numerical simulations were conducted to predict flow field characteristics supporting the design of the flow loop facility. Preliminary experiments were conducted to characterize the flow field using velocity and temperature profile measurements. In addition, infrared (IR) thermography methods were developed to measure surface temperatures on the hot side of the test plate. The IR measurement methods including calibration of the IR camera is explained in detail. It was concluded that appropriate hot gas path flow conditioning could be achieved using a strainer-like tube, a perforated plate, and a honeycomb-mesh screen system upstream of the test section. Flow field measurements from preliminary experiments showed that the boundary layer profile follows the law of the wall.

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