Study on the Obstruction Effect for a Large Explosion Shock Wave Tube Used the Ideal Nozzle Theory

2013 ◽  
Vol 378 ◽  
pp. 87-90
Author(s):  
Chao Cheng Wang ◽  
Hui Qi Ren ◽  
Hai Lu Wang
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Wave Tube ◽  
Large Explosion ◽  
Fundamental Equations ◽  
The Given ◽  
The Relationship ◽  
The Ideal

This paper presents a calculation on the obstruction effects for the given large explosion shock wave tube using the ideal nozzle the theory. The relationship among Mach number, Mach number ratio, dynamic pressure ratio in the nozzle throat and blocking area ratio are established according to the fundamental equations of one-dimensional steady flow, which can be taken as the reference of blocking limit design.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Vol 126 (4) ◽  
pp. 473-481 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Turbulent jet in crossflow analysis with LES approach

2016 ◽  
Vol 88 (6) ◽  
pp. 717-728 ◽  
Author(s):  
Mojtaba Tahani ◽  
Mohammad Hojaji ◽  
Seyed Vahid Mahmoodi Jezeh
Keyword(s):  
Mach Number ◽  
Free Stream ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Content Type ◽  
Order Polynomial ◽  
Pressure Area ◽  
Barrel Shock ◽  
Free Stream Mach Number ◽  
Shock Zone

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.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

2004 ◽  
Author(s):  
Hirotaka Higashimori ◽  
Kiyoshi Hasagawa ◽  
Kunio Sumida ◽  
Tooru Suita
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Centrifugal Compressor ◽  
Transonic Flow ◽  
Reverse Flow ◽  
Pressure Ratio ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and CFD. This report presents the measured results of the high transonic flow at the impeller inlet using LDV and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460m/s and the relative Mach number reaches about 1.6. Using an LDV, about 500m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Boundary Layer Suction via a Slot in a Transonic Compressor: Numerical Parameter Study and First Experiments

2004 ◽  
Author(s):  
K. Hubrich ◽  
A. Bo¨lcs ◽  
P. Ott
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Parameter Study ◽  
Flow Angle ◽  
Boundary Layer Suction ◽  
Transonic Compressor ◽  
Shock Mach Number

In the present paper a numerical and experimental study aiming at the enhancement of the working range of a transonic compressor via boundary layer suction (BLS) is presented. The main objective of the investigation is to study the influence of BLS on the interference between shock wave and boundary layer and to identify the possible benefit of BLS on the compressor working characteristics. An extensive numerical study has been carried out for the DATUM blade and for 2 different suction location configurations for one speed line and varying back-pressure levels, ranging from choked conditions to stall. It was found that the working range of the transonic compressor with a nominal inlet Mach number of 1.2 and a nominal pre-shock Mach number of 1.35 could be increased by sucking 2% of flow on the SS away, in such a way that the maximum pressure ratio and maximum diffusion could both be increased by 10%, when compared to the DATUM case. For smaller pressure ratios with respect to the design pressure ratio, the BLS is located in a supersonic flow region and thus creates additional losses due to a more divergent flow channel, which additionally accelerates the flow and results in a higher pre-shock Mach number creating higher losses. First measurements carried out in LTTs annular cascade, do show reasonable agreement with the computations in terms of inlet Mach number, flow angle, main shock location and stall limit. The most pronounced difference between measurements and computations is the occurrence of a terminal normal channel shock behind a bowed detached shock wave and a separation on the SS of the blade, which were both not predicted by the CFD.

Investigation of Transient Shock Wave in Supersonic Wind Tunnel

2012 ◽  
Vol 232 ◽  
pp. 228-233
Author(s):  
Behnam Ghadimi ◽  
Mojtaba Dehghan Manshadi ◽  
Mehrdad Bazazzadeh
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Wind Tunnel ◽  
Pressure Ratio ◽  
Wind Tunnels ◽  
Supersonic Wind Tunnel ◽  
Blow Down ◽  
Fortran Code ◽  
Nozzle Contour ◽  
Fluent Software

Wind tunnels are the experimental apparatuses which provide an airstream flowing under controlled conditions so that interesting items in aerospace engineering such as pressure and velocity can be tested. In this work, Shock wave passes through the intermittent blow-down wind tunnel at Mach=2,3,4 has been investigated. The shape of the nozzle contour for a given Mach number was determined using the method of characteristics. For this purpose MATLAB code was developed and this code was verified with Osher’s and AUSM methods, FORTRAN code and FLUENT software was used for these two methods, respectively. Dimensions of different parts of wind tunnel are determined and minimum pressure ratio for the starting condition has been founded using FLUENT software. Good agreement was considered compared with the data from eleven tunnels over their range of Mach number.

Numerical analysis of multiple jets discharging into a confined cylindrical crossflow

2002 ◽  
Vol 216 (3) ◽  
pp. 173-180 ◽  
Author(s):  
Y Tao ◽  
W Adler ◽  
E Specht
Keyword(s):  
Penetration Depth ◽  
Control Volume ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Design Procedure ◽  
Multiple Jets ◽  
Geometric Variables ◽  
The Relationship ◽  
Upstream Flow ◽  
Nozzle Spacing

A row of jets discharging normally into a confined cylindrical crossflow is numerically investigated using the control-volume-based finite difference method. Interest is focused on determining the relationship between the temperature trajectory and the upstream flow and geometric variables. Parameter variations studied include nozzle diameter, number of nozzles, duct radius, jet and mainstream volume-flow, temperature ratio, and dynamic pressure ratio. The dynamic pressure ratio, the number of nozzles, and nozzle spacing are found to be significant variables. A logarithmic function describing the relationship between penetration depth and dynamic pressure divided by the square of the number of nozzles is derived by fitting the data of the computation results. The values for penetration depth and nozzle spacing are described for optimum mixing. A suggested design procedure is presented, which can be used as a first approach in configuration design.

Fugitive Listening: Sounds from the Undercommons

2020 ◽  
Vol 37 (6) ◽  
pp. 25-45
Author(s):  
Andrew Navin Brooks
Keyword(s):  
The Self ◽  
Ideal Object ◽  
Black Performance ◽  
The Given ◽  
The Relationship ◽  
The Voice ◽  
The Ideal

This essay builds on various critiques of the relationship between the voice and autonomous individual subjectivity, briefly tracking the specific history through which the voice transformed into an ideal object representing the liberal subject of post-Enlightenment thought. This paper asks: what are we to make of those enfleshed voices that do not conform to the ideal voice of the self-possessed liberal subject? What are we to make of those voices that refuse the imperative of improvement that underpins social and economic contractualism? How might we attend to the sonicity of those voices that refuse to individuate, possess, and accumulate? And what fugitive modes of speech might be transmitted by such un-formed and un-organized voices? Against the idealized voice of liberalism, and the gendered and racialized exclusions that this voice implies, I propose a mode of fugitive listening that allows us to open our ears to the noisy voices and modes of speech that sound outside the locus of politics proper. Indebted to the Black radical tradition, fugitive listening attends to sonic practices that refuse the given grounds of representation. I argue that fugitive listening is a practice that can be situated in what Stefano Harney and Fred Moten call ‘the undercommons’. The essay concludes by turning to gossip, figuring this noisy modality of speech as central to undercommon spaces shaped by Black performance.

On Study of Free Metal Forming Using Underwater Shock Wave

2002 ◽  
Author(s):  
Shigeru Itoh ◽  
Hirofumi Iyama ◽  
K. Raghukandan ◽  
Shiro Nagano ◽  
Ryo Matsumura ◽  
...  
Keyword(s):  
Shock Wave ◽  
High Explosive ◽  
Dynamic Pressure ◽  
Powder Materials ◽  
Underwater Shock Wave ◽  
Powder Consolidation ◽  
Underwater Shock ◽  
Overdriven Detonation ◽  
The Given ◽  
Basic Configuration

In the material processing such as shock synthesis and powder consolidation by shock waves the method for generating dynamic pressure is of very importance for the final recovered materials. A general and convenient way for producing shock wave needed in such field is to take advantage of the explosion effect from high explosive. Therefore, it becomes an interest subject how to produce dynamic pressure as high as possible under the given high explosive. Starting from this motivation, we put forward a method of high-pressure generation by using the overdriven detonation of high explosive. The basic configuration for this device is summarized in the following. A metal flyer accelerated by the high explosive is used to impact another layer of high explosive to incur an overdriven detonation in this layer of explosive. The overdriven detonation of high explosive acts on the powder materials, providing the high dynamic pressure for it. To examine the efficiency of this combination, a numerical computation is performed to this system. The details on the illustration of this system and numerical treatment will be given.

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