Estimation of Mach Number of Shock Wave Propagating in Pipe

2003 ◽  
Author(s):  
Takanori Yamazaki ◽  
Masaki Endo
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Shock Tube ◽  
Dynamic Characteristic ◽  
Pressure Transducer ◽  
Shock Strength ◽  
Pressure History ◽  
Boundary Layer Interaction ◽  
Shock Mach Number

A useful way to estimate the local Mach number of shock propagating in a pipe is proposed in this paper. The shock Mach number, or the shock strength gradually decreases or increases as the shock propagates downstream in pipe due to the shock-boundary layer interaction. In general, the shock Mach number is estimated through the measurement of the time, which it takes for the shock to propagate between any two points along the pipe. This technique is very useful if the decreasing rate of the shock strength is given and it, after all, yields the average Mach number between the two points. In this paper the measurement of the local Mach number of shock in the shock tube is examined using one pressure transducer. The pressure history after the shock reaches the pressure tap is analyzed. The method to estimate the local Mach number is discussed considering the dynamic characteristic of the pressure transducer.

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.

Experimental investigation of diaphragm material combinations in a small-scale shock tube

2020 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Anugya Singh ◽  
Aravind Satheesh Kumar ◽  
Kannan B.T.
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Tube ◽  
Production Capacity ◽  
Incident Shock ◽  
Small Scale ◽  
Content Type ◽  
Shock Mach Number ◽  
Reinforcing Material ◽  
Shock Wave Mach Number

Purpose The purpose of this study is to experimentally investigate the trends in shock wave Mach number that were observed when different diaphragm material combinations were used in the small-scale shock tube. Design/methodology/approach A small-scale shock tube was designed and fabricated having a maximum Mach number production capacity to be 1.5 (theoretically). Two microphones attached in the driven section were used to calculate the shock wave Mach number. Preliminary tests were conducted on several materials to obtain the respective bursting pressures to decide the final set of materials along with the layered combinations. Findings According to the results obtained, 95 GSM tracing paper was seen to be the strongest reinforcing material, followed by 75 GSM royal executive bond paper and regular 70 GSM paper for aluminium foil diaphragms. The quadrupled layered diaphragms revealed a variation in shock Mach number based on the position of the reinforcing material. In quintuple layered combinations, the accuracy of obtaining a specific Mach number was seen to be increasing. Optimization of the combinations based on the production of the shock wave Mach number was carried out. Research limitations/implications The shock tube was designed taking maximum incident shock Mach number as 1.5, the experiments conducted were found to achieve a maximum Mach number of 1.437. Thus, an extension to further experiments was avoided considering the factor of safety. Originality/value The paper presents a detailed study on the effect of change in the material and its position in the layered diaphragm combinations, which could lead to variation in Mach numbers that are produced. This could be used to obtain a specific Mach number for a required study accurately, with a low-cost setup.

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.

Numerical Investigation on Suppression of 3-Dimensional Shock Wave/Turbulent Boundary Layer Interaction by Passive Cavity

2003 ◽  
Author(s):  
Kazuyuki Toda ◽  
Shinsuke Dambara ◽  
Makoto Yamamoto ◽  
Shinji Honami ◽  
Nobuyuki Akahoshi
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Boundary Layer ◽  
Passive Control ◽  
Interaction Region ◽  
Mean Flow ◽  
3 Dimensional ◽  
Layer Interaction ◽  
Boundary Layer Interaction

Suppression of three-dimensional shock wave/turbulent boundary layer interaction is one of the important subjects on supersonic air intake. In the present study, the passive control of 2- and 3-dimensional shock wave/turbulent boundary layer interactions is considered. First, computations are performed for two-dimensional flow field at freestream Mach number of 2.46 with various passive cavities beneath the interaction region. The results suggest that the parallel blowing from a cavity to the mean flow with a guide plate can highly keep the interaction region narrow. Next, the most suitable cavity shape clarified in the 2-dimensional computations is applied to the 3-dimensional swept shock wave/turbulent boundary layer interaction at Mach number of 3.11. It is exhibited that the blowing direction is important, and the effect of passive cavity is nearly the same as the bleeding.

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