A24 Effect of Pressure Ratio for Hysteresis Phenomenon of Shock Wave in Axisymmetric Supersonic Jet

The unsteady supersonic jet formed by the shock tube with small high-pressure section was used as a simple alternative system of pulsed laser ablation. The dynamic of the supersonic jet impinging upon a flat plate are discussed by comparing experimental and calculated results. The experiment and numerical calculation were carried out by schlieren method and by solving the axisymmetric two-dimensional compressible Navier-Stokes equations, respectively. The main parameters are distance between the open end of the shock tube and the flat plate, L/D, and the pressure ratio of the shock tube, Ph/Pb. Where, L, D, Ph and Pb are the distance between the open end of the shock tube and the flat plate, the diameter of the shock tube, pressure of the high and low section of the shock tube, respectively. Collision between the shock wave reflected at the flat plate and the head of supersonic jet takes place. Computational results well predict the experimental dynamic behavior of the shock wave and the supersonic jet. Marked increase in the static pressure on the flat plate under high Ph/Pb and short L/D is observed due to interaction between the shock wave and the unsteady jet flow.

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Stall Behavior in an Ultra-High-Pressure-Ratio Centrifugal Compressor: Backward-Traveling Rotating Stall

Journal of Turbomachinery ◽

10.1115/1.4050918 ◽

2021 ◽

pp. 1-51

Author(s):

Yingjie Zhang ◽

Xingen Lu ◽

Yanfeng Zhang ◽

Ziqing Zhang ◽

Xu Dong ◽

...

Keyword(s):

Boundary Layer ◽

Shock Wave ◽

Shock Intensity ◽

Incidence Angle ◽

Pressure Ratio ◽

Leading Edge ◽

Boundary Layer Separation ◽

Rotating Stall ◽

High Pressure Ratio ◽

Ultra High Pressure

Abstract This paper describes the stall mechanism in an ultra-high-pressure-ratio centrifugal compressor. A model comprising all impeller and diffuser blade passages is used to conduct unsteady simulations that trace the onset of instability in the compressor. Backward-traveling rotating stall waves appear at the inlet of the radial diffuser when the compressor is throttled. Six stall cells propagate circumferentially at approximately 0.7% of the impeller rotation speed. The detached shock of the radial diffuser leading edge and the number of stall cells determine the direction of stall propagation, which is opposite to the impeller rotation direction. Dynamic mode decomposition is applied to instantaneous flow fields to extract the flow structure related to the stall mode. This shows that intensive pressure fluctuations concentrate in the diffuser throat as a result of changes in the detached shock intensity. The diffuser passage stall and stall recovery are accompanied by changes in incidence angle and shock wave intensity. When the diffuser passage stalls, the shock-induced boundary-layer separation region near the diffuser vane suction surface gradually expands, increasing the incidence angle and decreasing the shock intensity. The shock is pushed from the diffuser throat toward the diffuser leading edge. When the diffuser passage recovers from stall, the shock wave gradually returns to the diffuser throat, with the incidence angle decreasing and the shock intensity increasing. Once the shock intensity reaches its maximum, the diffuser passage suffers severe shock-induced boundary-layer separation and stalls again.

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Shock Tube Simulation of the Transient Surge Phase Inside an Axial Compressor

Volume 2A: Turbomachinery ◽

10.1115/gt2018-75052 ◽

2018 ◽

Author(s):

Paul Xiubao Huang ◽

Robert S. Mazzawy

Keyword(s):

Shock Wave ◽

Shock Tube ◽

High Speed ◽

Transient Flow ◽

Pressure Ratio ◽

Axial Compressor ◽

Dynamic Effect ◽

Initiation Site ◽

Flow Reversal ◽

Expansion Waves

This paper is a continuing work from one author on the same topic of the transient aerodynamics during compressor stall/surge using a shock tube analogy by Huang [1, 2]. As observed by Mazzawy [3] for the high-speed high-pressure (HSHP) ratio compressors of the modern aero-engines, surge is an event characterized with the stoppage and reversal of engine flow within a matter of milliseconds. This large flow transient is accomplished through a pair of internally generated shock waves and expansion waves of high strength. The final results are often dramatic with a loud bang followed by the spewing out of flames from both the engine intake and exhaust, potentially damaging to the engine structure [3]. It has been demonstrated in the previous investigations by Marshall [4] and Huang [2] that the transient flow reversal phase of a surge cycle can be approximated by the shock tube analogy in understanding its generation mechanism and correlating the shock wave strength as a function of the pre-surge compressor pressure ratio. Kurkov [5] and Evans [8] used a guillotine analogy to estimate the inlet overpressure associated with the sudden flow stoppage associated with surge. This paper will expand the progressive surge model established by the shock tube analogy in [2] by including the dynamic effect of airflow stoppage using an “integrated-flow” sequential guillotine/shock tube model. It further investigates the surge formation (characterized by flow reversal) and propagation patterns (characterized by surge shock and expansion waves) after its generation at different locations inside a compressor. Calculations are conducted for a 12-stage compressor using this model under various surge onset stages and compared with previous experimental data [3]. The results demonstrate that the “integrated-flow” model closely replicates the fast moving surge shock wave overpressure from the stall initiation site to the compressor inlet.

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Flow With Shock Waves in a Transonic Centrifugal Compressor With a Low-Solidity Cascade Diffuser Using PIV

Volume 5: Turbo Expo 2004, Parts A and B ◽

10.1115/gt2004-53268 ◽

2004 ◽

Cited By ~ 5

Author(s):

Hiroshi Hayami ◽

Masahiro Hojo ◽

Norifumi Hirata ◽

Shinichiro Aramaki

Keyword(s):

Shock Wave ◽

Shock Waves ◽

Centrifugal Compressor ◽

Gas Flow ◽

Pressure Ratio ◽

Transonic Centrifugal Compressor ◽

Image Velocimetry ◽

Rotor Stator Interaction ◽

Rotating Impeller ◽

Low Solidity Cascade Diffuser

A single-stage transonic centrifugal compressor with a pressure ratio greater than six was tested in a closed loop with HFC134a gas. Flow at the inducer of a rotating impeller as well as flow in a stationary low-solidity cascade diffuser was measured using a double-pulse and double-frame particle image velocimetry (PIV). Shock waves in both flows were clearly observed. The effect of flow rate on a 3D configuration of shock wave at the inducer and a so-called rotor-stator interaction between a rotating impeller and a stationary cascade were discussed based on a phase-averaged measurement technique. Furthermore, the unsteadiness of inducer shock wave and the flow in a cascade diffuser during surge were discussed based on instantaneous velocity vector maps.

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Approximate Analytical Models of Shock-Wave Structure at Steady Mach Reflection

Fluids ◽

10.3390/fluids6090305 ◽

2021 ◽

Vol 6 (9) ◽

pp. 305

Author(s):

Mikhail V. Chernyshov ◽

Karina E. Savelova ◽

Anna S. Kapralova

Keyword(s):

Experimental Data ◽

Shock Wave ◽

Comparative Analysis ◽

Analytical Model ◽

Mach Reflection ◽

Supersonic Jet ◽

Wave Structure ◽

Analytical Models ◽

Analytical Prediction ◽

Shock Wave Structure

In this study, we obtain the comparative analysis of methods of quick approximate analytical prediction of Mach shock height in planar steady supersonic flows (for example, in supersonic jet flow and in narrowing channel between two wedges), that are developed since the 1980s and being actively modernized now. A new analytical model based on flow averaging downstream curved Mach shock is proposed, which seems more accurate than preceding models, comparing with numerical and experimental data.

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