Aero-Thermodynamic Loss Analysis in Cases of Normal Shock Wave-Turbulent Shear Layer Interaction

1997 ◽  
Vol 119 (2) ◽  
pp. 297-303 ◽  
Author(s):  
J. K. Kaldellis
Keyword(s):  
Shear Layer ◽  
Practical Interest ◽  
Flow Simulation ◽  
Strong Shock ◽  
Turbulent Shear ◽  
Complex Flow ◽  
Loss Model ◽  
Layer Interaction ◽  
Loss Analysis ◽  
Shock Loss

Transonic-supersonic decelerating flow cases appearing in modern turbomachines are some of the most complex flow cases in fluid mechanics which also present practical interest. In the present work, a closed and coherent shock loss model is proposed based on the complete viscous flow simulation using some fast and reliable computational tools. The resulting model describes accurately the entropy rise and the total pressure loss in cases of strong shock-shear layer interaction and cancels the need to use low speed correlations modified for compressibility effects and extrapolated to transonic-supersonic flow cases. The accuracy and the reliability of the proposed shock-loss model are verified using detailed experimental data concerning various flow cases which present either flow separation or industrial interest.

Parametrical Investigation of the Interaction Between Turbulent Wall Shear Layers and Normal Shock Waves, Including Separation

1993 ◽  
Vol 115 (1) ◽  
pp. 48-55 ◽  
Author(s):  
J. K. Kaldellis
Keyword(s):  
Shock Waves ◽  
Flow Field ◽  
Shear Layer ◽  
Integral Method ◽  
Strong Shock ◽  
Interaction Region ◽  
Normal Shock ◽  
Turbulent Shear ◽  
Zone Model ◽  
Displacement Thickness

The existence of strong shock waves plays a major role in the performance of modern aero-mechanical devices, since it is primarily responsible not only for the shock induced total pressure drop, but also for the increased shear layer losses due to flow separation. In this paper a fast energy-type integral method along with an approximate shock-turbulent shear layer interaction procedure are presented. This integral method, based on the two-zone model, is able to predict attached and fully detached shear flows. An extended turbulence model is also used in order to take the influence of the turbulence inside the interaction region better into account. The external flow pressure distribution results from an improved and extended form of an approximate small disturbance theory. A detailed investigation is carried out to estimate the influence of the inlet Mach number, the shear layer characteristics and the confinement of the geometry upon the static pressure field. The resulting method has been successfully applied to several test cases including ones where separation appears. Comparison between results of previous calculations, experimental data and results of the proposed method is also presented, along with the convergence history of the shear layer—shock wave interaction procedure. Finally, the method has been applied to one-stage high pressure supersonic flow compressor with normal shock appearance inside the rotor of the machine. The major conclusion drawn from the present work is that the shear layer characteristics (e.g., displacement thickness and form factor) have a dominant effect upon the flow field near the interaction region. Additionally, the proposed method requires no more than five overall iterations to reproduce the real flow field for all cases examined.

Features of a reattaching turbulent shear layer in divergent channelflow

AIAA Journal ◽  
1985 ◽  
Vol 23 (2) ◽  
pp. 163-171 ◽  
Author(s):  
David M. Driver ◽  
H. Lee Seegmiller
Keyword(s):  
Shear Layer ◽  
Turbulent Shear ◽  
Turbulent Shear Layer

scholarly journals A Three-Dimensional Shock Loss Model Applied to an Aft-Swept, Transonic Compressor Rotor

1996 ◽  
Author(s):  
Steven L. Puterbaugh ◽  
William W. Copenhaver ◽  
Chunill Hah ◽  
Arthur J. Wennerstrom
Keyword(s):  
Flow Rate ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Numerical Results ◽  
Design Flow ◽  
Design System ◽  
Loss Model ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Shock Loss

An analysis of the effectiveness of a three-dimensional shock loss model used in transonic compressor rotor design is presented. The model was used during the design of an aft-swept, transonic compressor rotor. The demonstrated performance of the swept rotor, in combination with numerical results, is used to determine the strengths and weaknesses of the model. The numerical results were obtained from a fully three-dimensional Navier-Stokes solver. The shock loss model was developed to account for the benefit gained with three-dimensional shock sweep. Comparisons with the experimental and numerical results demonstrated that shock loss reductions predicted by the model due to the swept shock induced by the swept leading edge of the rotor were exceeded. However, near the tip the loss model under-predicts the loss because the shock geometry assumed by the model remains swept in this region while the numerical results show a more normal shock orientation. The design methods and the demonstrated performance of the swept rotor is also presented. Comparisons are made between the design intent and measured performance parameters. The aft-swept rotor was designed using an inviscid axisymmetric streamline curvature design system utilizing arbitrary airfoil blading geometry. The design goal specific flow rate was 214.7 kg/sec/m2 (43.98 lbm/sec/ft2), the design pressure ratio goal was 2.042, and the predicted design point efficiency was 94.0. The rotor tip sped was 457.2 m/sec (1500 ft/sec). The design flow rate was achieved while the pressure ratio fell short by 0.07. Efficiency was 3 points below prediction, though at a very high 91 percent. At this operating condition the stall margin was 11 percent.

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

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