Blade Loading and Shock Wave in a Transonic Circular Cascade Diffuser

1992 ◽  
Author(s):  
H. Hayami ◽  
T. Nakamura ◽  
M. Sawae ◽  
N. Kawaguchi
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Lift Coefficient ◽  
Side Wall ◽  
Suction Surface ◽  
Contour Map ◽  
Blade Loading ◽  
Straight Line ◽  
Transonic Centrifugal Compressor ◽  
Wide Range

Low-solidity circular cascade, conformally transformed from high-stagger linear cascade of double-circular-arc vanes with solidity 0.69, was tested as a part of diffuser systems of a transonic centrifugal compressor and the static pressures were measured around a vane of the cascade and on the side wall between cascade vanes in detail. The blade loading of cascade vane was discussed by integrating the pressure distribution around the vane. The experimental data for lift-coefficient of vane were almost on a single straight line with positive gradient against angle-of-attack over a wide range of inflow Mach number and inflow angle. The maximum lift-coefficient reached about 1.5 and the vane worked well to the surge condition of the compressor. The structure of shock wave was also discussed by drawing a contour map of flow Mach number between cascade vanes. The normal shock wave was observed on the suction surface of vane and it moved upstream along the suction surface with the decrease in inflow angle. The vane did not fall in stall even though the Mach number upstream of the shock wave was over 1.4.

Blade Loading and Shock Wave in a Transonic Circular Cascade Diffuser

1993 ◽  
Vol 115 (3) ◽  
pp. 560-564 ◽  
Author(s):  
H. Hayami ◽  
M. Sawae ◽  
T. Nakamura ◽  
N. Kawaguchi
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Lift Coefficient ◽  
Side Wall ◽  
Suction Surface ◽  
Contour Map ◽  
Blade Loading ◽  
Straight Line ◽  
Transonic Centrifugal Compressor ◽  
Wide Range

A low-solidity circular cascade, conformally transformed from a high-stagger linear cascade of double-circular-arc vanes with solidity 0.69, was tested as a part of diffuser systems of a transonic centrifugal compressor and the static pressures were measured around a vane of the cascade and on the side wall between cascade vanes in detail. The blade loading of cascade vane was discussed by integrating the pressure distribution around the vane. The experimental data for lift-coefficient of vane were almost on a single straight line with positive gradient against angle-of-attack over a wide range of inflow Mach number and inflow angle. The maximum lift coefficient reached about 1.5 and the vane worked well to the surge condition of the compressor. The structure of shock wave was also discussed by drawing a contour map of the flow Mach number between cascade vanes. The normal shock wave was observed on the suction surface of vane and it moved upstream along the suction surface with the decrease in inflow angle. The vane did not fall in stall even though the Mach number upstream of the shock wave was over 1.4.

Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

2020 ◽  
Author(s):  
Hoshio Tsujita ◽  
Masanao Kaneko
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Flow Behavior ◽  
Turbine Cascade ◽  
Blade Loading ◽  
Passage Vortex ◽  
Exit Mach Number

Abstract Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.

Behavior of Secondary Flow in a Low Solidity Tandem Cascade Diffuser

2012 ◽  
Author(s):  
Daisaku Sakaguchi ◽  
Hironobu Ueki ◽  
Masahiro Ishida ◽  
Hiroshi Hayami
Keyword(s):  
Secondary Flow ◽  
Flow Rate ◽  
Flow Separation ◽  
Lift Coefficient ◽  
Pressure Recovery ◽  
Suction Surface ◽  
Flow Rates ◽  
Blade Loading ◽  
Small Flow Rate ◽  
Small Flow

Low solidity circular cascade diffuser abbreviated by LSD was proposed by Senoo et al. showing a high blade loading or a high lift coefficient without stall even under small flow rate conditions. These high performances were achieved by that the flow separation on the suction surface of the LSD blade was successfully suppressed by the secondary flow formed along the side walls. The higher performance of the LSD was achieved in both pressure recovery and operating range by adopting the tandem cascade because the front blade of the tandem cascade was designed suitably for small flow rates while the rear blade of the tandem cascade was designed suitably for large flow rates. In order to clarify the reason why the tandem cascade could achieve a high pressure recovery in a wide range of flow rate, the flow in the LSD with the tandem cascade is analyzed numerically in the present study by using the commercial CFD code of ANSYS-CFX 13.0. The behavior of the secondary flow is compared between the cases with the single cascade and the tandem one. It is found that the high blade loading of the front blade is achieved at the small flow rate by formation of the favorable secondary flow which suppresses the flow separation on suction surface of the front blade, and the flow separation on pressure surface of the front blade appeared at the design flow rate can be suppressed by the accelerated flow in the gap between the trailing edge of the front blade and the leading edge of the rear blade, resulting in the positive lift coefficient in spite of a large negative angle of attack.

scholarly journals Aerodynamic Investigation of Datum and Slotted Blade Profiles under Different Mach Number Conditions

Energies ◽  
2020 ◽  
Vol 13 (7) ◽  
pp. 1673
Author(s):  
Yumeng Tang ◽  
Yangwei Liu
Keyword(s):  
Mach Number ◽  
Fluid Dynamic ◽  
Mechanical Energy ◽  
Adverse Pressure Gradient ◽  
Dissipation Function ◽  
Suction Surface ◽  
Trailing Edge ◽  
Loss Mechanism ◽  
Wide Range ◽  
Downstream Flow

Mach number effects on loss and loading are evaluated in both the datum and slotted compressor profiles under a wide range of incidences based on two-dimensional (2D) computational fluid dynamic (CFD) simulations. First, total pressure loss and loading abilities are compared. Then, three kinds of deficit thickness are defined and evaluated, and a correlation is made between the loading and the momentum deficit thickness at the profile trailing edge. Finally, the nondimensionalized destruction of mean mechanical energy and dissipation function are employed to analyze the loss mechanism. The slotted profile broadens the low loss range towards the positive incidence range. The slotted profile allows a higher diffusion factor (DF) than the datum profile. It is hard to distinguish failure simply based on the DF values, whereas the Zweifel loading coefficient connects well with the low momentum deficit in the profile trailing edge. The peak of the V-shaped distributions in the Ψ - θ d e f plot could better suggest the design condition and determine the correct operating range despite the occurrence of bulk separation. The slotted profile gains the ability of the boundary layer flow near the suction surface to resist the adverse pressure gradient, hence, a reduced shear thickness and a uniformed downstream flow field is obtained.

Criterion Prediction of Changes of Shock-Wave Structure at Transonic Flow Around Airfoils with Local Impulse Energy

2014 ◽  
Vol 9 (1) ◽  
pp. 39-48
Author(s):  
Sergey Aulchenko ◽  
Vladimir Zamuraev ◽  
Anna Kalinina
Keyword(s):  
Shock Wave ◽  
Numerical Modeling ◽  
Mach Number ◽  
Pulse Energy ◽  
Transonic Flow ◽  
Energy Supply ◽  
Wave Structure ◽  
Shock Wave Structure ◽  
Wide Range ◽  
Comparison Of The Results

The results of numerical modeling of transonic streamline of profile with a local pulse energy supply were generalized in terms of criteria analysis for a wide range of parameters. Defined set of criteria allows predicting not only the degree, but the features of the flow transformation. A comparison of the results of the criteria analysis with numerical calculations of transonic streamline (Mach number M∞ = 0,85) of a symmetrical profile NASA-0012 were performed for energy supply in areas such as narrow and compact form

scholarly journals Experimental Study of the Weak Shock Wave Action on the Boundary Layer of a Plate at the Mach Number 2.5

2019 ◽  
Vol 14 (2) ◽  
pp. 46-55 ◽  
Author(s):  
V. L. Kocharin ◽  
A. A. Yatskikh ◽  
A. D. Kosinov ◽  
Yu. G. Yermolaev ◽  
N. V. Semionov
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Experimental Study ◽  
Mach Number ◽  
Side Wall ◽  
Leading Edge ◽  
Weak Shock ◽  
Supersonic Boundary Layer ◽  
Weak Shock Wave ◽  
Frequency Spectra

Experimental study of the effect of a weak shock wave from the protuberance of two-dimensional roughness installed on the side wall of the test section of the wind tunnel on the supersonic boundary layer of the blunted flat plate at the Mach number 2.5 was carried out. The measurements were performed by a constant temperature hot-wire anemometer in the region of stream wise vortices generated by the shock wave from the protuberance during interaction with the flow in the vicinity of the leading edge of the model. The spectral and statistical analyses of the measured disturbances in the boundary layer were carried out. The amplitude-frequency spectra of mass flow pulsations and statistical diagrams of the measured disturbances in the supersonic part of the boundary layer were obtained.

The Flow Over a “High” Aspect Ratio Gothic Wing at Supersonic Speeds

1975 ◽  
Vol 26 (3) ◽  
pp. 189-201 ◽  
Author(s):  
K Yegna Narayan
Keyword(s):  
Shock Wave ◽  
Delta Wing ◽  
Lift Coefficient ◽  
Leading Edge ◽  
Two Dimensional ◽  
Delta Wings ◽  
Pressure Drag ◽  
Wide Range ◽  
Detached Shock Wave ◽  
The One

SummaryResults are presented of an experimental investigation on a non-conical wing which supports an attached shock wave over a region of the leading edge near the vertex and a detached shock elsewhere. The shock detachment point is determined from planform schlieren photographs of the flow field and discrepancies are shown to exist between this and the one calculated by applying the oblique shock equations normal to the leading edge. On a physical basis, it is argued that shock detachment has to obey the two-dimensional law normal to the leading edges. From this, and from other measurements on conical wings, it is thought that the planform schlieren technique may not be particularly satisfactory for detecting shock detachment. Surface pressure distributions are presented and are explained in terms of the flow over related delta wings which are identified as a vertex delta wing and a local delta wing. The forces acting on the wing are calculated and are shown to be very close to the two-dimensional wedge values over a wide range of incidence. In particular, it is shown that this wing, compared to one which supports a fully detached shock wave, generates a higher lift/(pressure drag) ratio at a given lift coefficient.

Mechanism of Blockage Generation in Transonic Centrifugal Compressor at Design and Off-Design Conditions

2015 ◽  
Author(s):  
Masanao Kaneko ◽  
Hoshio Tsujita
Keyword(s):  
Shock Wave ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Leading Edge ◽  
Suction Surface ◽  
Tip Leakage Vortex ◽  
Stall Inception ◽  
Aerodynamic Loss ◽  
Tip Leakage ◽  
Transonic Centrifugal Compressor

In a transonic centrifugal compressor, the loss generation is intensified by the formation of the shock wave and consequently the blockage is expected to increase. The blockage is considered to influence not only the flow rate and the increase of the static pressure but also the stall inception. However, the detailed mechanism of the blockage generation in the transonic centrifugal compressor has not been fully clarified. In this study, in order to clarify the mechanisms of loss and blockage generations in the transonic centrifugal compressor which are expected to be strongly influenced by the operating condition, the flows in the compressor at the off-design condition as well as at the design condition were analyzed numerically. The verifications of the computed results were carried out by comparing with available experimental results. The computed result clarified that the loss generation near the impeller inlet at design condition was mainly caused by the interactions of the shock wave with the tip leakage vortex appearing from the leading edge of the main blade as well as the boundary layer on the suction surface of the main blade. Moreover, these interactions were intensified by the decrease of the flow rate, and consequently enhanced the blockage effects by the tip leakage vortex from the leading edge of the main blade and resulted in the increase of the aerodynamic loss especially along the shroud surface in the impeller passage. On the other hand, the decrease of the blockage effects by the tip leakage vortex from the main blade with the increase of the flow rate formed the shock wave on the suction surface of the splitter blade at near-choke condition. This shock wave interacted with the tip leakage vortex from the splitter blade and consequently increased the aerodynamic loss.

Numerical Simulation of Transonic Fan Flutter With 3D N-S CFD Code

2008 ◽  
Author(s):  
Mizuho Aotsuka ◽  
Naoki Tsuchiya ◽  
Yasuo Horiguchi ◽  
Osamu Nozaki ◽  
Kazuomi Yamamoto
Keyword(s):  
Shock Wave ◽  
Stokes Equations ◽  
Navier Stokes ◽  
Suction Surface ◽  
Navier Stokes Equations ◽  
Blade Loading ◽  
Flow Solver ◽  
Flutter Boundary ◽  
Transonic Fan ◽  
Volume Method

This paper describes the calculation of transonic stall flutter of a fan. A new CFD code has been developed and validated. The code is an unsteady 3D multi-block flow solver. The Reynolds-Averaged Navier-Stokes equations are solved using a finite volume method with Spallart-Allmaras 1 equation turbulence model. A grid deforming system is applied, so the new code is capable of simulating an oscillating blade row. This grid deforming system produces less grid distortion and the code has robustness for a blade oscillating calculation. The code has validated on an IHI’s research transonic fan rig test, and the result was in good agreement with the test data in the prediction of the flutter boundary. In the rig test at part-speed condition, stall-side flutter was experienced. In that condition, the inlet relative Mach number in the tip region is about unity. The aerodynamic work by the CFD at the near flutter condition is positive, which means that the flutter characteristic is unstable, while at other conditions the aerodynamic work is negative. The aerodynamic work increases rapidly just before the zero damping point with the increase of the blade loading. From the detailed CFD result, the shock wave on the suction surface contributes to the excitement of the blade oscillation, and the aerodynamic work of the shock wave has large value at the flutter condition.

Designing Low-Pressure Turbine Blades With Integrated Flow Control

2005 ◽  
Author(s):  
Jeffrey P. Bons ◽  
Laura C. Hansen ◽  
John P. Clark ◽  
Peter J. Koch ◽  
Rolf Sondergaard
Keyword(s):  
Reynolds Number ◽  
Flow Control ◽  
Turbine Blade ◽  
Separation Bubble ◽  
Low Pressure ◽  
Suction Surface ◽  
Flow Measurements ◽  
Blade Loading ◽  
Low Pressure Turbine ◽  
Wide Range

A low pressure turbine blade was designed to produce a 17% increase in blade loading over an industry-standard airfoil using integrated flow control to prevent separation. The design was accomplished using two-dimensional CFD predictions of blade performance coupled with insight gleaned from recently published work in transition modeling and from previous experiments with flow control using vortex generator jets (VGJs). In order to mitigate the Reynolds number lapse in efficiency associated with LPT airfoils, a mid-loaded blade was selected. Also, separation predictions from the computations were used to guide the placement of control actuators on the blade suction surface. Three blades were fabricated using the new design and installed in a two-passage linear cascade facility. Flow velocity and surface pressure measurements taken without activating the VGJs indicate a large separation bubble centered at 68% axial chord on the suction surface. The size of the separation and its growth with decreasing Reynolds number agree well with CFD predictions. The separation bubble reattaches to the blade over a wide range of inlet Reynolds numbers from 150,000 down to below 20,000. This represents a marked improvement in separation resistance compared to the original blade profile which separates without reattachment below a Reynolds number of 40,000. This enhanced performance is achieved by increasing the blade spacing while simultaneously adjusting the blade shape to make it less aft-loaded but with a higher peak cp. This reduces the severity of the adverse pressure gradient in the uncovered portion of the modified blade passage. With the use of pulsed VGJs, the design blade loading was achieved while providing attached flow over the entire range of Re. Detailed phase-locked flow measurements using three-component PIV show the trajectory of the jet and its interaction with the unsteady separation bubble. Results illustrate the importance of integrating flow control into the turbine blade design process and the potential for enhanced turbine performance.

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