scholarly journals On the Effect of Single Mode Bending/Torsion Coupling on the Flutter Behavior in a Transonic Fan

1998 ◽  
Author(s):  
Kousuke Isomura
Keyword(s):  
Shock Wave ◽  
Vibration Mode ◽  
Amplitude Ratio ◽  
Single Mode ◽  
Blade Vibration ◽  
Pressure Surface ◽  
Torsion Mode ◽  
Bending Mode ◽  
Torsional Component ◽  
Transonic Fan

The effect of the coupling of the torsion mode blade vibration to the bending mode flutter in transonic fans has been studied by quasi-3D viscous unsteady calculations. The type of flutter in this research is that of a highly loaded blade with a tip relative Mach number just above unity, for which the mechanism is that the instability of the passage shock wave when it unstarts generates the dominant blade exciting aerodynamic work at its foot on the pressure surface of the blade. The dependence of such flutter on the blade vibration mode, i.e. the amplitude ratio and the phase difference of the bending and torsional components has been studied. The study showed that the blade exciting aerodynamic work reduced when the torsional component is added in-phase (torsional motion noses up during the upward bending motion) to the bending oscillation. The amplitude of the torsional component was shown to have an optimum amplitude. It was also shown that this tendency would switch when the shock wave is fully detached, and the blade exciting aerodynamic work would increase by adding the torsional component in-phase to the bending oscillation.

A Numerical Study of Flutter in a Transonic Fan

1997 ◽  
Author(s):  
Kousuke Isomura ◽  
Michael B. Giles
Keyword(s):  
Cfd Simulation ◽  
Numerical Study ◽  
Point Model ◽  
Blade Vibration ◽  
Blade Loading ◽  
Pressure Surface ◽  
Dominant Component ◽  
Bending Mode ◽  
The Stability ◽  
Transonic Fan

The bending mode flutter of a modern transonic fan has been studied using a quasi-3D viscous unsteady CFD code. The type of flutter in this research is that of a highly loaded blade with a tip relative Mach number just above unity, commonly referred to as transonic stall flutter. This type of flutter is often encountered in modern wide chord fans without a part span shroud. The CFD simulation uses an upwinding scheme with Roe’s 3rd-order flux differencing, and Johnson and King’s turbulence model with the later modification due to Johnson and Coakley. A dynamic transition point model is developed using the en method and Schubauer and Klebanoff’s experimental data. The calculations of the flow in this fan reveal that the source of the flutter of IHI transonic fan is an oscillation of the passage shock, rather than a stall. As the blade loading increases, the passage shock moves forward. Just before the passage shock unstarts, the stability of the passage shock decreases, and a small blade vibration causes the shock to oscillate with a large amplitude between unstarted and started positions. The dominant component of the blade excitation force is due to the foot of the oscillating passage shock on the blade pressure surface.

A Numerical Study of Flutter in a Transonic Fan

1998 ◽  
Vol 120 (3) ◽  
pp. 500-507 ◽  
Author(s):  
K. Isomura ◽  
M. B. Giles
Keyword(s):  
Cfd Simulation ◽  
Numerical Study ◽  
Three Dimensional ◽  
Third Order ◽  
Blade Vibration ◽  
Pressure Surface ◽  
Dominant Component ◽  
Bending Mode ◽  
The Stability ◽  
Transonic Fan

The bending mode Flutter of a modern transonic fan has been studied using a quasi-three-dimensional viscous unsteady CFD code. The type of flutter in this research is that of a highly loaded blade with a tip relative Mach number just above unity, commonly referred to as transonic stall flutter. This type of Flutter is often encountered in modern wide chord fans without a part span shroud. The CFD simulation uses an upwinding scheme with Roe’s third-order flux differencing, and Johnson and King’s turbulence model with the later modification due to Johnson and Coakley. A dynamic transition point model is developed using the en method and Schubauer and Klebanoff’s experimental data. The calculations of the flow in this fan reveal that the source of the flutter of IHI transonic fan is an oscillation of the passage shock, rather than a stall. As the blade loading increases, the passage shock moves forward. Just before the passage shock unstarts, the stability of the passage shock decreases, and a small blade vibration causes the shock to oscillate with a large amplitude between unstarted and started positions. The dominant component of the blade excitation force is due to the foot of the oscillating passage shock on the blade pressure surface.

Unsteady Three-Dimensional Analysis of Inlet Distortion in a Transonic Fan

2006 ◽  
Author(s):  
Peng Sun ◽  
Guotal Feng
Keyword(s):  
Shock Wave ◽  
Total Pressure ◽  
Large Scale ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Inlet Distortion ◽  
Flow Loss ◽  
Large Scale Vortex ◽  
Total Temperature ◽  
Transonic Fan

A time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan was carried out using "Fluent-parallel" in a parallel supercomputer. The numerical simulation focused on a transonic fan with inlet square wave total pressure distortion and the analysis of result consisted of three aspects. The first was about inlet parameters redistribution and outlet total temperature distortion induced by inlet total pressure distortion. The pattern and causation of flow loss caused by pressure distortion in rotor were analyzed secondly. It was found that the influence of distortion was different at different radial positions. In hub area, transportation-loss and mixing-loss were the main loss patterns. Distortion not only complicated them but enhanced them. Especially in stator, inlet total pressure distortion induced large-scale vortex, which produced backflow and increased the loss. While in casing area, distortion changed the format of shock wave and increased the shock loss. Finally, the format of shock wave and the hysteresis of rotor to distortion were analyzed in detail.

Effect of Shock Wave Behavior on Unsteady Aerodynamic Characteristics of Oscillating Transonic Compressor Cascade

2021 ◽  
Author(s):  
Jiuliang Gan ◽  
Toshinori Watanabe ◽  
Takehiro Himeno
Keyword(s):  
Shock Wave ◽  
Aerodynamic Force ◽  
Aerodynamic Characteristics ◽  
Fast Response ◽  
Influence Coefficient ◽  
Compressor Cascade ◽  
Blade Vibration ◽  
Unsteady Pressure ◽  
Transonic Compressor ◽  
Unsteady Aerodynamic

Abstract The unsteady behavior of the shock wave was studied in an oscillating transonic compressor cascade. The experimental measurement and corresponding numerical simulation were conducted on the cascade with different shock patterns based on influence coefficient method. The unsteady pressure distribution on blade surface was measured with fast-response pressure-sensitive paint (PSP) to capture the unsteady aerodynamic force as well as the shock wave movement. It was found that the movement of shock waves in the neighboring flow passages of the oscillating blade was almost anti-phase between the two shock patterns, namely, the double shocks pattern and the merged shock pattern. It was also found that the amplitude of the unsteady pressure caused by the passage shock wave was very large under the merged shock pattern compared with the double shocks pattern. The stability of blade vibration was also analyzed for both shock patterns including 3-D flow effect. These findings were thought to shed light on the fundamental understanding of the unsteady aerodynamic characteristics of oscillating cascade caused by the shock wave behavior.

The Unsteady Effect of Passing Shocks on Pressure Surface Versus Suction Surface Heat Transfer in Film-Cooled Transonic Turbine Blades

2003 ◽  
Author(s):  
A. C. Smith ◽  
A. C. Nix ◽  
T. E. Diller ◽  
W. F. Ng
Keyword(s):  
Heat Transfer ◽  
Shock Wave ◽  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Surface Heat ◽  
Suction Surface ◽  
Time Resolved ◽  
Pressure Surface ◽  
Transonic Turbine

This paper documents the measurement of the unsteady effects of passing shock waves on film cooling heat transfer on both the pressure and suction surfaces of first stage transonic turbine blades with leading edge showerhead film cooling. Experiments were performed for several cooling blowing ratios with an emphasis on time-resolved pressure and heat flux measurements on the pressure surface. Results without film cooling on the pressure surface demonstrated that increases in heat flux were a result of shock heating (the increase in temperature across the shock wave) rather than shock interaction with the boundary layer or film layer. Time-resolved measurements with film cooling demonstrated that the relatively strong shock wave along the suction surface appears to retard coolant ejection there and causes excess coolant to be ejected from pressure surface holes. This actually causes a decrease in heat transfer on the pressure surface during a large portion of the shock passing event. The magnitude of the decrease is almost as large as the increase in heat transfer without film cooling. The decrease in coolant ejection from the suction surface holes did not appear to have any effects on suction surface heat transfer.

scholarly journals Coupled 3-D Aeroelastic Stability Analysis of Bladed Disks

1992 ◽  
Author(s):  
Georg A. Gerolymos
Keyword(s):  
Finite Element ◽  
Stability Analysis ◽  
Traveling Wave ◽  
Wave Mode ◽  
Aeroelastic Stability ◽  
Blade Vibration ◽  
Modification Technique ◽  
Transonic Fan ◽  
Aerodynamic Simulation

In the present work an algorithm for the coupled aeromechanical computation of 3-D compressor cascades vibrating in a traveling-wave mode is presented and applied to the determination of aeroelastic stability of a transonic fan rotor. The initial vibratory modes are computed using a finite-element structural analysis code. The unsteady flowfield response to blade vibration is estimated by numerical integration of the 3-D unsteady Euler equations. Coupling relations are formulated in the frequency domain, using a mode-modification technique, based on modal projection. The vibratory mode is updated at the end of the aerodynamic simulation of each period, and the updated mode is used for the simulation of the next period. A number of results illustrate the method’s potential.

Development of Design Method for Supersonic Turbine Aerofoils Near the Tip of Long Blades in Steam Turbines: Part 2 — Configuration Details and Validation

2013 ◽  
Author(s):  
Shigeki Senoo ◽  
Hideki Ono
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Method Development ◽  
Design Space ◽  
Design Method ◽  
Flow Angle ◽  
Flow Passage ◽  
Pressure Surface ◽  
Outlet Flow ◽  
Stagger Angle

Both inflow and outflow velocities near the blade tip become supersonic when the blade length exceeds a threshold limit. The aerofoil near the tip of such a long blade has four features that demand an original supersonic turbine aerofoil design: supersonic flow in the entire field, high reaction, large stagger angle, and large pitch-to-chord ratio. This paper describes design method development for the supersonic turbine aerofoil. First, the aerofoil shape is defined using a curve with continuity in the gradient of the curvature. Second, six loss generation mechanisms are clarified by turbulent flow analysis. Third, an allowable design space between the pitch-to-chord ratio, the stagger angle and the axial-chord-to-pitch ratio is clarified by formulating three geometrical constraints to accelerate supersonic flow smoothly. When there is no solution in the theoretically allowable design space because of the large pitch-to-chord ratio, methods to reduce shock wave losses are proposed. Increasing the outlet metal angle of the pressure surface by around 10 deg from the theoretical outlet flow angle reduces the loss caused by the trailing shock wave. The physical mechanism for this is as follows: the increased outlet metal angle increases the outlet flow passage area so that the overexpansion is suppressed downstream from the flow passage. Fourth, both a cusped leading edge and an upstream pressure surface which has both an angle corresponding to the inflow angle and near-zero curvature can reduce the loss caused by the leading shock wave and satisfy the unique incidence relation. Finally, the aerodynamic performance of the supersonic turbine cascade and the design method are validated by supersonic cascade wind tunnel tests.

scholarly journals The Effect of the Inlet Boundary Layer on the Tip Flow Field in a Transonic Fan Rotor

1998 ◽  
Author(s):  
Junji Takado ◽  
Toyotaka Sonoda ◽  
Satoshi Nakamura
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Interaction Region ◽  
Operating Conditions ◽  
Laser Doppler Velocimeter ◽  
Intensity Level ◽  
Pressure Surface ◽  
Transonic Fan

Experimental and numerical investigations have been carried out to understand the effects of the inlet boundary layer (IBL) on the tip flow field including the aerodynamic performance in a transonic fan rotor. Both the steady and the unsteady phenomena in the tip flow field have been investigated for operating conditions near peak efficiency and near stall with the two types of tip IBL. In order 10 study these phenomena, high response pressure data with Kulite transducers and laser doppler velocimeter (LDV) data have been acquired around the tip region. Furthermore, three-dimensional Navier-Stokes numerical simulations have been compared with the measured results. The results indicate that the tip IBL significantly influences the spanwise distribution of pressure ratio around the tip region and the stall characteristics including the passage shock / tip leakage vortex interaction, the blockage generation, the wake structure, and the unsteadiness of the tip flow field. In particular, at a near stall condition for the thick IBL with high turbulence intensity level, the tip diffusion level is increased due to a larger blockage, which is generated downstream of a much stronger interaction region. These phenomena are a consequence of the low momentum fluid in the tip IBL, and significantly reduce the stall margin. Furthermore, the unsteadiness drastically increases around the interaction region and around the pressure surface where the blockage migrates. These unsteady phenomena are distinctive features near stall. Downstream of the rotor, the larger and more unsteady blockage is discharged from the pressure surface side, and complicates the three-dimensional rotor exit flow field around the tip region.

Radial Decomposition of Blade Vibration to Identify a Stall Flutter Source in a Transonic Fan

2019 ◽  
Vol 141 (10) ◽  
Author(s):  
Q. Rendu ◽  
M. Vahdati ◽  
L. Salles
Keyword(s):  
Pressure Fluctuations ◽  
Navier Stokes ◽  
Local Vibration ◽  
Blade Vibration ◽  
Nonlinear Solver ◽  
Low Mass ◽  
Stall Flutter ◽  
Fan Blade ◽  
Unstable Oscillation ◽  
Transonic Fan

Abstract This paper investigates the three dimensionality of the unsteady flow responsible for stall flutter instability. Nonlinear unsteady Reynolds-averaged Navier–Stokes (RANS) computations are used to predict the aeroelastic behavior of a fan blade at part speed. Flutter is experienced by the blades at low mass flow for the first flap mode at nodal diameter 2. The maximal energy exchange is located near the tip of the blade, at 90% span. The modeshape is radially decomposed to investigate the main source of instability. This decomposition method is validated for the first time in 3D using a time-marching nonlinear solver. The source of stall flutter is finally found at 65% span where the local vibration induces an unstable oscillation of the shock-wave of large amplitude. This demonstrates that the radial migration of the pressure fluctuations must be taken into account to predict stall flutter.

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