Prediction of Aerodynamically Induced Vibrations in Turbomachinery Blading

1983 ◽  
Vol 105 (4) ◽  
pp. 375-381 ◽  
Author(s):  
D. Hoyniak ◽  
S. Fleeter
Keyword(s):  
Energy Balance ◽  
Incidence Angle ◽  
Forced Response ◽  
Aerodynamic Coefficients ◽  
Unsteady Aerodynamic ◽  
Reduced Frequency ◽  
Balance Technique ◽  
Energy Dissipated ◽  
Airfoil Cascade ◽  
Interblade Phase Angle

To predict the aerodynamically forced response of an airfoil, an energy balance between the unsteady aerodynamic work and the energy dissipated through the airfoil structural and aerodynamic damping is performed. Theoretical zero incidence unsteady aerodynamic coefficients are then utilized in conjunction with this energy balance technique to predict the effects of reduced frequency, inlet Mach number, cascade geometry and interblade phase angle on the torsion mode aerodynamically forced response of the cascade. In addition, experimental unsteady aerodynamic gust data for flat plate and cambered cascaded airfoils are used together with these theoretical cascade unsteady self-induced aerodynamic coefficients to indicate the effects of incidence angle and airfoil camber on the forced response of the airfoil cascade.

The Aerodynamics of an Oscillating Cascade in a Compressible Flow Field

1990 ◽  
Vol 112 (4) ◽  
pp. 759-767 ◽  
Author(s):  
D. H. Buffum ◽  
S. Fleeter
Keyword(s):  
Incidence Angle ◽  
Major Effect ◽  
Cascade Model ◽  
Nasa Lewis Research ◽  
Airfoil Surface ◽  
Torsion Mode ◽  
Unique Data ◽  
Reduced Frequency ◽  
Airfoil Cascade ◽  
Interblade Phase Angle

Fundamental experiments are performed in the NASA Lewis Research Center Transonic Oscillating Cascade Facility to investigate and quantify the aerodynamics of a cascade of bioconvex airfoils executing torsion mode oscillations at realistic reduced frequency values. Both steady and unsteady airfoil surface pressures are measured at two inlet Mach numbers, 0.65 and 0.80. and two incidence angles, 0 and 7 deg, with the harmonic torsional airfoil cascade oscillations at realistic high reduced frequency and unsteady data obtained at several interbladephase angle values. The time-variant pressures are analyzed by means of discrete Fourier transform techniques, with these unique data compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle has a major effect on the chordwise distributions of the airfoil surface unsteady pressure, with the effects of reduced frequency, incidence angle, and Mach number somewhat less significant.

scholarly journals The Aerodynamics of an Oscillating Cascade in a Compressible Flow Field

1989 ◽  
Author(s):  
Daniel H. Buffum ◽  
Sanford Fleeter
Keyword(s):  
Phase Angle ◽  
Incidence Angle ◽  
Major Effect ◽  
Nasa Lewis Research ◽  
Airfoil Surface ◽  
Torsion Mode ◽  
Unique Data ◽  
Reduced Frequency ◽  
Airfoil Cascade ◽  
Interblade Phase Angle

Fundamental experiments are performed in the NASA Lewis Research Center Transonic Oscillating Cascade Facility to investigate and quantify the aerodynamics of a cascade of biconvex airfoils executing torsion mode oscillations at realistic reduced frequency values. Both steady and unsteady airfoil surface pressures are measured at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 degrees, with the harmonic torsional airfoil cascade oscillations at realistic high reduced frequency and unsteady data obtained at several interblade phase angle values. The time-variant pressures are analyzed by means of discrete Fourier transform techniques, with these unique data compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle has a major effect on the chordwise distributions of the airfoil surface unsteady pressure, with the effects of reduced frequency, incidence angle, and Mach number somewhat less significant.

An Experimental Determination of the Unsteady Aerodynamics in a Controlled Oscillating Cascade

1977 ◽  
Vol 99 (1) ◽  
pp. 88-96 ◽  
Author(s):  
S. Fleeter ◽  
A. S. Novick ◽  
R. E. Riffel ◽  
J. E. Caruthers
Keyword(s):  
Phase Angle ◽  
Unsteady Aerodynamics ◽  
Time Dependent ◽  
Phase Lag ◽  
Reduced Frequency ◽  
On Line ◽  
Torsional Frequency ◽  
Airfoil Cascade ◽  
Interblade Phase Angle

A unique supersonic inlet flow field unsteady cascade experiment is described wherein the time-dependent pressure distribution within an harmonically oscillating airfoil cascade is quantitatively determined. The torsional frequency of oscillation and the inter-blade phase angle are precisely controlled by means of on-line digital computers. The dynamic data obtained include the chordwise distribution of the unsteady pressure magnitude and its phase lag as referenced to the airfoil motion. Parameters varied include the cascade inlet Mach number, the interblade phase angle, and the reduced frequency. The time-dependent data are correlated with state-of-the-art analytical predictions.

Rotor Wake Generated Unsteady Aerodynamic Response of a Compressor Stator

1978 ◽  
Vol 100 (4) ◽  
pp. 664-675 ◽  
Author(s):  
S. Fleeter ◽  
R. L. Jay ◽  
W. A. Bennett
Keyword(s):  
Pressure Difference ◽  
Second Harmonic ◽  
Incidence Angle ◽  
Dynamic Pressure ◽  
Pressure Differential ◽  
Phase Lag ◽  
Trailing Edge ◽  
Unsteady Pressure ◽  
Forcing Function ◽  
Reduced Frequency

An experimental investigation was conducted to determine the fluctuating pressure distribution on a stationary vane row, with the primary source of excitation being the wakes from the upstream rotor blades. This was accomplished in a large scale, low speed, single stage research compressor. The forcing function, the velocity defect created by the rotor wakes, was measured with a crossed hot-wire probe. The aerodynamic response on the vanes was measured by means of flush mounted high response dynamic pressure transducers. The dynamic data were analyzed to determine the chordwise distribution of the dynamic pressure coefficient and aerodynamic phase lag as referenced to a transverse gust at the vane leading edge. Vane suction and pressure surface data as well as the pressure difference across the vane were obtained for reduced frequency values ranging from 3.65 to 16.80 and for an incidence angle range of 35.5 deg. The pressure difference data were correlated with a state-of-the-art aerodynamic cascade transverse gust analysis. The correlation was quite good for all reduced frequency values for small values of incidence. For the more negative incidence angle data points, it was shown that a convected wake phenomena not modeled in the analysis existed. Both the first and second harmonic unsteady pressure differential magnitude data decrease in the chordwise direction. The second harmonic magnitude data attains a value very nearly zero at the vane trailing edge transducer location, while the first harmonic data is still finite, albeit small, at this location. That the magnitude of the unsteady pressure differential data approaches zero near to the trailing edge, particularly the second harmonic data which has reduced frequency values to 16.8, is significant in that it reflects upon the validity of the Kutta condition for unsteady flows.

Aerodynamic Asymmetry Analysis of Unsteady Flows in Turbomachinery

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
Kivanc Ekici ◽  
Robert E. Kielb ◽  
Kenneth C. Hall
Keyword(s):  
Harmonic Balance ◽  
Coupling Effect ◽  
Unsteady Flows ◽  
Unsteady Aerodynamic ◽  
Blade Row ◽  
Balance Technique ◽  
Turbomachinery Blades ◽  
Aerodynamic Asymmetry ◽  
Frequency Domain Approach ◽  
Harmonic Balance Technique

A nonlinear harmonic balance technique for the analysis of aerodynamic asymmetry of unsteady flows in turbomachinery is presented. The present method uses a mixed time-domain/frequency-domain approach that allows one to compute the unsteady aerodynamic response of turbomachinery blades to self-excited vibrations. Traditionally, researchers have investigated the unsteady response of a blade row with the assumption that all the blades in the row are identical. With this assumption the entire wheel can be modeled using complex periodic boundary conditions and a computational grid spanning a single blade passage. In this study, the steady/unsteady aerodynamic asymmetry is modeled using multiple passages. Specifically, the method has been applied to aerodynamically asymmetric flutter problems for a rotor with a symmetry group of 2. The effect of geometric asymmetries on the unsteady aerodynamic response of a blade row is illustrated. For the cases investigated in this paper, the change in the diagonal terms (blade on itself) dominated the change in stability. Very little mode coupling effect caused by the off-diagonal terms was found.

scholarly journals Determination of Aerodynamic Damping at High Reduced Frequencies

2018 ◽  
Author(s):  
Minghao Pan ◽  
Paul Petrie-Repar ◽  
Hans Mårtensson ◽  
Tianrui Sun ◽  
Tobias Gezork
Keyword(s):  
Acoustic Resonance ◽  
Forced Response ◽  
Economic Consequences ◽  
Navier Stokes ◽  
Flow Solver ◽  
Aerodynamic Damping ◽  
Acoustic Modes ◽  
Reduced Frequency ◽  
Standard Configuration

In turbomachines, forced response of blades is blade vibrations due to external aerodynamic excitations and it can lead to blade failures which can have fatal or severe economic consequences. The estimation of the level of vibration due to forced response is dependent on the determination of aerodynamic damping. The most critical cases for forced response occur at high reduced frequencies. This paper investigates the determination of aerodynamic damping at high reduced frequencies. The aerodynamic damping was calculated by a linearized Navier-Stokes flow solver with exact 3D non-reflecting boundary conditions. The method was validated using Standard Configuration 8, a two-dimensional flat plate. Good agreement with the reference data at reduced frequency 2.0 was achieved and grid converged solutions with reduced frequency up to 16.0 were obtained. It was concluded that at least 20 cells per wavelength is required. A 3D profile was also investigated: an aeroelastic turbine rig (AETR) which is a subsonic turbine case. In the AETR case, the first bending mode with reduced frequency 2.0 was studied. The 3D acoustic modes were calculated at the far-fields and the propagating amplitude was plotted as a function of circumferential mode index and radial order. This plot identified six acoustic resonance points which included two points corresponding to the first radial modes. The aerodynamic damping as a function of nodal diameter was also calculated and plotted. There were six distinct peaks which occurred in the damping curve and these peaks correspond to the six resonance points. This demonstrates for the first time that acoustic resonances due to higher order radial acoustic modes can affect the aerodynamic damping at high reduced frequencies.

Aerodynamic Asymmetry Analysis of Unsteady Flows in Turbomachinery

2008 ◽  
Author(s):  
Kivanc Ekici ◽  
Robert E. Kielb ◽  
Kenneth C. Hall
Keyword(s):  
Harmonic Balance ◽  
Coupling Effect ◽  
Unsteady Flows ◽  
Unsteady Aerodynamic ◽  
Blade Row ◽  
Balance Technique ◽  
Turbomachinery Blades ◽  
Aerodynamic Asymmetry ◽  
Frequency Domain Approach ◽  
Harmonic Balance Technique

A nonlinear harmonic balance technique for the analysis of aerodynamic asymmetry of unsteady flows in turbomachinery is presented. The present method uses a mixed time-domain/frequency-domain approach that allows one to compute the unsteady aerodynamic response of turbomachinery blades to self-excited vibrations. Traditionally, researchers have investigated the unsteady response of a blade row with the assumption that all the blades in the row are identical. With this assumption the entire wheel can be modeled using complex periodic boundary conditions and a computational grid spanning a single blade passage. In this study, the steady/unsteady aerodynamic asymmetry is modeled using multiple passages. Specifically, the method has been applied to aerodynamically asymmetric flutter problems for a rotor with a symmetry group of two. The effect of geometric asymmetries on the unsteady aerodynamic response of a blade row is illustrated. For the cases investigated in this paper, the change in the diagonal terms (blade on itself) dominated the change in stability. Very little mode coupling effect caused by the off-diagonal terms was found.

Forced Response Sensitivity of a Mistuned Rotor From an Embedded Compressor Stage

2015 ◽  
Author(s):  
Fanny M. Besem ◽  
Robert E. Kielb ◽  
Nicole L. Key
Keyword(s):  
Experimental Data ◽  
Forced Response ◽  
Symmetric System ◽  
Response Analysis ◽  
Cfd Simulations ◽  
Forcing Function ◽  
Wear And Tear ◽  
The Individual ◽  
The Impact ◽  
Interblade Phase Angle

The frequency mistuning that occurs due to manufacturing variations and wear and tear of the blades can have a significant effect on the flutter and forced response behavior of a blade row. Similarly, asymmetries in the aerodynamic or excitation forces can tremendously affect the blade responses. When conducting CFD simulations, all blades are assumed to be tuned (i.e. to have the same natural frequency) and the aerodynamic forces are assumed to be the same on each blade except for a shift in interblade phase angle. The blades are thus predicted to vibrate at the same amplitude. However, when the system is mistuned or when asymmetries are present, some blades can vibrate with a much higher amplitude than the tuned, symmetric system. In this research, we first conduct a deterministic forced response analysis of a mistuned rotor and compare the results to experimental data from a compressor rig. It is shown that tuned CFD results cannot be compared directly with experimental data because of the impact of frequency mistuning on forced response predictions. Moreover, the individual impact of frequency, aerodynamic, and forcing function perturbations on the predictions is assessed, leading to the conclusion that a mistuned system has to be studied probabilistically. Finally, all perturbations are combined and Monte-Carlo simulations are conducted to obtain the range of blade response amplitudes that a designer could expect.

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