Unsteady Aerodynamics and Gapwise Periodicity of Oscillating Cascaded Airfoils

1983 ◽  
Vol 105 (3) ◽  
pp. 565-574 ◽  
Author(s):  
F. O. Carta
Keyword(s):  
Unsteady Flow ◽  
Fourier Coefficient ◽  
Unsteady Aerodynamics ◽  
Leading Edge ◽  
Experimental Measurements ◽  
Pressure Data ◽  
Linear Cascade ◽  
Aerodynamic Damping ◽  
The Stability ◽  
Interblade Phase Angle

Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade and over the chord of the center blade. The pressure data were reduced to Fourier coefficient form for direct comparison and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. In addition, results from two unsteady theories for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results that emerged from this investigation were: (a) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested, (b) as before, the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades, and (c) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.

scholarly journals Unsteady Gapwise Periodicity of Oscillating Cascaded Airfoils

1982 ◽  
Author(s):  
F. O. Carta
Keyword(s):  
Unsteady Flow ◽  
Fourier Coefficient ◽  
Phase Angle ◽  
Leading Edge ◽  
Experimental Measurements ◽  
Pressure Data ◽  
Linear Cascade ◽  
Aerodynamic Damping ◽  
The Stability ◽  
Interblade Phase Angle

Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade and over the chord of the center blade. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. In addition, results from two unsteady theories for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results that emerged from this investigation were: (a) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested, (b) as before, the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades, and (c) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.

Experimental Study of Aerodynamic Damping of an Annular Compressor Cascade With Large Mean Incidences

2019 ◽  
Vol 141 (6) ◽  
Author(s):  
M. C. Keerthi ◽  
Abhijit Kushari
Keyword(s):  
Torsional Oscillation ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Aerodynamic Damping ◽  
Reduced Frequency ◽  
Incidence Studies ◽  
Stable Performance ◽  
Phase Variations ◽  
The Stability ◽  
Interblade Phase Angle

This study addresses flutter that can occur in compressors when operating at high relative incidence. Studies are performed on a subsonic annular compressor cascade containing a sector of blades that can be subjected to controlled torsional oscillation. Measurements taken on the centrally located blade are used to study the unsteady surface pressures developed. Three large mean incidences are considered to characterize the aeroelastic performance. Aerodynamic damping is calculated for each test condition and its variation due to interblade phase angle (IBPA), reduced frequency, and incidence is studied. The source of stability or instability is traced to the nature of unsteady pressures. When the incidence is below the static-stall limit, an increasing incidence is found to exhibit aeroelastically more stable performance, whereas beyond the limit, the stability worsens. For the latter, the amount of improvement in stability by increasing reduced frequency is less compared to the former and its variation with IBPA is not as regular owing to the associated large uncertainty. The nonlinearity effects were found to be relatively higher for this case, especially from the aft region of the suction surface. It is also found that the phase of the local fluctuating pressure and its location on the chord has a decisive influence on the aerodynamic damping and its trends with respect to various parameters are discussed. The results are expected to be useful in the assessing aerodynamic damping trends in relation to the pressure phase variations in specific regions along the chord.

Effect of Adjacent Blade Motion on the Aerodynamics of a Linear Cascade Blade

2012 ◽  
Author(s):  
K. Naresh Babu ◽  
A. Kushari ◽  
C. Venkatesan
Keyword(s):  
Phase Angle ◽  
Unsteady Aerodynamics ◽  
Major Effect ◽  
Compressor Blades ◽  
Unsteady Forces ◽  
Linear Cascade ◽  
Reduced Frequency ◽  
The Stability ◽  
Lift And Drag ◽  
Phase Angles

Due to the trend of increasing power and reducing weight, the fan and compressor blades of turbo machinery might be more sensitive to flutter, which must strictly be avoided in the design process. In order to increase our understanding of the flutter phenomena for fan and compressor cascades, aero-elastic investigations are essential. In the present work experiments were performed in the specifically designed Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics forces and moments acting on a blade in a linear cascade of blades when the instrumented blade is stationary and the two adjacent blades on both sides of the instrumented blade are executing torsion-mode oscillations about their mid-chord. A 5-component strain gage balance was used to measure the unsteady aerodynamic forces on the model blade. The forces were measured for six inter-blade phase angles (i.e., the phase angle between the moving blade motions at a given frequency where the central blade is stationary) at low subsonic conditions, different reduced frequencies and different stagger. The time-variant forces were analyzed and variation of lift and drag coefficients for different inter-blade phase angles and reduced frequencies were plotted. The experimental results indicate that the inter-blade phase angle had a major effect on the variation of the unsteady forces and that reduced frequency had a somewhat less significant effect. Also in order to investigate the influence of the reduced frequency and inter-blade phase angles on the global stability of the cascade and its local contributions, experiments were performed for different reduced frequencies and phase angles. At the higher inter-blade phase angles (180°) the blade remains aerodynamically stable at 0° stagger, but the stability reduces at higher stagger angles. The blade is usually unstable when the interblade phase angle is 0°. At different flow conditions, some of the inter-blade phase angles appear to be aerodynamically unstable.

Flow criticality governs leading-edge-vortex initiation on finite wings in unsteady flow

2021 ◽  
Vol 910 ◽  
Author(s):  
Yoshikazu Hirato ◽  
Minao Shen ◽  
Ashok Gopalarathnam ◽  
Jack R. Edwards
Keyword(s):  
Unsteady Flow ◽  
Leading Edge ◽  
Leading Edge Vortex ◽  
Edge Vortex

Abstract

Numerical Investigation of the Solidity Effect on Linear Compressor Cascades

2014 ◽  
Author(s):  
J. Sans ◽  
M. Resmini ◽  
J.-F. Brouckaert ◽  
S. Hiernaux
Keyword(s):  
Numerical Investigation ◽  
State Of The Art ◽  
Leading Edge ◽  
Operating Conditions ◽  
Compressor Cascade ◽  
Minimum Loss ◽  
Low Speed ◽  
High Mach ◽  
The Stability ◽  
The Impact

Solidity in compressors is defined as the ratio of the aerodynamic chord over the peripheral distance between two adjacent blades, the pitch. This parameter is simply the inverse of the pitch-to-chord ratio generally used in turbines. Solidity must be selected at the earliest design phase, i.e. at the level of the meridional design and represents a crucial step in the whole design process. Most of the existing studies on this topic rely on low-speed compressor cascade correlations from Carter or Lieblein. The aim of this work is to update those correlations for state-of-the-art controlled diffusion blades, and extend their application to high Mach number flow regimes more typical of modern compressors. Another objective is also to improve the physical understanding of the solidity effect on compressor performance and stability. A numerical investigation has been performed using the commercial software FINE/Turbo. Two different blade profiles were selected and investigated in the compressible flow regime as an extension to the low-speed data on which the correlations are based. The first cascade uses a standard double circular arc profile, extensively referenced in the literature, while the second configuration uses a state-of-the-art CDB, representative of low pressure compressor stator mid-span profile. Both profiles have been designed with the same inlet and outlet metal angles and the same maximum thickness but the camber and thickness distributions, the stagger angle and the leading edge geometry of the CDB have been optimized. The determination of minimum loss, optimum incidence and deviation is addressed and compared with existing correlations for both configurations and various Mach numbers that have been selected in order to match typical booster stall and choke operating conditions. The emphasis is set on the minimum loss performance at mid-span. The impact of the solidity on the operating range and the stability of the cascade are also studied.

Investigation of the stability of boundary layers by a finite-difference model of the Navier—Stokes equations

1976 ◽  
Vol 78 (2) ◽  
pp. 355-383 ◽  
Author(s):  
H. Fasel
Keyword(s):  
Finite Difference ◽  
Stability Theory ◽  
Stokes Equations ◽  
Time Dependent ◽  
Linear Stability Theory ◽  
Navier Stokes ◽  
Experimental Measurements ◽  
Navier Stokes Equations ◽  
The Stability ◽  
Small Periodic

The stability of incompressible boundary-layer flows on a semi-infinite flat plate and the growth of disturbances in such flows are investigated by numerical integration of the complete Navier–;Stokes equations for laminar two-dimensional flows. Forced time-dependent disturbances are introduced into the flow field and the reaction of the flow to such disturbances is studied by directly solving the Navier–Stokes equations using a finite-difference method. An implicit finitedifference scheme was developed for the calculation of the extremely unsteady flow fields which arose from the forced time-dependent disturbances. The problem of the numerical stability of the method called for special attention in order to avoid possible distortions of the results caused by the interaction of unstable numerical oscillations with physically meaningful perturbations. A demonstration of the suitability of the numerical method for the investigation of stability and the initial growth of disturbances is presented for small periodic perturbations. For this particular case the numerical results can be compared with linear stability theory and experimental measurements. In this paper a number of numerical calculations for small periodic disturbances are discussed in detail. The results are generally in fairly close agreement with linear stability theory or experimental measurements.

Modeling of External Pressure Gradient Influence on the Stability of Disturbances in the Boundary Layers of Compressed Gas

2013 ◽  
Vol 8 (4) ◽  
pp. 64-75
Author(s):  
Sergey Gaponov ◽  
Natalya Terekhova
Keyword(s):  
Mach Number ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Leading Edge ◽  
External Pressure ◽  
Transition To Turbulence ◽  
Compressed Gas ◽  
High Mach ◽  
The Stability ◽  
Very High

This work continues the research on modeling of passive methods of management of flow regimes in the boundary layers of compressed gas. Authors consider the influence of pressure gradient on the evolution of perturbations of different nature. For low Mach number M = 2 increase in pressure contributes to an earlier transition of laminar to turbulent flow, and, on the contrary, drop in the pressure leads to a prolongation of the transition to turbulence. For high Mach number M = 5.35 found that the acoustic disturbances exhibit a very high dependence on the sign and magnitude of the external gradient, with a favorable gradient of the critical Reynolds number becomes smaller than the vortex disturbances, and at worst – boundary layer is destabilized directly on the leading edge

Experimental and Theoretical Investigation of the Supersonic Boundary Layer Stability on the Swept Wing

2008 ◽  
Vol 3 (3) ◽  
pp. 34-38
Author(s):  
Sergey A. Gaponov ◽  
Yuri G. Yermolaev ◽  
Aleksandr D. Kosinov ◽  
Nikolay V. Semionov ◽  
Boris V. Smorodsky
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Leading Edge ◽  
Supersonic Boundary Layer ◽  
Mean Flow ◽  
Swept Wing ◽  
Wing Model ◽  
Dimensional Boundary ◽  
The Stability ◽  
Good Agreement

Theoretical and an experimental research results of the disturbances development in a swept wing boundary layer are presented at Mach number М = 2. In experiments development of natural and small amplitude controllable disturbances downstream was studied. Experiments were carried out on a swept wing model with a lenticular profile at a zero attack angle. The swept angle of a leading edge was 40°. Wave parameters of moving disturbances were determined. In frames of the linear theory and an approach of the local self-similar mean flow the stability of a compressible three-dimensional boundary layer is studied. Good agreement of the theory with experimental results for transversal scales of unstable vertices of the secondary flow was obtained. However the calculated amplification rates differ from measured values considerably. This disagreement is explained by the nonlinear processes observed in experiment

A Study of Influences of Inlet Total Pressure Distortions On Clearance Flow in an Axial Compressor

2021 ◽  
Author(s):  
Guoming Zhu ◽  
Xiaolan Liu ◽  
Bo Yang ◽  
Moru Song
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Stall Margin ◽  
Flow Cells ◽  
Static Pressure Distribution ◽  
Speed Flow ◽  
Clearance Flow ◽  
The Stability

Abstract The rotating distortion generated by upstream wakes or low speed flow cells is a kind of phenomenon in the inlet of middle and rear stages of an axial compressor. Highly complex inflow can obviously affect the performance and the stability of these stages, and is needed to be considered during compressor design. In this paper, a series of unsteady computational fluid dynamics (CFD) simulations is conducted based on a model of an 1-1/2 stage axial compressor to investigate the effects of the distorted inflows near the casing on the compressor performance and the clearance flow. Detailed analysis of the flow field has been performed and interesting results are concluded. The distortions, such as total pressure distortion in circumferential and radial directions, can block the tip region so that the separation loss and the mixing loss in this area are increased, and the efficiency and the total pressure ratio are dropped correspondingly. Besides, the distortions can change the static pressure distribution near the leading edge of the rotor, and make the clearance flow spill out of the rotor edge more easily under near stall condition, especially in the cases with co-rotating distortions. This phenomenon can be used to explain why the stall margin is deteriorated with nonuniform inflows.

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