Experimental Investigation of an Aerodynamically Mistuned Oscillating Compressor Cascade

2019 ◽  
Vol 141 (7) ◽  
Author(s):  
Leonie Malzacher ◽  
Christopher Schwarze ◽  
Valentina Motta ◽  
Dieter Peitsch
Keyword(s):  
Test Facility ◽  
Random Pattern ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Angle Variation ◽  
Linear Cascade ◽  
Speed Test ◽  
The One ◽  
Phase Angles ◽  
Stagger Angle

In this paper, the effect of aerodynamic mistuning on stability of a compressor cascade is studied. The experiments have been carried out at a low-speed test facility of the Technische Universität Berlin. The test section contains a linear cascade with compressor blades that are forced to oscillate in sinusoidal pitching motion. The aerodynamic mistuning is realized by a blade-to-blade stagger angle variation, and three mistuning patterns have been investigated: one-blade mis-staggering, alternating mis-staggering, and random mis-staggering. Mis-staggering can have a stabilizing or destsabilizing effect, but depends strongly on the amount of detuning that alters the flow passage. For positive stagger angle variation for the one-blade and alternating mis-staggering, the trend of the damping curve was maintained, in the sense that the unstable interblade phase angles (IBPAs) remained unstable. For negative stagger angle variation, one IBPA shifted from stable to unstable. For the random pattern, only very moderate changes are observed. The cascade stability was not noticeably affected by the aerodynamic mistuning.

Experimental Investigation of an Aerodynamically Mistuned Oscillating Compressor Cascade

2018 ◽  
Author(s):  
Leonie Malzacher ◽  
Valentina Motta ◽  
Christopher Schwarze ◽  
Dieter Peitsch
Keyword(s):  
Test Facility ◽  
Random Pattern ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Angle Variation ◽  
Linear Cascade ◽  
Speed Test ◽  
The One ◽  
Phase Angles ◽  
Stagger Angle

In this paper, the effect of aerodynamic mistuning on stability of a compressor cascade is studied. The experiments have been carried out at a low speed test facility of the Technische Universität Berlin. The test section contains a linear cascade with compressor blades that are forced to oscillate in sinusoidal pitching motion. The aerodynamic mistuning is realized by a blade-to-blade stagger angle variation, three mistuning patterns have been investigated: one-blade mis-staggering, alternating mis-staggering and random mis-staggering. Mis-staggering can have stabilizing or destsabilizing effect, but depends strongly on the amount of the detuning that alters the flow passage. For positive stagger angle variation for the one-blade and alternating mis-staggering, the trend of the damping curve was maintained, in the sense that the unstable interblade phase angles (IBPA) remained unstable. For negative stagger angle variation, one IBPA shifted from stable to unstable. For the random pattern only very moderate changes are observed. The cascade stability was not noticeably effected by the aerodynamic mistuning.

Experimental Investigation of the Forcing Function and Forced Pitching Blade Oscillations of an Annular Compressor Cascade in Transonic Flow

2010 ◽  
Author(s):  
Joachim Belz ◽  
Holger Hennings ◽  
Gerhard Kahl
Keyword(s):  
Transonic Flow ◽  
Turbine Blades ◽  
Rotor Blades ◽  
Test Facility ◽  
Experimental Investigations ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Forcing Function ◽  
Annular Cascade

The interaction between rotor blades and non-rotating stator blades is the most significant blade excitation mechanism in turbomachines. It is well documented in various numerical and experimental investigations for turbine cascades. Like turbine blades, also compressor blades are excited as well by potential fields of the following stator, the downstream flowfield of the stator of the previous stage or struts and incoming flow distortions. In this paper, experimental investigations of the excitation of a transonic compressor cascade due to gust generating struts upstream are presented. The experiments were performed in the test facility of non-rotating annular cascades at EPFL using a compressor cascade, which consists of 20 blades (NACA3506 profile) mounted on elastic spring suspensions for torsional motions at the midchord. For the non-rotating annular cascade, relative flow conditions similar to those present in a rotating cascade are generating by swirling the flow in front of the test test section. The struts are rotating in order to create a periodic excitation upstream of the cascade. The so generated pressure distribution on the cascade’s profiles as well as the measured vibration response of the blades are presented and compared for a pure subsonic and a transonic flow case.

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.

A Low Speed Compressor Test Rig for Flutter Investigations

2016 ◽  
Author(s):  
Leonie Malzacher ◽  
Silvio Geist ◽  
Dieter Peitsch ◽  
Holger Hennings
Keyword(s):  
Test Facility ◽  
Test Rig ◽  
Compressor Blades ◽  
Low Speed ◽  
Surface Pressure Distribution ◽  
Oil Flow ◽  
Oil Flow Visualization ◽  
Aero Engines ◽  
Flutter Testing ◽  
Phase Angles

A test facility for aereolastic investigations has been installed at the chair of Aero Engines at the Technische Universität Berlin. The test rig provides data for tool and code validation and is used for basic aeroelastic experiments. It is a low speed wind tunnel which allows free and controlled flutter testing. The test section contains a linear cascade with eleven compressor blades. Nine of them are elastically suspended. The paper presents a detailed description of the test facility, results to evaluate the overall flow quality and an aeroelastic model to predict the flutter velocity and critical interblade phase angles. Hot-wire anemometry has been applied to examine the inlet flow for several Mach- and Reynolds numbers. The results show small turbulence intensities. The blade surface pressure distribution and the flow field of the blade’s suction and pressure side has been accessed by oil flow visualization.

The Effect of Four Part Gap Geometry Configurations for Variable Stator Vanes in a Compressor Cascade

2012 ◽  
Author(s):  
Marcel Gottschall ◽  
Konrad Vogeler ◽  
Ronald Mailach
Keyword(s):  
Flow Field ◽  
Separated Flow ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Field Development ◽  
Operating Range ◽  
Rear Part ◽  
Axis Position ◽  
Stagger Angle ◽  
Variable Stator

The article describes numerical investigations on the influence of four different endwall clearance topologies for variable stator vanes to secondary flow field development and the performance of high pressure compressors. The aim of this work is to quantify the characteristics of different clearance configurations depending on the penny-axis position and the penny diameter for a typical operating range. All clearance configurations were implemented to a linear cascade of modern stator profiles. The analysis was introduced using a relative clearance size of 1.3% chord at three stagger angles and two characteristic Reynolds numbers to model the operating range on aircraft engines. 3D numerical calculations were carried out to gain information about the flow field inside the cascade. They were compared with measurements of a 5-hole-probe as well as pressure tappings on the airfoil and the endwall. The CFD shows the clearance characteristics in good agreement with the measurements for the lower and the nominal stagger angle. Small gaps in the rear part of the vane have a beneficial effect on the flow field. In contrast, a clearance in the higher loaded front part of the vane always resulted in increased losses. Otherwise, the significant enhanced performance of a rear part gap, which was measured at the higher stagger angle, was not reflected by the CFD. The reduced mixing losses and the higher averaged flow turning even compared to a configuration without a clearance are not verified with the calculations. Large flow separations at the high stagger angle result in a two to four times higher underturning of the CFD in comparison to the experiments. The clearance effects to the characteristic radial loss distribution up to 40 % bladeheight also deviate from the measurements due to heavy mixing of clearance and reversed separated flow.

A Physically Consistent Reduced Order Model for Plasma Aeroelastic Control on Compressor Blades

2018 ◽  
Author(s):  
Valentina Motta ◽  
Leonie Malzacher ◽  
Dieter Peitsch ◽  
Giuseppe Quaranta
Keyword(s):  
Suction Side ◽  
Reduced Order Model ◽  
Plasma Actuators ◽  
Pitching Moment ◽  
Compressor Cascade ◽  
Order Model ◽  
Compressor Blades ◽  
Reduced Order ◽  
Thin Airfoil ◽  
Pressure Suction

Plasma actuators may be successfully employed as virtual control surfaces, located at the trailing edge of blades, both on the pressure and on the suction side, to control the aeroelastic response of a compressor cascade. Actuators generate an induced flow against the direction of the freestream. As a result, actuating on the pressure side yields an increase in lift and nose down pitching moment, whereas the opposite is obtained by operating on the suction side. A properly phased alternate pressure/suction side actuation allows to reduce vibration and to delay the flutter onset. This paper presents the development of a linear frequency domain reduced order model for lift and pitching moment of the plasma-equipped cascade. Specifically, an equivalent thin airfoil model is used as a physically consistent basis for the model. Modifications in the geometry of the thin airfoil are generated to account for the effective chord and camber changes induced by the plasma actuators, as well as for the effects of the neighboring blades. The model reproduces and predicts correctly the mean and the unsteady loads, along with the aerodynamic damping on the plasma equipped cascade. The relationship between the parameters of the reduced order model with the flow physics is highlighted.

Large-Eddy Simulation of Roughened NACA65 Compressor Cascade

2017 ◽  
Author(s):  
Jongwook Joo ◽  
Gorazd Medic ◽  
Om Sharma
Keyword(s):  
Surface Roughness ◽  
Large Eddy Simulations ◽  
Surface Imaging ◽  
Potential Solution ◽  
Discrete Elements ◽  
Compressor Cascade ◽  
Eddy Simulation ◽  
Large Eddy ◽  
Rans Modeling ◽  
The One

Large eddy simulations over a NACA65 compressor cascade with roughness were performed for multiple roughness heights. The experiments show flow separation as airfoil roughness is increased. In LES computations, surface roughness was represented by regularly arranged discrete elements using guidelines from Schlichting. Results from wall-resolved LES indicate that specifying an equivalent sandgrain roughness height larger than the one in experiments is required to reproduce the same effects observed in experiments. This highlights the persisting uncertainty with matching the experimental roughness geometry in LES computations, pointing towards surface imaging and digitization as a potential solution. Some initial analysis of flow physics has been conducted with the aim of guiding the RANS modeling for roughness.

Experimental Investigation of the Effect of the Probe Support Tail Structure on the Compressor Cascade Flow Field

2019 ◽  
Vol 36 (1) ◽  
pp. 9-18
Author(s):  
Honghui Xiang ◽  
Ning Ge ◽  
Jie Gao ◽  
Rongfei Yang ◽  
Minjie Hou
Keyword(s):  
Experimental Investigation ◽  
Flow Field ◽  
Flow Condition ◽  
Test Facility ◽  
Axial Distance ◽  
Compressor Cascade ◽  
Inlet Flow ◽  
Disturbance Intensity ◽  
Cascade Flow ◽  
Tail Structure

Abstract Aiming at resolving the problem of measuring probe blockage effect in the performance experiments of high loaded axial flow compressors, an experimental investigation of the probe support disturbance effect on the compressor cascade flow field was conducted on a transonic plane cascade test facility. The influence characteristics of the probe support tail structure on the cascade downstream flow field under different operation conditions were revealed through the detailed analysis of the test data. The results show that the aerodynamic coupling effect between the upstream probe support wake and the downstream cascade flow field is very intense. Some factors, i. e. inlet Mach number, probe support tail structure, circumferential installing position of probe, and axial distance from the probe support trailing edge to the downstream cascade, are found to have the most impact on the probe disturbance intensity. Under high speed inlet flow condition, changing probe support tail structure can’t inhibit probe support disturbance intensity effectively. Whereas under low speed inlet flow condition, compared with the cylindrical probe, the elliptic probe can inhibit probe support wake loss and reduce disturbance effects on the downstream cascade flow field.

Vortex Simulation of Propagating Stall in a Linear Cascade of Airfoils

1986 ◽  
Vol 108 (3) ◽  
pp. 304-312 ◽  
Author(s):  
C. G. Speziale ◽  
F. Sisto ◽  
S. Jonnavithula
Keyword(s):  
Experimental Studies ◽  
Vortex Method ◽  
Reynolds Numbers ◽  
Vortex Sheets ◽  
Discrete Vortex ◽  
Linear Cascade ◽  
Attached Flow ◽  
Tracing Algorithm ◽  
High Reynolds ◽  
Stagger Angle

A numerical simulation of propagating stall in a linear cascade of airfoils at high Reynolds numbers is conducted using a vortex method which was first developed by Spalart [7] for this problem. In this approach, the vorticity is discretized into a large collection of vortex blobs whose motion is tracked in time by the use of a well-known vortex tracing algorithm based on the Euler equation. The near-wall effects of viscosity are accounted for by the creation of discrete vortex sheets at the boundaries of the airfoils consistent with the no-slip condition. These boundary vortices are then released into the flow field downstream of the separation points which are obtained from a boundary-layer routine. Calculations are presented for a variety of flow geometries. It is demonstrated that (for a given cascade of airfoils, disturbance wavelength, and stagger angle) several different flow regimes are obtained: Attached flow at lower angles of attack and a chaotic deep stall configuration at larger angles of attack with a narrow intermediate range of such angles where propagating stall occurs. The physical characteristics of this propagating stall are parameterized and a quantitative study of the effects of camber and imposed wavelength is conducted. Comparisons are made with previous theoretical and experimental studies.

Sign in / Sign up

Export Citation Format

Share Document