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.

Aeroelastic Investigation of an Annular Transonic Compressor Cascade: Experimental Results

2015 ◽  
Author(s):  
Virginie Anne Chenaux ◽  
Achim Zanker ◽  
Peter Ott
Keyword(s):  
Spatial Resolution ◽  
Transonic Flow ◽  
High Spatial Resolution ◽  
Experimental Results ◽  
Quality Data ◽  
Test Facility ◽  
Flow Conditions ◽  
Compressor Cascade ◽  
High Quality ◽  
Transonic Compressor

A reliable determination of the unsteady aerodynamic loads acting on the blades is essential to predict the aeroelastic stability of vibrating compressor cascades with accuracy. At transonic flow conditions, the vibration of the shock may change the blade aeroelastic behavior. Numerical tools still have difficulties to capture the physics associated to this effect. In order to increase the prediction’s accuracy, high quality experimental data at high spatial resolution is therefore required to enable the calibration and validation of these tools. Within the frame of the European project FUTURE, experimental aeroelastic investigations were performed on a transonic compressor cascade in the Non-Rotating Annular Test Facility at EPFL. Associated to the measurements, the numerical flutter prediction procedure was applied. This paper focuses on the experimental results. The experimental database gained during the project is presented and aims at helping the aeroelastic community to develop and improve their flutter prediction capabilities. The test model consists of twenty prismatic blades. Each blade of the cascade assembly was mounted on an elastic spring element enabling harmonic bending vibrations in the twenty possible cascade’s travelling wave modes. Large efforts were made to improve the measuring techniques and to provide high quality data at relatively high spatial resolution. For various sub- and transonic flow conditions, steady-state and unsteady blade surface pressure distributions were measured to evaluate the local contributions to the blade stability in terms of local aerodynamic work. The blade global aerodynamic stability is determined applying an integration of all unsteady pressure signals measured over the airfoil.

Comparison of Theoretical and Experimental Data for an Oscillating Transonic Compressor Cascade

1999 ◽  
Author(s):  
Volker Carstens ◽  
Stefan Schmitt
Keyword(s):  
Experimental Data ◽  
Transonic Flow ◽  
Operating Conditions ◽  
Experimental Investigations ◽  
Test Case ◽  
Flow Conditions ◽  
Compressor Cascade ◽  
Blade Passage ◽  
Aerodynamic Damping ◽  
Transonic Compressor

Numerical and experimental results are compared for a compressor cascade performing harmonic oscillations in transonic flow. The flow field was calculated by a Q3D Navier Stokes code, the basic features of which are the use of an upwind flux difference scheme for the convective terms, the implementation of an effective one-equation turbulence model and the use of deforming multi-block grids. The experimental investigations were performed in an annular cascade windtunnel where unsteady blade pressures were measured for two different operating conditions of the cascade. The present data were all obtained for tuned torsional modes where the blades performed pitching oscillations with the same frequency and amplitude, but with a constant interblade phase angle. In the first test case the steady flow around the blades was purely subsonic. For the second test case the compressor cascade was run under transonic flow conditions where a normal shock in the front part of the blades’ suction side is followed by a blade passage shock. It becomes apparent that under subsonic flow conditions the predicted aerodynamic damping coefficients are in resonable agreement with the experimental data, although the numerical pressure amplitudes are much higher than the measured ones. In transonic flow significant discrepancies between computed and experimentally determined pressure amplitudes are observed, whereas the accuracy of the pressure phase prediction is comparable to the subsonic test case. Another important result of these investigations is that oscillations of the blade passage shock lead to strong variations of the local aerodynamic damping of the blades, but do not significantly change the global damping coefficient of the tested compressor cascade.

On the Importance of Transition Control at Transonic Compressor Blades

2019 ◽  
Author(s):  
A. Hergt ◽  
J. Klinner ◽  
S. Grund ◽  
C. Willert ◽  
W. Steinert ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
High Speed ◽  
Experimental Investigations ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Transition Control ◽  
Transonic Compressor ◽  
Air Jet ◽  
Loss Level

Abstract The flow through a transonic compressor cascade is characterized by high unsteadiness and a high loss level. This results from the shock waves in the blade cascade and their interaction with the blade suction side boundary layer. In the case of a laminar shock wave boundary layer interaction the loss level is higher due to the occurrence of a laminar separation bubble below the shock wave compared to the shock wave interaction with a turbulent boundary layer. In addition, the oscillation of the shock position in both cases influences the working range concerning the point of stall onset as well as leading to an unsteady interaction with the blade, called buffeting. The reduction of losses and of unsteadiness in the shock wave oscillation, connected to a decrease of the blade buffeting effect, are the aims of the current investigation. Therefore, experimental investigations using a roughness patch as well as air jet vortex generators in order to control the transition in a transonic compressor cascade have been conducted at the transonic cascade wind tunnel of DLR at Cologne. At an inflow Mach number of 1.21 a loss reduction for both transition control cases is achieved. In spite of a nearly uninfluenced fluctuation range of the passage shock wave compared to the reference cascade, the oscillation spectra of the transition control cases show a reduction of the shock movement amplitude at a frequency below 500 Hz and above 1 kHz. In the closing section of the paper a detailed discussion on the reasons for the resulting flow behaviour based on PIV and High Speed Shadowgraphy data is given. The resulting conclusion of the study is that the consideration of transition control at transonic compressor blades is very important in order to reduce losses and flow unsteadiness which directly influences blade buffeting and the numerical prediction quality of the stall onset.

Experimental Investigation of the Aerodynamic Stability of an Annular Compressor Cascade Performing Tuned Pitching Oscillations in Transonic Flow

1999 ◽  
Author(s):  
H. Hennings ◽  
J. Belz
Keyword(s):  
Transonic Flow ◽  
Suction Side ◽  
Test Facility ◽  
Feedback Signal ◽  
Detailed Knowledge ◽  
Compressor Cascade ◽  
Aerodynamic Stability ◽  
Pressure Transducers ◽  
Unsteady Pressure ◽  
Unsteady Aerodynamic

A prerequisite for aeroelastic stability investigations on vibrating compressor cascades is the detailed knowledge of the unsteady aerodynamic loads acting on the blades. In order to obtain precise insight into the aerodynamic damping of a vibrating blade assembly, a basic experiment was performed where unsteady pressure distributions were measured for subsonic and transonic flow conditions. The experiments were performed on a non-rotating, two-dimensional section of a compressor cascade in an annular test facility. The cascade consists of 20 blades (NACA3506 profile) mounted on elastic spring suspensions. In order to measure the unsteady pressure distribution, the cascade was set to tuned pitching oscillations (traveling wave modes). Each blade was driven to controlled harmonic torsional motions around midchord by a magnetic excitation system and by inductive displacement probes which measure the feedback signal of the motion. Steady and unsteady pressures were measured by steady pressure taps and piezo-electric pressure transducers, respectively. The measurement of the unsteady aerodynamic response to a shock vibrating on the suction side of the blades was enabled by a dense spacing of transducers in this region. The global aerodynamic stability is assessed by a damping coefficient evaluated from the out-of-phase parts of the unsteady moment coefficients and by the contributions from the local work coefficient, using the measured pressure data.

An Experimentally Derived Model to Predict the Water Film in a Compressor Cascade With Droplet Laden Flow

2019 ◽  
Vol 141 (9) ◽  
Author(s):  
Niklas Neupert ◽  
Janneck Christoph Harbeck ◽  
Franz Joos
Keyword(s):  
Film Thickness ◽  
Gas Turbines ◽  
Three Dimensional ◽  
Water Film ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
System A ◽  
Industrial Gas Turbines ◽  
Measured Flow

In recent years, overspray fogging has become a powerful means for power augmentation of industrial gas turbines. Despite the positive thermodynamic effect on the cycle, droplets entering the compressor increase the risk of water droplet erosion and deposition of water on the blades leading to an increase of required torque and profile loss. Due to this, detailed information about the structure and the amount of water on the surface is key for compressor performance. Experiments were conducted with a droplet laden flow in a transonic compressor cascade focusing on the film formed by the deposited water. Two approaches were taken. In the first approach, the film thickness on the blade was directly measured using white light interferometry. Due to significant distortion of the flow caused by the measurement system, a transfer of the measured film thickness to the undisturbed case is not possible. Therefore, a film model is adapted to describe the film flow in terms of height averaged film parameters. In the second approach, experiments were conducted in an undisturbed cascade setup and the water film pattern was measured using a nonintrusive quantitative image processing tool. Utilizing the measured flow pattern in combination with findings from the literature, the rivulet flow structure is resolved. From continuity of the water flow, a film thickness is derived showing good agreement with the previously calculated results. Using both approaches, a three-dimensional (3D) reconstruction of the water film pattern is created giving first experimental results of the film forming on stationary compressor blades under overspray fogging conditions.

Influence of Gap Detailing on Calculated Unsteady Non-Adjacent Blade Row Aero-Forcing in a Transonic Compressor Stage

2020 ◽  
Vol 142 (11) ◽  
Author(s):  
Tobias Gezork ◽  
Paul Petrie-Repar
Keyword(s):  
Guide Vane ◽  
Computational Cost ◽  
Life Time ◽  
Forced Response ◽  
Inlet Guide Vane ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Forcing Function ◽  
Gap Flow ◽  
Blade Row

Abstract Resonant or close to resonant forced response excitation of compressor blades limits component life time and can potentially lead to high-cycle fatigue failure if the exciting forces are large and damping is insufficient. When numerically quantifying the forcing function by means of simulations, simplifications are typically made in the analysis to reduce complexity and computational cost. In this paper, we numerically investigate how the blade forcing function is influenced by the rotor tip gap flow and by flow across gaps in the upstream variable inlet guide vane row. Unsteady simulations are made using a test rig geometry where a forcing crossing with an excitation from a non-adjacent blade row had previously been measured. The effects of the gaps on the forcing function for the first torsion mode are presented for both the non-adjacent blade row excitation (changes compared with a case without gaps indicating a 20% reduction) and an adjacent excitation (changes indicating an 80% increase in terms of forcing function amplitude comparing with a case without gaps).

Effect of Shock Wave Movement on Aerodynamic Instability of Annular Cascade Oscillating in Transonic Flow

1988 ◽  
Author(s):  
Hiroshi Kobayashi
Keyword(s):  
Shock Wave ◽  
Phase Angle ◽  
Test Facility ◽  
Blade Surface ◽  
Compressor Cascade ◽  
Reduced Frequency ◽  
Wide Range ◽  
Cascade Flutter ◽  
Annular Cascade ◽  
Interblade Phase Angle

Effects attributable to shock wave movement on cascade flutter were examined for both turbine and compressor blade rows, using a controlled-oscillating annular cascade test facility and a method for accurately measuring time-variant pressures on blade surfaces. Nature of the effects and blade surface extent influenced by the shock movement were clarified in a wide range of Mach number, reduced frequency and interblade phase angle. Remarkable unsteady aerodynamic force was generated by the shock movement and it significantly affected the occurrence of compressor cascade flutter as well as turbine one. For turbine cascade the interblade phase angle remarkably controlled the effect of the force, while for compressor one the reduced frequency dominated it. The chordwise extent on blade surface influenced by the shock movement was suggested to be about 6% chord length.

scholarly journals Comparison of Experimental and Theoretical Results for Unsteady Transonic Cascade Flow at Design and Off-Design Conditions

1993 ◽  
Author(s):  
V. Carstens ◽  
A. Bölcs ◽  
H. Körbacher
Keyword(s):  
Transonic Flow ◽  
Incidence Angle ◽  
Design Flow ◽  
The Other ◽  
Test Facility ◽  
Experimental Investigations ◽  
Suction Surface ◽  
Flow Angle ◽  
Flux Vector Splitting ◽  
Theoretical Results

This paper presents experimental and theoretical results for turbine cascades performing harmonic oscillations in transonic flow at design and off-design conditions. The experimental investigations were performed in an annular test facility where unsteady blade pressures were measured in two different test cascades, one operating at the nominal inlet flow angle, the other at an incidence angle exceeding the normal value by more than 20 degrees. The corresponding theoretical results were computed with a 2D Euler code which makes use of flux vector splitting in combination with a time-dependent grid generation. The present data were all obtained for tuned bending modes where the blades performed heaving oscillations with the same frequency and amplitude, but with a constant interblade phase angle. For the cascade operating at design conditions, the steady flow was purely subsonic. The other test cascade was run in transonic flow, and a normal shock appeared on the rear part of the blade’s suction surface. It was found that measured unsteady pressure and damping coefficients are well reproduced by the computed results for the first test cascade. In the case of steady off-design flow (the second test cascade), significant differences between experimental and theoretical results are observed.

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.

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