Analysis of mistuned forced response in an axial high-pressure compressor rig with focus on Tyler–Sofrin modes

2019 ◽  
Vol 123 (1261) ◽  
pp. 356-377
Author(s):  
F. Figaschewsky ◽  
A. Kühhorn ◽  
B. Beirow ◽  
T. Giersch ◽  
S. Schrape
Keyword(s):  
Axial Compressor ◽  
Maximum Amplitude ◽  
Forced Response ◽  
Operating Conditions ◽  
Cfd Simulations ◽  
High Pressure Compressor ◽  
Engine Order ◽  
Stator Vane ◽  
Vibration Responses ◽  
The Impact

ABSTRACTThis paper aims at contributing to a better understanding of the effect of Tyler–Sofrin Modes (TSMs) on forced vibration responses by analysing a 4.5-stage research axial compressor rig. The first part starts with a brief review of the involved physical mechanisms and necessary prerequisites for the generation of TSMs in multistage engines. This review is supported by unsteady CFD simulations of a quasi 2D section of the studied engine. It is shown that the amplitude increasing effect due to mistuning can be further amplified by the presence of TSMs. Furthermore, the sensitivity with respect to the structural coupling of the blades and the damping as well as the shape of the expected envelope is analysed.The second part deals with the Rotor 2 blisk of the research compressor rig. The resonance of a higher blade mode with the engine order of the upstream stator is studied in two different flow conditions realised by different variable stator vane (VSV) schedules which allows to separate the influence of TSMs from the impact of mistuning. A subset of nominal system modes representation of the rotor is used to describe its mistuned vibration behaviour, and unsteady CFD simulations are used to characterise the present strength of the TSMs in the particular operating conditions. Measured maximum amplitude vs blade pattern and frequency response functions are compared against the predictions of the aeromechanical models in order to assess the strength of the TSMs as well as its influence on vibration levels.

Design and Analysis of an Intentional Mistuning Experiment Reducing Flutter Susceptibility and Minimizing Forced Response of a Jet Engine Fan

2017 ◽  
Author(s):  
Felix Figaschewsky ◽  
Arnold Kühhorn ◽  
Bernd Beirow ◽  
Jens Nipkau ◽  
Thomas Giersch ◽  
...  
Keyword(s):  
Computational Model ◽  
Paint Layer ◽  
Forced Vibrations ◽  
Forced Response ◽  
Operating Conditions ◽  
Jet Engines ◽  
Jet Engine ◽  
Aspect Ratios ◽  
Engine Order ◽  
Stall Flutter

Recent demands for a reduction of specific fuel consumption of jet engines have been opposed by increasing propulsive efficiency with higher bypass ratios and increased engine sizes. At the same time the challenge for the engine development is to design safe and efficient fan blades of high aspect ratios. Since the fan is the very first rotor stage, it experiences significant distortions in the incoming flow depending on the operating conditions. Flow distortions do not only lead to a performance and stall margin loss but also to remarkable low engine order (LEO) excitation responsible for forced vibrations of fundamental modes. Additionally, fans of jet engines typically suffer from stall flutter, which can be additionally amplified by reflections of acoustic pressure waves at the intake. Stall flutter appears before approaching the stall line on the fan’s characteristic and limits its stable operating range. Despite the fact that this “flutter bite” usually affects only a very narrow speed range, it reduces the overall margin of safe operation significantly. With increasing aspect ratios of ultra-high bypass ratio jet engines the flutter susceptibility will probably increase further and emphasizes the importance of considering aeromechanical analyses early in the design phase of future fans. This paper aims at proving that intentional mistuning is able to remove the flutter bite of modern jet engine fans without raising issues due to heavily increased forced vibrations induced by LEO excitation. Whereas intentional mistuning is an established technology in mitigating flutter, it is also known to amplify the forced response. However, recent investigations considering aeroelastic coupling revealed that under specific circumstances mistuning can also reduce the forced response due to engine order excitation. In order to allow a direct comparison and to limit costs as well as effort at the same time, the intentional mistuning is introduced in a non-destructive way by applying heavy paint to the blades. Its impact on the blade’s natural frequencies is estimated via finite element models with an additional paint layer. In parallel, this procedure is experimentally verified with painted fan blades in the laboratory. A validated SNM (subset of nominal system modes) representation of the fan is used as a computational model to characterize its mistuned vibration behavior. Its validation is done by comparing mistuned mode shape envelopes and frequencies of an experimental modal analysis at rest with those obtained by the updated computational model. In order to find a mistuning pattern minimizing the forced response of mode 1 and 2 at the same time and satisfying stability and imbalance constraints, a multi-objective optimization has been carried out. Finally, the beneficial properties of the optimized mistuning pattern are verified in a rig test of the painted rotor.

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.

Effect of Mistuning and Damping on the Forced Response of a Compressor Blisk Rotor

2015 ◽  
Author(s):  
Bernd Beirow ◽  
Arnold Kühhorn ◽  
Felix Figaschewsky ◽  
Jens Nipkau
Keyword(s):  
Forced Response ◽  
Influence Coefficient ◽  
Response Level ◽  
High Pressure Compressor ◽  
Engine Order ◽  
Technical Realization ◽  
Bending Mode ◽  
Random Mistuning ◽  
Longitudinal Bending ◽  
System Mode

The forced response of an E3E-type high pressure compressor blisk front rotor is analyzed with regard to intentional mistuning and its robustness towards additional random mistuning. Both a chosen alternating mistuning pattern and artificial mistuning patterns optimized concerning the forced response are considered. Focusing on three different blade modes, subset of nominal system mode-based reduced order models are employed to compute the forced response. The disk remains unchanged while the Young’s modulus of each blade is used to define the particular mistuning pattern. The well established aerodynamic influence coefficient technique is employed to model aeroelastic coupling and hence to consider the strongly mode- and inter blade phase angle-dependent aerodynamic damping contribution. It has been found that a reduction of the maximum forced response beyond that of the tuned reference can be achieved for particular mistuning patterns and all modes considered. This implies an exciting engine order which would cause a low nodal diameter mode in case of a tuned blisk. At best a nearly 50% reduction of maximum response magnitudes is computed for the fundamental bending mode and large mistuning. The solution proved to be robust towards additional random mistuning of reasonable magnitude, which is of particular interest with regard to a potential technical realization. In case of small mistuning as assumed for the first torsion and the longitudinal bending mode the advantage of achieving response magnitudes beyond the tuned reference gets lost indeed, if random mistuning is superimposed. However, mostly a lower response level is calculated compared to responses obtained from models adjusted to mistuning determined by experiment.

Development of a Novel Axial Compressor Generation for Industrial Applications: Part 2 — Blade Mechanics

2014 ◽  
Author(s):  
Dirk Anding ◽  
Henning Ressing ◽  
Klaus Hörmeyer ◽  
Roland Pisch ◽  
Kai Ziegler
Keyword(s):  
High Efficiency ◽  
Axial Compressor ◽  
Forced Response ◽  
Operating Conditions ◽  
Strain Gauges ◽  
Blade Design ◽  
Operating Range ◽  
Resonance Frequencies ◽  
Wide Range ◽  
Blade Vibrations

Blade vibrations resulting in alternating stresses are often the critical factor in determining blade life. Indeed, many of the failures experienced by turbomachinery blades occur due to high-cycle fatigue caused by blade vibrations. These vibrations can arise either through self-excited oscillations known as flutter or through aerodynamic forcing of the blades from factors such as periodic wakes from up and/or downstream vanes or unsteady flow phenomena such as compressor surge. The current paper deals with the design and the analytical and experimental verification of the axial blading for a new generation of industrial compressors, a hybrid axial compressor that combines the advantages of conventional industrial compressors — broad operating range and high efficiency — with the advantages of gas turbine compressors — high power-density and high stage pressure ratios. Additionally, the surge robustness of this novel compressor blading has been greatly improved. During the development phase extensive efforts were made to ensure safe operation for future service life. This was achieved by designing blades that will not flutter, do not have high resonance amplitudes throughout their entire operating range and are extremely robust against surge. This strongly increased robustness of the new compressor blading was achieved by the implementation of a “wide-chord” blade design in all rotor blade rows in combination with a proper tuning of resonance frequencies throughout the entire operating range. For the verification of the new blading well-established methods accepted by industry were used such as CFD and FEA. Furthermore, coupling of the two into a method referred to as Fluid Structure Interaction (FSI) was used to more closely investigate the interaction of flow and structural dynamics phenomena. These analytical techniques have been used in conjunction with extensive testing of a scaled test compressor, which was operated at conditions of dynamic similitude (matching of scaled blade vibration frequencies, flow conditions, and Mach number) with full-scale operational conditions. Strain gauges placed on the blades and a state of the art technique known as “tip timing” were used to verify blade vibrations over a wide range of combinations of guide vane positions and rotational speeds. No propensity was found of any of the blades to develop high vibration amplitudes at any of the operating conditions investigated in the rig tests. The comparison of non-linear forced response analyses and the rig test results from strain gauges and tip timing showed close agreement, verifying the analysis techniques used. In conclusion it can be stated that the blade design exhibits a very high level of safety against vibrations within the entire operating range and during surge.

scholarly journals Effects of Stator Indexing on Performance in a Low Speed Multistage Axial Compressor

1997 ◽  
Author(s):  
Wendy S. Barankiewicz ◽  
Michael D. Hathaway
Keyword(s):  
Peak Pressure ◽  
Axial Compressor ◽  
Loss Coefficient ◽  
Operating Conditions ◽  
Nasa Lewis Research ◽  
Performance Measurements ◽  
Peak Efficiency ◽  
Blade Row ◽  
The Impact ◽  
Multistage Axial Compressor

The results of an experimental investigation to determine the impact of stator row indexing or clocking on multistage axial compressor performance are presented. Testing was conducted in the NASA Lewis Research Center’s Four-Stage Axial Compressor Facility. The impact of stator row indexing on both the overall and stator 3 blade element performance is presented for both the peak efficiency and peak pressure operating conditions. The change in overall performance due to stator indexing is 0.2% for both operating conditions. Indexing resulted in a 5% change in stator 3 mass averaged loss coefficient at the peak efficiency condition and a 10% change at the peak pressure condition. Since the mass-averaged stator 3 loss coefficient is on the order of 7%, the changes in loss coefficient due to indexing are on the order of 0.35–0.7%. These changes are considered to be small and are of the same order of magnitude as the passage-to-passage differences in loss coefficient due to manufacturing and assembly tolerances in the test compressor. The effects of stator-stator wake interactions are also shown and indicate that for rows with unequal blade counts it may be necessary to survey across more than one blade row pitch for accurate blade row performance measurements.

scholarly journals An Experimental Investigation of Vibration Localization in Bladed Disks: Part II — Forced Response

1997 ◽  
Author(s):  
Marlin J. Kruse ◽  
Christophe Pierre
Keyword(s):  
Experimental Investigation ◽  
Forced Response ◽  
Mode Shapes ◽  
Bladed Disks ◽  
Structural Coupling ◽  
Vibration Localization ◽  
Bladed Disk ◽  
Engine Order ◽  
Localized Mode ◽  
The Impact

The results of an experimental investigation on the effects of random blade mistuning on the forced dynamic response of bladed disks are reported. Two experimental specimens are considered: a nominally periodic twelve-bladed disk with equal blade lengths, and the corresponding mistuned bladed disk, which features slightly different blades of random lengths. Both specimens are subject to traveling-wave excitations delivered by piezo-electric actuators. The primary aim of the experiment is to demonstrate the occurrence of an increase in forced response blade amplitudes due to mistuning, and to verify analytical predictions about the magnitude of these increases. In particular, the impact of localized mode shapes, engine order excitation, and disk structural coupling on the sensitivity of forced response amplitudes to blade mistuning is reported. This work reports one of the first systematic experiments carried out to demonstrate and quantify the effect of mistuning on the forced response of bladed disks.

Impact of Turbine Center Frame Wakes on Downstream Rows in Heavy-Duty Low-Pressure Turbine

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
Sara Biagiotti ◽  
Juri Bellucci ◽  
Michele Marconcini ◽  
Andrea Arnone ◽  
Gino Baldi ◽  
...  
Keyword(s):  
Experimental Data ◽  
Low Pressure ◽  
Numerical Approach ◽  
Forced Response ◽  
Operating Conditions ◽  
Response Analysis ◽  
Test Case ◽  
Low Pressure Turbine ◽  
Harmonic Content ◽  
The Impact

Abstract In this work, the effects of turbine center frame (TCF) wakes on the aeromechanical behavior of the downstream low-pressure turbine (LPT) blades are numerically investigated and compared with the experimental data. A small industrial gas turbine has been selected as a test case, composed of a TCF followed by the two low-pressure stages and a turbine rear frame (TRF) before the exhaust plenum. Full annulus unsteady computations of the whole low-pressure module have been performed. Two operating conditions, full (100%) and partial (50%) load, have been investigated with the aim of highlighting the impact of TCF wakes convection and diffusion through the downstream rows. Attention was paid to the harmonic content of rotors’ blades. The results show a slower decay of the wakes through the downstream rows in off-design conditions compared with the design point. The analysis of the rotors’ frequency spectrum reveals that moving from design to off-design conditions, the effect of the TCF does not change significantly. The harmonic contribution of all turbine components has been extracted, highlighting the effect of statoric parts on the last LPT blade. The TCF harmonic content remains the most relevant from an aeromechanic point of view as per experimental evidence, and it is considered for an forced response analysis (FRA) on the last LPT blade itself. Finally, aerodynamic and aeromechanic predictions have been compared with the experimental data to validate the numerical approach. Some general design solutions aimed at mitigating the TCF wakes impact are discussed.

The Mechanism on Prediction of Transient Maximum Amplitude for Tuned and Mistuned Blisks

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Luohui Ouyang ◽  
Hai Shang ◽  
Hua Chen ◽  
Qingzhen Bi ◽  
Li-Min Zhu
Keyword(s):  
Damping Ratio ◽  
Maximum Amplitude ◽  
Element Model ◽  
Forced Response ◽  
Lumped Parameter Model ◽  
Fundamental Parameter ◽  
Reduced Order ◽  
Lumped Parameter ◽  
Engine Order ◽  
Acceleration And Deceleration

Abstract Blisks are subjected to frequent acceleration and deceleration, which leads to a transient forced response; however, there is limited understanding of this response. In this work, the mechanism on prediction of transient maximum amplitude is found. An analytical link is proposed between the transient maximum amplitude and a fundamental dimensionless parameter which combines the damping ratio, natural frequency, acceleration, and engine order of the system to reveal the mechanism of the transient maximum amplitude. Therefore, the transient maximum amplitudes of tuned and mistuned blisks are predicted analytically. First, a lumped parameter model is used to study the mechanism of the transient maximum amplitude for a tuned blisk, and an approximated analytical expression is derived between the fundamental parameter and the transient amplification factor of a 1DOF (degree-of-freedom) model. The relationship is also applicable to a reduced order, tuned finite element model (FEM). Second, the mechanism of the transient response for a mistuned blisk is studied in the decoupled modal space of the blisk, based on the 1DOF transient relationship. The transient maximum amplitude in a reduced order, mistuned FEM is predicted. Two lumped parameter models and a FEM are employed to validate the prediction.

Impact of Turbine Center Frame Wakes on Downstream Rows in Heavy Duty Low Pressure Turbine

2019 ◽  
Author(s):  
Sara Biagiotti ◽  
Juri Bellucci ◽  
Michele Marconcini ◽  
Andrea Arnone ◽  
Gino Baldi ◽  
...  
Keyword(s):  
Experimental Data ◽  
Low Pressure ◽  
Forced Response ◽  
Point Of View ◽  
Operating Conditions ◽  
Response Analysis ◽  
Low Pressure Turbine ◽  
Harmonic Content ◽  
Partial Load ◽  
The Impact

Abstract In this work, the effects of Turbine Center Frame (TCF) wakes on the aeromechanical behavior of the downstream Low Pressure Turbine (LPT) blades are numerically investigated and compared with experimental data. A small industrial gas turbine has been selected as a test case, composed of a TCF followed by the two low pressure stages and a Turbine Rear Frame (TRF) before the exhaust plenum. Full annulus unsteady computations of the whole low-pressure module have been performed. Two operating conditions, full (100%) and partial (50%) load, have been investigated with the aim of highlighting the impact of TCF wakes convection and diffusion through the downstream rows. Attention was paid to the harmonic content of rotors’ blades. From an aerodynamic point of view, the results show a slower decay of the wakes through the downstream rows in off-design conditions as compared to the design point. The wakes generated by the struts at partial load persist throughout the domain outlet, while they are chopped and circumferentially transported by the rotors motion. This is due to the strong incidence variation at which the TCF works, which induces the growth of wide regions of separated flow on the rear part of the struts. Nevertheless, the analysis of the rotors’ frequency spectrum reveals that moving from design to off-design conditions, the effect of the TCF does not change significantly, thanks to the filtering action of the first LPT stage movable Nozzle Guide Vane (NGV). From unsteady calculations the harmonic contribution of all turbine components has been extracted, highlighting the effect of statoric parts on the last LPT blade. Anyhow the TCF harmonic content remains the most relevant from an aeromechanic point of view as per experimental evidence, and it is considered for a Forced Response Analysis (FRA) on the last LPT blade itself. Finally, aerodynamic and aeromechanic predictions have been compared with the experimental data to validate the numerical approach. In the last part of this paper some general design solutions, that can help mitigation of the TCF wakes impact, are discussed.

scholarly journals A Navier-Stokes Analysis of the Effect of Tip Clearance on Compressor Stall Margin

1995 ◽  
Author(s):  
Robert P. Dring ◽  
William D. Sprout ◽  
Harris D. Weingold
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Stall Margin ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Practical Tool ◽  
The Impact ◽  
Pressure Compressor

A three-dimensional Navier-Stokes calculation was used to analyze the impact of rotor tip clearance on the stall margin of a multi-stage axial compressor. This paper presents a summary of: (1) a study of the sensitivity of the results to grid refinement, (2) an assessment of the calculation’s ability to predict stall margin when the stalling row was the first rotor in a multi-stage rig environment, (3) an analysis of the impact of including the effects of the downstream stator through body force effects on the upstream rotor, and (4) the ability of the calculation to predict the impact of tip clearance on stall margin through a calculation of the rear seven airfoil rows of an eleven stage high pressure compressor rig. The result of these studies was that a practical tool is available which can predict stall margin, and the impact of tip clearance, with reasonable accuracy.

Sign in / Sign up

Export Citation Format

Share Document