Experimental Reduction of Transonic Fan Forced Response by IGV Flow Control

2004 ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Forced Response ◽  
Resonant Response ◽  
Compressor Blades ◽  
Resonant Amplitude ◽  
Engine Order ◽  
Blade Vibrations ◽  
Fan Blade ◽  
Transonic Fan ◽  
Bending Response

The main contributor to the high-cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds can be especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced fan blade vibrations was evaluated in a modern single-stage transonic fan rig. Data was collected for multiple uniform full-span TEB conditions over a range of rotor speed including multiple modal resonance crossings. Resonant response sensitivity was generally characterized by a robust region of strong attenuation. The baseline resonant amplitude of the first torsion mode, which exceeded the endurance limit on the critical blade, was reduced by more than 80% with TEB at 1.0% of the total rig flow. The technique was also found to be modally robust; similar reductions were achieved for all tested modal crossings, including more than 90% reduction of the second LE bending response using 0.7% of the rig flow.

Experimental Reduction of Transonic Fan Forced Response by Inlet Guide Vane Flow Control

2009 ◽  
Vol 132 (2) ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Guide Vane ◽  
Leading Edge ◽  
Forced Response ◽  
Inlet Guide Vane ◽  
Resonant Response ◽  
Compressor Blades ◽  
Resonant Amplitude ◽  
Blade Vibrations ◽  
Transonic Fan

The main contributor to the high cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds can be especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced fan blade vibrations was evaluated in a modern single-stage transonic fan rig. Data were collected for multiple uniform full-span TEB conditions over a range of rotor speeds including multiple modal resonance crossings. Resonant response sensitivity was generally characterized by a robust region of strong attenuation. The baseline resonant amplitude of the first torsion mode, which exceeded the endurance limit on the critical blade, was reduced by more than 80% with TEB at 1.0% of the total rig flow. The technique was also found to be modally robust; similar reductions were achieved for all tested modal crossings, including more than 90% reduction in the second leading-edge bending response using 0.7% of the rig flow.

Effects of Flow Control on Forced Response and Performance of a Transonic Compressor

2002 ◽  
Author(s):  
S. Todd Bailie ◽  
Wing F. Ng ◽  
Alfred L. Wicks ◽  
William W. Copenhaver
Keyword(s):  
Flow Control ◽  
Rotor Blade ◽  
Forced Response ◽  
Strain Gages ◽  
Complex Flow ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Engine Order ◽  
Blade Vibrations ◽  
And Performance

The main contributor to the high-cycle fatigue of compressor blades is the response to aerodynamic forcing functions generated by an upstream row of stators or inlet guide vanes. Resonant response to engine order excitation at certain rotor speeds is especially damaging. Studies have shown that flow control by trailing edge blowing (TEB) can reduce stator wake strength and the amplitude of the downstream rotor blade vibrations generated by the unsteady stator-rotor interaction. In the present study, the effectiveness of TEB to reduce forced blade vibrations was evaluated in a modern transonic compressor rig. A row of wake generator (WG) vanes with TEB capability was installed upstream of the rotor, which was instrumented with strain gages. Data was collected with and without TEB at various rotor speeds involving resonance crossings. Using 0.8% of the compressor core flow for TEB along the full WG-span, rotor blade strain was reduced by 66% at the first torsional resonance crossing. Substantial reductions were also achieved with only partial span TEB. The results demonstrate the effectiveness of the TEB technique for reducing rotor vibrations in the complex flow environment of a closely-spaced transonic stage row. Moderate increases in stage performance were also measured.

scholarly journals Using Shock Control Bumps to Improve Transonic Fan/Compressor Blade Performance

2019 ◽  
Vol 141 (8) ◽  
Author(s):  
Alistair John ◽  
Ning Qin ◽  
Shahrokh Shahpar
Keyword(s):  
Compressor Blade ◽  
3D Analysis ◽  
Aerodynamic Optimization ◽  
Compressor Blades ◽  
Stall Inception ◽  
Stall Margin ◽  
Shock Control ◽  
Fan Blade ◽  
Transonic Fan ◽  
Critical Regions

Shock control bumps can help to delay and weaken shocks, reducing loss generation and shock-induced separation and delaying stall inception for transonic turbomachinery components. The use of shock control bumps on turbomachinery blades is investigated here for the first time using 3D analysis. The aerodynamic optimization of a modern research fan blade and a highly loaded compressor blade is carried out using shock control bumps to improve their performance. Both the efficiency and stall margin of transonic fan and compressor blades may be increased through the addition of shock control bumps to the geometry. It is shown how shock-induced separation can be delayed and reduced for both cases. A significant efficiency improvement is shown for the compressor blade across its characteristic, and the stall margin of the fan blade is increased by designing bumps that reduce shock-induced separation near to stall. Adjoint surface sensitivities are used to highlight the critical regions of the blade geometries, and it is shown how adding bumps in these regions improves blade performance. Finally, the performance of the optimized geometries at conditions away from where they are designed is analyzed in detail.

scholarly journals Using Shock Control Bumps to Improve Transonic Fan/Compressor Blade Performance

2018 ◽  
Author(s):  
Alistair John ◽  
Ning Qin ◽  
Shahrokh Shahpar
Keyword(s):  
Compressor Blade ◽  
3D Analysis ◽  
Compressor Blades ◽  
Stall Inception ◽  
Stall Margin ◽  
Shock Control ◽  
Turbomachinery Blades ◽  
Fan Blade ◽  
Transonic Fan ◽  
Critical Regions

Shock control bumps can help to delay and weaken shocks, reducing loss generation and shock-induced separation and delaying stall inception for transonic turbomachinery components. The use of shock control bumps on turbomachinery blades is investigated here for the first time using 3D analysis. The aerodynamic optimisation of a modem research fan blade and a highly loaded compressor blade are carried out using shock control bumps to improve their performance. Both the efficiency and stall margin of transonic fan and compressor blades may be increased through the addition of shock control bumps to the geometry. It is shown how shock induced separation can be delayed and reduced for both cases. A significant efficiency improvement is shown for the compressor blade across its characteristic, and the stall margin of the fan blade is increased by designing bumps that reduce shock-induced separation near to stall. Adjoint surface sensitivities are used to highlight the critical regions of the blade geometries, and it is shown how adding bumps in these regions improves blade performance. Finally, the performance of the optimised geometries at conditions away from where they are designed is analysed in detail.

Low Pressure Compression System Effects on Fan Assembly Forced Response

2007 ◽  
Author(s):  
Abdulnaser I. Sayma ◽  
Mehdi Vahdti ◽  
Mehmet Imregun ◽  
John Marshal
Keyword(s):  
Large Diameter ◽  
Pressure Fluctuations ◽  
Element Model ◽  
Low Pressure ◽  
Forced Response ◽  
Blade Vibration ◽  
Compression System ◽  
Engine Order ◽  
Modelling Methodology ◽  
Fan Blade

This paper describes a numerical modelling methodology for fan blade forced response calculations by considering the low-pressure compression system (LPCS) as a whole in order to include flow distortions caused by the asymmetric flight intake upstream, and the pylon downstream. Emphasis also is placed on blade mistuning or mis-placement which may be due to inherent manufacturing and assembly tolerances, or to small inservice displacements. Several levels of geometric complexity were used in the analysis, ranging from an isolated fan bladerow to a complete LPCS of a large-diameter aero-engine, consisting of the intake duct, the fan assembly, the outflow guide vanes, the pylon and a downstream nozzle. The aerodynamic model was coupled to a finite element model of the fan assembly for computing the blade vibration levels. The study revealed two major findings. The first is the unsteady forcing under one engine-order (1EO) excitation is found to be linked to the mean shock position on the fan blade, the highest forcing occurring when the shock is just swallowed since this position is particularly sensitive to pressure fluctuations. The second finding is that the 1EO fan assembly forcing resulting from an asymmetric intake and the pylon are of comparable magnitude but their relative phasing is the key parameter in determining the overall fan forced response levels.

Piezoelectric Networks for Vibration Suppression of Mistuned Bladed Disks

2007 ◽  
Vol 129 (5) ◽  
pp. 559-566 ◽  
Author(s):  
Hongbiao Yu ◽  
K. W. Wang
Keyword(s):  
Vibration Suppression ◽  
Periodic Structures ◽  
Electric Circuit ◽  
Forced Response ◽  
Bladed Disks ◽  
Vibration Localization ◽  
Network Concept ◽  
Engine Order ◽  
The Individual ◽  
Local Circuits

Extensive investigations have been conducted to study the vibration localization phenomenon and the excessive forced response that can be caused by mistuning in bladed disks. Most previous researches have focused on analyzing∕predicting localization or attacking the mistuning issue via mechanical tailoring. Few have focused on developing effective vibration control methods for such systems. This study extends the piezoelectric network concept, which has been utilized for mode delocalization in periodic structures, to the control of mistuned bladed disks under engine order excitation. A piezoelectric network is synthesized and optimized to effectively suppress vibration in bladed disks. One of the merits of such an approach is that the optimum design is independent of the number of spatial harmonics, or engine orders. Local circuits are first formulated by connecting inductors and resistors with piezoelectric patches on the individual blades. Although these local circuits can function as conventional damped absorber when properly tuned, they do not perform well for bladed disks under all engine order excitations. To address this issue, capacitors are introduced to couple the individual local circuitries. Through such networking, an absorber system that is independent of the engine order can be achieved. Monte Carlo simulation is performed to investigate the effectiveness of the network for a bladed disk with a range of mistuning level of its mechanical properties. The robustness issue of the network in terms of detuning of the electric circuit parameters is also studied. Finally, negative capacitance is introduced and its effect on the performance and robustness of the network is investigated.

Transonic Fan Blade Redesign Approach to Attenuate Nonsynchronous Vibration

2021 ◽  
Author(s):  
Yaozhi Lu ◽  
Bharat Lad ◽  
Mehdi Vahdati
Keyword(s):  
Fan Blade ◽  
Transonic Fan

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.

Modal Analyses of an Axial Turbine Blisk With Intentional Mistuning

2017 ◽  
Author(s):  
Bernd Beirow ◽  
Felix Figaschewsky ◽  
Arnold Kühhorn ◽  
Alfons Bornhorn
Keyword(s):  
Aerodynamic Performance ◽  
Forced Response ◽  
Mode Shapes ◽  
Axial Turbine ◽  
The Real ◽  
Operational Conditions ◽  
Aerodynamic Damping ◽  
Optimization Study ◽  
Engine Order ◽  
Turbine Blisk

The potential of intentional mistuning to reduce the maximum forced response is analyzed within the development of an axial turbine blisk for ship diesel engine turbocharger applications. The basic idea of the approach is to provide an increased aerodynamic damping level for particular engine order excitations and mode shapes without any significant distortions of the aerodynamic performance. The mistuning pattern intended to yield a mitigation of the forced response is derived from an optimization study applying genetic algorithms. Two blisk prototypes have been manufactured a first one with and another one without employing intentional mistuning. Hence, the differences regarding the real mistuning and other modal properties can be experimentally determined and evaluated as well. In addition, the experimental data basis allows for updating structural models which are well suited to compute the forced response under operational conditions. In this way, the real benefit achieved with the application of intentional mistuning is demonstrated.

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