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.
The use of blisks is becoming more common in modern aero-engine designs. Such structures have very low mechanical damping and hence are more susceptible to flutter instability. Therefore, accurate prediction of blade aerodynamic damping in a multi-row environment becomes vital. The aim of this work is to understand the effects of acoustic reflections on flutter stability of an embedded rotor in a multi-stage compressor. To achieve this goal, flutter analysis of an embedded rotor blade in a high pressure compressor is undertaken using a validated 3D unsteady RANS solver with mesh movement. In the first part of this work, flutter computations are performed on the rotor with and without the presence of its adjacent blade rows and the influence of reflections on aerodynamic damping of the embedded row is studied. In the second part of this work, reflective boundary conditions are used to reflect pressure waves from known locations. A wave-splitting procedure is performed to split the unsteady pressure into an outgoing wave and a reflected wave. Using this technique a relationship between the phase of the reflected wave and the susceptibility to flutter is established.
The effect of blade frequency mistuning on the forced response of HPC-blisks is studied by means of experimental and numerical investigations applying discrete mechanical low degree of freedom models. Besides the mistuning resulting from manufacturing and inhomogeneous material also strain gauge (S/G) induced mistuning is considered. Blade by blade measurements supported by numerical calculations are used to determine mistuning distributions within an iterative approach. Due to the stiffness contribution of high temperature S/G, a significant increase of blade alone frequencies can be proved. It is shown within laser scanning measurements that this S/G induced mistuning can cause strongly localized mode shapes. Since S/G signals are used to monitor also non-instrumented blade resonances in engine-tests, it is reasonable to consider the S/G contribution within model-updates. The numerical models introduced in this paper are adjusted to experimentally determined blade alone frequency distributions. Within simulations of the forced response it is shown in principle, that the S/G-instrumentation also affects the response of non-instrumented blades which is important with regard to the S/G calibration process. Additional investigations are addressed to the consequences of small variations in measured mistuning distributions on the maximum forced response, i. e. resulting from a changing ambient temperature while measurement or a limited frequency resolution. In this context, a strong dependence on the engine order excited, the damping level and thus the flow conditions could be proved. As an example all investigations presented in this paper are carried out for two stages of a research compressor.
Experimental tests were performed to investigate flutter behaviour of the transonic rotor in a high-speed multistage compressor test facility. Besides the acquisition of overall performance parameters the rig was equipped with special instrumentation such as strain gauges, tip-timing system and dynamic pressure transducers. The 4-stage compressor comprises 3 variable vanes. The instrumentation was able to measure stall and flutter and forced responses of the rotor blades. The experimental data was used to validate the in-house aeroelastic solver. During testing flutter was triggered intentionally at part speed conditions with malscheduled variable vanes. The malschedule changed the flow incidences for the relevant rotor and induced flutter. An aeroelastic simulation was set up according to the flutter conditions found during the experiment. The measured boundary conditions were used to set up the aeroelastic simulation. The calculation shows the torsion mode with a negative aerodynamic damping which confirms the test results. The only difference between test and prediction is the nodal diameter, at which the flutter occurs.
The forced response of an E3E-type HPC-blisk front rotor is analyzed with regard to varying mistuning and the consideration of the fluid-structure interaction (FSI). For that purpose, a reduced order model is used in which the disk remains unchanged and mechanical properties of the blades namely stiffness and damping are adjusted to measured as well as intentional blade frequency mistuning distributions. The aerodynamic influence coefficient technique is employed to model the aeroelastics. Depending on the blade mode, the exciting engine order and aerodynamic influences it is sought for the worst mistuning distributions with respect to the maximum blade displacement based on optimization analyses. Genetic algorithms using blade alone frequencies as design variables are applied. The validity of the Whitehead-limit is assessed in this context. In particular, the question is addressed if and how far aeroelastic effects, mainly caused by aerodynamic damping, combined with mistuning can even cause a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the inter-blade phase angle is the main driver for a possible response attenuation considering the fundamental as well as a higher blade mode. Furthermore, the differences to the blisk vibration response without a consideration of the flow and an increase of the disk’s stiffness are discussed. Closing, the influence of pure damping mistuning is analyzed again using optimization.
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