J. S. Green

Friction Damper Optimization: Simulation of Rainbow Tests

Friction dampers have been used to reduce turbine blade vibration levels for a considerable period of time. However, optimal design of these dampers has been quite difficult due both to a lack of adequate theoretical predictions and to difficulties in conducting reliable experiments. One of the difficulties of damper weight optimization via the experimental route has been the inevitable effects of mistuning. Also, conducting separate experiments for different damper weights involves excessive cost. Therefore, current practice in the turbomachinery industry has been to conduct so-called rainbow tests where friction dampers with different weights are placed between blades with a predefined configuration. However, it has been observed that some rainbow test results have been difficult to interpret and have been inconclusive for determining the optimum damper weight for a given bladed-disk assembly. A new method of analysis-a combination of the harmonic balance method and structural modification approaches-is presented in this paper for the analysis of structures with friction interfaces and the method is applied to search for qualitative answers about the so-called rainbow tests in turbomachinery applications. A simple lumped-parameter model of a bladed-disk model was used and different damper weights were modeled using friction elements with different characteristics. Resonance response levels were obtained for bladed disks with various numbers of blades under various engine-order excitations. It was found that rainbow tests, where friction dampers with different weights are used on the same bladed-disk assembly, can be used to find the optimum damper weight if the mode of vibration concerned has weak blade-to-blade coupling (the case where the disk is almost rigid and blades vibrate almost independently from each other). Otherwise, it is very difficult to draw any reliable conclusion from such expensive experiments.

Whole-Assembly Flutter Analysis of a Low Pressure Turbine Blade

This paper reports the findings of a flutter investigation on a low pressure turbine blade using a 3D non-linear integrated aeroelasticity method. The approach has two important features: (i) the calculations are performed in a time-accurate and integrated fashion whereby the structural and fluid domains are linked via an exchange of boundary conditions at each time step, and (ii) the analysis is performed on the entire bladed-disk assembly, thus removing the need to assume a critical interblade phase angle. Although such calculations are both CPU and in-core memory intensive, they do not require pre-knowledge of the flutter mode and hence they allow a better understanding of the aeroelasticity phenomena involved.

Forced Response of a Large Civil Fan Assembly

Forced response analysis has become commonplace for predicting the vibration amplitude of turbomachinery blading. These analyses are usually limited because they rely on predicting a well defined source of flow distortion, such as blade wakes and shocks etc. However, the sources of excitation of civil fans are not well defined and yet are able to produce high levels of force. The objective of the work described in this paper is to investigate the forced response of a large civil fan assembly using CFD. An unsteady, time accurate, 3D CFD model of the complete low pressure compression system has been used to calculate the modal response of a large civil fan. The mesh consists of the ground plane, intake, fan, OGV, bypass duct and compressor inlet stator, with every aerofoil passage modelled. The analysis tool allows calculation of a time history of modal response for a range of modes simultaneously to provide a description of the overall vibration behaviour. The results of the analyses have been used to investigate the modal contributions to the off-resonant first engine order response at a range of operating conditions to assess the contribution of various geometric features. The response is shown to compare well with measured strain gauge data for both ground and altitude conditions. The response of the majority of resonances was found to be heavily influenced by the presence of the ground plane, which is consistent with the available experimental data.Copyright

A non-linear integrated aeroelasticity method for the prediction of turbine forced response with friction dampers

An integrated aeroelasticity model was described for turbine blade forced response predictions. An iterative procedure was developed to determine the resonance shift under the effects of both unste ...

A Numerical Study of Labyrinth Seal Flutter

A numerical study of a labyrinth-type turbine seal flutter in a large turbofan engine is described. The flutter analysis was conducted using a coupled fluid-structure interaction code, which was originally developed for turbomachinery blade applications. The flow model is based on an unstructured, implicit Reynolds-averaged Navier-Stokes solver. The solver is coupled to a modal model for the structure obtained from a standard structural finite element code. During the aeroelasticity computations, the aerodynamic grid is moved at each time step to follow the structural motion, which is due to unsteady aerodynamic forces applied onto the structure by the fluid. Such an integrated time-domain approach allows the direct computation of aeroelastic time histories from which the aerodynamic damping, and hence, the flutter stability, can be determined. Two different configurations of a large-diameter aeroengine labyrinth seal were studied. The first configuration is the initial design with four fins, which exhibited flutter instability during testing. The second configuration is a modified design with three fins and a stiffened ring. The steady-state flow was computed for both configurations, and good agreement was reached with available reference data. An aeroelasticity analysis was conducted next for both configurations, and the model was able to predict the observed flutter behavior in both cases. A flutter mechanism is proposed, based on the matching of the structural frequencies to the frequencies of waves traveling in the fluid, in the interfin cavities and in the high- and low-pressure cavities.

Numerical and Experimental Investigation of an Underplatform Damper Test Rig

During operation mechanical structures can experience large vibration amplitudes. One of the challenges encountered in gas-turbine blade design is avoiding high-cycle fatigue failure usually caused by large resonance stresses driven by aeroelastic excitation. A common approach to control the amplitude levels relies on increasing friction damping by incorporating underplatform dampers (UPD). An accurate prediction of the dynamics of a blade-damper system is quite challenging, due to the highly nonlinear nature of the friction interfaces and detailed validation is required to ensure that a good modelling approach is selected. To support the validation process, a newly developed experimental damper rig will be presented, based on a set of newly introduced non-dimensional parameters that ensure a similar dynamic behaviour of the test rig to a real turbine blade-damper system. An ini- tial experimental investigation highlighted the sensitivity of the measured response with regards to settling and running in of the damper, and further measurements identified a strong dependence of the nonlinear behaviour to localised damper motion. Numerical simulations of the damper rig with a simple macroslip damper model were performed during the preliminary design, and a comparison to the measured data highlighted the ability of the basic implicit model to capture the resonance frequencies of the system accuratelyю

Low-Engine-Order Excitation Mechanisms in Axial-Flow Turbomachinery

One particular regime of forced response, occurring at much lower frequencies than the blade-passing frequency, is the so-called low-engine-order (LEO) forced response. The source of LEO excitation is a loss of symmetry in the flow, such as that caused by stator blade throat width variations, flow exit angle variations, perturbations in the passage cooling flow,or temperature distortions. Using a 3D integrated time-domain aeroelasticity code, parametric forced response studies were conducted for a high-pressure turbine stage with 36 stator and 90 rotor blades. Both whole-annulus and sector models were used to investigate the effects of individual and combined LEO parameters. For individual parameters, the amplitude of the excitation was, proportional to the imposed variation. For combined cases, the total excitation was found to be determined by the phasing between the individual excitations. A ballpark comparison suggested that LEO and blade-passing forced response vibration amplitudes were similar for typical variations of the controlling LEO parameters.

A Resonance Tracking Algorithm for the Prediction of Turbine Forced Response With Friction Dampers

This paper presents an iterative method for determining the resonant speed shift when non-linear friction dampers are included in turbine blade roots. Such a need arises when conducting response calculations for turbine blades where the unsteady aerodynamic excitation must be computed at the exact resonant speed of interest. The inclusion of friction dampers is known to raise the resonant frequencies by up to 20% from the standard assembly frequencies. The iterative procedure uses a viscous, time-accurate flow representation for determining the aerodynamic forcing, a look-up table for evaluating the aerodynamic boundary conditions at any speed, and a time-domain friction damping module for resonance tracking. The methodology was applied to an HP turbine rotor test case where the resonances of interest were due to the 1T and 2F blade modes under 40 engine-order excitation. The forced response computations were conducted using a multi-stage approach in order to avoid errors associated with “linking” single stage computations since the spacing between the two bladerows was relatively small. Three friction damper elements were used for each rotor blade. To improve the computational efficiency, the number of rotor blades was decreased by 2 to 90 in order to obtain a stator/rotor blade ratio of 4/9. However, the blade geometry was skewed in order to match the capacity (mass flow rate) of the components and the condition being analysed. Frequency shifts of 3.2% and 20.0% were predicted for the 1T/40EO and 2F/40EO resonances in about 3 iterations. The predicted frequency shifts and the dynamic behaviour of the friction dampers were found to be within the expected range. Furthermore, the measured and predicted blade vibration amplitudes showed a good agreement, indicating that the methodology can be applied to industrial problems.© 2000 ASME

Lip Stall Suppression in Powered Intakes

This work describes a computational study into lip stall in subsonic civil aircraft intakes and its alleviation by action of the fan. Beyond a certain flow incidence, the lower lip of a civil aircraft intakes stalls. This phenomenon causes entropy and vortical distortions at the fan face. Consequently, it has detrimental effects on vibration levels and performance of the low-pressure compressor system. The most important parameters influencing lip separation are the flight Mach number, the Reynolds number, and altitude. Fully three-dimensional simulations have been performed on a flight intake in current service for which the experimental data are available. Steady and time-resolved simulations have been performed. Distortion coefficients have been evaluated as functions of incidence and have been compared with experimental results. A comparison between an isolated intake and a powered intake shows that the fan stage has the beneficial effect of increasing tolerance to flow incidence and decreasing distor...

Test Method Development for Nonlinear Damping Extraction of Dovetail Joints

The traditional measurement techniques to acquire the linear dynamic response of a single component are well established and have been in use for many decades to provide reliable input data for model updating. The measurement of assembled structures normally follows a very similar approach, although the presence of joints can introduce a nonlinear dynamic behaviour, which impacts the measurement results. Applying traditional linear test methods to a highly nonlinear structure, such as a dovetail joint in an aircraft engine blade-disk connection, does not necessarily take the special features of the nonlinear response into account and may lead to unreliable data. This is particularly true, if modal information such as damping are required. The influence of the measurement setup and the test procedures must be well understood for an accurate measurement. In this paper the influence of the different measurement components on a simple clamped beam and a compressor blade dove tail test rig will be investigated. A particular focus will be on the support of the test rig, the location and control of the excitation and the influence of the accelerometer on the response. Based on the findings an approach will be recommended that allows a reliable measurement of the dynamic behaviour of this heavily nonlinear structure.