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Francois Maurice Garcin - One of the best experts on this subject based on the ideXlab platform.

  • Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions
    Journal of Engineering for Gas Turbines and Power-transactions of The Asme, 2014
    Co-Authors: Alain Batailly, Mathias Legrand, Antoine Millecamps, Sébastien Cochon, Francois Maurice Garcin
    Abstract:

    Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural Blade/casing contacts at the design stage of a high-pressure Compressor Blade. The Blade of interest underwent structural divergence after Blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the Blade stacking law while maintaining aerodynamic performance of the original Blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final Blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.Copyright © 2014 by ASME

  • Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions
    Journal of Engineering for Gas Turbines and Power, 2014
    Co-Authors: Alain Batailly, Mathias Legrand, Antoine Millecamps, Sébastien Cochon, Francois Maurice Garcin
    Abstract:

    Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural Blade/casing contacts at the design stage of a high-pressure Compressor Blade. The Blade of interest underwent structural divergence after Blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the Blade stacking law while maintaining aerodynamic performance of the original Blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final Blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

Alain Batailly - One of the best experts on this subject based on the ideXlab platform.

  • A program to compute Compressor Blade geometries from multiple-circular-arc parameters with sweep and lean (v1.0)
    2019
    Co-Authors: Elsa Piollet, Alain Batailly
    Abstract:

    This repository contains the source code used to generate Compressor Blade geometries for the following publication (currently under minor revisions): Piollet, E., Nyssen, F. and Batailly, A. "Blade/casing rubbing interactions in aircraft engines: numerical benchmark and design guidelines based on NASA rotor 37". This code computes the cross-section coordinates of a Compressor Blade composed of multiple-circular-arc elements including variable sweep and lean angles. The goal is to study the design of Blades that are robust to contact interactions, and to serve as a basis for comparative work between different research teams on Blade dynamics. All input and output files related to the seven Blades studied in the aforementioned publication are provided along with the program. This includes input parameters (multiple-circular-arc profile parameters and sweep and lean parameters) and output for all seven Blades ((x, y, z) coordinates of the suction side and pressure side for each stacked profile -- CAD (Computer-Aided Design) files in STEP format -- finite element mesh files).

  • Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions
    Journal of Engineering for Gas Turbines and Power-transactions of The Asme, 2014
    Co-Authors: Alain Batailly, Mathias Legrand, Antoine Millecamps, Sébastien Cochon, Francois Maurice Garcin
    Abstract:

    Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural Blade/casing contacts at the design stage of a high-pressure Compressor Blade. The Blade of interest underwent structural divergence after Blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the Blade stacking law while maintaining aerodynamic performance of the original Blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final Blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.Copyright © 2014 by ASME

  • Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions
    Journal of Engineering for Gas Turbines and Power, 2014
    Co-Authors: Alain Batailly, Mathias Legrand, Antoine Millecamps, Sébastien Cochon, Francois Maurice Garcin
    Abstract:

    Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural Blade/casing contacts at the design stage of a high-pressure Compressor Blade. The Blade of interest underwent structural divergence after Blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the Blade stacking law while maintaining aerodynamic performance of the original Blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final Blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

Angelos Filippatos - One of the best experts on this subject based on the ideXlab platform.

  • Design and testing of composite Compressor Blades with focus on the vibration behaviour
    Composites Part A: Applied Science and Manufacturing, 2017
    Co-Authors: Thomas Wollmann, Michael Dannemann, Stefan Nitschke, Niels Modler, Albert Langkamp, Angelos Filippatos
    Abstract:

    Increasing demands on the performance of rotating components as well as on the noise reduction in jet engines have led to higher dynamic requirements. Composite materials can contribute to a reduction of vibration amplitudes due to their inherent advantageous damping behaviour and their high specific stiffness. A numerical prediction of the vibration behaviour is necessary for an efficient use of composite materials in rotors, especially since the modal parameters such as the eigenfrequency and the modal damping are influenced by the fibre orientation. For this investigation, a composite Compressor Blade was designed with a focus on its vibration behaviour. The modal parameters were numerically predicted and experimentally validated. The damping of composites was taken into account using the strain energy method. Utilising this method, the modal damping parameters of the Compressor Blade were calculated based on direction-dependent material properties of single unidirectional carbon fibre-reinforced epoxy layers.

J. Loellbach - One of the best experts on this subject based on the ideXlab platform.

  • Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows
    Journal of Turbomachinery, 1999
    Co-Authors: Chunill Hah, J. Loellbach
    Abstract:

    A detailed investigation has been performed to study hub corner stall phenomena in Compressor Blade rows. Three-dimensional flows in a subsonic annular Compressor stator and in a transonic Compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of Compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the Blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the Blade by this vortex. On the Blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic Compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced Blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the Blade row.

  • development of hub corner stall and its influence on the performance of axial Compressor Blade rows
    Volume 1: Aircraft Engine; Marine; Turbomachinery; Microturbines and Small Turbomachinery, 1997
    Co-Authors: Chunill Hah, J. Loellbach
    Abstract:

    A detailed investigation has been performed to study hub corner stall phenomena in Compressor Blade rows. Three-dimensional flows in a subsonic annular Compressor stator and in a transonic Compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of Compressor hub corner stall. Both the measurements and the numerical solutions indicate that a strong twister-like vortex is formed near the rear part of the Blade suction surface. Low momentum fluid inside the hub boundary layer is transported toward the suction side of the Blade by this vortex. On the Blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic Compressor rotor studied in this paper, the numerical solution and the measured data indicate that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced Blade loading which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the Blade row.Copyright © 1997 by ASME

A. L. Stel’makh - One of the best experts on this subject based on the ideXlab platform.

  • Rapid method of Predicting the Subsonic Flutter Stability of AGTE Axial-Flow Compressor Blade Cascades. Part 2. Mathematical Implementation of Method and its Potentional Application
    Strength of Materials, 2019
    Co-Authors: A. P. Zinkovskii, S. N. Kabannik, A. L. Stel’makh
    Abstract:

    The paper considers the implementation of the rapid method of predicting the dynamic stability of the Compressor Blade assemblies against subsonic flutter. The first flexural mode of Blade vibrations is analyzed at the design stage for a wide range of the angle of attack based on the developed database of the critical values of the reduced vibration frequency in straight cascades of Blade airfoils. The multiple regression equation is developed that depends on the relative Blade spacing and stagger angle of the straight cascade of Blade airfoils, coefficient of the flexural-torsional coupling of the Blade, and angle of attack with the correlation factor of the studied parameters in the range of 0.92–0.98. Using the obtained equation, the numerical program has been developed for the determination of the dynamic stability limit against subsonic flutter for the first flexural mode of the Blade vibrations. The program allows one to find the numerical values of critical reduced frequencies as the characteristics of its dynamic stability and determine their dependence on the angle of attack. The results of practical application of the developed program are presented using the assessment of the dynamic stability of the flexural mode of axial Compressors in four modern aircraft gas-turbine engines. It is shown that at the design engine stage it is possible to select the reduced vibration frequencies of the Blade assembly for the specified geometry of its peripheral sections and angle of attack of the inflow upon the condition of the occurrence of subsonic flutter.

  • Rapid Method of Predicting the Subsonic Flutter Stability of AGTE Axial-Flow Compressor Blade Cascades. Part 1. Physical Backgrounds of the Method
    Strength of Materials, 2019
    Co-Authors: A. L. Stel’makh, A. P. Zinkovskii, S. N. Kabannik
    Abstract:

    Generalization of experimental investigation results for direct Compressor cascades of Blade profiles at a subsonic (continuous and separated) gas flow with bending, torsional, and bending-torsional vibrations created the basis for defining the physical mechanisms of subsonic flutter initiation in AGTE axial-flow Compressor Blade cascades. Such a combination of reduced frequencies and angles of attack is possible when the aerodynamic Blade vibration decrement equals zero. It corresponds to the critical reduced vibration frequency value below which the aeroexcitation of Blade vibrations and an increase in its level are observed, i.e., the dynamic subsonic cascade flutter loss is taking place. Known methods of evaluating these critical loss conditions are analyzed. The rapid method of predicting the dynamic subsonic flutter stability for Compressor Blade cascades is described. The scheme of critical reduced Blade vibration frequency data base generation is tabulated as the critical values at fixed geometric cascade parameters (pitch-chord ratio and deflection angle), angles of attack, and coefficients of bending-torsional coupling. An example of such a data base for specific Compressor Blade cascades is given.