Aeroelastic Flutter Analysis of Linear Cascade Blades: STC5

2017 ◽  
Author(s):  
Kirubakaran Purushothaman ◽  
Sankar Kumar Jeyaraman ◽  
Sasikanta Parida ◽  
Kishore Prasad Deshkulkarni
Keyword(s):  
Test Data ◽  
Aerodynamic Force ◽  
Damping Ratio ◽  
Pressure Coefficient ◽  
Axial Compressor ◽  
Sensitivity Study ◽  
Standard Test ◽  
Test Configuration ◽  
Linear Cascade ◽  
Aerodynamic Damping

Study of aerodynamic flow and aeroelastic stability in vibrating blades of cascade is the main objective of this study. Standard test configuration (STC-5) was chosen for this study as it involves transonic flow regime in compressor blade cascades. CFD analysis were carried out for 11 test cases of STC-5 configuration and pressure coefficient values were compared with test data. The range of incidence angles vary from 2° to 10° and reduced frequency varies from 0.14 to 1.02. Inflow Mach number was fixed at 0.5 and Reynolds number was fixed at 1.4 × 106. Analysis of vibrating blades and comparison of test data results of axial compressor with linear cascade stator blades of fifth standard configuration at high subsonic speed is compared with CFD results. While doing this vibration of only the center blade is concerned when all the other blades in the cascade are fixed. Fluid structure interaction approach is used here to evaluate the unsteady aerodynamic force and work done for a vibrating blade in CFD domain. Energy method and work per cycle approach is adapted for aerodynamic damping prediction. A framework has been developed to estimate the work per cycle and aerodynamic damping ratio. Final sensitivity study was carried out to evaluate the influence of blade incidence and frequency on blade damping values.

Aeroelastic Flutter Investigation and Stability Enhancement of a Transonic Axial Compressor Rotor Using Casing Treatment

2017 ◽  
Author(s):  
Kirubakaran Purushothaman ◽  
Sankar Kumar Jeyaraman ◽  
Ajay Pratap ◽  
Kishore Prasad Deshkulkarni
Keyword(s):  
Flow Separation ◽  
Aerodynamic Force ◽  
Damping Ratio ◽  
Axial Compressor ◽  
Casing Treatment ◽  
Aerodynamic Damping ◽  
Compressor Rotor ◽  
Aeroelastic Flutter ◽  
Blade Tip ◽  
Transonic Axial Compressor

This study discusses in detail the aeroelastic flutter investigation of a transonic axial compressor rotor using computational methods. Fluid structure interaction approach is used in this method to evaluate the unsteady aerodynamic force and work done of a vibrating blade in CFD domain. Energy method and work per cycle approach is adapted for this flutter prediction. A framework has been developed to estimate the work per cycle and aerodynamic damping ratio. Based on the aerodynamic damping ratio, occurrence of flutter is estimated for different inter blade phase angles. Initially, the baseline rotor blade design was having negative aerodynamic damping at part speed conditions. The main cause for this flutter occurrence was identified as large flow separation near blade tip region due to high incidence angles. The unsteadiness in the flow was leading to aerodynamic force fluctuation matching with natural frequency of blade, resulting in excitation of the blades. Hence axially skewed slot casing treatment was implemented to reduce the flow separation at blade tip region to alleviate the onset of flutter. By this method, the stall margin and aerodynamic damping of the test compressor was improved and flutter was avoided.

Simplified Estimation of Aerodynamic Damping for Bladed Rotors: Part 2 — Experimental Validation During Operation

2016 ◽  
Author(s):  
Felix Figaschewsky ◽  
Arnold Kühhorn ◽  
Bernd Beirow ◽  
Thomas Giersch ◽  
Jens Nipkau ◽  
...  
Keyword(s):  
Aerodynamic Force ◽  
Axial Compressor ◽  
Industrial Applications ◽  
Mean Value ◽  
Vibration Response ◽  
Simplified Approach ◽  
Aerodynamic Damping ◽  
Reduced Frequency ◽  
Significant Difference ◽  
The Mean

Due to increasing requirements of future engine projects, much effort has been spent on the design of more efficient turbomachinery blades in the recent years. Besides aerodynamic efficiency constraints, these designs need to meet structural criteria ensuring that they are safe and robust with respect to High Cycle Fatigue (HCF). The estimation of the resonant vibration amplitude is done based on the aerodynamic force and the overall damping level. Since, for many applications the contribution of mechanical damping is often rather low compared to the aerodynamic counterpart, the determination of the aerodynamic damping is vital for the estimation of the forced vibration response. This second part is meant to contribute to a simplified computation of the aerodynamic damping during operation by making additional assumptions: The investigated mode family shall not suffer from flutter, has a high reduced frequency and the influence of adjacent blades is negligible. Under these circumstances a simplified approach can be introduced that allows for the computation of the mean value of the aerodynamic damping based on a steady state CFD solution of the regarded stage. It is well known, that the aerodynamic damping of a blade mode family depends on the inter blade phase angle (IBPA) and its direction of propagation, which is not covered by the simplified approach. For higher modes the difference between the minimum and maximum damping is often low and the mean value is a good approximation, whereas for fundamental modes there is often a significant difference. However, it is shown that considering a mistuned vibration response of the rotor, the expected value of the mistuned damping exhibits the mean value of IBPA-dependent aerodynamic damping. CFD simulations of an oscillating airfoil indicate a certain validity range of the simplified approach based on a modified reduced frequency and inlet Mach number, which allows to determine for which industrial applications the approach is most suitable. Finally, this range of validity is verified with experimentally determined overall damping values from strain gauge measurements during operation for 2 different industrial applications, an axial compressor stage of a jet engine and a radial turbine stage of a turbocharger.

Efficient Identification of Naval High-Speed Craft Shock Mitigation Seat Modal Parameters From Drop-Test Data

2017 ◽  
Vol 139 (3) ◽  
Author(s):  
Zuneid Alam ◽  
Fred Afagh ◽  
Robert Langlois
Keyword(s):  
Test Data ◽  
High Speed ◽  
Dynamic Properties ◽  
Damping Ratio ◽  
Standard Test ◽  
Drop Test ◽  
Modal Parameters ◽  
Shock Attenuation ◽  
Shock Mitigation ◽  
Defence Research

Naval high-speed craft (HSC) operating in moderate to high seas experiences high-g and repeated shock loading at the seat–deck interface. These conditions are known to pose a serious potential for injury to the occupants. While various shock-mitigating seats are commercially available; their designs are in many cases quite different, and quantifying their shock attenuation characteristics can be challenging. The need for a standard test platform and experimental analysis methodology to investigate HSC seat effectiveness is a major objective of research being conducted by Carleton University's Applied Dynamics Laboratory (ADL) in partnership with Defence Research and Development Canada-Atlantic (DRDC Atlantic). A drop tower was designed and manufactured for testing HSC seats in order to characterize their shock-mitigating effectiveness by simulating the severe conditions of a slam impact at sea. Further, in order to identify seat dynamic parameters from drop-test data, the eigensystem realization algorithm (ERA), a modal-analysis-based system identification method, was applied to efficiently extract the modal parameters. The technique was shown to successfully extract the damping ratio as well as the damped and undamped natural frequencies of the seats from impact test data. The evaluated dynamic properties of the seats can subsequently inform decisions related to the design and/or procurement of commercially available seats.

scholarly journals Aerodynamic Damping Characteristics of a Transonic Turbine Cascade

1999 ◽  
Author(s):  
Mizuho Aotsuka ◽  
Toshinori Watanabe ◽  
Yasuo Machida
Keyword(s):  
Phase Angle ◽  
Aerodynamic Force ◽  
Turbine Blades ◽  
Aerodynamic Characteristics ◽  
Influence Coefficient ◽  
Suction Surface ◽  
Aerodynamic Damping ◽  
Unsteady Aerodynamic ◽  
Design Case ◽  
Oscillation Instability

The unsteady aerodynamic characteristics of oscillating thin turbine blades were studied both experimentally and numerically to obtain the comprehensive knowledge on the aerodynamic damping of the blades operating in transonic flows. The experiment was carried out in a linear cascade tunnel by use of the influence coefficient method. The two flow conditions were adopted, namely, a near-design condition and an off-design condition with a higher back pressure. In the results for the near-design case, a strong vibration instability was observed in the positive side of the interblade phase angle. In the off-design case, however, the instability did not appear for almost all the interblade phase angles. A drastic change was found in the phase angle of unsteady aerodynamic force between the two cases, which change was a governing factor for the oscillation instability. Numerical simulation based on 2-D Euler equation revealed that the phase change came from the change in phase of the unsteady surface pressure across the shock impingement point on the blade suction surface in the off-design case. The numerical results also showed that the aerodynamic damping increased with increasing reduced frequency, and that the oscillation instability disappeared.

scholarly journals Unsteady Gapwise Periodicity of Oscillating Cascaded Airfoils

1982 ◽  
Author(s):  
F. O. Carta
Keyword(s):  
Unsteady Flow ◽  
Fourier Coefficient ◽  
Phase Angle ◽  
Leading Edge ◽  
Experimental Measurements ◽  
Pressure Data ◽  
Linear Cascade ◽  
Aerodynamic Damping ◽  
The Stability ◽  
Interblade Phase Angle

Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade and over the chord of the center blade. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. In addition, results from two unsteady theories for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results that emerged from this investigation were: (a) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested, (b) as before, the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades, and (c) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.

Test & Research on the Shearing Strength of the Continuous Seams in the Mortar-Less Fabricated Block & Masonry

2015 ◽  
Vol 777 ◽  
pp. 194-200
Author(s):  
Xi Xi He ◽  
Ye Lin
Keyword(s):  
Test Data ◽  
Concrete Block ◽  
Standard Test ◽  
Shear Loading ◽  
Shearing Strength ◽  
Loading Methods

Test & research on the shearing strength of the hollow mortar-less fabricated concrete block & masonry has been done in this article. A proposed formula has been put forward in this article by analyzing the influence of the shear loading methods on the shearing strength using the shearing test data of 27 standard test specimens’ continuous seams divided into 9 groups with 100% concrete infill ratio.

Heat Transfer and Aerodynamics of Turbine Blade Tips in a Linear Cascade

2004 ◽  
Vol 128 (2) ◽  
pp. 300-309 ◽  
Author(s):  
P. J. Newton ◽  
G. D. Lock ◽  
S. K. Krishnababu ◽  
H. P. Hodson ◽  
W. N. Dawes ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Flow Visualization ◽  
Turbine Blade ◽  
Pressure Coefficient ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Linear Cascade ◽  
Side Edge ◽  
High Heat Transfer

Local measurements of the heat transfer coefficient and pressure coefficient were conducted on the tip and near tip region of a generic turbine blade in a five-blade linear cascade. Two tip clearance gaps were used: 1.6% and 2.8% chord. Data was obtained at a Reynolds number of 2.3×105 based on exit velocity and chord. Three different tip geometries were investigated: A flat (plain) tip, a suction-side squealer, and a cavity squealer. The experiments reveal that the flow through the plain gap is dominated by flow separation at the pressure-side edge and that the highest levels of heat transfer are located where the flow reattaches on the tip surface. High heat transfer is also measured at locations where the tip-leakage vortex has impinged onto the suction surface of the aerofoil. The experiments are supported by flow visualization computed using the CFX CFD code which has provided insight into the fluid dynamics within the gap. The suction-side and cavity squealers are shown to reduce the heat transfer in the gap but high levels of heat transfer are associated with locations of impingement, identified using the flow visualization and aerodynamic data. Film cooling is introduced on the plain tip at locations near the pressure-side edge within the separated region and a net heat flux reduction analysis is used to quantify the performance of the successful cooling design.

Numerical Analysis on Axial Compressor Stage Performance With Vortex Generators

2015 ◽  
Author(s):  
Ruchika Agarwal ◽  
Sridharan R. Narayanan ◽  
Shraman N. Goswami ◽  
Balamurugan Srinivasan
Keyword(s):  
Secondary Flow ◽  
Cross Flow ◽  
Axial Compressor ◽  
Axial Flow ◽  
Vortex Generator ◽  
Separated Flow ◽  
Sensitivity Study ◽  
Vortex Generators ◽  
Test Case ◽  
Compressor Stage

The performance of axial flow compressor stage can be improved by minimizing the effects of secondary flow and by avoiding flow separation. At higher blade loading, interaction of tip secondary flow and separated flow on blade surface is more near the tip of the rotor. This results in stall and hence decreases compressor performance. A previous study [1] was carried out to improve the performance of a rotating row of blades with the help of Vortex Generators (VGs) and considerable effects were observed. The current investigation is carried out to find out the effect of Vortex Generator (VG) on the performance of a compressor stage. NASA Rotor 37 with NASA Stator 37 (stage) is used as a test case for the current numerical investigation. VGs are placed at different chord wise as well as span wise locations. A mesh sensitivity study has been done so that simulation result is mesh independent. The results of numerical simulation with different geometrical forms and locations of VGs are presented in this paper. The investigation includes a description of the secondary flow effect and separation zone in compressor stage based on numerical as well as experimental results of NASA Rotor 37 with Stator 37 (without VG). It is also observed that the shape and location of the VG impacts the end wall cross flow and flow deflection [1], which result in enhanced stall range.

Analysis of a Platform for Measuring Moments and Products of Inertia of Large Vehicles

1976 ◽  
Vol 98 (2) ◽  
pp. 186-195
Author(s):  
D. Orne ◽  
T. Schmitz
Keyword(s):  
System Identification ◽  
Test Data ◽  
Normal Modes ◽  
Free Vibration Analysis ◽  
Identification Algorithm ◽  
Test Configuration ◽  
Truncation Errors ◽  
Sample Test ◽  
Instantaneous Center ◽  
Cable Lines

A rigid platform symmetrically supported by four sloping cables is proposed for measuring the center-of-gravity coordinates and the moments and products of inertia of large vehicles such as buses, trucks, and trailers. In addition to a torsional degree-of-freedom, the system undergoes pitch and roll motions about axes through the system instantaneous center which lies directly below the center of the platform at the intersection of the cable lines-of-action under quiescent conditions. The natural frequencies and normal modes of the freely vibrating loaded platform are used as inputs to a linearized System Identification Algorithm for computing the inertia properties of the test vehicle. Hypothetical test data generated from the Free Vibration Analysis of a sample test configuration are used to evaluate the sensitivity of the System Identification Algorithm to inaccuracies in test data or to truncation errors in computation.

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