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Artykuły w czasopismach na temat "Transonic Cascade Flutter Study"
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Lu, P. J., i S. K. Chen. "Evaluation of Acoustic Flutter Suppression for Cascade in Transonic Flow". Journal of Engineering for Gas Turbines and Power 124, nr 1 (1.11.2000): 209–19. http://dx.doi.org/10.1115/1.1365933.
Pełny tekst źródłaStreszczenie:
Flutter suppression via actively excited acoustic waves is a new idea proposed recently. The high flutter frequency (typically 50–500 Hz for a fan blade) and stringent space constraint make conventional mechanical type flutter suppression devices difficult to implement for turbomachines. Acoustic means arises as a new alternative which avoids the difficulties associated with the mechanical methods. The objective of this work is to evaluate numerically the transonic flutter suppression concept based on the application of sound waves to two-dimensional cascade configuration. This is performed using a high-resolution Euler code based on a dynamic mesh system. The concept has been tested to determine the effectiveness and limitations of this acoustic method. In a generic bending-torsion flutter study, trailing edge is found to be the optimal forcing location and the control gain phase is crucial for an effective suppression. The P&W fan rig cascade was used as the model to evaluate the acoustic flutter suppression technique. With an appropriate selection of the control logic the flutter margin can be enlarged. Analogous to what were concluded in the isolated airfoil study, for internal excitation, trailing-edge forcing was shown to be optimal since the trailing-edge receptivity still works as the dominant mechanism for generating the acoustically induced airloads.
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Kobayashi, H. "Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter". Journal of Turbomachinery 112, nr 4 (1.10.1990): 768–77. http://dx.doi.org/10.1115/1.2927720.
Pełny tekst źródłaStreszczenie:
Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement, due to blade oscillation, and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting on an oscillating blade, were joined and, then, the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semichord from 0.0375 to 0.547, six interblade phase angles, and inlet flow velocities from subsonic to supersonic flow. The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle, and inlet flow velocity was clarified, including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles that caused flutter were in the range from 40 to 160 deg for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.
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Ott, P., A. Bo¨lcs i T. H. Fransson. "Experimental and Numerical Study of the Time-Dependent Pressure Response of a Shock Wave Oscillating in a Nozzle". Journal of Turbomachinery 117, nr 1 (1.01.1995): 106–14. http://dx.doi.org/10.1115/1.2835625.
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Investigations of flutter in transonic turbine cascades have shown that the movement of unsteady normal shocks has an important effect on the excitation of blades. In order to predict this phenomenon correctly, detailed studies concerning the response of unsteady blade pressures versus different parameters of an oscillating shock wave should be performed, if possible isolated from other flow effects in cascades. In the present investigation the correlation between an oscillating normal shock wave and the response of wall-mounted time-dependent pressure transducers was studied experimentally in a nozzle with fluctuating back pressure. Excitation frequencies between 0 Hz and 180 Hz were investigated. For the measurements, various measuring techniques were employed. The determination of the unsteady shock position was made by a line scan camera using the Schlieren flow visualization technique. This allowed the simultaneous use of unsteady pressure transducers to evaluate the behavior of the pressure under the moving shock. A numerical code, based on the fully unsteady Euler equations in conservative form, was developed to simulate the behavior of the shock and the pressures. The main results of this work were: (1) The boundary layer over an unsteady pressure transducer has a quasi-steady behavior with respect to the phase lag. The pressure amplitude depends on the frequency of the back pressure. (2) For the geometry investigated the shock amplitude decreased with increasing excitation frequency. (3) The pressure transducer sensed the arriving shock before the shock had reached the position of the pressure transducer. (4) The computed unsteady phenomena agree well with the results of the measurements.
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Lepicovsky, J., R. V. Chima, E. R. McFarland i J. R. Wood. "On Flowfield Periodicity in the NASA Transonic Flutter Cascade". Journal of Turbomachinery 123, nr 3 (1.02.2000): 501–9. http://dx.doi.org/10.1115/1.1378300.
Pełny tekst źródłaStreszczenie:
A combined experimental and numerical program was carried out to improve the flow uniformity and periodicity in the NASA transonic flutter cascade. The objectives of the program were to improve the periodicity of the cascade and to resolve discrepancies between measured and computed flow incidence angles and exit pressures. Previous experimental data and some of the discrepancies with computations are discussed. In the present work surface pressure taps, boundary layer probes, shadowgraphs, and pressure-sensitive paints were used to measure the effects of boundary layer bleed and tailboard settings on flowfield periodicity. These measurements are described in detail. Two numerical methods were used to analyze the cascade. A multibody panel code was used to analyze the entire cascade and a quasi-three-dimensional viscous code was used to analyze the isolated blades. The codes are described and the results are compared to the measurements. The measurements and computations both showed that the operation of the cascade was heavily dependent on the endwall configuration. The endwalls were redesigned to approximate the midpassage streamlines predicted using the viscous code, and the measurements were repeated. The results of the program were that: (1) Boundary layer bleed does not improve the cascade flow periodicity. (2) Tunnel endwalls must be shaped like predicted cascade streamlines. (3) The actual flow incidence must be measured for each cascade configuration rather than using the tunnel geometry. (4) The redesigned cascade exhibits excellent periodicity over six of the nine blades.
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Kobayashi, H. "Unsteady Aerodynamic Damping Measurement of Annular Turbine Cascade With High Deflection in Transonic Flow". Journal of Turbomachinery 112, nr 4 (1.10.1990): 732–40. http://dx.doi.org/10.1115/1.2927716.
Pełny tekst źródłaStreszczenie:
Unsteady aerodynamic forces acting on oscillating blades of a transonic annular turbine cascade were investigated in both aerodynamic stable and unstable domains, using a Freon gas annular cascade test facility. In the facility, whole blades composing the cascade were oscillated in the torsional mode by a high-speed mechanical drive system. In the experiment, the reduced frequency K was changed from 0.056 to 0.915 with a range of outlet Mach number M2 from 0.68 to 1.39, and at a constant interblade phase angle. Unsteady aerodynamic moments obtained by two measuring methods agreed well. Through the moment data the phenomenon of unstalled transonic cascade flutter was clarified as well as the significance of K and M2 for the flutter. The variation of flutter occurrence with outlet flow velocity in the experiments showed a very good agreement with theoretical analysis.
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Lepicovsky, Jan, David Šimurda, Jindřich Hála, Petr Šidlof i Martin Štěpán. "Blade pressure loading and torque measurement in a transonic linear cascade". Journal of Physics: Conference Series 2511, nr 1 (1.05.2023): 012030. http://dx.doi.org/10.1088/1742-6596/2511/1/012030.
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Abstract Experimental results of a transonic compressor blade pressure loadings and blade shaft torque measurements are presented in this paper. Data were acquired for the cascade middle blade being set to a number of incidence angle offsets to simulate phases of a blade flutter oscillatory motion. This paper should be viewed as a progress report on the ongoing larger research effort on blade flutter in transonic flow regimes.
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Bakhle, Milind A., T. S. R. Reddy i Theo G. Keith. "Subsonic/Transonic Cascade Flutter Using a Full-Potential Solver". AIAA Journal 31, nr 7 (lipiec 1993): 1347–49. http://dx.doi.org/10.2514/3.49072.
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Kobayashi, H. "Effects of Shock Waves on Aerodynamic Instability of Annular Cascade Oscillation in a Transonic Flow". Journal of Turbomachinery 111, nr 3 (1.07.1989): 222–30. http://dx.doi.org/10.1115/1.3262259.
Pełny tekst źródłaStreszczenie:
The effects of shock waves on the aerodynamic instability of annular cascade oscillation were examined for rows of both turbine and compressor blades, using a controlled-oscillating annular cascade test facility and a method for accurately measuring time-variant pressures on blade surfaces. The nature of the effects and blade surface extent affected by shock waves were clarified over a wide range of Mach number, reduced frequency, and interblade phase angle. Significant unsteady aerodynamic forces were found generated by shock wave movement, which markedly affected the occurrence of compressor cascade flutter as well as turbine cascade flutter. For the turbine cascade, the interblade phase angle significantly controlled the effect of force, while for the compressor cascade the reduced frequency controlled it. The chordwise extent of blade surface affected by shock movement was estimated to be approximately 6 percent chord length.
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McBean, Ivan, Kerry Hourigan, Mark Thompson i Feng Liu. "Prediction of Flutter of Turbine Blades in a Transonic Annular Cascade". Journal of Fluids Engineering 127, nr 6 (29.05.2005): 1053–58. http://dx.doi.org/10.1115/1.2060731.
Pełny tekst źródłaStreszczenie:
A parallel multiblock Navier-Stokes solver with the k‐ω turbulence model is used to solve the unsteady flow through an annular turbine cascade, the transonic Standard Test Case 4, Test 628. Computations are performed on a two- and three-dimensional model of the blade row with either the Euler or the Navier-Stokes flow models. Results are compared to the experimental measurements. Comparisons of the unsteady surface pressure and the aerodynamic damping are made between the three-dimensional, two-dimensional, inviscid, viscous simulations, and experimental data. Differences are found between the stability predictions by the two- and three-dimensional computations, and the Euler and Navier-Stokes computations due to three-dimensionality of the cascade model and the presence of a boundary layer separation, respectively.
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Cinnella, P., P. De Palma, G. Pascazio i M. Napolitano. "A Numerical Method for Turbomachinery Aeroelasticity". Journal of Turbomachinery 126, nr 2 (1.04.2004): 310–16. http://dx.doi.org/10.1115/1.1738122.
Pełny tekst źródłaStreszczenie:
This work provides an accurate and efficient numerical method for turbomachinery flutter. The unsteady Euler or Reynolds-averaged Navier-Stokes equations are solved in integral form, the blade passages being discretised using a background fixed C-grid and a body-fitted C-grid moving with the blade. In the overlapping region data are exchanged between the two grids at every time step, using bilinear interpolation. The method employs Roe’s second-order-accurate flux difference splitting scheme for the inviscid fluxes, a standard second-order discretisation of the viscous terms, and a three-level backward difference formula for the time derivatives. The dual-time-stepping technique is used to evaluate the nonlinear residual at each time step. The state-of-the-art second-order accuracy of unsteady transonic flow solvers is thus carried over to flutter computations. The code is proven to be accurate and efficient by computing the 4th Aeroelastic Standard Configuration, namely, the subsonic flow through a turbine cascade with flutter instability in the first bending mode, where viscous effect are found practically negligible. Then, for the very severe 11th Aeroelastic Standard Configuration, namely, transonic flow through a turbine cascade at off-design conditions, benchmark solutions are provided for various values of the inter-blade phase angle.
Rozprawy doktorskie na temat "Transonic Cascade Flutter Study"
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Roy, Arnab. "Experimental Study of Gas Turbine Endwall Cooling with Endwall Contouring under Transonic Conditions". Diss., Virginia Tech, 2014. http://hdl.handle.net/10919/25801.
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The effect of global warming due to increased level of greenhouse gas emissions from coal fired thermal power plants and crisis of reliable energy resources has profoundly increased the importance of natural gas based power generation as a major alternative in the last few decades. Although gas turbine propulsion system had been primarily developed and technological advancements over the years had focused on application in civil and military aviation industry, use of gas turbine engines for land based power generation has emerged as the most promising candidate due to higher thermal efficiency, abundance of natural gas resources, development in generation of hydrogen rich synthetic fuel (Syngas) using advanced gasification technology for further improved emission levels and strict enforcement in emission regulations on installation of new coal based power plants. The fundamental thermodynamic principle behind gas turbine engines is Brayton cycle and higher thermal efficiency is achieved through maximizing the Turbine Inlet Temperature (TIT). Modern gas turbine engines operate well beyond the melting point of the turbine component materials to meet the enhanced efficiency requirements especially in the initial high pressure stages (HPT) after the combustor exit. Application of thermal barrier coatings (TBC) provides the first line of defense to the hot gas path components against direct exposure to high temperature gases. However, a major portion of the heat load to the airfoil and passage is reduced through injection of secondary air from high pressure compressor at the expense of a penalty on engine performance. External film cooling comprises a significant part of the entire convective cooling scheme. This can be achieved injecting coolant air through film holes on airfoil and endwall passages or utilizing the high pressure air required to seal the gaps and interfaces due to turbine assembly features. The major objective is to maximize heat transfer performance and film coverage on the surface with minimum coolant usage.
Endwall contouring on the other hand provides an effective means of minimizing heat load on the platform through efficient control of secondary flow vortices. Complex vortices form due to the interaction between the incoming boundary layer and endwall-airfoil junction at the leading edge which entrain the hot gases towards the endwall, thus increasing surface heat transfer along its trajectory. A properly designed endwall profile can weaken the effects of secondary flow thereby improving the aerodynamic and associated heat transfer performance.
This dissertation aims to investigate heat transfer characteristics of a non-axisymmetric contoured endwall design compared to a baseline planar endwall geometry in presence of three major endwall cooling features – upstream purge flow, discrete hole film cooling and mateface gap leakage under transonic operating conditions. The preliminary design objective of the contoured endwall geometry was to minimize stagnation and secondary aerodynamic losses. Upstream purge flow and mateface gap leakage is necessary to prevent ingestion to the turbine core whereas discrete hole cooling is largely necessary to provide film cooling primarily near leading edge region and mid-passage region. Different coolant to mainstream mass flow ratios (MFR) were investigated for all cooling features at design exit isentropic Mach number (0.88) and design incidence angle. The experiments were performed at Virginia Tech's quasi linear transonic blow down cascade facility. The airfoil span increases in the mainstream flow direction in order to match realistic inlet/exit airfoil surface Mach number distribution. A transient Infrared (IR) thermography technique was employed to measure the endwall surface temperature and a novel heat transfer data reduction method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness (ETA), assuming a 1D semi-infinite transient conduction. An experimental study on endwall film cooling with endwall contouring at high exit Mach numbers is not available in literature.
Results indicate significant benefits in heat transfer performance using the contoured endwall in presence of individual (upstream slot, discrete hole and mateface gap) and combined (upstream slot with mateface gap) cooling flow features. Major advantages of endwall contouring were observed through reduction in heat transfer coefficient and increase in coolant film coverage by weakening the effects of secondary flow and cross passage pressure differential. Net Heat Flux Reduction (NHFR) analysis was carried out combining the effect of heat transfer coefficient and film cooling effectiveness on both endwall geometries (contoured and baseline) where, the contoured endwall showed major improvement in heat load reduction near the suction side of the platform (upstream leakage only and combined upstream with mateface leakage) as well as further downstream of the film holes (discrete hole film cooling). Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed.Ph. D.
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Kullberg, James C. "An experimental and numerical study of secondary flows and film cooling effectiveness in a transonic cascade". Honors in the Major Thesis, University of Central Florida, 2011. http://digital.library.ucf.edu/cdm/ref/collection/ETH/id/454.
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Experimental tests on a transonic annular rig are time-consuming and expensive, so it is desirable to use experimental results to validate a computational model which can then be used to extract much more information. The purpose of this work is to create a numerical model that can be used to simulate many different scenarios and then to apply these results to experimental data.; In the modern world, gas turbines are widely used in aircraft propulsion and electricity generation. These applications represent a massive use of energy worldwide, so even a very small increase in efficiency would have a significant beneficial economic and environmental impact. There are many ways to optimize the operation of a gas turbine, but a fundamental approach is to increase the turbine inlet temperature to increase the basic thermodynamic efficiency of the turbine. However, these temperatures are already well above the melting temperature of the components. A primary cooling methodology, called film cooling, creates a blanket of cool air over the surface and is an effective way to help protect these components from the hot mainstream gasses. This paper focuses on the effect of the film holes upstream of the first row of blades in the turbine because this is the section that experiences the highest thermal stresses. Many factors can determine the effectiveness of the film cooling, so a complete understanding can lead to effective results with the minimum flow rate of coolant air. Many studies have been published on the subject of film cooling, but because of the difficulty and expense of simulating turbine realistic conditions, many authors introduce vast simplifications such as low speed conditions or linear cascades. These simplifications do not adequately represent the behavior of a turbine and therefore their results are of limited use. This study attempts to eliminate many of those simplifications. The test rig used in this research is based on the NASA-GE E³ design, which stands for Energy Efficient Engine. It was introduced into the public domain to provide an advanced platform from which open-literature research could be performed.B.S.M.E.
Bachelors
Engineering and Computer Science
Mechanical Engineering
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Prahallada, J. "Blade Flutter in a Linear Cascade: Unsteady Loads and Flow Features in Subsonic and Transonic Flows". Thesis, 2018. https://etd.iisc.ac.in/handle/2005/4108.
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Vibration related issues like flutter can significantly limit the performance of aircraft engines and cause unwarranted cost and time overruns. The increased demand for more powerful yet compact engines has resulted in the use of relatively thin and long blade rows, which are more susceptible to such vibration related issues. Flutter refers to an aeroelastic instability in which the motion of the blade interacts with the flow to generate the unsteady fluid loads that can sustain or possibly grow its oscillations, which can ultimately lead to structural failure. The severe consequences that follow the phenomenon of flutter has triggered substantial interest in flutter studies. Linear cascades that represent a particular span section of the rotor have proven to be a reliable tool for such flutter studies. Although several studies pertaining to flutter have been reported in the literature, only a few of them have concentrated on linking the flutter characteristics to the corresponding unsteady flow field around the blade. With this in mind, in the present work, detailed experimental study of bending mode flutter in two representative cascades, one operating in stalled conditions at low Reynolds number (∼ 104) incompressible flow, and the other at transonic conditions has been carried out. The main focus of the work is on simultaneous measurements of flutter characteristics and the unsteady flow field around the blade that can help in understanding the unsteady flow features that contribute to flutter. The blade oscillations are forced and the flutter characteristics are deduced in terms of the energy transfer to the blade from the measured unsteady loads, and the flow field is measured with the help of PIV in both the cases, in addition to high-speed shadowgraphy for the transonic case.
In the low Reynolds number cascade case, three blades in the cascade were oscillated with arbitrary phase difference between adjacent blades referred to as Inter Blade Phase Angle (IBPA). The response of the flow to these imposed oscillations is measured in terms of unsteady loads on the central blade in the cascade, and this is used to quantify the mean energy transfer to the blade from the fluid over an oscillation cycle. The parameters varied in this case include the reduced frequency (k) (up to 0.1) and the Inter Blade Phase Angle (IBPA), the latter being varied from +180◦ to −180◦ in steps of 45◦. The experiments were conducted at three different post stall incidence angles of the blades in the cascade to assess the influence of blade loading on flutter behaviour. In each case, IBPA and k have been
varied and contours plots of the excitation have been obtained in the plane of IBPA and k, showing the region of excitation, with the results indicating that most of the excitation occurs around IBPA of +90◦. Fluid excitation at lower k values for specific IBPA cases of +45◦ and +90◦ was observed, indicating the influence of reduced frequency (k) and Inter Blade Phase Angle (IBPA) on cascade stability. Also, an increased blade loading is observed to significantly increase the extent of excitation or damping. To understand the contribution from oscillating adjacent blades, experiments involving a single blade oscillating in a cascade have also been performed. PIV measurements at different IBPA values show that there exists significant differences in the phase of the separated unsteady shear layer dynamics with respect to the blade motion between excitation (flutter) and damping (no flutter) cases. The PIV measurements also clearly show the effect of the time-varying inter-blade spacing on the shear layer dynamics, with the shear layer tending to separate at instances when there is a large inter-blade spacing, compared to the instances when the inter-blade spacing is small, this being important for cases with different IBPA values.
For the transonic cascade case, in order to facilitate flutter studies at flow conditions that are realistic to aircraft engine components, a new blow-down transonic cascade facility has been developed as part of the present work. The facility is equipped with a mechanism that can oscillate the central blade in the cascade at realistic reduced frequency (k) of about 0.1. The parameters varied in this case include the reduced frequency (k) up to 0.1 and the static pressure ratio (SPR) across the cascade. The SPR in these transonic cases alters the passage shock position, which is seen to have a large effect on the corresponding flutter characteristics of the blade. Four SPR cases of 1.05, 1.25, 1.35 and 1.55 are considered for flutter studies of which the first three have a passage shock, while the SPR = 1.55 case corresponds to an unstarted cascade with a detached shock at the leading edge. The experimental results indicate striking differences in the flutter behaviour between the started and unstarted cascade cases. While both the cases show excitation at lower k values (k ≈ 0.05), in the unstarted cascade, an additional regime of huge excitation is observed at relatively higher k values, with the excitation values in this case being about an order of magnitude higher that that at lower k values. A large PIV data set has been obtained simultaneously with unsteady loads in select cases. High-speed shadowgraphy visualizations have also been carried out at different reduced frequencies for both a started cascade case with passage shock, and an unstarted cascade case with a detached leading edge shock. The results indicate that the shocks oscillate in response to the blade motions, with the phase between the shock motion and the blade motion being dependent on the reduced frequency and SPR. In particular, differences are seen in the phase of the shock motion with respect to the blade motion between the excitation and damping cases, and also between started and
unstarted cascade cases. These measurements also show striking differences in the shock phase between the two excitation cases of the unstarted cascade case, indicating differences in the underlying mechanisms responsible for the two excitation regimes observed at lower and higher frequencies, respectively. Specifically, when the blade is at the suction side extreme location, the detached leading edge shock is found to be located upstream of its mean location in the low frequency excitation case, while the shock is located downstream of the mean location at the higher frequency excitation regime.
In summary, the measurements indicate that the unsteady shear layer and its phase relation with the blade motion is the deciding factor for stall flutter at low Reynolds numbers, while it is the phase of the unsteady shock motions with respect to the blade motion that is crucial in the transonic cascade case.GATET initiative of AR&DB and GTRE
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Yao, Lo-shan, i 姚樂山. "The Experiment of SDOF Transonic Cascade Flutter Suppression Using Active Acoustic Control ─ Cascade Flutter Experiment". Thesis, 2000. http://ndltd.ncl.edu.tw/handle/90266539357495716146.
Pełny tekst źródłaStreszczenie:
碩士國立成功大學
航空太空工程學系
88
Flutter is a dynamic instability due to mutual interaction among inertia, elastic, and unsteady aerodynamic forces. Single-degree-of-freedom (SDOF) flutter is a typical flutter type occurring in the transonic flow regime. The objective of the present work is to study the SDOF transonic torsional flutter of cascade via experimental method. Method of image is applied to render the single blade testing in a solid wall wind tunnel a simulation that represents a cascade flow of 180 degrees shift in interblade phase angle. The experiment was conducted in a 200 mm ´ 200 mm blow-down type transonic pilot wind tunnel. The tested Mach number ranges from 0.4 to 0.8. The experimental works include the manufacturing of a fan blade, the design and construction of a model support and safety interlock system, as well as the design of data acquisition and processing system. The present flutter experiment successfully obtained blade flutter boundaries at Mach numbers 0.404, 0.564, 0.692, 0.728, and 0.817, respectively. The test data show that the flutter boundaries exhibit “transonic dip” phenomenon and the flutter frequency approaches the blade natural vibration frequency as Mach number increases. Comparing to the perforated wall flutter experiment, the present cascade flutter shows premature transonic dip at a smaller Mach number. Moreover, the flutter frequency approaches the structural natural vibration frequency via a different route of trend.
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Chen, Sen-Kuei, i 陳森貴. "ACOUSTIC FLUTTER SUPPRESSION OF CASCADE ININVISCID AND VISCOUS TRANSONIC FLOWS". Thesis, 1999. http://ndltd.ncl.edu.tw/handle/57245433871899532585.
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博士國立成功大學
航空太空工程學系
87
The objective of this research is to evaluate numerically the transonic flutter suppression concept based on the application of sound waves to a cascade configuration. Both Euler and Navier-Stokes equations are used. The emphasis is placed on finding the basic suppression mechanism and the important physical and control parameters involved. Of particular interest is the viscous effect on the present acoustic method, which has not yet been attempted before. A high-resolution Modified Osher-Chakravarthy MUSCL type third order upwind total variation diminishing (TVD) scheme based on dynamic mesh is used as the analysis tool. Modification on the reconstruction of the cell interface values makes the present method more suitable for treating viscous and highly stretched cascade grid system. A non-reflected boundary condition capable of treating reflected outgoing and imposed incoming waves is implemented in the present code. Monopole sound source treatment is numerically designed to model the acoustic waves emitted from the solid wall. This time-accurate flow solver is validated first using various kinds of acoustic, flutter and viscous flow model problems. In a generic bending-torsion flutter study, trailing-edge is found to be the optimal forcing location and the control gain phase is crucial for an effective suppression. The P&W TS33 test cascade was used as the model to evaluate the acoustic flutter suppression technique. The analysis of blade row instability versus some discrete interblade phase angles was presented. The cases with interblade phase angle are selected to test the active acoustic method for the cascade flow. Both internal and external active excitations were applied. With an appropriate selection of the control logic the flutter margin can be enlarged. For external excitation enforced at the downstream exit plane, it is found that the critical angle of acoustic incidence inhibits certain interblade phase angle range from being controlled acoustically. Analogous to what were concluded in the isolated airfoil study, for internal excitation, trailing-edge forcing is shown to be optimal since the trailing-edge receptivity still works as the dominant mechanism for generating the acoustically-induced airloads. The influence of boundary layer effect on the acoustic excitation was explored. The vorticity blown across the boundary layer by acoustic excitation is found to be the contributing factor which causes the discrepancy of the acoustically-induced airloads between inviscid and viscous results. The effect of these ejected and induced vortices has been studied by both analytical and numerical methods. A subsequent limit cycle type transonic airfoil flutter in both inviscid and viscous flows was examined using active acoustic control technique. These results show that flutter can be suppressed at the beginning stage of the unstable motion, and the ability to suppress the instability is dependent on the energy (acoustic strength) supplied by the acoustic actuator. Reversed trend is found in viscous flutter suppression study as the gain amplitude is increased to strengthen the acoustic actuator. This is conjectured that too strong forcing might cause boundary layer separation, leading to a more complex and unstable situation than what is anticipated.
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I, Su Hung, i 蘇鴻義. "The Experiment of SDOF Transonic Cascade Flutter Suppression Using Active Acoustic Control". Thesis, 2001. http://ndltd.ncl.edu.tw/handle/01966449700219802516.
Pełny tekst źródłaStreszczenie:
碩士國立成功大學
航空太空工程學系
89
Flutter analysis is a required step that has to be validated in the design loop of a new turboengine. Flutter suppression technology can be either passive or active. This research adopts the active method. Acoustic control technique is used here to suppress flutter instability occurring in the transonic flow. For acoustic flutter control, the primary field is that induced by the structural vibration or elastic deformation, whereas the secondary field comes from the active acoustic devices. The aerodynamic damping which is induced by the trailing-edge receptivity effect of the secondary fields holds a key role in suppressing the flutter. The experiment was conducted in a 200mm 200mm blow-down type transonic pilot wind tunnel. Flutter boundary is found in the first step, and active acoustic control is employed to suppress the flutter motion. Acoustic flutter control calls for sensitive sensor, fast-responsive sound actuator, and correct control logic including gain phase and amplitude. It has been shown previously that only when flutter instability is detected early in the beginning infinitesimal stage and a sufficient strength of sound wave can be generated with appropriate phase shift could a flutter motion be suppressed acoustically. The present work include 1) the hardware improvement of the control valve and the front setting chamber, 2) the calibration of the flutter flow field using new equipments, 3) the search of flutter boundary in transonic flow, and 4) to confirm the operation range of acoustic equipment. The objective of the present research is to study the feasibility of suppressing single-degree-of-freedom torsional flutter of a cascad flow.
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Panthi, Niraj. "Shock Dynamics due to Downstream Pressure Perturbations: Idealization of Transonic Unstarted Cascade Flutter". Thesis, 2019. https://etd.iisc.ac.in/handle/2005/4723.
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Shock wave unsteadiness is an important phenomenon in high-speed aerodynamics. This
phenomenon is observed in many locations of high-speed vehicles, such as supersonic inlets,
ramjet engines, transonic airfoils, high-speed fans, and compressors. In supersonic inlets and
ramjet engines, unsteady shock motions can lead to large undesirable local fluctuations in
properties such as pressure and heat transfer rate, besides overall thrust fluctuations. Shock
unsteadiness in transonic airfoils can induce structural vibrations known as buffeting, while
in gas turbine fans/compressors, shock oscillations can lead to blade vibrations known as
flutter. Motivated by the above problems, the purpose of the present experimental study is
to understand the response of a normal shock subjected to downstream pressure perturbations.
Although several studies pertaining to shock dynamics due to downstream pressure
perturbations have been reported in the literature, only a few of them have concentrated on
the effect of downstream perturbation to the normal shock behavior in a constant area duct,
with detailed flow field measurements that can help to understand the flow physics that drive
these shock oscillations. With this in mind, in the present work, a detailed experimental
study of shock dynamics due to downstream pressure perturbations within a constant area
duct have been done in two different configurations. The first configuration is one in which
the downstream pressure perturbations are generated in the far field region by rotating a
triangular cross-sectional shaft, this being an idealization of inlet shock dynamics in ramjets
caused by downstream combustion chamber pressure fluctuations. The second configuration
is one in which the downstream pressure perturbations are generated in the near field
region by heaving an airfoil, and this may be considered as an idealization of the unstarted
cascade flutter of high-speed compressors. In both cases, the normal shock is induced and
stabilized at low Mach numbers (M∞ ∼ 1.3) within a supersonic/transonic tunnel, and the
shock dynamics in response to the downstream pressure perturbations are visualized using
high-speed shadowgraphy. In addition to the high-speed shadowgraphy, high-speed wall
pressure measurements have been carried out in the first configuration, and in the second
configuration, unsteady force measurements have been carried out to understand the influence
of shock oscillations on airfoil flutter.
In the far field pressure perturbation case, pressure perturbations are generated by rotating
a triangular cross-sectional shaft which is 580 mm downstream from the normal shock.
The normal shock is induced and stabilized in the constant area section of a supersonic
wind tunnel which is operated at M∞ = 1.4. The main parameter varied in this case is the
perturbation frequency ( f ), which is varied from low frequencies to 60 Hz in steps of 10
Hz. High-speed shadowgraphy visualizations indicate that the shock oscillates in response
to the exciter perturbation frequency, with a phase difference between exciter motion and
the shock displacement. The shock shows large streamwise motions (up to 60 mm), with
distinct differences in the shock structure and velocity during its upstream and downstream
motions. It is also observed that the amplitude of shock motion decreases with increase in
perturbation frequency, while the shock velocity is almost independent of the perturbation
frequency. The results from-high speed pressure measurements indicate that the downstream
pressure fluctuations are nearly 3-5% of the mean static pressure at the exciter region.
In the near field pressure perturbation case, pressure perturbations are generated by
heaving an airfoil (at frequency f ) with its leading edge being 0.1 chord length downstream
from the normal shock. The normal shock is induced and stabilized in the constant area
section of a transonic wind tunnel which is operated at M∞ = 1.3. The parameter varied
in this case is the reduced frequency (k = π f c/U), which is varied from low values up to
0.264. Flutter characteristics of the airfoil are deduced in terms of the energy transfer to
the heaving airfoil from the measured unsteady loads, and it indicates that there are two
excitation regions, one corresponding to lower reduced frequency and other corresponding to
higher reduced frequency, which is similar to the case of unstarted cascade flutter observed
by Jutur (2018). High-speed shadowgraphy visualizations have been carried out at different
airfoil heave frequencies, and the results indicate that the shock oscillates in response to the
airfoil heave motions, with the phase between the shock motion and the airfoil motion being
dependent on the reduced frequency. The correlation between the shock motion and airfoil
position indicate a negative correlation value at k = 0.049, and for all cases with k ≥ 0.117,
it is positively correlated.
In summary, measurements from both configurations indicate that the shock oscillates in
response to the exciter perturbation frequency, with a phase difference between shock motion
and exciter motion. This phase difference observed between the shock displacement and the
exciter for variation in perturbation frequency in the first configuration may be attributed to
shock wave boundary layer interactions, while in the second configuration it is the phase of
the unsteady shock motions with respect to the airfoil motion that is important in deciding
the flutter characteristics of the downstream airfoil. Further, in both the configurations, the
amplitude of shock motion is found to be decreasing with increase in perturbation frequency
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Lu, Feng-Tai, i 呂鋒泰. "The Experiment of SDOF Transonic Cascade Flutter Suppression - Construction of the Active Acoustic Control Facilities". Thesis, 2000. http://ndltd.ncl.edu.tw/handle/96607107588269794092.
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碩士國立成功大學
航空太空工程學系
88
Flutter analysis is a required step that has to be validated in the design loop of a new engine. Active acoustic control technique is a potential candidate that could be considered for suppressing engine blade flutter. Acoustic flutter control calls for sensitive sensor, fast-responsive sound actuator, and correct control logic including gain phase and amplitude. It has been shown previously that only when flutter instability can be detected early in the infinitesimal stage and sufficient strength of sound wave can be generated with appropriate phase shift could a flutter motion be suppressed acoustically. The objective of the present work is to develop an acoustic control device that can be applied to suppress blade flutter in a transonic wind tunnel experiment. The main design guideline lies in the direct amplification and phase shifting of the sensed flutter signals, and the use of this amplified signal to drive a loudspeaker system for flutter suppression. The tasks accomplished in the present design work include 1) a flutter test section design and manufacturing, 2) a design and validation of the flutter signal detection system and the phase changer system, and 3) the design and fabrication of a loudspeaker and the adapter system to the wind tunnel wall.
Książki na temat "Transonic Cascade Flutter Study"
1
Methodology of blade unsteady pressure measurement in the NASA transonic flutter cascade. [Cleveland, Ohio]: National Aeronautics and Space Administration, Glenn Research Center, 2002.
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2
National Aeronautics and Space Administration (NASA) Staff. Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic Flutter Cascade. Independently Published, 2018.
Znajdź pełny tekst źródłaCzęści książek na temat "Transonic Cascade Flutter Study"
1
Kaji, Shojiro. "Transonic Cascade Flutter in Combined Bending-Chordwise Translational Mode". W Unsteady Aerodynamics and Aeroelasticity of Turbomachines, 783–95. Dordrecht: Springer Netherlands, 1998. http://dx.doi.org/10.1007/978-94-011-5040-8_51.
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Prasad, Chandra Shekhar, i Ludek Pesek. "Classical flutter study in turbomachinery cascade using boundary element method for incompressible flows". W Advances in Mechanism and Machine Science, 4055–64. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-030-20131-9_404.
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Qiu, Ju, i Chaofeng Liu. "Verification and Validation of Supersonic Flutter of Rudder Model for Experiment". W Optimization Problems in Engineering [Working Title]. IntechOpen, 2021. http://dx.doi.org/10.5772/intechopen.98384.
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The abrupt and explosive nature of flutter is a dangerous failure mode, which is closely related to the structural modes. In this work, the principal goal of the study is to produce the model, which is used very accurately for flutter predictions. Mode correctness of the model can correct the test deflects by the optimization technique----Sequential Quadratic Programming (SQP). The optimization of two finite element models for two flight conditions, transonic and supersonic speeds, had the different objectives which were defined by the nonlinear and linear eigenvector errors. The first and second frequencies were taken as constraints. And the stiffness of the rotation shaft was also restricted to some limits. The stiffness of the rudder axle was defined as design variables. Experiments were performed for considering springs both in plunge and in torsion of the rudder shaft. When the comparison between experimental information and analyzed calculations is described, generally excellent agreement is obtained between experimental and calculated results, and aeroelastic instability is predicted that agrees with experimental observations. Comments are also given concerning improvements of the flutter speed to be made to the model with changing stiffness of the rudder axle. Most importantly, V&V Method is used to provide the confidence in the results from simulation in this paper. Firstly, it introduces experimental data from Ground Vibration Test to build up or modify the Finite Element Model, during the Verification phase, which makes simulated models closer to the real world and guarantees satisfaction of final computed results to requirements, such as airworthiness. Secondly, the flutter consequence is validated by wind tunnel test. These enhancements could find potential applications in industrial problems.
Streszczenia konferencji na temat "Transonic Cascade Flutter Study"
1
Chenaux, Virginie Anne, i Björn Grüber. "Aeroelastic Investigation of an Annular Transonic Compressor Cascade: Numerical Sensitivity Study for Validation Purposes". W ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-43297.
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The accuracy of flutter or forced response analyses of turbomachinery blade assemblies strongly depends on the correct prediction of the unsteady aerodynamic loads acting on the vibrating blades. This paper presents the aeroelastic numerical results of an annular transonic compressor cascade subjected to harmonic oscillation conditions. The measurements associated were performed in an annular test facility for non-rotating cascades. The aim of this investigation is to get a deeper understanding of the specific characteristics of this test facility as well as improving the flutter prediction procedure and accuracy. For a subsonic and a transonic flow condition, the steady-state blade surface pressure distributions were predicted with two mesh configurations and results were compared to the experimental results. The first configuration omits the geometrical complexity of the experimental model and only models the blade passage. The second mesh configuration includes the cascade’s detailed geometry and cavities. The presence of leakage flows arisen due to the cascade’s slits and cavities are identified and their impact on the main flow field is analyzed and discussed. For the flutter computations, two mesh resolutions were investigated. The global damping predicted with a fine and a coarse mesh was compared, as well as the local pressure amplitudes and phases predicted with both configurations. Results show that even though similar global damping curves are predicted with both mesh resolution, for some IBPAs, local differences exist on the pressure amplitudes and phases. This highlights that only comparing the global damping coefficient, is not sufficient for code validation.
2
Lu, Pong-Jeu, i Sen-Kuei Chen. "Evaluation of Acoustic Flutter Suppression for Cascade in Transonic Flow". W ASME 1998 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1998. http://dx.doi.org/10.1115/98-gt-065.
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Flutter suppression via actively excited acoustic waves is a new idea proposed recently. The high flutter frequency (typically 50–500 Hz for a fan blade) and stringent space constraint makes conventional mechanical type flutter suppression devices difficult to implement for turbomachines. Acoustic means arises as a new alternative which avoids the difficulties associated with the mechanical methods. The objective of this work is to evaluate numerically the transonic flutter suppression concept based on the application of sound waves to 2D cascade configuration. This is performed using a high-resolution Euler code based on a dynamic mesh system. The concept has been tested to determine the effectiveness and limitations of this acoustic method. In a generic bending-torsion flutter study, trailing-edge is found to be the optimal forcing location and the control gain phase is crucial for an effective suppression. The P&W fan rig cascade was used as the model to evaluate the acoustic flutter suppression technique. With an appropriate selection of the control logic the flutter margin can be enlarged. Analogous to what were concluded in the isolated airfoil study, for internal excitation, trailing-edge forcing was shown to be optimal since the trailing-edge receptivity still works as the dominant mechanism for generating the acoustically-induced airloads.
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Lepicovsky, J., E. R. McFarland, R. V. Chima i J. R. Wood. "On Flowfield Periodicity in the NASA Transonic Flutter Cascade: Part I — Experimental Study". W ASME Turbo Expo 2000: Power for Land, Sea, and Air. American Society of Mechanical Engineers, 2000. http://dx.doi.org/10.1115/2000-gt-0572.
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An extensive study to improve flow uniformity and periodicity in the NASA Transonic Flutter Cascade is presented here. The results are reported in two independent parts dealing with the experimental approach and the analytical approach. The first part, the Experimental Study, focuses first on the data sets acquired in this facility in the past and explains several discrepancies, particularly the questions of actual flow incidence and cascade backpressure levels. Next, available means for control and modifications of the cascade flowfield, boundary layer bleed and tailboard settings are presented in detail. This is followed by experimental data sets acquired in modified test facility configurations that were based on analytical predictions of the cascade flowfield. Finally, several important conclusions about improving the cascade flowfield uniformity and blade load periodicity are summarized. The most important conclusions are: (1) boundary layer bleed system does not improve the cascade flow periodicity; (2) carefully match the tunnel wall contours to the expected shape of cascade streamlines; (3) measure actual flow incidence for each cascade configuration rather than rely on the tunnel geometry; and (4) the current cascade configuration exhibits a very high blade load uniformity over six blades from blade #2 to blade #7, and the facility is now ready for unsteady pressure data acquisition.
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Lepicovsky, J., V. R. Capece i C. T. Ford. "Resonance Effects in the NASA Transonic Flutter Cascade Facility". W ASME Turbo Expo 2003, collocated with the 2003 International Joint Power Generation Conference. ASMEDC, 2003. http://dx.doi.org/10.1115/gt2003-38344.
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Investigations of unsteady pressure loadings on the blades of fans operating near the stall flutter boundary are carried out under simulated conditions in the NASA Transonic Flutter Cascade facility (TFC). It has been observed that for inlet Mach numbers of about 0.8, the cascade flowfield exhibits intense low-frequency pressure oscillations. The origins of these oscillations were not clear. It was speculated that this behavior was either caused by instabilities in the blade separated flow zone or that it was a tunnel resonance phenomenon. It has now been determined that the strong low-frequency oscillations, observed in the TFC facility, are not a cascade phenomenon contributing to blade flutter, but that they are solely caused by the tunnel resonance characteristics. Most likely, the self-induced oscillations originate in the system of exit duct resonators. For sure, the self-induced oscillations can be significantly suppressed for a narrow range of inlet Mach numbers by tuning one of the resonators. A considerable amount of flutter simulation data has been acquired in this facility to date, and therefore it is of interest to know how much this tunnel self-induced oscillations influences the experimental data at high subsonic Mach numbers since this facility is being used to simulate flutter in transonic fans. In short, can this body of experimental data still be used reliably to verify computer codes for blade flutter and blade life predictions? To answer this question a study on resonance effects in the NASA TFC facility was carried out. The results, based on spectral and ensemble averaging analysis of the cascade data, showed that the interaction between self-induced oscillations and forced blade motion oscillations is very weak and can generally be neglected. The forced motion data acquired with the mistuned tunnel, when strong self-induced oscillations were present, can be used as reliable forced pressure fluctuations provided that they are extracted from raw data sets by an ensemble averaging procedure.
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Chima, R. V., E. R. McFarland, J. R. Wood i J. Lepicovsky. "On Flowfield Periodicity in the NASA Transonic Flutter Cascade: Part II — Numerical Study". W ASME Turbo Expo 2000: Power for Land, Sea, and Air. American Society of Mechanical Engineers, 2000. http://dx.doi.org/10.1115/2000-gt-0573.
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The transonic flutter cascade facility at NASA Glenn Research Center was redesigned based on a combined program of experimental measurements and numerical analyses. The objectives of the redesign were to improve the periodicity of the cascade in steady operation, and to better quantify the inlet and exit flow conditions needed for CFD predictions. Part I of this paper describes the experimental measurements, which included static pressure measurements on the blade and endwalls made using both static taps and pressure sensitive paints, cobra probe measurements of the endwall boundary layers and blade wakes, and shadowgraphs of the wave structure. Part II of this paper describes three CFD codes used to analyze the facility, including a multibody panel code, a quasi-three-dimensional viscous code, and a fully three-dimensional viscous code. The measurements and analyses both showed that the operation of the cascade was heavily dependent on the configuration of the sidewalls. Four configurations of the sidewalls were studied and the results are described. For the final configuration, the quasi-three-dimensional viscous code was used to predict the location of mid-passage streamlines for a perfectly periodic cascade. By arranging the tunnel sidewalls to approximate these streamlines, side-wall interference was minimized and excellent periodicity was obtained.
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Watanabe, Toshinori, Junichi Kazawa, Seiji Uzawa i Benjamin Keim. "Numerical and Experimental Study of Active Flutter Suppression With Piezoelectric Device for Transonic Cascade". W ASME Turbo Expo 2008: Power for Land, Sea, and Air. ASMEDC, 2008. http://dx.doi.org/10.1115/gt2008-51467.
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Possibility of active suppression for transonic cascade flutter with piezoelectric device was studied both numerically and experimentally. In the numerical study, a previously proposed control method in which the blade trailing edges were actively oscillated was analyzed in detail toward realistic application by a developed numerical method with flow-structure coupling. From the results, the effect of the control was confirmed, and the suppression was revealed to come from the appropriate change in the oscillatory behavior of the passage shock. Experimental study was conducted in linear cascade wind tunnel under transonic flow condition to verify that the method realized substantial effect on stability of the blade oscillation. Unsteady aerodynamic forces induced by the active oscillation of a blade on which piezoelectric devices were glued were measured and superposed with the unsteady induced force causing flutter instability. The results showed a distinctive stabilization effect of flutter suppression in the case with appropriate phase difference between original blade vibration and the active oscillation of the piezoelectric device. The active oscillation was, however, found to generate destabilization effect if the phase was inappropriate.
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Jutur, Prahallada, i Raghuraman N. Govardhan. "Effect of Pressure Ratio on Bending Mode Flutter in a Transonic Linear Cascade". W ASME 2017 Gas Turbine India Conference. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gtindia2017-4569.
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Vibration related issues such as flutter have always been a cause of concern for aircraft engine designers. They not only incur unwarranted cost and time overruns, but also significantly compromise performance and can cause structural damage. This phenomenon has become more relevant for the modern aircraft engines, which employ relatively thin, long blade rows to satisfy ever growing demand for a powerful yet compact engine. The tip sections of such blade rows operate with supersonic relative velocity, where prediction of flutter can get challenging due to unsteady flow features like oscillating shocks and their interaction with the blade motion. Linear cascades that represent a specific radial location of the rotor have proven to be a reliable tool for flutter studies. To facilitate flutter experiments at flow Mach numbers realistic to the aircraft engine components, a transonic cascade facility operating at a Mach Number (M) of 1.3 with the ability to oscillate the central blade in the cascade has been developed. The cascade consists of 5 blades and two false blades of which the central blade is oscillated in heave, which represents the bending mode of the rotor. The typical reduced frequencies associated with this kind of flutter in practice (k ∼ 0.1) correspond to a high dimensional frequency of 200 Hz for the present case. A barrel cam mechanism is used to provide such high frequency oscillations. The parameters varied in the present study include the reduced frequency (k) and the static pressure ratio (SPR) across the cascade, which is varied with the help of tailboard and flap arrangement located at the back end of the cascade. Three SPR cases of 1.05, 1.25, and 1.35 are considered and at each of these pressure ratio cases, the reduced frequency is varied. The unsteady loads are measured on the oscillating central blade during the oscillation cycle to quantify the energy transfer from flow to blade and shadowgraphy is used to visualize the shocks. The results from these experiments indicate flutter at lower k values for all the SPR cases tested, while the higher k values are damped. The magnitude of excitation or damping at any particular frequency is also observed to increase with increasing SPR.
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Kobayashi, Hiroshi. "Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter". W ASME 1989 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1989. http://dx.doi.org/10.1115/89-gt-297.
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Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement due to blade oscillation and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting, on an oscillating blade were joined, and then the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semi-chord from 0.0375 to 0.547, 6 interblade phase angles and inlet flow velocities from subsonic to supersonic flow. The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle and inlet flow velocity was clarified including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles which caused flutter were in the range from 40° to 160° for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.
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Purushothaman, Kirubakaran, Sankar Kumar Jeyaraman, Sasikanta Parida i Kishore Prasad Deshkulkarni. "Aeroelastic Flutter Analysis of Linear Cascade Blades: STC5". W ASME 2017 Gas Turbine India Conference. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gtindia2017-4773.
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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.
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Széchényi, Edmond. "Fan Blade Flutter: Single Blade Instability or Blade to Blade Coupling?" W ASME 1985 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-gt-216.
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Different types of fan blade flutter occur at the various compressor flow regimes. Sub/transonic stall flutter and two forms of supersonic started flow flutter have been studied in a straight cascade wind tunnel. Results show clearly that these three common forms of flutter can exist as single-degree-of-freedom (single-blade instabilities). Cascade effects, though at times important, are never the only flutter mechanism: flutter limits are essentially controlled by single-blade aeroelastic coefficients, though blade-to-blade coupling arising from cascade effects can modify these limits according to the mode order. Thus, contrary to widespread practice, the fundamental approach to flutter problems should lie at least as much in the study of single blade flutter as in that of unsteady cascade effects. The two should anyhow best be considered separately when searching for a better physical insight.