Relevant bibliographies by topics / Transonic Axial Compressor

Academic literature on the topic 'Transonic Axial Compressor'

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Published: 6 September 2023

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Journal articles on the topic "Transonic Axial Compressor"

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Moumen Idres and Muhamad Adi Muqri Saiful Azmi. "Computational Prediction of the Performance Map of a Transonic Axial Flow Compressor." CFD Letters 14, no. 3 (April 1, 2022): 11–21. http://dx.doi.org/10.37934/cfdl.14.3.1121.

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Aviation fuel efficiency is an important target for aviation industry. Aircraft engine compression ratio is a key factor to improve fuel consumption. Compression ratio can be increased using transonic compressor. In this study, performance prediction of a transonic axial compressor at design and off-design operating conditions is investigated numerically using ANSYS-CFX software. The compressor is NASA Rotor 37. Firstly, the performance at design point is predicted, where mesh independence study is performed to determine suitable mesh size. Three-dimensional flow details for meridional plane, blade-to-blade plane and airfoil surface are explored. The design point study successfully captured flow features such as shock waves and flow separation regions. When compared with experimental data, the predicted compressor pressure ratio deviation error is less than 5%. 3D flow details show that shock wave strength increases from hub to tip. The shock wave moves backward as we move from hub to tip indicating that the flow separation covers lesser portion of the blade. Secondly, off-design performance is predicted for various rotational speeds. A simple procedure is utilized to predict surge and choke limits. The predicted compressor map is compared with experimental data and it shows overall root mean square error less than 5%. The success of the method developed in this research make it a viable method to be used in the design phase of transonic compressors to evaluate the effect of design modifications for both design and off-design operating conditions.
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Li, Zhihui, and Yanming Liu. "Optimization of rough transonic axial compressor." Aerospace Science and Technology 78 (July 2018): 12–25. http://dx.doi.org/10.1016/j.ast.2018.03.031.

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Meng, Bo, Zhiping Li, Haihui Wang, and Qiushi Li. "An improved wavelet adaptive logarithmic threshold denoising method for analysing pressure signals in a transonic compressor." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 229, no. 11 (September 4, 2014): 2023–30. http://dx.doi.org/10.1177/0954406214550512.

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This paper proposes an improved wavelet threshold denoising algorithm based on orthogonal wavelet transform. The algorithm, called adaptive logarithmic threshold denoising algorithm based on wavelet (ALTDAW), processes the dynamic pressure signals generated by a transonic axial compressor. Combined with an adaptive logarithmic threshold function, it sets the optimal threshold for each decomposition level. In this way, noise is effectively identified and eliminated. Since ALTDAW is adaptable, it reduces the maximum decomposition level, thereby decreasing the processing time and improving the computational efficiency. In numerical experiments, the performance of ALTDAW was compared with that of the classical soft and hard threshold algorithms. Relative to the classical algorithms, ALTDAW increased the signal-to-noise ratio (SNR) by 27.9% and 44.2%, and reduced the processing time by 38.5% and 37.6%, respectively. The practicality of the algorithm was validated on the complex dynamic pressure signals of a transonic axial compressor. When processing these signals, the algorithm depicted that disturbance passes through the tips and the spike three revolutions before the compressor stalls, consistent with the physical properties of transonic axial compressors.
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Calvert, W. J., and R. B. Ginder. "Transonic fan and compressor design." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 213, no. 5 (May 1, 1999): 419–36. http://dx.doi.org/10.1243/0954406991522671.

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Transonic fans and compressors are now widely used in gas turbine engines because of their benefits in terms of compactness and reduced weight and cost. However, careful and precise design is essential if high levels of performance are to be achieved. In this paper, the evolution of transonic compressor designs and methods is outlined, followed by more detailed descriptions of current compressor configurations and requirements and modern aerodynamic design methods and philosophies. Current procedures employ a range of methods to allow the designer to refine a new design progressively. Overall parameters, such as specific flow and mean stage loading, the axial matching between the stages at key operating conditions and the radial matching between the blade rows are set in turn, using one- and two-dimensional techniques. Finally, detailed quasi-three-dimensional and three-dimensional computational fluid dynamics (CFD) analyses are employed to refine the design. However, it is important to appreciate that the methods all have significant limitations and designers must take this into account if successful compressors are to be produced.
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Biollo, Roberto, and Ernesto Benini. "Recent advances in transonic axial compressor aerodynamics." Progress in Aerospace Sciences 56 (January 2013): 1–18. http://dx.doi.org/10.1016/j.paerosci.2012.05.002.

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Liu, Hanru, Yangang Wang, Songchuan Xian, and Wenbin Hu. "Effect of inlet distortion on the performance of axial transonic contra-rotating compressor." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 1 (September 30, 2016): 42–54. http://dx.doi.org/10.1177/0954410016670421.

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The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.
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Редин, И. И., and М. А. Шевченко. "СИСТЕМАТИЗАЦІЯ І УЗАГАЛЬНЕННЯ ТЕОРЕТИЧНИХ ТА ЕКСПЕРИМЕНТАЛЬНИХ ДАНИХ ПО ЕФЕКТИВНОСТІ НАДРОТОРНОГО ПРИСТРОЮ В ОСЬОВОМУ КОМПРЕСОРІ." Open Information and Computer Integrated Technologies, no. 87 (June 30, 2020): 181–99. http://dx.doi.org/10.32620/oikit.2020.87.11.

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The systematization of physical flow models at the peripheral region of the rotors of axial compressor is carried out. Based on the experimental and numerical studies, the flow features in subsonic and transonic rotors are analyzed. Similar features of the flow near the wall at the periphery of subsonic and transonic rotors are formulated. The characteristic areas and individual features of the near-wall flow in them, which are obtained in experimental studies of the flow structure at the periphery of the blade rows, are reflected. The analysis of the influence of annular grooves in the axial compressor case on the flow in the airfoil channel of the subsonic and transonic rotors is presented. The hypothetical mechanism of the flow effect in the cavity of the annular groove on the main flow at the tip region of the airfoil channel of axial compressor rotor is described. An approach to generalize the experimental data of the axial compressor stages with the casing treatment based on the selected fundamental system of dimensionless parameters characterizing the main features of the flow at the rotor tip region are proposed. Using the approach, the dependences of the casing treatment effect on the gas-dynamic stability limit and efficiency are obtained. It was found, when Reynolds numbers ReΔr> 400 increase, the efficiency of annular groove casing treatment in the axial compressor wall on the gas-dynamic stability limit of the compressor decreases. The existence of the region of aerodynamic efficiency modes of the annular groove casing treatment in the case is shown. In this area, there is an optimal mode when the maximum effect of efficiency from install annular groove casing treatment is achieved. The obtained generalization al-lows us, at the step of making design decisions, to evaluate the effectiveness of the annular groove casing treatment in the case when it is used in subsonic and transonic rotors of the axial compressor stages.
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Law, C. H., and A. J. Wennerstrom. "Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl." Journal of Turbomachinery 109, no. 1 (January 1, 1987): 142–48. http://dx.doi.org/10.1115/1.3262059.

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A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.
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Rabe, D. C., A. J. Wennerstrom, and W. F. O’Brien. "Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor." Journal of Turbomachinery 110, no. 3 (July 1, 1988): 386–92. http://dx.doi.org/10.1115/1.3262208.

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The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.
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Kodama, H. "Performance of Axial Compressor With Nonuniform Exit Static Pressure." Journal of Turbomachinery 108, no. 1 (July 1, 1986): 76–81. http://dx.doi.org/10.1115/1.3262027.

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An analytical model has been developed to predict the performance of axial compressors with an exit static pressure perturbation. The model uses a two-dimensional compressible semi-actuator disk model. This method can be applied to the compressor with known circumferential variation in exit static pressure which is measured or predicted by an analytical method. The analytical results are found to be in good agreement with experiments carried out on two transonic fans.
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Dissertations / Theses on the topic "Transonic Axial Compressor"

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Grossman, Bart L. "Testing and analysis of a transonic axial compressor." Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1997. http://handle.dtic.mil/100.2/ADA341113.

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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1997.
"September 1997." Thesis advisor(s): Raymond P. Shreeve. Includes bibliographical references (p. 57). Also available online.
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Sadek, Joseph. "Effect of Axial Gap Distance on Transonic Compressor Performance." Thesis, KTH, Kraft- och värmeteknologi, 2015. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-172994.

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The modern trend of gas turbines design is towards lighter, highly efficient,and more compact engines. Such situation imposes on engineers to continuouslysearch for improved and optimum designs. The thesis presented aims at researching possible performance improvements regarding axial gapdistance in transonic compressors. Decreasing the axial gap would result inlighter engines and achieve design goals. The influence of decreasing the axialgap on performance and structure integrity should be throughly analyzed. This thesis work includes numerical investigations on the axial gap distance effect on performance efficiency and related unsteady aerodynamics phenomena. The first one and a half compressor stages of a Siemens Gas Turbine are modeled in ANSYS CFX. Different axial gap models are simulated for differentconfigurations. The steady state solution is obtained to be initialized for transient time marching calculations. Furthermore, the computational cost of transient calculations is reduced through a geometry scaling technique. The unsteady behavior is further analyzed by a Harmonic Balance solver implemented in STAR-CCM+ software, and compared to a reference case transient calculations. The results obtained supports the presence of an optimalaxial gap distance for maximum efficiency in transonic compressors. Further, the harmonic balance method shows good possibilities for cost and time reductions in transonic compressors performance calculations.
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Reid, William D. "Transonic axial compressor design case study and preparations for testing." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1995. http://handle.dtic.mil/100.2/ADA306259.

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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1995.
Thesis advisor(s): Raymond P. Shreeve. "September 1995." Includes bibliographical references. Also available online.
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Cahill, Joseph E. "Identification and Evaluation of Loss and Deviation Models for use in Transonic Compressor Stage Performance Prediction." Thesis, Virginia Tech, 1997. http://hdl.handle.net/10919/37041.

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The correlation of cascade experimental data is one method for obtaining compressor stage characteristics. These correlations specify pressure loss and flow turning caused by the blades. Current open literature correlations used in streamline curvature codes are inadequate for general application to high-speed transonic axial-flow compressors. The objective of this research was to investigate and evaluate the available correlations and ultimately discover sets of correlations which best fit the empirical data to be used in streamline curvature codes. Correlations were evaluated against experimental data from NASA Rotor 1-B and NASA Stage 35. It was found that no universal set of correlations was valid for minimum-loss point predictions. The Bloch shock loss model showed promising results in the stall regime for supersonic relative inlet Mach numbers. The Hearsey and Creveling off-minimum-loss deviation angle prediction performed consistently better than all other correlations tested.
Master of Science
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Drayton, Scott. "Design, test, and evaluation of a transonic axial compressor rotor with splitter blades." Thesis, Monterey, California: Naval Postgraduate School, 2013. http://hdl.handle.net/10945/37616.

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Approved for public release; distribution is unlimited
A new design procedure was developed and documented that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected at the NPS TPL utilizing the Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.
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Weigl, Harald Jürgen. "Active stabilization of rotating stall and surge in a transonic single stage axial compressor." Thesis, Massachusetts Institute of Technology, 1997. http://hdl.handle.net/1721.1/10353.

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Ryman, John Franklin. "Prediction of Inlet Distortion Transfer Through the Blade Rows in a Transonic Axial Compressor." Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/43207.

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Inlet total pressure non-uniformities in axial flow fans and compressors can contribute to the loss of component structural integrity through high cycle fatigue (HCF) induced by the excitation of blade vibratory modes. As previous research has shown total pressure distortion to be the dominant HCF driver in aero engines [Manwaring et al, 1997], an understanding of its transfer through, and impact on, subsequent turbomachine stages and engine components is an important topic for assessment. Since current modeling techniques allow for total pressure distortion magnitudes to be directly related to blade vibratory response, the prediction of downstream distortion patterns from an upstream measurement would allow for the inference of the vibratory response of downstream blade rows to an inlet total pressure distortion. Nonlinear Volterra theory can be used to model any periodic nonlinear system as an infinite sum of multidimensional convolution integrals. A semi-empirical model has been developed using this theory by assuming that a distortion waveform is a periodic signal that is being presented to a nonlinear system, the compressor being the system. The use of Volterra theory in nonlinear system modeling relies on the proper identification of the Volterra kernels, which make up the transfer function that defines the systemâ s impulse response characteristics. Once the kernels of a system are properly identified, the systemâ s response can be calculated for any arbitrary input. This model extracts these kernels from upstream and downstream total pressure distortion measurements of a transonic rotor of modern design. The resulting transfer function is then applied to predict distortion transfer at new operating points on the same rotor and compared with the measured data. The judicious choice of distortion measurement data allows predictions of the downstream distortion content based on a measured non-uniform inlet flow at conditions different from those at which the transfer function was derived. This allows for the determination of downstream total pressure distortion that has the potential to excite blade vibratory modes that could lead to HCF under operating conditions other than those at which the data was taken, such as varying inlet distortion patterns, mass flow settings, rotational speeds, and inlet geometry. This report presents the creation of a Volterra model in order to predict distortion transfer in axial flow fans and compressors. This model, in three variations, is applied to a variety of distortions and compressor operating conditions as measured in the ADLARF tests at the Compressor Research Facility. Predictions are compared with data from the test and final results are also compared with two previous studies conducted at Virginia Tech using the same experimental data. Using the Volterra model it is shown that, with appropriate limitations, distortion transfer can be predicted for flow conditions different from those used for calibration. The model is considered useful for both performance and HCF investigations.
Master of Science
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Jones, James A. "A multidisciplinary algorithm for the 3-D design optimization of transonic axial compressor blades." Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02Jun%5FJones%5FJames.pdf.

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Thesis (Ph. D. in Aeronautics and Astronautics)--Naval Postgraduate School, June 2002.
Dissertation supervisor: Raymond P. Shreeve. Includes bibliographical references (p. 157-161). Also available online.
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Heinlein, Gregory S. "Aerodynamic Behavior of Axial Flow Turbomachinery Operating in Transient Transonic Flow Regimes." The Ohio State University, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=osu1573149943024303.

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Figueiredo, Josué Sanches. "Determination of stall and choke limits of a transonic axial flow compressor using the straemline curvature method." Instituto Tecnológico de Aeronáutica, 2010. http://www.bd.bibl.ita.br/tde_busca/arquivo.php?codArquivo=2090.

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The necessity for high performance axial compressors for various applications is shown. The basic theory associated with the study of these machines is presented. A method for calculating the properties of the flow along the compressor is presented. Depending on the design requirements the number of stages is selected and an appropriate shape of the channel studied. Then the compressor blading is made, row-by-row, from the calculation of the flow using a non-viscous flow associated with a loss model that allows, from the known flow at the leading edge, calculate the flow at the trailing edge. The calculations are made on streamlines initially positioned by the criterion of the same area and then repositioned according to the flow calculation in each row until the streamlines do not change their position anymore. Then the boundary-layer on hub and casing walls is evaluated, calculating the blockage coefficients, initially arbitrated. An iterative procedure is done until the blockage coefficients no longer vary. At the end, there are all the dimensions of the compressor and the properties of the flow on the streamlines at the leading and trailing edges of each row. With the geometry fixed and varying the inlet mass flow the maximum mass flow that would result in choke and the minimum one that would result in stall, or even surge, of the compressor can be determined. The computer program was written aiming at being used in a design optimization research.
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Books on the topic "Transonic Axial Compressor"

1

United States. National Aeronautics and Space Administration., ed. Blockage development in a transonic, axial compressor rotor. [Washington, D.C: National Aeronautics and Space Administration, 1997.

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United States. National Aeronautics and Space Administration., ed. Blockage development in a transonic, axial compressor rotor. [Washington, D.C: National Aeronautics and Space Administration, 1997.

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Grossman, Bart L. Testing and analysis of a transonic axial compressor. Monterey, Calif: Naval Postgraduate School, 1997.

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Neuhoff, F. Modifications to the inlet flow field of a transonic compressor rotor. Monterey, Calif: Naval Postgraduate School, 1985.

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United States. National Aeronautics and Space Administration., ed. Experimental determination of aerodynamic damping in a three-syage transonic axial-flow compressor. [Washington, DC: National Aeronautics and Space Administration, 1988.

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Newman, Frederick A. Experimental vibration damping characteristics of the third-stage rotor of a three-stage transonic axial-flow compressor. [Washington, DC]: National Aeronautics and Space Administration, 1988.

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L, Suder Kenneth, and United States. National Aeronautics and Space Administration., eds. The effect of adding roughness and thickness to a transonic axial compressor rotor. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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L, Celestina Mark, and United States. National Aeronautics and Space Administration., eds. Experimental and computational investigation of the tip clearance flow in a transonic axial compressor rotor. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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Chung-hua, Wu. A general theory of two-and three-dimensional rotational flow in subsonic and transonic turbomachines. [Washington, D.C.]: National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1993.

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Chung-hua, Wu. A general theory of two-and three-dimensional rotational flow in subsonic and transonic turbomachines. [Washington, D.C.]: National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1993.

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Book chapters on the topic "Transonic Axial Compressor"

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Barthmes, Sebastian, Jakob P. Haug, Andreas Lesser, and Reinhard Niehuis. "Unsteady CFD Simulation of Transonic Axial Compressor Stages with Distorted Inflow." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 303–21. Cham: Springer International Publishing, 2015. http://dx.doi.org/10.1007/978-3-319-21127-5_18.

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Iseler, Jens, Andreas Lesser, and Reinhard Niehuis. "Numerical Investigation of a Transonic Axial Compressor Stage with Inlet Distortions." In High Performance Computing in Science and Engineering, Garching/Munich 2009, 185–95. Berlin, Heidelberg: Springer Berlin Heidelberg, 2010. http://dx.doi.org/10.1007/978-3-642-13872-0_16.

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Nha, Tuong-Linh, Van-Hoang Nguyen, Xuan-Truong Le, and Cong-Truong Dinh. "Aerodynamic Performance of Single-Stage Transonic Axial Compressor with Multi-Bleed Airflow." In Lecture Notes in Mechanical Engineering, 109–16. Cham: Springer Nature Switzerland, 2023. http://dx.doi.org/10.1007/978-3-031-31824-5_14.

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Yu, S., G. H. Schnerr, U. Dohrmann, and O. Sadi. "Passive control of shock-boundary layer interaction in transonic axial compressor cascade flow." In Fluid- and Gasdynamics, 207–17. Vienna: Springer Vienna, 1994. http://dx.doi.org/10.1007/978-3-7091-9310-5_24.

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Huang, N. Z., X. Zhao, and Y. H. Zhang. "The Swept and Leaned Blade Influence on the Aerodynamic Performance of a Transonic Axial Compressor Rotor." In Lecture Notes in Electrical Engineering, 227–35. Singapore: Springer Singapore, 2019. http://dx.doi.org/10.1007/978-981-13-3305-7_18.

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Zhao, Xuning, Xuhui Zhou, Jinxin Cheng, and Jiang Chen. "Numerical Investigation on Arbitrary Polynomial Blade Model for a Transonic Axial-Flow Compressor Rotor with Multi-parameter Optimization." In Lecture Notes in Electrical Engineering, 19–33. Singapore: Springer Singapore, 2019. http://dx.doi.org/10.1007/978-981-13-3305-7_2.

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Hofmann, Willy, and Josef Ballmann. "Tip-Vortices in Transonic Axial-Compressors." In New Results in Numerical and Experimental Fluid Mechanics III, 341–49. Berlin, Heidelberg: Springer Berlin Heidelberg, 2002. http://dx.doi.org/10.1007/978-3-540-45466-3_41.

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Pitroda, Darshan P., Dilipkumar Bhanudasji Alone, and Harish S. Choksi. "Understanding of an Effect of Plenum Volume of a Low Porosity Bend Skewed Casing Treatment on the Performance of Single-Stage Transonic Axial Flow Compressor." In Proceedings of the National Aerospace Propulsion Conference, 3–26. Singapore: Springer Singapore, 2020. http://dx.doi.org/10.1007/978-981-15-5039-3_1.

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Du, W. H., H. Wu, and L. Zhang. "Off-design Performance Analysis of Multi-Stage Transonic Axial Compressors." In New Trends in Fluid Mechanics Research, 504. Berlin, Heidelberg: Springer Berlin Heidelberg, 2007. http://dx.doi.org/10.1007/978-3-540-75995-9_167.

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Dalbanjan, Manjunath S., and Niranjan Sarangi. "Sensitivity Study of Stagger Angle on the Aerodynamic Performance of Transonic Axial Flow Compressors." In Proceedings of the National Aerospace Propulsion Conference, 3–14. Singapore: Springer Nature Singapore, 2022. http://dx.doi.org/10.1007/978-981-19-2378-4_1.

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Conference papers on the topic "Transonic Axial Compressor"

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Suder, Kenneth L. "Blockage Development in a Transonic, Axial Compressor Rotor." In ASME 1997 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1997. http://dx.doi.org/10.1115/97-gt-394.

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A detailed experimental investigation to understand and quantify the development of blockage in the flow field of a transonic, axial flow compressor rotor (NASA Rotor 37) has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. The impact of the shock on the blockage development, pertaining to both the shock / boundary layer interactions and the shock / tip clearance flow interactions, is discussed. The results indicate that for this rotor the blockage in the endwall region is 2–3 times that of the core flow region, and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer.
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Kumar, S. Satish, Ranjan Ganguli, S. B. Kandagal, and Soumendu Jana. "Flow Behavior in a Transonic Axial Compressor Stage." In ASME 2015 Gas Turbine India Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gtindia2015-1231.

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The steady and unsteady flow characteristics typically vary along and across the axial compressor stage. This coupled with asymmetric rotor tip clearance that occurs in practice makes flow even more complex. Understanding the complex flow behavior inside the transonic compressor stage will aid in developing flow control devices that are meant for purposes such as improving the rotating stall margin, flutter margin, etc. Here, a detailed time averaged numerical analysis is performed on the single stage transonic axial compressor with averaged rotor tip clearance (1.75% of rotor tip axial chord). An attempt is made to study the compressor stall phenomenon. Computational Fluid Dynamics (CFD) helps in resolving the complex flow features involved in a turbomachinery component and at transonic Mach numbers fairly well. Commercial tool ANSYS CFX is used for solving the 3D compressible Reynolds Averaged Navier-Stokes (RANS) equation with Shear Stress Transport (SST) turbulence model. Grid independency is carried out for three different mesh size models. All mesh models chosen have fine mesh near wall boundary regions to capture the boundary layer effects. Overall performance maps of the compressor are generated for 50% to 100% rated design speeds in steps of 10% for the chosen optimum grid. Flow variations along the blade annulus are studied for three different operating conditions: choke/free flow, peak efficiency and near stall flow conditions and for different speeds. Flow parameters such as Mach number, static and total pressure variations, etc. are studied at the inlet to rotor, exit to rotor and exit to stator for the various flow conditions and speeds. The boundary layer growth is clearly captured when the flow is throttled from choke/free flow conditions to near stall condition for all the speeds investigated. Mach number variation along blade height clearly shows decrease in Mach number as stall is approached. Blade loading distribution of the rotor at hub, mean and tip sections are clearly captured. Shock motion from around mid-chord region at free flow condition to towards the leading edge at near stall condition is clearly highlighted. Velocity streamlines near the tip section show the complex interaction of the tip leakage and clearance flows. Velocity vectors near the blade tip shows, the backflow near the trailing edge and tendency for leading edge spillage as the back pressure is increased. The flow blockage region is captured in the meridional plot and the motion of vortex core region as stall is approached is demarcated in the r-θ plots. Tangential velocity variation across the annulus for the two flow conditions investigated shows stall initiating from the tip section of the blade as compressor is throttled. Flow compensation at near stall conditions is explained.
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Siller, Ulrich, Christian Voß, and Eberhard Nicke. "Automated Multidisciplinary Optimization of a Transonic Axial Compressor." In 47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2009. http://dx.doi.org/10.2514/6.2009-863.

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Li, Tao, Yadong Wu, Hua Ouyang, and Xiaoqing Qiang. "Axial Compressor Performance Prediction Using Improved Streamline Curvature Approach." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75450.

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This paper presents in detail the improved streamline curvature approach (SLC) to the performance evaluation and internal flow field calculation of subsonic and transonic axial compressors. Based on previous research, the diverse incidence, deviation and total pressure loss models, generally existing in the form of fitting curves and semi-empirical correlations, are discussed respectively. Typically, transonic flow in axial compressor results in the variation of several flow parameters and particularly the appearance of shock waves compared with subsonic flow. In this paper, the revision and improvement of loss models are applied to reach higher accuracy, especially considering the loss component due to actual incidence angle. Several modifications have been made as well considering the influence of three-dimensional flow. For the purpose of validating this approach, two test cases, including a single-stage transonic axial compressor NASA Stage37 and a 3-stage subsonic axial compressor P&W 3S1, are calculated. The overall characteristics and spanwise aerodynamic parameters for blade rows are demonstrated at both design and off-design conditions. Furthermore, the results agree well with both experimental data and computational fluid dynamic (CFD) results. This throughflow method is verified as an applicable and convenient tool for aerodynamic analysis and performance prediction of subsonic and transonic axial compressors.
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Law, C. Herbert, and Arthur J. Wennerstrom. "Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl." In ASME 1986 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1986. http://dx.doi.org/10.1115/86-gt-33.

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A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performances of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.
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Kaur, Baljeet, Nitin B. Balsaraf, and Ajay Pratap. "A Study of Existing Multistage Transonic Axial Compressor Design for Surge Margin Improvement." In ASME 2013 Gas Turbine India Conference. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gtindia2013-3616.

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The design of multistage axial flow compressors has been revolutionised in recent years by the development of three dimensional multistage viscous calculations (CFD) and the availability of the computational power to allow these methods to be used extensively in design process. Such a multistage turbomachinery was used to redesign the existing three stage transonic compressor for improved aerodynamic performance in terms of SM limit. The redesign activity of compressor configuration was carried out as surge margin obtained with hardware testing of existing machinery was not sufficient to meet the desired design goals. The higher limit of surge margin in accordance with design specification is required to maintain the successful and stable operation of aircraft engine. As at stall point of compressor, aerodynamic instabilities would be initiated resulting surge or rotating stall which potentially leading to a complete mechanical failure of the compression system as well as of the whole engine. Maximizing the SM of multistage compressors is particularly a complex process especially alongwith achieving higher efficiency. Outlined in this paper are the details of how advanced design techniques were incorporated using traditional 2D and CFD methods into redesign activities for compressor performance improvement. The approach used in this work was to modify compressor annulus flowpath and rotor and stator blade geometries based on output of 2D calculations and 3D N-S analysis for SM enhancement of existing design. While carrying out redesign activities for compressor, the constraints of retaining existing inlet and outlet flow area, axial length as well as design parameters i.e. inlet mass flow, rotational speed and operating point pressure ratio were taken care. The advanced design techniques like 3D blading, wide chord blade performance and many others were studied in detailed manner for incorporating them into compressor redesign procedure. Simultaneous resigned compressor configuration based on in-house design procedure showed improvement of SM by 17% at design speed with maintaining mentioned design constraints. Subsequently, the detailed analysis was also performed at off design speeds to have satisfactory performance.
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Zhang, Xiawen, Yaping Ju, Chuhua Zhang, Cong Li, and Dejun Meng. "Geometry Scaling Technique For Aerodynamic Redesign Of Multistage Transonic Compressor." In GPPS Xi'an21. GPPS, 2022. http://dx.doi.org/10.33737/gpps21-tc-240.

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The aerodynamic design of multi-stage transonic axial-flow compressors plays a critical role in the overall performance of aeronautical engine, gas turbine and industry compression unit. Despite a series of new design methods have been successfully proposed in the research and development of advanced axial-flow compressors, redesign methods are still one of the most effective measures in the development of available advanced products to meet new requirements. In this work, a geometry scaling technique is proposed in which 1D mean-line analysis and 3D parametric geometric modeling are used to define a series of key redesign criteria with as minor variations as possible in the aerodynamic performance compared against the original design. The motivation behind this technique is to develop a CFD validation rig while keep both the aerodynamic performance and CFD prediction accuracy unchanged with reduced partial annulus model. The proposed scaling technique is then verified by the redesign of a 3.5-stage transonic axial-flow compressor with geometry scaling of blade/vane numbers and IGV-rotor-stator axial spacing. The CFD prediction shows that the variation in overall performance of redesigned compressors is generally within one percent, verifying the effectiveness of the proposed scaling technique.
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Manfredi, Marco, Cedric Babin, and Fabrizio Fontaneto. "Transonic Axial Compressors Loss Correlations: Part II — Loss Models Validation." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-16131.

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Abstract The quest for greener, more efficient aircraft engines is the main driver for the development of innovative compression system designs. Reduced order design tools rely nevertheless on semi-empirical loss models, whose validity range is often not net or in general not verified. The present work aims at defining a set of loss correlations, which could readily be employed in the analysis and design process of modern transonic axial compressors. In Part I, various loss correlations were deeply described and, in some cases, updated to enhance both their generality and their prediction capability. In Part II, the effectiveness of both original and updated models will be tested for one specific low aspect ratio axial compressor stage. Experimental and numerical data will be used at such extent.
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Shahpar, Shahrokh, Andrey Polynkin, and Vassili Toropov. "Large Scale Optimization of Transonic Axial Compressor Rotor Blades." In 49th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference
16th AIAA/ASME/AHS Adaptive Structures Conference
10t. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2008. http://dx.doi.org/10.2514/6.2008-2056.

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Li, Qiushi, Tianyu Pan, Zhiping Li, Tailu Sun, and Yifang Gong. "Experimental Study of Compressor Instability Inception in a Transonic Axial Flow Compressor." In ASME Turbo Expo 2014: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gt2014-25190.

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An experimental investigation is conducted to study the details of instability inception in a transonic axial flow compressor. The experimental results indicate that the compressor instability is initiated through the development of a low-frequency axisymmetric disturbance. The frequency of this disturbance is approximately the Helmholtz frequency of the test facility. The low-frequency disturbance can be detected over 3000 rotor revolutions before the compressor becomes unstable. Further experimental investigations illustrate that this low-frequency axisymmetric disturbance is initiated at the hub region of the compressor. This new kind of instability inception is termed “partial surge.” Examination of the design parameters of the compressor indicates that a high diffusion factor in the rotor root region might be the cause of the partial surge type instability inception.
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Reports on the topic "Transonic Axial Compressor"

1

Chen, Jen-Ping, Michael D. Hathaway, and Gregory P. Herrick. Prestall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation. Fort Belvoir, VA: Defense Technical Information Center, October 2008. http://dx.doi.org/10.21236/ada500494.

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