Статті в журналах: "Transonic Axial Compressor"

1

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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3

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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4

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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5

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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6

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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7

Редин, И. И., та М. А. Шевченко. "СИСТЕМАТИЗАЦІЯ І УЗАГАЛЬНЕННЯ ТЕОРЕТИЧНИХ ТА ЕКСПЕРИМЕНТАЛЬНИХ ДАНИХ ПО ЕФЕКТИВНОСТІ НАДРОТОРНОГО ПРИСТРОЮ В ОСЬОВОМУ КОМПРЕСОРІ". Open Information and Computer Integrated Technologies, № 87 (30 червня 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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8

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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9

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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10

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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11

Smith,, Leroy H. "Axial Compressor Aerodesign Evolution at General Electric." Journal of Turbomachinery 124, no. 3 (July 1, 2002): 321–30. http://dx.doi.org/10.1115/1.1486219.

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This paper traces the origins of the GE Design System and how it has evolved from early methods to underlie and supplement present CFD methods, which are not themselves discussed herein. The two main elements of the detailed aero design process are vector diagram establishment and airfoil design. Their evolution is examined, and examples of how they were used to design some early GE compressors of interest are given. By the late 1950s, some transonic airfoil shapes were being custom tailored using internal blade station data from more complete radial equilibrium solutions. In the 1960s, rules for shaping transonic passages were established, and by the 1970s, custom tailoring was done for subsonic blading as well. The preliminary design layout process for a new compressor is described. It involves selecting an annulus shape and blading overall proportions that will allow a successful detailed design to follow. This requires establishment of stage loading limits that permit stall-free operation, and an efficiency potential prediction method for state-of-the-art blading. As design methods evolved, the newer approaches were calibrated with data-match experience, a process that is expected to always be needed.

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12

Suder, K. L. "Blockage Development in a Transonic, Axial Compressor Rotor." Journal of Turbomachinery 120, no. 3 (July 1, 1998): 465–76. http://dx.doi.org/10.1115/1.2841741.

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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 percent 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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13

Ahmad, Naseem, Qun Zheng, Hamza Fawzy, Tasneem Yaqoob, Salman Abdu Ahmad, and Jiang Bin. "Axial transonic compressor performance enhancement with circumferential grooves." Mechanical Sciences 11, no. 1 (May 12, 2020): 153–61. http://dx.doi.org/10.5194/ms-11-153-2020.

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Abstract. Casing treatment has been broadly used as a prudent passive flow control technique to improve the stall margin with a small drop in efficiency. The effect of grooves in certain details is overlooked however, different grooves shape (angels and location) has a remarkable effect on the controlling impact. In the current research work, an investigation on the effect of fillet and chamfer corners of rectangular circumferential grooves with various tip gap height on the performance of casing treatment is carried out with the help of CFD simulation. The performance of different models of grooves with various tip gap height on NASA rotor 37 is investigated by discretizing 3D RANS equations based on finite volume technique. Rectangular circumferential grooves casing treatment (CGCT) profile and smooth wall casing performances are evaluated. Moreover, the adiabatic efficiencies and the stall margins of smooth wall casing, rectangular grooves and rectangular grooves with fillet and chamfer corners are compared to assess the impacts of profiles of grooves on the stability and performance of axial flow compressor with different tip gaps. The stall margin of models 1–6 increased by 4.39 %, 2.52 %, 2.16 %, 1.75 %, 1.69 % and 2.06 % respectively. While the adiabatic efficiency of the models 1–6 decreased by 0.9 %, 1.01 %, 1.08 %, 1.12 %, 1.22 % and 1.16 % respectively.

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14

Manwell, A. R. "Unsteady transonic flows in an axial flow compressor." Quarterly of Applied Mathematics 43, no. 3 (October 1, 1985): 369–83. http://dx.doi.org/10.1090/qam/814234.

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15

Beheshti, Behnam H., Joao A. Teixeira, Paul C. Ivey, Kaveh Ghorbanian, and Bijan Farhanieh. "Parametric Study of Tip Clearance—Casing Treatment on Performance and Stability of a Transonic Axial Compressor." Journal of Turbomachinery 126, no. 4 (October 1, 2004): 527–35. http://dx.doi.org/10.1115/1.1791643.

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The control of tip leakage flow through the clearance gap between the moving and stationary components of rotating machines is still a high-leverage area for improvement of stability and performance of aircraft engines. Losses in the form of flow separation, stall, and reduced rotor work efficiency are results of the tip leakage vortex (TLV) generated by interaction of the main flow and the tip leakage jet induced by the blade pressure difference. The effects are more detrimental in transonic compressors due to the interaction of shock TLV. It has been previously shown that the use of slots and grooves in the casing over tip of the compressor blades, known as casing treatment, can substantially increase the stable flow range and therefore the safety of the system but generally with some efficiency penalties. This paper presents a numerical parametric study of tip clearance coupled with casing treatment for a transonic axial-flow compressor NASA Rotor 37. Compressor characteristics have been compared to the experimental results for smooth casing with a 0.356 mm tip clearance and show fairly good agreement. Casing treatments were found to be an effective means of reducing the negative effects of tip gap flow and vortex, resulting in improved performance and stability. The present work provides guidelines for improvement of steady-state performance of the transonic axial-flow compressors and improvement of the stable operating range of the system.

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16

Lu¨, Pan-Ming, and Chung-Hua Wu. "Computation of Potential Flow on S2 Stream Surface for a Transonic Axial-Flow Compressor Rotor." Journal of Engineering for Gas Turbines and Power 107, no. 2 (April 1, 1985): 323–28. http://dx.doi.org/10.1115/1.3239722.

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A set of conservative full potential function equations governing the fluid flow along a given S2 streamsurface in a transonic axial compressor rotor was obtained. By the use of artificial density and a potential function/density iteration, this set of equations can be solved, and the passage shock on the S2 streamsurface can be captured. A computer program for this analysis problem has been developed and used to compute the flow field along a mean S2 streamsurface in the DFVLR transonic axial compressor rotor. A comparison of computed results with DFVLR L2F measurement at 100 percent design speed shows fairly good agreement.

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17

Lohmberg, A., M. Casey, and S. Ammann. "Transonic radial compressor inlet design." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 217, no. 4 (January 1, 2003): 367–74. http://dx.doi.org/10.1243/095765003322315423.

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The design of radial compressor inlets for transonic flow is examined. A theoretical model [1] quantifies the losses in the tip sections caused by the choke margin (incidence) and the blockage of the blades. It identifies clear design rules for the tip sections: to achieve the highest efficiency, these require minimum blockage (low blade thickness and splitter vanes) and low choke margin (close to the unique-incidence condition). Simulations of the NASA rotor 37 transonic axial compressor (with CFX-TASCflow) are used to validate the use of three-dimensional viscous computational fluid dynamics (CFD) for transonic compressor inlets and to demonstrate that the key performance features suggested by the simple model are also modelled in three-dimensional viscous flow simulations. The simple model together with CFD simulations has been used for the design of tip sections at the inlet of a transonic radial compressor. CFD simulations were used to select the position of the shock to give a low choke margin, to reduce the preshock Mach number and also to optimize the shape and position of the leading edge of the splitter vanes.

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18

Abbasi, Sarallah, and Maryam Alizadeh. "Flow behavior in a contra-rotating transonic axial compressor." Journal of Mechanical Engineering and Sciences 15, no. 3 (September 19, 2021): 8440–49. http://dx.doi.org/10.15282/jmes.15.3.2021.20.0664.

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This study investigated a three-dimensional flow analysis on a two-stage contra-rotating axial compressor using the Navier–Stokes, continuity, and energy equations with Ansys CFX commercial software. In order to validate the obtained results, the absolute and relative flow angles curves for each rotor in radial direction were extracted and compared with the other investigation results, indicating good agreement. The compressor efficiency curve also was extracted by varying the compressor pressure ratio and compressor efficiency against mass flow rate. The flow results revealed that further distortion of the flow structure in the second rotor imposed a greater increase in the amount of entropy, especially at near-stall conditions. The increase of entropy in the second rotor is due to the interference of the tip leakage flow with the main flow which consequently caused more drops in the second rotor, suggesting that more efficacy of flow control methods occurred in the second rotor than in the first rotor.

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19

Rechter, H., W. Steinert, and K. Lehmann. "Comparison of Controlled Diffusion Airfoils With Conventional NACA 65 Airfoils Developed for Stator Blade Application in a Multistage Axial Compressor." Journal of Engineering for Gas Turbines and Power 107, no. 2 (April 1, 1985): 494–98. http://dx.doi.org/10.1115/1.3239758.

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In their transonic cascade wind tunnel, DFVLR has done measurements on a conventional NACA 65, as well as on a controlled diffusion airfoil, designed for the same velocity triangle at supercritical inlet condition. These tested cascades represent the first stator hub section of a three-stage axial/one-stage radial combined compressor developed by MTU with the financial aid of the German Ministry of Research and Technology. One aspect of this project was the verification of the controlled diffusion concept for axial compressor blade design, in order to demonstrate the capabilities of some recent research results which are now available for industrial application. The stator blades of the axial compressor section were first designed using NACA 65 airfoils. In the second step, the controlled diffusion technique was applied for building a new stator set. Both stator configurations were tested in the MTU compressor test facility. Cascade and compressor tests revealed the superiority of the controlled diffusion airfoils for axial compressors. In comparison to the conventional NACA blades, the new blades obtained a higher efficiency. Furthermore, a closer matching of the compressor performance data to the design requirements was possible due to a more precise prediction of the turning angle.

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20

Park, Kun, In Jung, Sung You, Seung Lee, Ali Zamiri, and Jin Chung. "Influences of the Flow Cut and Axial Lift of the Impeller on the Aerodynamic Performance of a Transonic Centrifugal Compressor." Energies 12, no. 23 (November 27, 2019): 4503. http://dx.doi.org/10.3390/en12234503.

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In this study, the influences of the flow cut and axial lift of the impeller on the aerodynamic performance of a transonic centrifugal compressor were analyzed. The flow cut is a method to reduce the flow rate by decreasing the impeller passage height. The axial lift is a method of increasing the impeller passage height in the axial direction, which increases the impeller exit width (B2) and increases the total pressure. A NASA CC3 transonic centrifugal compressor with a backswept angle was used as a base compressor. After applying the flow cut, the total pressure at the target flow rate was lower than the total pressure at the design point due to the increase in the relative velocity at the impeller exit. After applying the axial lift, the total pressure at the design flow rate was increased, which was caused by the reduction in the relative velocity as the passage area at the impeller exit was increased. By applying the flow cut and axial lift methods, it was shown that the variation in relative velocity at the impeller exit has a significant effect on the variation in total pressure. In addition, it was found that the relative velocity at the impeller exit of the target flow rate is maintained similar to the base impeller when the flow cut and the axial lift are combined. Therefore, by combining the flow cut and the axial lift, three transonic centrifugal impellers with flow fractions of 0.7, 0.8, and 0.9 compared to the design flow rate were newly designed.

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21

Dinh, Cong-Τruong, and Kwang-Υong Kim. "Effects of Air Injection on Aerodynamic Performance of a Single-Stage Transonic Axial Compressor." DESIGN, CONSTRUCTION, MAINTENANCE 1 (March 8, 2021): 24–32. http://dx.doi.org/10.37394/232022.2021.1.4.

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This paper presents a parametric study of stator air injection on aerodynamic performances of a single-stage transonic axial compressor, NASA Stage 37, using three-dimensional Reynolds-averaged Navier- Stokes equations with the k-ε turbulence model and scalable wall functions. The curvature, location and width of the stator injector and the mass flow rate of air injection were tested to find their effects on the aerodynamic performances, such as total pressure ratio, peak adiabatic efficiency, stall margin and stable range extension. The numerical results for adiabatic efficiency and total pressure ratio were validated with experimental data for smooth casing. The results of parametric study show that the aerodynamic performances of the single-stage transonic axial compressor improve greatly the peak adiabatic efficiency compared to the compressor with smooth casing.

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Zhang, Yingying, and Shijie Zhang. "Performance prediction of transonic axial multistage compressor based on one-dimensional meanline method." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 235, no. 6 (February 25, 2021): 1355–69. http://dx.doi.org/10.1177/0957650921998819.

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This study proposes a 1D meanline program for the modeling of modern transonic axial multistage compressors. In this method, an improved blockage factor model is proposed. Work-done factor that varies with the compressor performance conditions is added in this program, and at the same time a notional blockage factor is kept. The coefficient of deviation angle model is tuned according to experimental data. In addition, two surge methods that originated from different sources are chosen to add in and compare with the new method called mass flow separation method. The salient issues presented here deal first with the construction of the compressor program. Three well-documented National Aerodynamics and Space Administration (NASA) axial transonic compressors are calculated, and the speedlines and aerodynamic parameters are compared with the experimental data to verify the reliability and robustness of the proposed method. Results show that consistent agreement can be obtained with such a performance prediction program. It was also apparent that the two common methods of surge prediction, which rely upon either stage or overall characteristic gradients, gave less agreement than the method called mass flow separation method.

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23

Paulon, J., Zhifang Zhang, Pingfang Jia, and Jingfei Meng. "Influence of Unsteady Effects on the Measurements in a Transonic Axial Compressor." Journal of Turbomachinery 114, no. 3 (July 1, 1992): 510–16. http://dx.doi.org/10.1115/1.2929174.

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Interaction phenomena between rotor and stator are unavoidable in advanced compressors and their effects increase with the performance of the turbomachines. Until now, it was not possible to quantify the interaction effects, but with the development of three-dimensional unsteady computation codes in a complete stage, it is possible to know, in detail, the flow field through the machine and to make evident and to explain the difficulties encountered in measuring the flow parameters. A study has been conducted in this way at ONERA, on an axial transonic compressor stage. The computations have been made with a simulation of the losses; in this manner, the overall computed and measured performances of the compressor are the same. A detailed analysis of the unsteady computation results makes evident, between rotor and stator, large variations of some parameters of the flow as a function of time, but also as a function of the axial and tangential relative position of steady probes and stator blades. Unsteady measurements made on another transonic machine confirm the indications given by these computations.

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Hembera, M., H. P. Kau, and E. Johann. "Simulation of Casing Treatments of a Transonic Compressor Stage." International Journal of Rotating Machinery 2008 (2008): 1–10. http://dx.doi.org/10.1155/2008/657202.

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This article presents the study of casing treatments on an axial compressor stage for improving stability and enhancing stall margin. So far, many simulations of casing treatments on single rotor or rotor-stator configurations were performed. But as the application of casing treatments in engines will be in a multistage compressor, in this study, the axial slots are applied to a typical transonic first stage of a high-pressure 4.5-stage compressor including an upstream IGV, rotor, and stator. The unsteady simulations are performed with a three-dimensional time accurate Favre-averaged Navier-stokes flow solver. In order to resolve all important flow mechanisms appearing through the use of casing treatments, a computational multiblock grid consisting of approximately 2.4 million nodes was used for the simulations. The configurations include axial slots in 4 different variations with an axial extension ranging into the blade passage of the IGV. Their shape is semicircular with no inclination in circumferential direction. The simulations proved the effectiveness of casing treatments with an upstream stator. However, the results also showed that the slots have to be carefully positioned relative to the stator location.

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Kim, Dae Woong, Jin Hyuk Kim, and Kwang Yong Kim. "Parametric Study on Aerodynamic Performance of a Transonic Axial Compressor with a Casing Groove and Tip Injection." Applied Mechanics and Materials 284-287 (January 2013): 872–77. http://dx.doi.org/10.4028/www.scientific.net/amm.284-287.872.

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This paper presents a parametric study on aerodynamic performance of a transonic axial compressor combined with a casing groove and tip injection using three-dimensional Reynolds-average Navier-Stokes equations. The front and rear lengths and height of the groove are selected as the geometric parameters to investigate their effects on the stall margin and peak adiabatic efficiency. These parameters are changed with constant injection. The validation of the numerical results is performed in comparison with experimental data for the total pressure ratio and adiabatic efficiency. As the results of the parametric study, the maximum stall margin and peak adiabatic efficiency are obtained in the axial compressor having 70% groove height of the reference groove. The stall margin and peak adiabatic efficiency in other cases are also improved in comparison with the axial compressors with the smooth casing and reference groove. The results show that both the stall margin and the peak adiabatic efficiency are considerably improved by the application of the casing groove combined with tip injection in an axial compressor.

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26

Mütschard, Silas, Jan Werner, Christian Kunkel, Maximilian Karl, Heinz-Peter Schiffer, Christoph Biela, and Sebastian Robens. "Investigation of surge in a 1.5-stage transonic axial compressor." Journal of the Global Power and Propulsion Society 6 (November 28, 2022): 304–17. http://dx.doi.org/10.33737/jgpps/156119.

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In this paper we give insight into characteristics of a 1.5-stage transonic axial compressor rig with focus on surge during a stalled operating point. The new compressor rig at TU Darmstadt is representative for the front stage of an industrial gas turbine. Transient throttling maneuvers were conducted for multiple operating points during the first test campaign of the TCD 2 (Transonic Compressor Darmstadt 2), providing an extensive set of unsteady structural and aerodynamic data beyond the stability limit. Enhanced analytical methods allow detailed studies including aerodynamic spectral analysis as well as determination of propagation speed and size of disturbances. The results differ from observations at comparable test rigs, revealing an interesting manifestation of stall: In a wide range of the stability limit it shows a periodicity. The stall emerges and vanishes recurrently, causing strong oscillations of the pressure ratio. Additional unsteady measurements of the mass flow indicate a surge. Regarding the compressor map, this results in staggering operating points, showing a hysteresis. However, due to a rather small plenum and experience with a similar test rig the TCD 2 was not expected to surge. Comprehensive analyses are carried out to characterize this phenomenon.

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27

Roberts, William B., Albert Armin, George Kassaseya, Kenneth L. Suder, Scott A. Thorp, and Anthony J. Strazisar. "The Effect of Variable Chord Length on Transonic Axial Rotor Performance." Journal of Turbomachinery 124, no. 3 (July 1, 2002): 351–57. http://dx.doi.org/10.1115/1.1459734.

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Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94–96% of nominal (new) blade chord.

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28

Wu, Xiaoxiong, Bo Liu, Nathan Ricks, and Ghader Ghorbaniasl. "Surrogate Models for Performance Prediction of Axial Compressors Using through-Flow Approach." Energies 13, no. 1 (December 30, 2019): 169. http://dx.doi.org/10.3390/en13010169.

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Two-dimensional design and analysis issues on the meridional surface, which is important in the preliminary design procedure of compressors, are highly dependent on the accuracy of empirical models, such as the prediction of total pressure loss model and turning flow angle. Most of the widely used models are derived or improved from experimental data of some specific cascades with low-loading and low-speed airfoil types. These models may work for most conventional compressors but are incapable of accurately estimating the performance for some specific cases like transonic compressors. The errors made by these models may mislead the final design results. Therefore, surrogate models are developed in this work to reduce the errors and replace the conventional empirical models in the through-flow calculation procedure. A group of experimental data considering a two-stage transonic compressor is used to generate the airfoils database for training the surrogate models. Sensitivity analysis is applied to select the most influential features. Two supervised learning approaches including support vector regression (SVR) and Gaussian process regression (GPR) are used to train the models with a Bayesian optimization algorithm to obtain the optimal hyper parameters. The trained models are integrated into the through-flow code based on streamline curvature method (SLC) to predict the overall performance and internal flow field of the transonic compressor on five rotational speed lines for validation. The predictions are compared with the experimental data and the results of conventional empirical models. The comparison shows that SVR and GPR respectively reduce the predicted error of empirical models by 62.2% and 55.2% for the total pressure ratio and 48.4% and 50.1% for adiabatic efficiency on average. This suggests that the surrogate models constitute an alternative way to predict the performance of airfoils in through-flow calculation where empirical models are inefficient.

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29

Shi, Hengtao. "A Parametric Blade Design Method for High-Speed Axial Compressor." Aerospace 8, no. 9 (September 18, 2021): 271. http://dx.doi.org/10.3390/aerospace8090271.

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The blade geometry design method is an important tool to design high performance axial compressors, expected to have large design space while limiting the quantity of design variables to a suitable level for usability. However, the large design space tends to increase the quantity of the design variables. To solve this problem, this paper utilizes the normalization and subsection techniques to develop a geometry design method featuring flexibility and local adjustability with limited design variables for usability. Firstly, the blade geometry parameters are defined by using the normalization technique. Then, the normalized camber angle f1(x) and thickness f2(x) functions are proposed with subsection techniques used to improve the design flexibility. The setting of adjustable coefficients acquires the local adjustability of blade geometry. Considering the usability, most of the design parameters have clear, intuitive meanings to make the method easy to use. To test this developed geometry design method, it is applied in the design of a transonic, two flow-path axial fan component for an aero engine. Numerical simulations indicate that the designed transonic axial fan system achieves good efficiency above 0.90 for the entire main-flow characteristic and above 0.865 for the bypass flow characteristic, while possessing a sufficiently stable operation range. This indicates that the developed design method has a large design space for containing the good performance compressor blade of different inflow Mach numbers, which is a useful platform for axial-flow compressor blade design.

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30

Dorney, D. J., and O. P. Sharma. "Evaluation of Flow Field Approximations for Transonic Compressor Stages." Journal of Turbomachinery 119, no. 3 (July 1, 1997): 445–51. http://dx.doi.org/10.1115/1.2841143.

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The flow through gas turbine compressors is often characterized by unsteady, transonic, and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor–stator interacting Navier–Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses, and performance quantities associated with axial compressor stages are presented. The methods studied include: (1) the unsteady fully coupled blade row technique, (2) the steady coupled blade row method, (3) the steady single blade row technique, and (4) the loosely coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.

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31

Alone, Dilipkumar Bhanudasji, S. Satish Kumar, Shobhavathy M. Thimmaiah, Janaki Rami Reddy Mudipalli, A. M. Pradeep, Srinivasan Ramamurthy, and Venkat S. Iyengar. "Improvement of Moderately Loaded Transonic Axial Compressor Performance Using Low Porosity Bend Skewed Casing Treatment." International Journal of Rotating Machinery 2014 (2014): 1–14. http://dx.doi.org/10.1155/2014/625876.

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This paper presents experimental results of a single stage transonic axial flow compressor coupled with low porosity bend skewed casing treatment. The casing treatment has a plenum chamber above the bend slots. The depth of the plenum chamber is varied to understand its impact on the performance of compressor stage. The performance of the compressor stage is evaluated for casing treatment and plenum chamber configurations at two axial locations of 20% and 40%. Experimental results reveal that the stall margin of the compressor stage increases with increase in the plenum chamber volume. Hot-wire measurements show significant reduction in the turbulence intensity with increase in the plenum chamber volume compared to that with the solid casing at the stall condition. At higher operating speeds of 80% and at 20% axial coverage, the stall margin of the compressor increases by 20% with half and full plenum depth. The improvement in the peak stage efficiency observed is 4.6% with half plenum configuration and 3.34% with the full plenum configuration. The maximum improvement in the stall margin of 29.16% is obtained at 50% operating speed with full plenum configurations at 40% axial coverage.

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32

Lu, Hanan, Qiushi Li, Tianyu Pan, and Ramesh Agarwal. "Analysis and application of shroud wall optimization to an axial compressor with upstream boundary layer to improve aerodynamic performance." International Journal of Numerical Methods for Heat & Fluid Flow 29, no. 11 (November 4, 2019): 4237–61. http://dx.doi.org/10.1108/hff-01-2019-0071.

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PurposeFor an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions.Design/methodology/approachThis paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition.FindingsAfter the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL.Originality/valueThe originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.

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33

Vuong, Tien-Dung, and Kwang-Yong Kim. "Casing treatment using oblique slots for a single-stage transonic axial compressor." Journal of Physics: Conference Series 2217, no. 1 (April 1, 2022): 012055. http://dx.doi.org/10.1088/1742-6596/2217/1/012055.

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Abstract This work presents a casing treatment design using oblique slots to improve the stability of a single-stage transonic axial compressor, NASA Stage 37. 36 individual slots are distributed uniformly on the rotors’ shroud. Numerical simulations were performed to find the effects of the casing treatment on the compressor’s stability as well as the adiabatic efficiency and pressure ratio. The oblique slots were found to effectively alleviate the negative influence of the tip leakage flow on the compressor’s performance, thus increasing the stall margin without significant sacrifices in efficiency and pressure rise. Three geometric parameters of the casing treatment were examined. The results showed that, with the oblique slots, the highest stall margin reached 13.59% - a 37% increase from the smooth casing case, with about 0.17% and 0.09% reductions in the peak efficiency and maximum pressure ratio, respectively.

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34

Thompson, D. W., P. I. King, and D. C. Rabe. "Experimental Investigation of Stepped Tip Gap Effects on the Performance of a Transonic Axial-Flow Compressor Rotor." Journal of Turbomachinery 120, no. 3 (July 1, 1998): 477–86. http://dx.doi.org/10.1115/1.2841743.

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The effects of stepped-tip gaps and clearance levels on the performance of a transonic axial-flow compressor rotor were experimentally determined. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested for an optimum tip clearance with variations in stepped gaps machined into the casing near the aft tip region of the rotor. Nine causing geometries were investigated consisting of three step profiles at each of three clearance levels. For small and intermediate clearances, stepped tip gaps were found to improve pressure ratio, efficiency, and flow range for most operating conditions. At 100 percent design rotor speed, stepped tip gaps produced a doubling of mass flow range with as much as a 2.0 percent increase in mass flow and a 1.5 percent improvement in efficiency. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an optimum casing profile.

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35

Xie, Fang, Chang Jiang Liu, and You Jun Wang. "Turbulence Model Influence on Numerical Investigation of Transonic Axial Compressor Rotor." Advanced Materials Research 308-310 (August 2011): 1519–22. http://dx.doi.org/10.4028/www.scientific.net/amr.308-310.1519.

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Numerical method using HI and HOH meshing combined B - L turbulent model and S - A turbulent model separately based on the Rotor 37 compressor Rotor was applied to the steady flow. results on pressure characteristic curve, stall point forecast etc were compared with related experimental data. This paper discussed calculation precision influenced by the turbulence model and numerical computation grid. This numerical investigation was basis for subsequent compressor internal flow field study.

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36

He, Xiao, Mingmin Zhu, Kailong Xia, Klausmann Fabian, Jinfang Teng, and Mehdi Vahdati. "Validation and verification of RANS solvers for TUDa-GLR-OpenStage transonic axial compressor." Journal of the Global Power and Propulsion Society 7 (January 27, 2023): 13–29. http://dx.doi.org/10.33737/jgpps/158034.

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This paper presents a comprehensive validation and verification study of turbomachinery Reynolds-averaged Navier-Stokes flow solvers on the transonic axial compressor TUDa-GLR-OpenStage. Two commercial solvers namely Ansys CFX and Numeca FineTurbo are adopted to provide the benchmark solutions, which can be used for verification of other RANS solvers in the future. Based on these solvers, five sets of grids, two advection schemes (i.e., central difference and second-order upwind), four turbulence models (i.e., SA, SA-RC, SST and EARSM) and two rotor-stator interface models (i.e., mixing plane and sliding plane) are investigated to quantify their effects on predicting the performance and the flow field of the compressor stage. Results show that the choices of grid density and turbulence model are most sensitive to the prediction, leading to 5% and 7% variation in compressor performance characteristics, respectively. Regarding the choice of grid density, a method to estimate the grid discretization error is demonstrated, which is transferrable to other cases. Regarding the choice of turbulence model, the EARSM model is found overall most accurate among the investigated models, and the limitations and deficiencies of the rest models are discussed in detail based on the analysis of the mean flow fields and the eddy viscosity fields. The grids and the major CFD results presented in this work are open-accessed to the community for further research. The results and discussions presented in this paper provide a useful reference for future practices of RANS simulations for compressors.

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37

Zhu, N. G., L. Xu, and M. Z. Chen. "Similarity Transformations for Compressor Blading." Journal of Turbomachinery 114, no. 3 (July 1, 1992): 561–68. http://dx.doi.org/10.1115/1.2929180.

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Improving the performance of high-speed axial compressors through low-speed model compressor testing has proved to be economical and effective (Wisler, 1985). The key to this technique is to design low-speed blade profiles that are aerodynamically similar to their high-speed counterparts. The conventional aerodynamic similarity transformation involves the small disturbance potential flow assumption; therefore, its application is severely limited and generally not used in practical design. In this paper, a set of higher order transformation rules are presented that can accommodate large disturbances at transonic speed and are therefore applicable to similar transformations between the high-speed high-pressure compressor and its low-speed model. Local linearization is used in the nonlinear equations and the transformation is obtained in an iterative process. The transformation gives the global blading parameters such as camber, incidence, and solidity as well as the blade profile. Both numerical and experimental validations of the transformation show that the nonlinear similarity transformations do retain satisfactory accuracy for highly loaded blades up to low transonic speeds. Further improvement can be made by only slightly modifying profiles numerically without altering the global similarity parameters.

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38

Vuong, Tien-Dung, Kwang-Yong Kim, and Cong-Truong Dinh. "Recirculation-groove coupled casing treatment for a transonic axial compressor." Aerospace Science and Technology 111 (April 2021): 106556. http://dx.doi.org/10.1016/j.ast.2021.106556.

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39

Liu, A., Y. P. Ju, and C. H. Zhang. "Parallel Simulation of Aerodynamic Instabilities in Transonic Axial Compressor Rotor." Journal of Propulsion and Power 34, no. 6 (November 2018): 1561–73. http://dx.doi.org/10.2514/1.b37038.

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40

Kang, Young-Seok, Tae-Choon Park, Oh-Sik Hwang, and Soo-Seok Yang. "Experimental Research on Multi Stage Transonic Axial Compressor Performance Evaluation." Journal of Fluid Machinery 14, no. 6 (December 1, 2011): 96–101. http://dx.doi.org/10.5293/kfma..2011.14.6.096.

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41

Koch, Peter J., Douglas P. Probasco, J. Mitch Wolff, William W. Copenhaver, and Randall M. Chriss. "Transonic Compressor Influences on Upstream Surface Pressures with Axial Spacing." Journal of Propulsion and Power 17, no. 2 (March 2001): 474–76. http://dx.doi.org/10.2514/2.5768.

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42

JANG, Choon-Man, Ping LI, and Kwang-Yong KIM. "Optimization of Blade Sweep in a Transonic Axial Compressor Rotor." JSME International Journal Series B 48, no. 4 (2005): 793–801. http://dx.doi.org/10.1299/jsmeb.48.793.

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43

LONG, Bingxiang, Jiming CHEN, Zhenhua CHEN, Zongzheng LIU, and Daxiong LIAO. "Aero-acoustic performance of a continuous transonic wind-tunnel axial-flow compressor." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 40, no. 4 (August 2022): 829–36. http://dx.doi.org/10.1051/jnwpu/20224040829.

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A three-stage axial-flow compressor was developed for a continuous transonic wind tunnel. The rotor-stator spacing and the ratio of number of rotor blades to number of stator vanes were properly chosen for the purpose of aero-acoustic noise suppression. The aero-acoustic performances of the axial-flow compressor were tested and analyzed. The test results show that the first BPF tonal noise is effectively suppressed even though some of it still exists; the noise amplitude frequency spectrum shows that the abnormal tonal noise, which is different from that related to BPF and its harmonics, exists under a wide range of working conditions and plays a dominant role in determining the sound pressure levels of the inlet and outlet of the axial-flow compressor. The comparison of test results shows that vibration is induced by the periodic non-uniform inlet flow condition and that the rotor blade that has a high aspect ratio is one of the main noise sources responsible for the abnormal tonal noise.

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44

Islam, Asad, and Hongwei Ma. "Numerical study of probe parameters on performance of a transonic axial compressor." PLOS ONE 16, no. 1 (January 25, 2021): e0245711. http://dx.doi.org/10.1371/journal.pone.0245711.

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The paper shows the effect of the probe on the performance of a transonic axial speed compressor. The unobstructed flow case with the experimental data was validated and used as a guide for all subsequent study cases. The aerodynamic performance for different probe parameters were calculated numerically using ANSYS-CFX. This covered the results on compressor output from changing probe axial positions, the radial immersion depths, the size of the probe, and the total number of probes. The findings were evaluated in relation to the total pressure ratio, performance, margin of deflation and stability. The velocity part distributions further showed that the probe block and raises the flow Mach value, which is the explanation why the compressor rotor’s total pressure ratio is lost. In fact, the parameters of the sample will significantly influence the calculation outcomes and affect the standard margin. The range of stability was also affected, which changes the performance trend from the choke to the stall. Consequently, the collection of correct probe parameters with fewer impact on compressor output is addressed.

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45

Pham, Ky-Quang, Xuan-Truong Le, and Cong-Truong Dinh. "Effects of stator splitter blades on aerodynamic performance of a single-stage transonic axial compressor." Journal of Mechanical Engineering and Sciences 14, no. 4 (December 17, 2020): 7369–78. http://dx.doi.org/10.15282/jmes.14.4.2020.05.0579.

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Splitter blades located between stator blades in a single-stage axial compressor were proposed and investigated in this work to find their effects on aerodynamic performance and operating stability. Aerodynamic performance of the compressor was evaluated using three-dimensional Reynolds-averaged Navier-Stokes equations using the k-e turbulence model with a scalable wall function. The numerical results for the typical performance parameters without stator splitter blades were validated in comparison with experimental data. The numerical results of a parametric study using four geometric parameters (chord length, coverage angle, height and position) of the stator splitter blades showed that the operational stability of the single-stage axial compressor enhances remarkably using the stator splitter blades. The splitters were effective in suppressing flow separation in the stator domain of the compressor at near-stall condition which affects considerably the aerodynamic performance of the compressor.

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46

Liu, ZX, HZ Diao, XC Zhu, and ZH Du. "Numerical investigation of the axial compressor performance with inlet distortions." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, no. 8 (March 28, 2017): 1434–41. http://dx.doi.org/10.1177/0954410017699853.

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In this paper, a three-dimensional body force model for predicting compressor performance and stability is implemented in the Ansys CFX. The influence of the blade rows on the flow field is represented by the source terms of CFX-solver equation. At first, a high-speed and high-pressure-ratio transonic compressor with the clean inlet is investigated. The overall performance and the flow fields are in agreement well with those of the experimental date, so the model is reliable and correct. Then, the effects of the circumferential distortions in the inlet total pressure and the total temperature on the compressor performance and flow field are also illustrated, respectively. In summary, the proposed body force model is suitable to investigate the flow field of the compressor with the inlet distortions.

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47

Kim, Jin Hyuk, Kwang Yong Kim, and Kyung Hun Cha. "Effects of Number of Circumferential Casing Grooves on Stall Flow Characteristics of a Transonic Axial Compressor." Applied Mechanics and Materials 284-287 (January 2013): 727–32. http://dx.doi.org/10.4028/www.scientific.net/amm.284-287.727.

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This work investigates the effects of circumferential casing grooves on stall flow characteristics of a transonic axial compressor. Numerical analysis is conducted by solving three-dimensional steady Reynolds-averaged Navier-Stokes equations with the shear stress transport turbulence model. The results of flow analysis for an axial compressor with smooth casing are validated in comparison with experimental data for the pressure ratio and adiabatic efficiency. The numerical stall inception point is identified from the last converged point by convergence criteria, and the stall margin is predicted numerically. The peak adiabatic efficiency point is also obtained by reducing the normalized mass flow in the high mass flow region. In order to explore the influence of number of the circumferential casing grooves on the performance of the compressor, the stall margins and peak adiabatic efficiencies are evaluated compared to the case smooth casing. The stability of the axial compressor with circumferential casing grooves is found to be sensitively influenced by the number of grooves.

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48

Zhang, Guochen, Tianyi Gao, Zhihui Xu, Pengcheng Liu, and Chengfeng Zhang. "Influences of slotted blade on performance and flow structure of a transonic axial compressor." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 235, no. 6 (February 17, 2021): 1344–54. http://dx.doi.org/10.1177/0957650921994382.

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Main reason of compressor instability is boundary layer separation on the surface of blades. As one of flow control methods of the compressor, slotted blade has attracted many researchers’ attention because of its simple geometric structure and remarkable flow control effect. In order to evaluate its availability in the compressor, a type of convergent slot is designed to implement in a single-stage transonic axial compressor. Three configurations, i.e. rotor slot, stator slot and rotor-stator combined slot, are introduced to study the aerodynamic performance of compressor by numerical simulations. Furthermore, flow structures have been analyzed to explain the corresponding mechanism. The results show that overall stability margin of the compressor has been improved by flow control with slotted blade. Behavior of the rotor slot is better than that of the stator slot, but due to mass flow leakage in the slot, peak efficiency and chocking mass flow rate of the compressor are decreased by 1.18% and 3.8% respectively. The low momentum flow on pressure surface is sucked into the jet slot of stator blade, which improves the overall stability margin of 0.63%. The combined scheme with slotted rotor and slotted stator has obtained the best aerodynamic behavior with the increase of the overall stability margin of 2.83%. During the future research, main goal will be improvement of the compressor performance and extension of the mass flow rate range.

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49

Kim, Dae-Woong, Jin-Hyuk Kim, and Kwang-Yong Kim. "AERODYNAMIC PERFORMANCE OF AN AXIAL COMPRESSOR WITH A CASING GROOVE COMBINED WITH INJECTION." Transactions of the Canadian Society for Mechanical Engineering 37, no. 3 (September 2013): 283–92. http://dx.doi.org/10.1139/tcsme-2013-0018.

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Aerodynamic performance of a transonic axial compressor with a casing groove combined with injection has been investigated in this work. Three-dimensional Reynolds-averaged Navier–Stokes equations with k-ε turbulence model are discretized by finite volume approximations and solved on hexahedral grids for the flow analyses. For parametric study, the front and rear lengths and height of the casing groove are selected as the geometric parameters and are changed with constant injection to investigate their effects on the stall margin and peak adiabatic efficiency. As a result of the parametric study, the maximum stall margin and peak adiabatic efficiency are found to be obtained in the axial compressor having 70% height of the reference groove. The results show that the application of the casing groove combined with injection to an axial compressor is effective for the simultaneous improvement of both the stall margin and peak adiabatic efficiency of the compressor.

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50

Srinivas, G., K. Raghunandana, and B. Satish Shenoy. "Flow blockage in a transonic axial flow compressor: simulation analysis under distorted conditions." International Journal of Engineering & Technology 7, no. 2.21 (April 20, 2018): 43. http://dx.doi.org/10.14419/ijet.v7i2.21.11833.

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Today the aircraft industry is looking for faster and safer engines for both civil and military applications. The performance of all different types of air breathing engines depends on the amount of mass flow rate of air entering and hot gas ejecting out from the engine. Thrust is the key role for any engine performance. To achieve more thrust all the turbo machinery components like axial fan, axial flow compressor and axial flow turbine should function effectively. This paper is primarily dealing about one of the turbo machinery component, axial flow compressor performance where the study is more focused on flow blockage formation under distorted phenomena. The complete blade boundary layer formation and related flow numerical theory are discussed in detail, accordingly the boundary conditions were set to have better numerical simulations using ANSYS tool. To find the flow blockage formation suitable turbulence model was coded using the well know compressible equations. The flow blockage between the rotor and stator of the compressor stage was calculated and also validated with that of experimental data effectively. The flow simulation results also revealed that the performance parameters under the modern engine transonic speed from Mach 0.8 to 1.2 under the distorted conditions are better for aeromechanical features.

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