Artigos de revistas: "Inlet distortions"

1

Ariga, I., S. Masuda e A. Ookita. "Inducer Stall in a Centrifugal Compressor With Inlet Distortion". Journal of Turbomachinery 109, n.º 1 (1 de janeiro de 1987): 27–35. http://dx.doi.org/10.1115/1.3262066.

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The effects of inlet distortion on the inducer stall in a centrifugal compressor are investigated. Cases of both radial and circumferential distortion are investigated. It is shown that the rotating stall onset is amplified by radial distortions, and restrained by circumferential distortions. These results are compared with calculations based on the small disturbance theory. The authors find that the stall onset is governed by the characteristic parameters related to the lower flow rate region for radial distortions, but affected by those of the higher flow rate region for circumferential distortion. It is shown that the process of stall is different for each distortion pattern. Existence of inlet distortion reduces compressor performance characteristics and strongly influences the stability margin.

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2

Lowe, K. Todd. "Laser velocimetry for turbofan inlet distortion applications". Aircraft Engineering and Aerospace Technology 92, n.º 1 (6 de janeiro de 2020): 20–26. http://dx.doi.org/10.1108/aeat-11-2018-0285.

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Purpose The purpose of this paper is to assess state-of-the-art techniques for quantifying flow distortion in the inlets of turbofan engines, particularly with respect to the prospects for future flight applications. Design/methodology/approach To adequately characterize the flow fields of complex aircraft inlet distortions, the author has incorporated laser velocimetry techniques, namely, stereoscopic particle image velocimetry (PIV) and Doppler velocimetry based on filtered Rayleigh scattering (FRS), into inlet distortion studies. Findings Overall, the results and experience indicate that the pathway for integration of FRS technologies into flight systems is clearer and more robust than that of PIV. Practical implications While always a concern, the topic of inlet distortion has grown in importance as contemporary airframe designers seek extremely compact and highly integrated inlets. This research offers a means for gaining new understanding of the in situ aerodynamic phenomena involved with complex inlet distortion. Originality/value This paper presents unique applications of turbofan inlet velocimetry methods while providing an original assessment of technological challenges involved with progressing advanced velocimetry techniques for flight measurements.

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3

Pazur, W., e L. Fottner. "The Influence of Inlet Swirl Distortions on the Performance of a Jet Propulsion Two-Stage Axial Compressor". Journal of Turbomachinery 113, n.º 2 (1 de abril de 1991): 233–40. http://dx.doi.org/10.1115/1.2929091.

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Aeroengine intakes containing S-shaped diffusers produce different types of inlet swirl distortions and essentially a combination of a twin swirl and a bulk swirl. The main object of this investigation was to assess the influence of inlet swirl distortions on the performance of a transonic two-stage axial compressor installed in a turbo jet bypass engine Larzac 04. A typical inlet swirl distortion was simulated by a delta-wing in front of the engine. An experimental method was investigated to measure the performance map of the installed low-pressure compressor for different engine operating lines. The influence of an inlet swirl distortion with different strengths on the performance map of the compressor was investigated experimentally. It is shown that the performance parameters decrease and a temperature distortion is generated behind the compressor. As the basis of the theoretical investigations of the performance map, including inlet swirl distortions, a computing model considering four compressors working in parallel was established. The model is based on the idea that an inlet swirl distortion can be substituted by two fundamental types of swirl components, i.e., a bulk swirl corotating, and a bulk swirl counterrotating to the revolution of the compressor. Computed performance maps of the compressor will be discussed and compared with the experimental data.

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4

Pečinka, Jiří, Gabriel Thomas Bugajski, Petr Kmoch e Adolf Jílek. "JET ENGINE INLET DISTORTION SCREEN AND DESCRIPTOR EVALUATION". Acta Polytechnica 57, n.º 1 (28 de fevereiro de 2017): 22–31. http://dx.doi.org/10.14311/ap.2017.57.0022.

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Total pressure distortion is one of the three basic flow distortions (total pressure, total temperature and swirl distortion) that might appear at the inlet of a gas turbine engine (GTE) during operation. Different numerical parameters are used for assessing the total pressure distortion intensity and extent. These summary descriptors are based on the distribution of total pressure in the aerodynamic interface plane. There are two descriptors largely spread around the world, however, three or four others are still in use and can be found in current references. The staff at the University of Defence decided to compare the most common descriptors using basic flow distortion patterns in order to select the most appropriate descriptor for future department research. The most common descriptors were identified based on their prevalence in widely accessible publications. The construction and use of these descriptors are reviewed in the paper. Subsequently, they are applied to radial, angular, and combined distortion patterns of different intensities and with varied mass flow rates. The tests were performed on a specially designed test bench using an electrically driven standalone industrial centrifugal compressor, sucking air through the inlet of a TJ100 small turbojet engine. Distortion screens were placed into the inlet channel to create the desired total pressure distortions. Of the three basic distortions, only the total pressure distortion descriptors were evaluated. However, both total and static pressures were collected using a multi probe rotational measurement system.

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5

Hah, C., D. C. Rabe, T. J. Sullivan e A. R. Wadia. "Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor". Journal of Turbomachinery 120, n.º 2 (1 de abril de 1998): 233–46. http://dx.doi.org/10.1115/1.2841398.

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The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of eight periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier–Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20 percent of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

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6

Fang, Yibo, Dakun Sun, Xu Dong e Xiaofeng Sun. "Effects of Inlet Swirl Distortion on a Multi-Stage Compressor with Inlet Guide Vanes and Stall Margin Enhancement Method". Aerospace 10, n.º 2 (2 de fevereiro de 2023): 141. http://dx.doi.org/10.3390/aerospace10020141.

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Inlet swirl distortion is generally considered as a type of velocity distortion, and inlet guide vanes (IGVs) are widely used in the multi-stage compressor of aero-engines to eliminate the tangential velocity of the swirl flow. However, few studies have explored whether there still exists some negative influence of inlet swirl distortion on the compressor, even after the installation of IGVs. Therefore, in this study, the influence of various types of inlet swirl distortions on a multi-stage compressor with the installation of IGVs is investigated. A swirl distortion generator installed in the inlet duct was designed to produce various types of swirl flow patterns. When the distortion intensity increased to some degree, there still existed a decrease in the compressive capability and an obvious additional efficiency loss. The inlet twin swirl distortion was accompanied by total pressure distortion, so even with the installation of IGVs, there was still a significantly negative influence on the performance of the multi-stage compressor, especially the stall margin. Subsequently, to improve the stall margin under inlet swirl distortion, the stall precursor-suppressed (SPS) casing treatment was installed in the first stage of the multi-stage compressor. It could enhance the stall margin of the compressor with no obvious change in the characteristic curves and no additional efficiency loss under various types of inlet swirl distortions, and its mechanism was verified by capturing the dynamic pressure characteristics.

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7

Leinhos, Dirk C., Norbert R. Schmid e Leonhard Fottner. "The Influence of Transient Inlet Distortions on the Instability Inception of a Low-Pressure Compressor in a Turbofan Engine". Journal of Turbomachinery 123, n.º 1 (1 de fevereiro de 2000): 1–8. http://dx.doi.org/10.1115/1.1330271.

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While studies on compressor flow instabilities under the presence of inlet distortions have been carried out with steady distortions in the past, the investigation presented here focuses on the influence of transient inlet distortions as generated by variable geometry engine intakes of super- and hypersonic aircraft on the characteristic and the nature of the instability inception of a LPC. The flow patterns (total pressure distortion with a superimposed co- or counterrotating swirl) of the distortions are adopted from a hypersonic concept aircraft. A LARZAC 04 twin-spool turbofan was operated with transient inlet distortions, generated by a moving delta wing, and steady total pressure distortions starting close to the LPC’s stability limit until it stalled. High-frequency pressure signals are recorded at different engine power settings. Instabilities are investigated with regard to the inception process and the early detection of stall precursors for providing data for a future stability control device. It turned out that the transient distortion does not have an influence on the surge margin of the LPC compared to the steady distortion, but that it changes the nature of stall inception. The pressure traces are analyzed in the time and frequency domain and also with tools like Spatial FFT, Power Spectral Density, and Traveling Wave Energy. A Wavelet Transformation algorithm is applied as well. While in the case of clean inlet flow, the compressor exhibits different types of stall inception depending on the engine speed, stall is always initiated by spike-type disturbances under the presence of steady or transient distortions. Modal disturbances are present in the mid-speed range that do not grow into stall, but rather interact with the inlet flow and produce short length scale disturbances. The obtained early warning times prior to stall are adversely affected by transient distortions in some cases. The problem of appropriate thresholding becomes evident. The best warning times have been acquired using a statistical evaluation of the Wavelet coefficients, which might be promising to apply in a staged active control system. This system could include different phases of detection and actuation depending on the current precursor.

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8

Ng, Eddie Yin-Kwee, Ningyu Liu, Hong Ngiap Lim e Daniel Tan. "An Improved Integral Method for Prediction of Distorted Inlet Flow Propagation in Axial Compressor". International Journal of Rotating Machinery 2005, n.º 2 (2005): 117–27. http://dx.doi.org/10.1155/ijrm.2005.117.

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An improved integral method is proposed and developed for the quantitative prediction of distorted inlet flow propagation through axial compressor. The novel integral method is formulated using more appropriate and practical airfoil characteristics, with less assumptions needed for derivation. The results indicate that the original integral method (Kim et al., 1996) underestimated the propagation of inlet flow distortion. The effects of inlet flow parameters on the propagation of inlet distortions as well as on the compressor performance and characteristic are simulated and analyzed. From the viewpoint of compressor efficiency, the propagation of inlet flow distortion is further described using a compressor critical performance and its associated critical characteristic. The results present a realistic physical insight to an axial-flow compressor behavior with a propagation of inlet distortion.

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9

Schmid, Norbert R., Dirk C. Leinhos e Leonhard Fottner. "Steady Performance Measurements of a Turbofan Engine With Inlet Distortions Containing Co- and Counterrotating Swirl From an Intake Diffuser for Hypersonic Flight". Journal of Turbomachinery 123, n.º 2 (1 de fevereiro de 2000): 379–85. http://dx.doi.org/10.1115/1.1343466.

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The influence of distorted inlet flow on the steady and unsteady performance of a turbofan engine, which is a component of an air-breathing combined propulsion system for a hypersonic transport aircraft, is reported in this paper. The performance and stability of this propulsion system depend on the behavior of the turbofan engine. The complex shape of the intake duct causes inhomogeneous flow at the engine inlet plane, where total pressure and swirl distortions are present. The S-bend intakes are installed axisymmetrically left and right into the hypersonic aircraft, generating axisymmetric mirror-inverted flow patterns. Since all turbo engines of the propulsion system have the same direction of rotation, one distortion corresponds to a corotating swirl at the low pressure compressor (LPC) inlet while the mirror-inverted image counterpart represents a counterrotating swirl. Therefore the influence of the distortions on the performance and stability of the ‘CO’ and ‘COUNTER’ rotating turbo engine are different. The distortions were generated separately by an appropriate simulator at the inlet plane of a LARZAC 04 engine. The results of low-frequency measurements at different engine planes yield the relative variations of thrust and specific fuel consumption and hence the steady engine performance. High-frequency measurements were used to investigate the different influence of CO and COUNTER inlet distortions on the development of LPC instabilities.

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10

Engeda, Abraham, Yunbae Kim, Ronald Aungier e Gregory Direnzi. "The Inlet Flow Structure of a Centrifugal Compressor Stage and Its Influence on the Compressor Performance". Journal of Fluids Engineering 125, n.º 5 (1 de setembro de 2003): 779–85. http://dx.doi.org/10.1115/1.1601255.

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The performance of centrifugal compressors can be seriously degraded by inlet flow distortions that result from an unsatisfactory inlet configuration. In this present work, the flow is numerically simulated and the flow details are analyzed and discussed in order to understand the performance behavior of the compressor exposed to different inlet configurations. In a previous work, complementary to this present work, experimental tests were carried out for the comparison of a centrifugal compressor stage performance with two different inlet configurations: one of which was a straight pipe with constant cross-sectional area and the other a 90-deg curved pipe with nozzle shape. The comparative test results indicated significant compressor stage performance difference between the two different inlet configurations. Steady-state compressor stage simulation including the impeller and diffuser with three different inlets has been carried out to investigate the influence of each inlet type on the compressor performance. The three different inlet systems included a proposed and improved inlet model. The flow from the bend inlet is not axisymmetric in the circumferential and radial distortion, thus the diffuser and the impeller are modeled with fully 360-deg passages.

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11

Yang, Bo, e Guoming Zhu. "Impacts of Inlet Circumferential Distortions on the Aerodynamic Performance of a Transonic Axial Compressor". Processes 11, n.º 7 (21 de julho de 2023): 2175. http://dx.doi.org/10.3390/pr11072175.

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Distortion, such as total pressure distortion, is a common phenomenon at the inlet of an axial compressor. Nonuniform inflow can greatly affect the aerodynamic performance and the stability of the compressor. In this paper, a sinusoidal distortion model is used at the inlet. Then, a series of unsteady computational fluid dynamics (CFD) simulations with a full-annulus model is conducted. Three kinds of the total pressure distortions, such as circumferentially covering 60, 120, and 180-degs sectors above 70% span of the blade at the inlet section, are adopted. Based on the results, how the rotating stall cells can be induced and developed under the different inlet conditions is understood during the stall process. It is found that the secondary stall cell is more easily triggered when the circumferential range of total pressure distortion is increased. Meanwhile, the influence of the rotating distortion on the compressor performance is also studied and a dangerous distortion rotating speed is observed.

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12

Baretter, Alberto, Benjamin Godard, Pierric Joseph, Olivier Roussette, Francesco Romanò, Raphael Barrier e Antoine Dazin. "Experimental and Numerical Analysis of a Compressor Stage under Flow Distortion". International Journal of Turbomachinery, Propulsion and Power 6, n.º 4 (23 de novembro de 2021): 43. http://dx.doi.org/10.3390/ijtpp6040043.

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On many occasions, fan or compressor stages have to face azimuthal flow distortion at inlet, which affects their performance and stability. These flow distortions can be caused by external events or by some particular geometrical features. The aim of this work is to propose a joined numerical and experimental analysis of the flow behavior in a single axial compressor stage under flow distortion. The distortions are generated by different grids that are placed upstream to the rotor. Experimentally, the flow analysis is based on the measurements obtained by a series of unsteady pressure sensors flush-mounted at the casing of the machine rotor. URANS computations are conducted using the elsA software. The flow distortion is simulated by a drop of stagnation pressure ratio at the inlet boundary condition. The study is focusing first on the ability of a pressure drop, imposed as an inlet boundary condition in CFD, to reproduce accurately the effect of a flow distortion. The analysis is conducted using singular value decomposition (SVD) and dynamic mode decomposition (DMD). A special attention is then paid, on the experimental level, to the arising of rotating stall, from the onset of the instability up to completely developed stall cells.

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13

Li, H. D., e L. He. "Single-Passage Analysis of Unsteady Flows Around Vibrating Blades of a Transonic Fan Under Inlet Distortion". Journal of Turbomachinery 124, n.º 2 (1 de abril de 2002): 285–92. http://dx.doi.org/10.1115/1.1450567.

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Computations of unsteady flows due to inlet distortion driven blade vibrations, characterized by long circumferential wavelengths, typically need to be carried out in multi-passage/whole-annulus domains. In the present work, a single-passage three-dimensional unsteady Navier-Stokes approach has been developed and applied to unsteady flows around vibrating blades of a transonic fan rotor (NASA Rotor-67) with inlet distortions. The phase-shifted periodic condition is applied using a Fourier series based method, “shape-correction,” which enables a single-passage solution to unsteady flows under influences of multiple disturbances with arbitrary interblade phase angles. The computational study of the transonic fan illustrates that unsteady flow response to an inlet distortion varies greatly depending on its circumferential wavelength. The response to a long wavelength (whole-annulus) distortion is strongly nonlinear with a significant departure of its time-averaged flow from the steady state, while that at a short wavelength (two passages) behaves largely in a linear manner. Nevertheless, unsteady pressures due to blade vibration, though noticeably different under different inlet distortions, show a linear behavior. Thus, the nonlinearity of the flow response to inlet distortion appears to influence the aerodynamic damping predominantly by means of changing the time-averaged flow. Good agreements between single-passage solutions and multi-passage solutions are obtained for all the conditions considered, which clearly demonstrates the validity of the phase-shifted periodicity at a transonic nonlinear distorted flow condition. For the present cases, typical CPU time saving by a factor of 5–10 is achieved by the single-passage solutions.

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Longley, J. P., H. W. Shin, R. E. Plumley, P. D. Silkowski, I. J. Day, E. M. Greitzer, C. S. Tan e D. C. Wisler. "Effects of Rotating Inlet Distortion on Multistage Compressor Stability". Journal of Turbomachinery 118, n.º 2 (1 de abril de 1996): 181–88. http://dx.doi.org/10.1115/1.2836624.

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In multispool engines, rotating stall in an upstream compressor will impose a rotating distortion on the downstream compressor, thereby affecting its stability margin. In this paper experiments are described in which this effect was simulated by a rotating screen upstream of several multistage low-speed compressors. The measurements are complemented by, and compared with, a theoretical model of multistage compressor response to speed and direction of rotation of an inlet distortion. For corotating distortions (i.e., distortions rotating in the same direction as rotor rotation), experiments show that the compressors exhibited significant loss in stability margin and that they could be divided into two groups according to their response. The first group exhibited a single peak in stall margin degradation when the distortion speed corresponded to roughly 50 percent of rotor speed. The second group showed two peaks in stall margin degradation corresponding to distortion speeds of approximately 25–35 percent and 70–75 percent of rotor speed. These new results demonstrate that multistage compressors can have more than a single resonant response. Detailed measurements suggest that the two types of behavior are linked to differences between the stall inception processes observed for the two groups of compressors and that a direct connection thus exists between the observed forced response and the unsteady flow phenomena at stall onset. For counterrotational distortions, all the compressors tested showed minimal loss of stability margin. The results imply that counterrotation of the fan and core compressor, or LP and HP compressors, could be a worthwhile design choice. Calculations based on the two-dimensional theoretical model show excellent agreement for the compressors, which had a single peak for stall margin degradation. We take this first-of-a-kind comparison as showing that the model, though simplified, captures the essential fluid dynamic features of the phenomena. Agreement is not good for compressors that had two peaks in the curve of stall margin shift versus distortion rotation speed. The discrepancy is attributed to the three-dimensional and short length scale nature of the stall inception process in these machines; this includes phenomena that have not yet been addressed in any model.

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15

Zhang, Mingming, e Anping Hou. "Investigation on stall inception of axial compressor under inlet rotating distortion". Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 231, n.º 10 (21 de dezembro de 2015): 1859–70. http://dx.doi.org/10.1177/0954406215623978.

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This paper applies a numerical approach to improve the understanding of reaction to various inflow conditions for the compressor system and the mechanism of stall inception under rotating inflow distortions. Full annulus, unsteady, three-dimensional computational fluid dynamics has been used to simulate an axial low-speed compressor operating under rotating distorted inflow conditions. The development of the flow through the rotor is then explained in terms of the redistribution of the flow field and the process of stall inception. The results suggest that the increased flow incidence close to the tip region under co-rotating inflow distortion plays an important role on the stall inception process. The presence of a strong modal wave is observed under co-rotating inflow distortions. This leads to a significant impact on the loss of stall margin, as compared with other distorted inflow conditions. There is a significant peak in the flow coefficient at stall for co-rotating inlet distortion. It can be interpreted as a resonant behavior of the compressor under a strong interaction between the flow field and inlet distortion. It indicates that the stall inception is triggered by the perturbation of the rotating distorted inflow through the long length scale disturbances.

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Guo, Jin, e Jun Hu. "Development of body force model for steady inlet distortions in high-speed multistage compressor". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, n.º 9 (7 de julho de 2016): 1650–59. http://dx.doi.org/10.1177/0954410016656880.

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This study aims at establishing a three-dimensional numerical model, compressor aerodynamic performance analysis model, to simulate the impact of complicated distorted flow on multistage axial flow compressor based on the body force model. The model solves the compressible three-dimensional Euler equations, which are modified to include source terms representing the effect of the blade rows. In this study, the association between blade source terms and entry Mach number together with attack angle could be established with the deviation angle model and loss model. In this paper, compressor aerodynamic performance analysis model is used to evaluate the effect of inlet circumferential total pressure distortion and swirl distortion on a five-stage high-pressure compressor. Calculated operating maps for compressor agree well with the experimental results. Meanwhile, the traveling process of inlet distortions in the multistage compressor is correctly revealed. The wide application prospect of the model can be seen in the area of inlet distortion problems.

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Verma, Vishwas, Gursharanjit Singh e AM Pradeep. "The effect of inlet distortion on low bypass ratio turbofan engines". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, n.º 8 (2 de março de 2020): 1395–413. http://dx.doi.org/10.1177/0954410020909190.

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Inlet flow non-uniformity, commonly known as inflow distortion, has been a long-standing problem in the history of gas turbine engines. Distortion can be present in the form of total pressure, total temperature or inflow incidence or any combinations of these. The search for better and robust performance requires engines that can sustain a large amount of inlet distortion without considerable loss in the thrust. In the present paper, the effect of total pressure distortion on a single-stage compressor and low bypass ratio fans are studied. Distortion near hub and tip in the form of step radial total pressure profiles is imposed at far upstream of the rotor leading edge. A systematic approach to qualitatively predict the performance maps in the presence of these distortions is discussed. Further, two extents of total pressure distortion are explored for constant inlet distortion intensity. Hub distortion is found to increase the stability margin, whereas tip distortion reduces it. On extending the distortion extent, hub distortion drastically reduces the stability margin, whereas a comparatively lower reduction in stability margin with tip distortion is observed. The critical distortion limit is observed by varying the inlet distortion extent. Also, it is found that downstream ducts in the bypass axial fan do not interact with the upstream fan. This can be exploited to perform independent simulations of the core engine from low bypass ratio fans. Hub distortion is found to drastically affect the duct performance owing to the presence of thicker upstream inlet boundary layer.

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Li, Yue Feng, Qing Zhen Yang e Xue Jiao Deng. "Aerodynamic Design Investigation on Inlet-Exhaust System Using Super-Ellipse Method". Applied Mechanics and Materials 246-247 (dezembro de 2012): 446–50. http://dx.doi.org/10.4028/www.scientific.net/amm.246-247.446.

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Generally, the shape index n is selected in a form when design the inlet-exhaust system using traditional super-ellipse method. Unfortunately, this selection process is time-consuming and not precise enough, so the cross-section designed by super-ellipse method may get distortions easily, which influences the inner flow and the total pressure of the inlet-exhaust system greatly. Associating the shape index n with the variation pattern of the inlet-exhaust cross-section, an improved super-ellipse method is developed to design the inlet-exhaust system. This method ensures the precision and uniqueness of shape index n for any cross-section in an adaptive way. The numerical simulation results show that the S-shape inlet designed using this method has high total pressure recovery coefficient and lower distortion coefficient, the S-shaped nozzle has high total pressure recovery coefficient and thrust coefficient.

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Cousins, W. T., K. K. Dalton, T. T. Andersen e G. A. Bobula. "Pressure and Temperature Distortion Testing of a Two-Stage Centrifugal Compressor". Journal of Engineering for Gas Turbines and Power 116, n.º 3 (1 de julho de 1994): 567–73. http://dx.doi.org/10.1115/1.2906857.

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Altitude pressure and temperature inlet distortion testing of the two-stage centrifugal compressor in the T800-LHT-800 engine is described. The test setup and the testing techniques are reviewed and the results of the test are presented. The generation of classical 180 deg patterns of both pressure and temperature distortion is discussed. Temperature distortion was created using a hydrogen burner system while pressure distortion was created in the classical manner, using screens. Results of both individual and combined temperature and pressure distortions in both opposed and concurrent patterns are shown.

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Taghavi Zenouz, Reza, Mehran Eshaghi Sir e Mohammad Hosein Ababaf Behbahani. "Performance of a Low Speed Axial Compressor Rotor Blade Row under Different Inlet Distortions". Mechanical Sciences 8, n.º 1 (31 de maio de 2017): 127–36. http://dx.doi.org/10.5194/ms-8-127-2017.

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Abstract. Responses of an axial compressor isolated rotor blade row to various inlet distortions have been investigated utilizing computational fluid dynamic technique. Distortions have been imposed by five screens of different geometries, but with the same blockage ratio. These screens were embedded upstream of the rotor blade row. Flow fields are simulated in detail for compressor design point and near stall conditions. Performance curves for distorted cases are extracted and compared to the undisturbed case. Flow simulations and consequent performance characteristics show that the worst cases belong to non-symmetric blockages, i.e., those of partial circumferential configurations. These cases produce the largest wakes which can disturb the flow, considerably. Superior performances correspond to the inner and outer continuous circumferential distortion screens. Since, they produce no significant disturbances to the main flow in comparison to the non-symmetric screens.

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Manwaring, S. R., e S. Fleeter. "Forcing Function Effects on Rotor Periodic Aerodynamic Response". Journal of Turbomachinery 113, n.º 2 (1 de abril de 1991): 312–19. http://dx.doi.org/10.1115/1.2929109.

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A series of experiments are performed in an extensively instrumented axial flow research compressor to investigate the effects of different low reduced frequency aerodynamic forcing functions and steady loading level on the gust-generated unsteady aerodynamics of a first-stage rotor blade row. Two different two-per-rev forcing functions are considered: (1) the velocity deficit from two 90 deg circumferential inlet flow distortions, and (2) the wakes from two upstream obstructions, which are characteristic of airfoil or probe excitations. The data show that the wake-generated rotor row first harmonic response is much greater than that generated by the inlet distortion, with the difference decreasing with increased steady loading.

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Giuliani, James E., e Jen-Ping Chen. "Fan Response to Boundary-Layer Ingesting Inlet Distortions". AIAA Journal 54, n.º 10 (outubro de 2016): 3232–43. http://dx.doi.org/10.2514/1.j054762.

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23

Levy, Y., e J. Pismenny. "Nonlinear Oscillations in Compressor under Rotating Inlet Distortions". International Journal of Turbo and Jet Engines 23, n.º 3 (janeiro de 2006): 165–72. http://dx.doi.org/10.1515/tjj.2006.23.3.165.

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Sohail, Muhammad Umer, Hossein Raza Hamdani, Asad Islam, Khalid Parvez, Abdul Munem Khan, Usman Allauddin, Muhammad Khurram e Hassan Elahi. "Prediction of Non-Uniform Distorted Flows, Effects on Transonic Compressor Using CFD, Regression Analysis and Artificial Neural Networks". Applied Sciences 11, n.º 8 (20 de abril de 2021): 3706. http://dx.doi.org/10.3390/app11083706.

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Non-uniform inlet flows frequently occur in aircrafts and result in chronological distortions of total temperature and total pressure at the engine inlet. Distorted inlet flow operation of the axial compressor deteriorates aerodynamic performance, which reduces the stall margin and increases blade stress levels, which in turn causes compressor failure. Deep learning is an efficient approach to predict catastrophic compressor failure, and its stability for better performance at minimum computational cost and time. The current research focuses on the development of a transonic compressor instability prediction tool for the comprehensive modeling of axial compressor dynamics. A novel predictive approach founded by an extensive CFD-based dataset for supervised learning has been implemented to predict compressor performance and behavior at different ambient temperatures and flow conditions. Artificial Neural Network-based results accurately predict compressor performance parameters by minimizing the Root Mean Square Error (RMSE) loss function. Computational results show that, as compared to the tip radial pressure distortion, hub radial pressure distortion has improved the stability range of the compressor. Furthermore, the combined effect of pressure distortion with the bulk flow has a qualitative and deteriorator effect on the compressor.

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Han, Fenghui, Zhe Wang, Yijun Mao, Jiajian Tan e Wenhua Li. "Flow Control of Radial Inlet Chamber and Downstream Effects on a Centrifugal Compressor Stage". Applied Sciences 11, n.º 5 (1 de março de 2021): 2168. http://dx.doi.org/10.3390/app11052168.

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Radial inlet chambers are widely used in various multistage centrifugal compressors, although they induce extra flow loss and inlet distortions. In this paper, the detailed flow characteristics inside the radial inlet chamber of an industrial centrifugal compressor have been numerically investigated for flow control and performance improvement. First, the numerical results are validated against the experimental data, and flow conditions inside the inlet chambers with different structures are compared. They indicate that, in the non-guide vane scheme, sudden expansions, tangential flows and flow separations in the spiral and annular convergent channels are the major causes of flow loss and distortions, while using guide vanes could introduce additional flow impacts, separations and wakes. Based on the flow analysis, structure improvements have been carried out on the radial inlet chamber, and an average increase of 4.97% has been achieved in the inlet chamber efficiencies over different operating conditions. However, the results further reveal that the increases in the performance and overall flow uniformity just in the radial inlet chamber do not necessarily mean a performance improvement in the downstream components, and the distribution of the positive tangential velocity at the impeller inlet might be a more essential factor for the efficiency of the whole compressor.

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Arshad, Ali, Qiushi Li, Simin Li e Tianyu Pan. "Effects of inlet radial distortion on the type of stall precursor in low-speed axial compressor". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, n.º 1 (30 de setembro de 2016): 55–67. http://dx.doi.org/10.1177/0954410016670679.

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Experimental investigations of the effect of inlet blade loading on the rotating stall inception process are carried out on a single-stage low-speed axial compressor. Temporal pressure signals from the six high response pressure transducers are used for the analysis. Pressure variations at the hub are especially recorded during the stall inception process. Inlet blade loading is altered by installing metallic meshed distortion screens at the rotor upstream. Three sets of experiments are performed for the comparison of results, i.e. uniform inlet flow, tip, and hub distortions, respectively. Regardless of the type of distortion introduced to the inflow, the compressor undergoes a performance drop, which is more severe in the hub distortion case. Under the uniform inlet flow condition, stall inception is caused by the modal type disturbance while the stall precursor switched to spike type due to the highly loaded blade tip. In the presence of high blade loading at the hub, spike disappeared and the compressor once again witnessed a modal type disturbance. Hub pressure fluctuations are observed throughout the process when the stall is caused by a modal wave while no disturbance is noticed at the hub in spike type stall inception. It is believed that the hub flow separation contributes to the modal type of stall inception phenomenon. Results are also supported by the recently developed signal processing techniques for the stall inception study.

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Tan, J., X. Wang, D. Qi e R. Wang. "The effects of radial inlet with splitters on the performance of variable inlet guide vanes in a centrifugal compressor stage". Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 225, n.º 9 (28 de junho de 2011): 2089–105. http://dx.doi.org/10.1177/0954406211407799.

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Variable inlet guide vanes (VIGVs) can regulate pressure ratio and mass flow at constant rotational speed in centrifugal compressors as a result of inducing a controlled prewhirl in front of impellers. Radial inlets and VIGVs are typical upstream components in front of the first-stage impellers in many industrial centrifugal compressors. However, previous investigations on VIGVs in centrifugal compressors were mostly conducted under the condition of axial inlets, and this study aims to focus on the effects of radial inlet on the VIGVs performance of a centrifugal compressor stage. The axial inlet stage model is compared with the radial inlet stage model with splitters using numerical flow simulation. The flow from the radial inlet was non-uniform in both circumferential and radial directions; thus, the VIGVs, the impeller, the vaneless diffuser, and the return vane channel are modelled with fully 360° passages. The three-dimensional (3D) flow field is numerically simulated at VIGVs setting angles ranging from - 20° to 60°. The overall stage performance parameters are obtained by integrating the field quantities. Though the splitters are equipped in the radial inlet, the overall stage polytropic efficiency decreases by an average of 4 per cent and total pressure ratio decreases by an average of 3.3per cent in comparison with the axial stage model. This can be attributed to the effect of both flow non-uniformity induced by radial inlet and flow loss in the radial inlet at different VIGV setting angles. The flow loss in the radial inlet with splitters is the main reason of the stage performance decrease compared with the flow non-uniformity. The simulation results show that the performance of VIGVs is degraded by its inlet flow distortions resulting from a radial inlet. The results in this study can be applied to centrifugal compressor design and optimization.

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Liu, ZX, HZ Diao, XC Zhu e 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, n.º 8 (28 de março de 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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Xu, Dengke, Chen He, Dakun Sun e Xiaofeng Sun. "Stall inception prediction of axial compressors with radial inlet distortions". Aerospace Science and Technology 109 (fevereiro de 2021): 106433. http://dx.doi.org/10.1016/j.ast.2020.106433.

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Klein, A. "The relation between losses and entry-flow conditions in short dump diffusers for combustors". Aeronautical Journal 92, n.º 920 (dezembro de 1988): 390–96. http://dx.doi.org/10.1017/s0001924000016523.

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SummaryAn experimental correlation is presented between the losses and the inlet flow conditions in short dump diffusers for turbojet combustors. Cascades of compressor blades upstream of the diffuser were used to make the flow field at inlet similar to that in a real jet engine. The flow field was altered in two ways — by varying the distance between the cascades and the diffuser inlet plane and by changing the blade aspect ratio. The measurements show clearly that distortions in the radial direction affect the losses to a much larger extent than non-uniformities in the circumferential direction. In consequence, the performance can be correlated to a satisfactory degree of accuracy simply by using the radial blockage factor at inlet.

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Han, Fenghui, Yijun Mao e Jiajian Tan. "Influences of flow loss and inlet distortions from radial inlets on the performances of centrifugal compressor stages". Journal of Mechanical Science and Technology 30, n.º 10 (outubro de 2016): 4591–99. http://dx.doi.org/10.1007/s12206-016-0930-y.

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Hynes, T. P., e E. M. Greitzer. "A Method for Assessing Effects of Circumferential Flow Distortion on Compressor Stability". Journal of Turbomachinery 109, n.º 3 (1 de julho de 1987): 371–79. http://dx.doi.org/10.1115/1.3262116.

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This paper describes the development of a new analysis to predict the onset of flow instability for an axial compressor operating in a circumferentially distorted inlet flow. A relatively simple model is used to examine the influence of various distortions in setting this instability point. It is found that the model reproduces known experimental trends for the loss of stability margin with increasing distortion amplitude and with changes in reduced frequency. In particular, there is a recognizable “critical sector angle” which characterizes loss of stability margin. To the authors’ knowledge, this is the first time the effects described herein have been theoretically demonstrated as the direct result of a fluid dynamic stability calculation.

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Voigt, Jonas, e Jens Friedrichs. "Development of a Multi-Segment Parallel Compressor Model for a Boundary Layer Ingesting Fuselage Fan Stage". Energies 14, n.º 18 (13 de setembro de 2021): 5746. http://dx.doi.org/10.3390/en14185746.

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The present methodological study aims to assess boundary layer ingestion (BLI) as a promising method to improve propulsion efficiency. BLI utilizes the low momentum inflow of the wing or fuselage boundary layer for thrust generation in order to minimize the required propulsive power for a given amount of thrust for wing or fuselage-embedded engines. A multi-segment parallel compressor model (PCM) is developed to calculate the power saving from full annular BLI as occurring at a fuselage tail center-mounted aircraft engine, employing radially subdivided fan characteristics. Applying this methodology, adverse effects on the fan performance due to varying inlet distortions depending on flight operating point as well as upstream boundary layer suction can be taken into account. This marks one step onto a further segmented PCM model for general cases of BLI-induced inlet distortion and allows the evaluation of synergies between combined BLI and active laminar flow control as a drag reduction measure. This study, therefore, presents one further step towards lower fuel consumption and, hence, a lower environmental impact of future transport aircraft.

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Zhang, Meijie, e Xinqian Zheng. "Criteria for the matching of inlet and outlet distortions in centrifugal compressors". Applied Thermal Engineering 131 (fevereiro de 2018): 933–46. http://dx.doi.org/10.1016/j.applthermaleng.2017.11.140.

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Bre´ard, C., M. Vahdati, A. I. Sayma e M. Imregun. "An Integrated Time-Domain Aeroelasticity Model for the Prediction of Fan Forced Response due to Inlet Distortion". Journal of Engineering for Gas Turbines and Power 124, n.º 1 (1 de fevereiro de 2000): 196–208. http://dx.doi.org/10.1115/1.1416151.

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The forced response of a low aspect-ratio transonic fan due to different inlet distortions was predicted using an integrated time-domain aeroelasticity model. A time-accurate, nonlinear viscous, unsteady flow representation was coupled to a linear modal model obtained from a standard finite element formulation. The predictions were checked against the results obtained from a previous experimental program known as “Augmented Damping of Low-aspect-ratio Fans” (ADLARF). Unsteady blade surface pressures, due to inlet distortions created by screens mounted in the intake inlet duct, were measured along a streamline at 85 percent blade span. Three resonant conditions, namely 1F/3EO, 1T & 2F/8EO and 2S/8EO, were considered. Both the amplitude and the phase of the unsteady pressure fluctuations were predicted with and without the blade flexibility. The actual blade displacements and the amount of aerodynamic damping were also computed for the former case. A whole-assembly mesh with about 2,000,000 points was used in some of the computations. Although there were some uncertainties about the aerodynamic boundary conditions, the overall agreement between the experimental and predicted results was found to be reasonably good. The inclusion of the blade motion was shown to have an effect on the unsteady pressure distribution, especially for the 2F/1T case. It was concluded that a full representation of the blade forced response phenomenon should include this feature.

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Manwaring, S. R., D. C. Rabe, C. B. Lorence e A. R. Wadia. "Structures and Dynamics Committee Best Paper of 1996 Award: Inlet Distortion Generated Forced Response of a Low-Aspect-Ratio Transonic Fan". Journal of Turbomachinery 119, n.º 4 (1 de outubro de 1997): 665–76. http://dx.doi.org/10.1115/1.2841176.

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This paper describes a portion of an experimental and computational program (ADLARF), which incorporates, for the first time, measurements of all aspects of the forced response of an airfoil row, i.e., the flow defect, the unsteady pressure loadings, and the vibratory response. The purpose of this portion was to extend the knowledge of the unsteady aerodynamics associated with a low-aspect-ratio transonic fan where the flow defects were generated by inlet distortions. Measurements of screen distortion patterns were obtained with total pressure rakes and casing static pressures. The unsteady pressure loadings on the blade were determined from high response pressure transducers. The resulting blade vibrations were measured with strain gages. The steady flow was analyzed using a three-dimensional Navier–Stokes solver while the unsteady flow was determined with a quasi-three-dimensional linearized Euler solver. Experimental results showed that the distortions had strong vortical, moderate entropic, and weak acoustic parts. The three-dimensional Navier–Stokes analyses showed that the steady flow is predominantly two-dimensional, with radially outward flow existing only in the blade surface boundary layers downstream of shocks and in the aft part of the suction surface. At near resonance conditions, the strain gage data showed blade-to-blade motion variations and thus, linearized unsteady Euler solutions showed poorer agreement with the unsteady loading data than comparisons at off-resonance speeds. Data analysis showed that entropic waves generated unsteady loadings comparable to vortical waves in the blade regions where shocks existed.

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Giel, P. W., J. R. Sirbaugh, I. Lopez e G. J. Van Fossen. "Three-Dimensional Navier–Stokes Analysis and Redesign of an Imbedded Bellmouth Nozzle in a Turbine Cascade Inlet Section". Journal of Turbomachinery 118, n.º 3 (1 de julho de 1996): 529–35. http://dx.doi.org/10.1115/1.2836699.

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Experimental measurements in the inlet of a transonic turbine blade cascade showed unacceptable pitchwise flow nonuniformity. A three-dimensional, Navier–Stokes computational fluid dynamics (CFD) analysis of the imbedded bellmouth inlet in the facility was performed to identify and eliminate the source of the flow nonuniformity. The blockage and acceleration effects of the blades were accounted for by specifying a periodic static pressure exit condition interpolated from a separate three-dimensional Navier–Stokes CFD solution of flow around a single blade in an infinite cascade. Calculations of the original inlet geometry showed total pressure loss regions consistent in strength and location to experimental measurements. The results indicate that the distortions were caused by a pair of streamwise vortices that originated as a result of the interaction of the flow with the imbedded bellmouth. Computations were performed for an inlet geometry that eliminated the imbedded bellmouth by bridging the region between it and the upstream wall. This analysis indicated that eliminating the imbedded bellmouth nozzle also eliminates the pair of vortices, resulting in a flow with much greater pitchwise uniformity. Measurements taken with an installed redesigned inlet verify that the flow nonuniformity has indeed been eliminated.

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Mirzaev, Z. N., M. S. Guseynov e T. G. Aigumov. "MATHEMATICAL MIXER MODEL WITH FORMATION OF HETERODYNE ANTIPHASE SIGNAL". Herald of Dagestan State Technical University. Technical Sciences 46, n.º 3 (24 de novembro de 2019): 97–107. http://dx.doi.org/10.21822/2073-6185-2019-46-3-97-107.

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Objectives To carry out calculations involved in the design of a microwave mixer with a diplexer with the formation of the antiphase heterodyne signal using a slot resonator.Method In order to calculate and optimise the characteristics, design and topological parameters of microwave mixers, the results of the design bandpass filter PF and low-pass filter (LPF) mixers of through-feed type were used. The characteristics of mixers and their structural elements were calculated using the Serenade software package intended for the automated calculation of microwave devices. A distinct feature of designing mixers (with a diplexer) involves the need to optimise the topology of the diplexer before optimising the mixer characteristics.Results The characteristics of nonlinear distortions show that the maximum power level at the mixer inlet should not exceed -15 – -20 dBm. In order to attenuate the intermodulation distortions of the 3rd order, this level should be higher that 50 dBs. The relatively low level of compression and suppression of harmonic and intermodulation distortions associated with the minimisation of the heterodyne power level at the calculation of characteristics of mixers of the required heterodyne power level (Rh ~ 5-7 dBm) is due to the minimum expenses at the realisation of sources of heterodyne signals. A noticeable improvement in the characteristics of mixers by nonlinear distortions can be achieved by shifting the operating point at the points on the current-voltage characteristic (VAC) diodes by an external voltage source with a simultaneous increase in Ph by several dB (up to Ph = 10 dBm).Conclusion A mode of increased nonlinear distortion suppression can be practically realised by switching on diodes through resistive-capacitive circuits (auto-shift) or using diodes with an increased potential barrier. The calculation shows that it is possible to realise sufficiently small conversion losses of 6.6-8.0 dB at low levels of Rh ~ 5-7 dBm.

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Rai, Man Mohan, e Robert P. Dring. "Navier-Stokes analyses of the redistribution of inlet temperature distortions in a turbine". Journal of Propulsion and Power 6, n.º 3 (maio de 1990): 276–82. http://dx.doi.org/10.2514/3.25431.

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Mao, Yinbo, e Thong Q. Dang. "A three-dimensional body-force model for nacelle-fan systems under inlet distortions". Aerospace Science and Technology 106 (novembro de 2020): 106085. http://dx.doi.org/10.1016/j.ast.2020.106085.

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Duvenhage, K., J. A. Vermeulen, C. J. Meyer e D. G. Kröger. "Flow distortions at the fan inlet of forced-draught air-cooled heat exchangers". Applied Thermal Engineering 16, n.º 8-9 (agosto de 1996): 741–52. http://dx.doi.org/10.1016/1359-4311(95)00063-1.

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MAAMMEUR, Moustafa, Abdallah BENAROUS, Ahmed BETTAHAR e Abdelkarim LIAZID. "Passive Flow Control for Centrifugal Compressors with Bents Intake Manifold". Mechanics 29, n.º 5 (18 de outubro de 2023): 369–76. http://dx.doi.org/10.5755/j02.mech.30987.

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Centrifugal compressors installations often require bents conduits at the intake due to space constraints. The intake bent manifold creates flow distortion affecting unfavorably the global performance of the compressor. The degradation caused by the inadequate inlet conditions due to turbulence and inlet flow disturbance, which negatively affects the flowing fluid through the curved intake manifold. Hence, significant distortions, within presence of single or twin core vortices whirling off center of the impeller intake alter flow incidence angle and consequently the impeller circumferential energy. Furthermore, inadequate bends configurations generate higher incidence angles, which lead to boundary layer separation and stalling. This paper deals with a 3D numerical investigations of dual bends located upstream of the centrifugal compressor. Several bend setting positions at the compressor intake describing various sequences of intake piping are considered. As first results, specific positions of dual bends are identified for fewer flow disturbance at the impeller eye and stable outflow controlling the incidence angel in a passive mode. it is less cumbersome and cost effective compared to additional devices. For practical industrial prediction, this investigation contribute also toward a mathematical correlation to check the adequacy of intermediate manifold configuration in order to predict the corresponding incidence angle at the impeller intake of the centrifugal compressor for enhanced performance and instantaneous control.

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Yamamoto, K., e W. F. Ng. "THE DEVELOPMENT OF FLOW DISTORTIONS AT THE FAN FACE FOR AN AXISYMMETRIC SUPERSONIC INLET". Journal of Sound and Vibration 203, n.º 1 (maio de 1997): 75–85. http://dx.doi.org/10.1006/jsvi.1996.0763.

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Sayma, A. I., M. Vahdati e M. Imregun. "Multi-bladerow fan forced response predictions using an integrated three-dimensional time-domain aeroelasticity model". Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 214, n.º 12 (1 de dezembro de 2000): 1467–83. http://dx.doi.org/10.1243/0954406001523425.

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A non-linear integrated aeroelasticity system to predict the forced vibration response of aero-engine fans is presented in this paper. The computational fluid dynamics (CFD) solver, which uses Favre-averaged Navier-Stokes equations on unstructured grids of mixed elements, is coupled to a modal model of the structure so that the effects of blade flexibility can be accommodated. The structural motion due to unsteady fluid forces is computed at every time step and the flow mesh is moved to follow the structure so that the resulting flow unsteadiness is determined in a time-accurate fashion. Two fan forced response case studies are reported in detail. The first one deals with a high-pressure ratio fan, the excitation being due to the upstream variable-angle inlet guide vanes (VIGVs). The unsteady flow analysis with blade motion was conducted using a sector of three VIGVs and four rotor blades. The wake predictions were found to be in good agreement with the corresponding laser measurements. The flow was observed to be completely separated for high VIGV angles and the excitation encompassed several harmonics. The predicted rotor blade vibration levels were generally found to be within 30 per cent of the measured values. The forced response to upstream obstructions was studied in the next fan case study. Three whole bladerows, consisting of 11 struts, 33 VIGVs and 26 rotor blades, were modelled in full. The model also included a prescribed inlet distortion pattern so that the combined effects of stator wakes and inlet distortion on the response of the rotor blades could be studied. The unsteady flow calculations were conducted using a time-accurate non-linear viscous flow representation. Blade motion was also included. Such an undertaking required about 4.2 million grid points to include all three bladerows in a complete stage calculation. To reduce the computational effort, a number of smaller computations were conducted by considering the stator and rotor domains separately: the outflow solution of the stator domain was used as an inflow boundary condition to the rotor domain. The results indicated that such isolated bladerow computations were likely to under-predict the response levels because of neglecting rotor-stator interactions. A number of low engine order (LEO) harmonics were identified from an inspection of the unsteady forcing created by the inlet distortions. Good agreement was obtained for cases where experimental data were available.

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Salta, C. A., e D. G. Kröger. "Effect of inlet flow distortions on fan performance in forced draught air-cooled heat exchangers". Heat Recovery Systems and CHP 15, n.º 6 (agosto de 1995): 555–61. http://dx.doi.org/10.1016/0890-4332(95)90065-9.

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Han, Fenghui, Zhe Wang, Yijun Mao, Jiajian Tan e Wenhua Li. "Experimental and numerical studies on the influence of inlet guide vanes of centrifugal compressor on the flow field characteristics of inlet chamber". Advances in Mechanical Engineering 12, n.º 11 (novembro de 2020): 168781402097490. http://dx.doi.org/10.1177/1687814020974909.

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Inlet chambers (IC) are the typical upstream component of centrifugal compressors, and inlet guide vanes in the IC have a great impact on its internal flow and aerodynamic loss, which will significantly influence the performance of the downstream compressor stages. In this paper, an experimental study was carried out on the flow characteristics inside a radial IC of an industrial centrifugal compressor, including five testing sections and 968 measuring points for two schemes with and without guide vanes. Detailed distributions of flow parameters on each section were obtained as well as the overall performance of the radial IC, and the causes of the flow loss inside the IC and the non-uniformity of flow parameters at the outlet section were investigated. Besides, numerical simulations were performed to further analyze the flow characteristics inside the radial IC. The experimental and numerical results indicate that, in the scheme without guide vanes, sudden expansions in the spiral channel and flow separations in the annular convergence channel are the major sources of flow loss and distortions generated in the radial IC; while in the scheme with guide vanes, the flow impacts, separations and wakes caused by the inappropriate design of guide vanes are the main reasons for the flow loss of the IC itself and the uneven flow distributions at the IC outlet.

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OOKITA, Akihiro, Shigeki MASUDA e Ichiro ARIGA. "Effects of inlet distortions on performance characteristics of a centrifugal compressor. Investigations concerning the unstable phenomenon." Transactions of the Japan Society of Mechanical Engineers Series B 51, n.º 472 (1985): 3970–79. http://dx.doi.org/10.1299/kikaib.51.3970.

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Mahalakshmi, N. V., G. Krithiga, S. Sandhya, J. Vikraman e V. Ganesan. "Experimental investigations of flow through conical diffusers with and without wake type velocity distortions at inlet". Experimental Thermal and Fluid Science 32, n.º 1 (outubro de 2007): 133–57. http://dx.doi.org/10.1016/j.expthermflusci.2007.02.008.

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Grenson, Pierre, e Eric Garnier. "Distortion reconstruction in S-ducts from wall static pressure measurements". International Journal of Numerical Methods for Heat & Fluid Flow 28, n.º 5 (8 de maio de 2018): 1134–55. http://dx.doi.org/10.1108/hff-06-2017-0232.

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Purpose This paper aims to report the attempts for predicting “on-the-fly” flow distortion in the engine entrance plane of a highly curved S-duct from wall static pressure measurements. Such a technology would be indispensable to trigger active flow control devices to mitigate the intense flow separations which occur in specific flight conditions. Design/methodology/approach Evaluation of different reconstruction algorithms is performed on the basis of data extracted from a Zonal Detached Eddy Simulation (ZDES) of a well-documented S-Duct (Garnier et al., AIAA J., 2015). Contrary to RANS methods, such a hybrid approach makes unsteady distortions available, which are necessary information for reconstruction algorithm assessment. Findings The best reconstruction accuracy is obtained with the artificial neural network (ANN) but the improvement compared to the classical linear stochastic estimation (LSE) is minor. The different inlet distortion coefficients are not reconstructed with the same accuracy. KA2 coefficient is finally identified as the more suited for activation of the control device. Originality/value LSE and its second-order variant (quadratic stochastic estimation [QSE]) are applied for reconstructing instantaneous stagnation pressure in the flow field. The potential improvement of an algorithm based on an ANN is also evaluated. The statistical link between the wall sensors and 40-Kulite rake sensors are carefully discussed and the accuracy of the reconstruction of the most used distortion coefficients (DC60, RDI, CDI and KA2) is quantified for each estimation technique.

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Hoyniak, D., e S. Fleeter. "Forced Response Analysis of an Aerodynamically Detuned Supersonic Turbomachine Rotor". Journal of Vibration and Acoustics 108, n.º 2 (1 de abril de 1986): 117–24. http://dx.doi.org/10.1115/1.3269311.

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High-performance aircraft engine fan and compressor blades are vulnerable to aerodynamically forced vibrations generated by inlet flow distortions due to wakes from upstream blade and vane rows, atmospheric gusts, and maldistributions in inlet ducts. In this paper, an analysis is developed to predict the flow-induced forced response behavior of an aerodynamically detuned rotor operating in a supersonic flow with a subsonic axial component. The aerodynamic detuning is achieved by alternating the circumferential spacing of adjacent rotor blades. The total unsteady aerodynamic loading acting on the blading, due to the convection of the transverse gust past the airfoil cascade and the resulting motion of the cascade, is developed in terms of influence coefficients. This analysis is then utilized to investigate the effect of aerodynamic detuning on the forced response characteristics of a 12-bladed rotor, with Verdon’s Cascade B flow geometry as a uniformly spaced baseline configuration. The results of this study indicate that for forward traveling wave gust excitations, aerodynamic detuning is generally very beneficial, resulting in significantly decreased maximum amplitude blade responses for many interblade phase angles.

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