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1
Guimaraes, Bucalo Tamara. "Fluid Dynamics of Inlet Swirl Distortions for Turbofan Engine Research." Diss., Virginia Tech, 2018. http://hdl.handle.net/10919/82921.
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Significant effort in the current technological development of aircraft is aimed at improving engine efficiency, while reducing fuel burn, emissions, and noise levels. One way to achieve these is to better integrate airframe and propulsion system. Tighter integration, however, may also cause adverse effects to the flow entering the engines, such as total pressure, total temperature, and swirl distortions. Swirl distortions are angular non-uniformities in the flow that may alter the functioning of specific components of the turbomachinery systems. To investigate the physics involved in the ingestion of swirl, pre-determined swirl distortion profiles were generated through the StreamVane method in a low-speed wind tunnel and in a full-scale turbofan research engine. Stereoscopic particle image velocimetry (PIV) was used to collect three-component velocity fields at discrete planes downstream of the generation of the distortions with two main objectives in mind: identifying the physics behind the axial development of the distorted flow; and describing the generation of the distortion by the StreamVane and its impact to the flow as a distortion generating device.
Analyses of the mean velocity, velocity gradients, and Reynolds stress tensor components in these flows provided significant insight into the driving physics. Comparisons between small-scale and full-scale results showed that swirl distortions are Mach number independent in the subsonic regime. Reynolds number independence was also verified for the studied cases. The mean secondary flow and flow angle profiles demonstrated that the axial development of swirl distortions is highly driven by two-dimensional vortex dynamics, when the flow is isolated from fan effects. As the engine fan is approached, the vortices are axially stretched and stabilized by the acceleration of the flow. The flow is highly turbulent immediately downstream of the StreamVane due to the presence of the device, but that vane-induced turbulence mixes with axial distance, so that the device effects are attenuated for distances greater than a diameter downstream, which is further confirmed by the turbulent length scales of the flow. These results provide valuable insight into the generation and development of swirl distortion for ground-testing environments, and establishes PIV as a robust tool for engine inlet investigations.<br>Ph. D.
2
Gong, Yifang 1964. "A computational model for rotating stall and inlet distortions in multistage compressors." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/9733.
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Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.<br>"February 1999."<br>Includes bibliographical references (p. 175-182).<br>This thesis presents the conceptualization and development of a computational model for describing three-dimensional non-linear disturbances associated with instability and inlet distortion in multistage compressors. Specifically, the model is aimed at simulating the non-linear aspects of short wavelength stall inception, part span stall cells, and compressor response to three-dimensional inlet distortions. The computed results demonstrated the first-of-a-kind capability for simulating short wavelength stall inception in multistage compressors. The adequacy of the model is demonstrated by its application to reproduce the following phenomena: (1) response of a compressor to a square-wave total pressure inlet distortion; (2) behavior of long wavelength small amplitude disturbances in compressors; (3) short wavelength stall inception in a multistage compressor and the occurrence of rotating stall inception on the negatively sloped portion of the compressor characteristic; ( 4) progressive stalling behavior in the first stage in a mismatched multistage compressor; (5) change of stall inception type (from modal to spike and vice versa) due to IGV stagger angle variation, and "unique rotor tip incidences at these points where the compressor stalls through short wavelength disturbances. The model has been applied to determine the parametric dependence of instability inception behavior in terms of amplitude and spatial distribution of initial disturbance, and intra-blade-row gaps. It is found that reducing the inter-blade row gaps suppresses the growth of short wavelength disturbances. It is also concluded from these parametric investigations that each local component group (rotor and its two adjacent stators) has its own instability point (i.e. conditions at which disturbances are sustained) for short wavelength disturbances, with the instability point for the compressor set by the most unstable component group. For completeness, the methodology has been extended to describe finite amplitude disturbances in high-speed compressors. Results are presented for the response of a transonic compressor subjected to inlet distortions.<br>by Yifang Gong.<br>Ph.D.
3
Eisemann, Kevin Michael. "A Computational Study of Compressor Inlet Boundary Conditions with Total Temperature Distortions." Thesis, Virginia Tech, 2005. http://hdl.handle.net/10919/35969.
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A three-dimensional CFD program was used to predict the flow field that would enter a downstream fan or compressor rotor under the influence of an upstream thermal distortion. Two distortion generation techniques were implemented in the model; (1) a thermal source and (2) a heated flow injection method. Results from the investigation indicate that both total pressure and velocity boundary conditions at the compressor face are made non-uniform by the upstream thermal distortion, while static pressure remains nearly constant. Total pressure at the compressor face was found to vary on the order of 10%, while velocity varies from 50-65%. Therefore, in modeling such flows, neither of these latter two boundary conditions can be assumed constant under these conditions.
The computational model results for the two distortion generation techniques were compared to one another and evaluations of the physical practicality of the thermal distortion generation methods are presented. Both thermal distortion methods create total temperature distortion magnitudes at the compressor face that may affect rotor blade vibration. Both analyses show that holding static pressure constant is an appropriate boundary condition for flow modeling at the compressor inlet. The analyses indicate that in addition to the introduction of a thermal distortion, there is a potential to generate distortion in total pressure, Mach number, and velocity. Depending on the method of thermally distorting the inlet flow, the flow entering the compressor face may be significantly non-uniform.
The compressor face boundary condition results are compared to the assumptions of a previous analysis (Kenyon et al., 2004) in which a 25 R total temperature distortion was applied to a computational fluid dynamics (CFD) model of a fan geometry to obtain unsteady blade pressure loading. Results from the present CFD analyses predict similar total temperature distortion magnitudes corresponding to the total temperature variation used in the Kenyon analyses. However, the results indicate that the total pressure and circumferential velocity boundary conditions assumed uniform in the Kenyon analyses could vary by the order of 2% in total pressure and approximately 8% in velocity distortion. This supports the previously stated finding that assuming a uniform total pressure profile at the compressor inlet may be an appropriate approximation with the presence of a weak thermal distortion, while assuming a constant circumferential velocity boundary condition is likely not sufficiently accurate for any thermal distortion.
In this work, the referenced Kenyon investigation and others related to the investigation of distortion-induced aeromechanical effects in this compressor rotor have assumed no aerodynamic coupling between the duct flow and the rotor. A full computational model incorporating the interaction between the duct flow and the fan rotor would serve to alleviate the need for assuming boundary conditions at the compressor inlet.<br>Master of Science
4
Giuliani, James Edward. "Jet Engine Fan Response to Inlet Distortions Generated by Ingesting Boundary Layer Flow." The Ohio State University, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=osu1468564279.
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Lambie, David. "Inlet distortion and turbofan engines." Thesis, University of Cambridge, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.305300.
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Longley, John Peter. "Inlet distortion and compressor stability." Thesis, University of Cambridge, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.304354.
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Hoopes, Kevin M. "A New Method for Generating Swirl Inlet Distortion for Jet Engine Research." Thesis, Virginia Tech, 2013. http://hdl.handle.net/10919/49545.
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Jet engines operate by ingesting incoming air, adding momentum to it, and exhausting it through a nozzle to produce thrust. Because of their reliance on an inlet stream, jet engines are very sensitive to inlet flow nonuniformities. This makes the study of the effects of inlet nonuniformities essential to improving jet engine performance. Swirl distortion is the presence of flow angle nonuniformity in the inlet stream of a jet engine. Although several attempts have been made to accurately reproduce swirl distortion profiles in a testing environment, there has yet to be a proven method to do so.<br /><br />A new method capable of recreating any arbitrary swirl distortion profile is needed in order to expand the capabilities of inlet distortion testing. This will allow designers to explore how an engine would react to a particular engine airframe combination as well as methods for creating swirl distortion tolerant engines. The following material will present such a method as well as experimental validation of its effectiveness.<br>Master of Science
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Dosne, Cyril. "Development and implementation of adjoint formulation of explicit body-force models for aero-propulsive optimizations." Electronic Thesis or Diss., Institut polytechnique de Paris, 2024. http://www.theses.fr/2024IPPAX026.
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Dans le domaine de l’aéronautique civile, les études de plus en plus nombreuses portant sur les nouveaux systèmes moteurs, tels que les turbofans à très haut taux de dilution et les open-rotors, ainsi que sur les architectures d'intégration motrice innovantes, telles que la propulsion distribuée ou les systèmes à ingestion de couche limite, nécessitent une modélisation couplée de l’aérodynamique externe et du système propulsif, et ce dès les premiers stades de la conception. Les modèles body-force se sont avérés capables de reproduire fidèlement la majeure partie des phénomènes de couplage aéro-propulsif, comme la réponse aérodynamique du moteur aux distorsions d’entrée d’air, et ce à un coût de calcul réduit. Cependant, ils manquent d'une formulation adjointe pour être employés efficacement dans des optimisations par gradient. Cette thèse de doctorat se concentre sur le développement d'une approche adjointe pour les modèles body-force explicites. Tout d'abord, plusieurs optimisations aéro-propulsives sont menées sur une configuration académique de propulsion distribuée, à l'aide d'un modèle body-force réduit. Malgré la simplicité de ce modèle (d’intérêt pour les études de conception amont), une réduction de 10,5 % de la consommation de puissance est obtenue. Le potentiel de cette nouvelle méthodologie est ensuite évalué pour l'optimisation préliminaire de compresseurs, d'abord sans distorsion d’entrée d’air. Le modèle body-force de Hall est considéré pour cette étude. Les gradients de forme des aubes calculés à l’aide de l’adjoint body-force sont comparés à ceux obtenus via des simulations de haute-fidélité. Les résultats obtenus révèlent une très bonne capacité de prédiction des gradients du rotor par l’adjoint body-force, pour une grande partie de la caractéristique du compresseur, et particulièrement pour les points de fonctionnement situés entre le pompage et la zone de fonctionnement nominal du compresseur. En revanche, la précision de ces gradients est réduite à proximité du blocage. Pour le stator, seuls les effets liés à la désadaptation de l’aube au flux incident peuvent être captés. L’optimisation conduite avec l’adjoint body-force au point de fonctionnement nominal a permis d’améliorer le rendement du compresseur, ce qui a été confirmé par des simulations de haute-fidélité. Sous distorsion radiale, la méthode adjointe du body-force s’est à nouveau révélée capable d’améliorer les performances du compresseur en adaptant la géométrie des aubages aux perturbations d’entrée d’air. Les analyses haute-fidélité conduites sur les géométries obtenues par optimisations utilisant l’adjoint body-force montrent une augmentation du rendement isentropique comprise entre 1,16 et 1,47%, selon la formulation du problème d’optimisation retenue. Enfin, une optimisation du compresseur a été conduite à l’aide de l’adjoint body-force dans le cas d’une distorsion s’étendant sur la totalité de la circonférence de l’entrée d’air. Ces résultats sont très prometteurs et les observations effectuées sont cohérentes avec celles disponibles dans la communauté scientifique et obtenue à l’aide de calcul de haute-fidélité<br>In civil aviation, the increasing exploration of innovative engine systems – such as ultra-high bypass ratio turbofan or open-rotor – and breakthrough engine-integration architectures – such as distributed propulsion or boundary-layer ingestion – require a coupled modeling of the aerodynamic and propulsion subsystems from the earliest design stages. Body-force models have proven capable of faithfully reproducing most of the coupling phenomena, such as the engine response to inlet flow distortions, at reduced computational cost. However, they lack an adjoint formulation to be efficiently used in gradient-based optimizations. The present PhD thesis focuses on the development of an adjoint approach for explicit body-force models. First, aero-propulsive optimizations of an academic distributed propulsion configuration are conducted using a lumped body-force model. Despite the simplicity of this model (of interest for conceptual design studies), 10.5% decrease in power consumption is achieved. Then the potential of this new methodology is investigated for the preliminary optimization of compressor stages, at first under clean inflow conditions. The Hall body-force model is considered for such purpose. The comparison of the blade shape gradients computed by the adjoint body-force with high-fidelity ones, obtained from blade-resolved computations, shows very good prediction for the rotor. This is observed over a large portion of the compressor characteristic, especially between near-design and surge operating conditions, while accuracy is reduced near the blockage. On the contrary, for stator shape gradients, only flow misalignment effects can be captured. At design conditions, the improvement of the compressor efficiency obtained by the adjoint body-force optimization has been confirmed through high-fidelity simulations. Optimization under radial inlet distortion are then investigated. Once again, the adjoint body-force approach is found capable of enhancing the compressor performances, by adapting its geometry to the off-design inflow conditions. According to high-fidelity analysis of the body-force optimized blade geometry, an increase in compressor isentropic efficiency between 1.16 and 1.47% is achieved, given the formulation of the optimization problem. Finally, an optimization of the compressor under full-annulus inlet distortion is conducted leading to very promising results, which are consistent with those found in the literature using advanced simulations
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Van, Schalkwyk Christiaan Mauritz. "Active control rotating stall with inlet distortion." Thesis, Massachusetts Institute of Technology, 1996. http://hdl.handle.net/1721.1/10827.
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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1996.<br>Includes bibliographical references (p. 179-182).<br>by Christiaan Mauritz Van Schalkwyk.<br>Ph.D.
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Papamarkos, Ioannis. "Inlet distortion generation for a transonic compressor." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2004. http://library.nps.navy.mil/uhtbin/hyperion/04Sep%5FPapamarkos.pdf.
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Shang, Tonghuo. "Influence of inlet temperature distortion on turbine heat transfer." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/47370.
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Ryman, John Franklin. "Prediction of Inlet Distortion Transfer Through the Blade Rows in a Transonic Axial Compressor." Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/43207.
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Inlet total pressure non-uniformities in axial flow fans and compressors can contribute to the loss of component structural integrity through high cycle fatigue (HCF) induced by the excitation of blade vibratory modes. As previous research has shown total pressure distortion to be the dominant HCF driver in aero engines [Manwaring et al, 1997], an understanding of its transfer through, and impact on, subsequent turbomachine stages and engine components is an important topic for assessment. Since current modeling techniques allow for total pressure distortion magnitudes to be directly related to blade vibratory response, the prediction of downstream distortion patterns from an upstream measurement would allow for the inference of the vibratory response of downstream blade rows to an inlet total pressure distortion.
Nonlinear Volterra theory can be used to model any periodic nonlinear system as an infinite sum of multidimensional convolution integrals. A semi-empirical model has been developed using this theory by assuming that a distortion waveform is a periodic signal that is being presented to a nonlinear system, the compressor being the system. The use of Volterra theory in nonlinear system modeling relies on the proper identification of the Volterra kernels, which make up the transfer function that defines the systemâ s impulse response characteristics. Once the kernels of a system are properly identified, the systemâ s response can be calculated for any arbitrary input. This model extracts these kernels from upstream and downstream total pressure distortion measurements of a transonic rotor of modern design. The resulting transfer function is then applied to predict distortion transfer at new operating points on the same rotor and compared with the measured data.
The judicious choice of distortion measurement data allows predictions of the downstream distortion content based on a measured non-uniform inlet flow at conditions different from those at which the transfer function was derived. This allows for the determination of downstream total pressure distortion that has the potential to excite blade vibratory modes that could lead to HCF under operating conditions other than those at which the data was taken, such as varying inlet distortion patterns, mass flow settings, rotational speeds, and inlet geometry.
This report presents the creation of a Volterra model in order to predict distortion transfer in axial flow fans and compressors. This model, in three variations, is applied to a variety of distortions and compressor operating conditions as measured in the ADLARF tests at the Compressor Research Facility. Predictions are compared with data from the test and final results are also compared with two previous studies conducted at Virginia Tech using the same experimental data. Using the Volterra model it is shown that, with appropriate limitations, distortion transfer can be predicted for flow conditions different from those used for calibration. The model is considered useful for both performance and HCF investigations.<br>Master of Science
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Zemp, Armin. "CFD investigation on inlet flow distortion in a centrifugal compressor." Zürich : Swiss Federal Institute of Technology, Turbomachinery Laboratory, 2007. http://e-collection.ethbib.ethz.ch/show?type=dipl&nr=308.
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Schulmeyer, Andreas. "Enhanced compressor distortion tolerance using asymmetric inlet guide vane stagger." Thesis, Massachusetts Institute of Technology, 1989. http://hdl.handle.net/1721.1/40144.
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Frohnapfel, Dustin Joseph. "Experimental Investigation of Fan Rotor Response to Inlet Swirl Distortion." Thesis, Virginia Tech, 2016. http://hdl.handle.net/10919/71323.
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Next generation aircraft design focuses on highly integrated airframe/engine architectures that exploit advantages in system level efficiency and performance. One such design concept incorporates boundary layer ingestion which locates the turbofan engine inlet near enough to the lifting surface of the aircraft skin that the boundary layer is ingested and reenergized. This process reduces overall aircraft drag and associated required thrust, resulting in fuel savings and decreased emissions; however, boundary layer ingestion also creates unique challenges for the turbofan engines operating in less than optimal inlet flow conditions.
The engine inlet flow profiles predicted from boundary layer ingesting aircraft architectures contain complex distortions that affect the engine operability, durability, efficiency, and performance. One component of these complex distortion profiles is off-axial secondary flow, commonly referred to as swirl. As a means to investigate the interactions of swirl distortion with turbofan engines, an experiment was designed to measure distorted flow profiles in an operating turbofan research engine.
Three-dimensional flow properties were measured at discrete planes immediately upstream and immediately downstream of the fan rotor, isolating the component for analysis. Constant speed tests were conducted under clean and distorted test conditions. For clean tests, a straight cylindrical inlet duct was attached to the fan case; for distorted tests, a StreamVane swirl distortion generator was inserted into the inlet duct. The StreamVane was designed to induce a swirl distortion matching results of computation fluid dynamics models of a conceptual blended wing body aircraft at a plane upstream of the fan. The swirl distortion was then free to develop naturally within the inlet duct before being ingested by the engine.
Results from the investigation revealed that the generated swirl profile developed, mixed, and dissipated in the inlet duct upstream of the fan. Measurements immediately upstream of the fan rotor leading edge revealed 50% reduction in measured flow angle magnitudes along with evidence of fanwise vortex convection when compared to the StreamVane design profile. The upstream measurements also indicated large amounts of secondary flow entered the fan rotor. Measurements immediately downstream of the fan rotor trailing edge demonstrated that the fan processed the distortion and further reduced the intensity of the swirl; however, non-uniform secondary flow persisted at this plane. The downstream measurements confirmed that off-design conditions entered the fan exit guide vanes, likely contributing to cascading performance deficiencies in downstream components and reducing the performance of the propulsor system.<br>Master of Science
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Lucas, James Redmond. "Effect of BLI-Type Inlet Distortion on Turbofan Engine Performance." Thesis, Virginia Tech, 2013. http://hdl.handle.net/10919/23272.
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Boundary Layer Ingestion (BLI) is currently being researched as a potential method to improve efficiency and decrease emissions for the next generation of commercial aircraft. While re-energizing the boundary layer formed over the fuselage of an aircraft has many system level benefits, ingesting the low velocity boundary layer flow through a serpentine inlet into a turbofan engine adversely affects the performance of the engine. The available literature has only yielded studies of the effects of this specific type of inlet distortion on engine performance in the form of numerical simulations. This work seeks to provide an experimental analysis of the effects of BLI-type distortion on a turbofan engine\'s performance. A modified JT15D-1 turbofan engine was investigated in this study. Inlet flow distortion was created by a layered wire mesh distortion screen designed to create a total pressure distortion profile at the aerodynamic interface plane (AIP) similar to NASA\'s Inlet A boundary layer ingesting inlet flow profile. Results of this investigation showed a 15.5% decrease in stream thrust and a 14% increase in TSFC in the presence of BLI-type distortion. <br /><br />Flow measurements at the AIP and the bypass nozzle exit plane provided information about the losses throughout the fan flow path. The presence of the distortion screen resulted in a 24% increase in mass-averaged entropy production along the entire fan flow path compared to the non-distorted test. A mass-averaged fan flow path efficiency was also calculated assuming an isentropic process as ideal. The non-distorted fan flow path efficiency was computed to be 60%, while the distorted fan flow path efficiency was computed to be 50.5%, a reduction in efficiency of 9.5%. The entropy generation between ambient conditions and the AIP was compared to the entropy production along the entire fan flow path. It was found that the majority of entropy generation occurred between the AIP and bypass nozzle exit. Based on flow measurements at the bypass nozzle exit plane, it was concluded that inlet flow distortion should be located away from the tip region of the fan in order to minimize losses in a very lossy region. It was also determined that the fan and bypass duct process the different regions of the total pressure distortion in different ways. In some regions the entropy production decreased for the distorted test compared to the clean test, while in other regions the entropy production increased for the distorted test compared to the clean test. Finally, it was found that small improvements in total pressure and total temperature variation at the bypass nozzle exit plane will greatly improve the fan flow path efficiency and entropy generation, thereby decreasing performance losses.<br /><br>Master of Science
17
Smith, Katherine Nicole. "New Methodology for the Estimation of StreamVane Design Flow Profiles." Thesis, Virginia Tech, 2018. http://hdl.handle.net/10919/82039.
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Inlet distortion research has become increasingly important over the past several years as demands for aircraft flight efficiency and performance has increased. To accommodate these demands, research progression has shifted the emphasis onto airframe-engine integration and improved understanding of engine operability in less than ideal conditions. Swirl distortion, which is considered a type of non-uniform inflow inlet distortion, is characterized by the presence of swirling flow in an inlet. The presence of swirling flow entering an engine can affect the compression systems performance and operability, therefore it is an area of current research.
A swirl distortion generation device created by Virginia Tech, identified as the StreamVane, has the ability to produce various swirl distortion flow profiles. In its current state, the StreamVane methodology generates a design swirl distortion at the trailing edge of the device. However, in many applications the plane at which the researcher wants a desired distortion is downstream of the StreamVane trailing edge. After the distortion is discharged from the StreamVane it develops as it moves downstream. Therefore, to more accurately replicate a desired swirl distortion at a given downstream plane, distortion development downstream of the StreamVane must be considered.
Currently Virginia Tech utilizes a numerical modeling design tool, designated StreamFlow, that generates predictions of how a StreamVane-generated distortion propagates downstream. However, due to the non-linear physics of the flow problem, StreamFlow cannot directly calculate an accurate inverse solution that can predict upstream conditions from a downstream boundary, as needed to design a StreamVane. To solve this problem, in this research, an efficient estimation process has been created, combining the use of the StreamFlow model with a Markov Chain Monte Carlo (MCMC) parameter estimation tool to estimate upstream flow profiles that will produce the desired downstream profiles. The process is designated the StreamFlow-MC2 Estimation Process.
The process was tested on four fundamental types of swirl distortions. The desired downstream distortion was input into the estimation process to predict an upstream profile that would create the desired downstream distortion. Using the estimated design profiles, 6-inch diameter StreamVanes were designed then wind tunnel tested to verify the distortion downstream. Analysis and experimental results show that using this method, the upstream distortion needed to create the desired distortion was estimated with excellent accuracy. Based on those results, the StreamFlow-MC2 Estimation Process was validated.<br>Master of Science
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Orme, Andrew Dallin. "Analysis of Inlet Distortion Patterns on Distortion Transfer and Generation Through a Highly Loaded Fan Stage." BYU ScholarsArchive, 2020. https://scholarsarchive.byu.edu/etd/8649.
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Characterization of distortion transfer and generation through fans with distorted inlet conditions enables progress towards designs with improved distortion tolerance. The abruptness of transition from undistorted to distorted total pressure regions at the inlet impacts the induced swirl profile and therefore the distortion transfer and generation. These impacts are characterized using URANS simulations of PBS Rotor 4 geometry under a variety of inlet distortion profiles. A 90° and a 135° sector, both of 15% total pressure distortion, are considered. Variants of each sector size, with decreasing levels of distortion transition abruptness, are each applied to the fan. Fourier-based distortion descriptors are used to quantify levels of distortion transfer and generation at axial locations through the fan, principally at the stator inlet. It is shown that a gradual transition in distortion at the inlet results in decreased levels of distortion transfer and generation. The flow physics resulting in this reduction are explored. URANS simulations involving turbomachinery are complex and often require simplifying assumptions to balance computational costs with accuracy. One assumption removes the need for a nozzle to control nozzle operation condition and replaces it with a static pressure boundary condition located at the stator exit. This assumption is challenged by conducting a series of distorted inlet simulations with a nozzle, which are then compared to a corresponding set of simulations conducted using the exit boundary assumption. Performance parameters for each set of simulations are compared. Performance was observed to be within 1% difference between the two methods, supporting the assumption that a static pressure boundary is adequate for controlling inlet distortion simulations. Finally, full annulus URANS simulations are presented to investigate distortion phase shift in a single stage transonic fan. The fan is subject to a 90° sector inlet total pressure distortion. Simulation results are presented for choke, design, and near-stall operating conditions. Circumferential profiles of swirl, total pressure, total temperature, power, and phase shift are analyzed at 10%, 30%, 50%, 70%, and 90% span. Several metrics for phase shift, which is a measure of the rotational translation of a distortion profile, are presented and compared. Each aims to assist understanding the translational motion of distortion as it passes through the fan. The different metrics used for phase shift are used to analyze distortion phase. Insights from each are presented alongside limitations for each method. A combination of methods is proposed to address their respective limitations.
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Baig, Aman uz zaman. "Adjoint Design Optimization for Boundary Layer Ingesting Inlet Guide Vanes with Distorted Inlet Profiles in SU2." University of Cincinnati / OhioLINK, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1613686047734509.
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Anderson, Jason Mitchell. "Non-Intrusive Sensing and Feedback Control of Serpentine Inlet Flow Distortion." Diss., Virginia Tech, 2003. http://hdl.handle.net/10919/27120.
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A technique to infer circumferential total pressure distortion intensity found in serpentine inlet airflow was established using wall-pressure fluctuation measurements. This sensing technique was experimentally developed for aircraft with serpentine inlets in a symmetric, level flight condition. The turbulence carried by the secondary flow field that creates the non-uniform total pressure distribution at the compressor fan-face was discovered to be an excellent indicator of the distortion intensity. A basic understanding of the secondary flow field allowed for strategic sensor placement to provide a distortion estimate with a limited number of sensors. The microphone-based distortion estimator was validated through its strong correlation with experimentally determined circumferential total pressure distortion parameter intensities (DPCP).
This non-intrusive DPCP estimation technique was then used as a DPCP observer in a distortion feedback control system. Lockheed Martin developed the flow control technique used in this control system, which consisted of jet-type vortex generators that injected secondary flow to counter the natural secondary flow inherent to the serpentine inlet. A proportional-integral-derivative (PID) based control system was designed that achieved a requested 66% reduction in DPCP (from a DPCP of 0.023 down to 0.007) in less than 1 second. This control system was also tested for its ability to maintain a DPCP level of 0.007 during a quick ramp-down and ramp-up engine throttling sequence, which served as a measure of system robustness. The control system allowed only a maximum peak DPCP of 0.009 during the engine ramp-up. The successful demonstrations of this automated distortion control system showed great potential for applying this distortion sensing scheme along with Lockheed Martinâ s flow control technique to military aircraft with serpentine inlets.
A final objective of this research was to broaden the non-intrusive sensing capabilities in the serpentine inlet. It was desired to develop a sensing technique that could identify control efforts that optimized the overall inlet aerodynamic performance with regards to both circumferential distortion intensity DPCP and average pressure recovery PR. This research was conducted with a new serpentine inlet developed by Lockheed Martin having a lower length-to-diameter ratio and two flow control inputs. A cost function based on PR and DPCP was developed to predict the optimal flow control efforts at several Mach numbers. Two wall-mounted microphone signals were developed as non-intrusive inlet performance sensors in response to the two flow control inputs. These two microphone signals then replaced the PR and DPCP metrics in the original cost function, and the new non-intrusive-based cost function yielded extremely similar optimal control efforts.<br>Ph. D.
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Warfield, Zachary (Zachary Greene) 1976. "Active control of separation induced distortion in a scaled tactial aircraft inlet." Thesis, Massachusetts Institute of Technology, 2001. http://hdl.handle.net/1721.1/82237.
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Walter, S. F. "Optimization of pressure probe placement and data analysis of engine-inlet distortion." Thesis, University of Colorado at Boulder, 2017. http://pqdtopen.proquest.com/#viewpdf?dispub=10244536.
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<p> The purpose of this research is to examine methods by which quantification of inlet flow distortion may be improved upon. Specifically, this research investigates how data interpolation effects results, optimizing sampling locations of the flow, and determining the sensitivity related to how many sample locations there are. The main parameters that are indicative of a "good" design are total pressure recovery, mass flow capture, and distortion. This work focuses on the total pressure distortion, which describes the amount of non-uniformity that exists in the flow as it enters the engine. All engines must tolerate some level of distortion, however too much distortion can cause the engine to stall or the inlet to unstart. Flow distortion is measured at the interface between the inlet and the engine.</p><p> To determine inlet flow distortion, a combination of computational and experimental pressure data is generated and then collapsed into an index that indicates the amount of distortion. Computational simulations generate continuous contour maps, but experimental data is discrete. Researchers require continuous contour maps to evaluate the overall distortion pattern. There is no guidance on how to best manipulate discrete points into a continuous pattern. Using one experimental, 320 probe data set and one, 320 point computational data set with three test runs each, this work compares the pressure results obtained using all 320 points of data from the original sets, both quantitatively and qualitatively, with results derived from selecting 40 grid point subsets and interpolating to 320 grid points. Each of the two, 40 point sets were interpolated to 320 grid points using four different interpolation methods in an attempt to establish the best method for interpolating small sets of data into an accurate, continuous contour map. Interpolation methods investigated are bilinear, spline, and Kriging in Cartesian space, as well as angular in polar space. Spline interpolation methods should be used as they result in the most accurate, precise, and visually correct predictions when compared results achieved from the full data sets.</p><p> Researchers were interested if fewer than the recommended 40 probes could be used – especially when placed in areas of high interest - but still obtain equivalent or better results. For this investigation, the computational results from a two-dimensional inlet and experimental results of an axisymmetric inlet were used. To find the areas of interest, a uniform sampling of all possible locations was run through a Monte Carlo simulation with a varying number of probes. A probability density function of the resultant distortion index was plotted. Certain probes are required to come within the desired accuracy level of the distortion index based on the full data set. For the experimental results, all three test cases could be characterized with 20 probes. For the axisymmetric inlet, placing 40 probes in select locations could get the results for parameters of interest within less than 10% of the exact solution for almost all cases. For the two dimensional inlet, the results were not as clear. 80 probes were required to get within 10% of the exact solution for all run numbers, although this is largely due to the small value of the exact result.</p><p> The sensitivity of each probe added to the experiment was analyzed. Instead of looking at the overall pattern established by optimizing probe placements, the focus is on varying the number of sampled probes from 20 to 40. The number of points falling within a 1\% tolerance band of the exact solution were counted as good points. The results were normalized for each data set and a general sensitivity function was found to determine the sensitivity of the results. A linear regression was used to generalize the results for all data sets used in this work. However, they can be used by directly comparing the number of good points obtained with various numbers of probes as well. The sensitivity in the results is higher when fewer probes are used and gradually tapers off near 40 probes. There is a bigger gain in good points when the number of probes is increased from 20 to 21 probes than from 39 to 40 probes.</p>
23
Wallace, Robert Malcolm. "Modal Response of a Transonic Fan Blade to Periodic Inlet Pressure Distortion." Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/35158.
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A new method for predicting forced vibratory blade response to total pressure distortion has been developed using modal and harmonic analysis. Total pressure distortions occur in gas turbine engines when the incoming airflow is partially blocked or disturbed. Distorted inlet conditions can have varying effects on engine performance and engine life. Short-term effects are often in the form of performance degradation where the distorted airflow causes a loss in pressure rise, and a reduction in mass flow and stall margin. Long-term effects are a result of vibratory blade response that can ultimately lead to high cycle fatigue (HCF), which in turn can quickly cause partial damage to a single blade or complete destruction of an entire compressor blade row, leading to catastrophic failure of the gas turbine engine. A better understanding and prediction of vibratory blade response is critical to extending engine life and reducing HCF-induced engine failures.
This work covers the use of finite element modeling coupled with computational fluid dynamics-generated pressure fields to create a generalized forcing function. The first three modes of a low-aspect-ratio, transonic, first stage blade of a two-stage fan were examined. The generalized forcing function was decomposed to the frequency domain to identify the dominant harmonic magnitude present, as well as other contributing harmonics. An attempt to define the relationship between modal force with varying total pressure distortion levels produced a sensitivity factor that describes the relationship in the form of a simple multiplier. A generalized force was applied to the blade and varied harmonically across a frequency range known to contain the first natural frequency. The mean rotor stress variation was recorded and compared to experimental results to validate the accuracy of the model and verify its ability to predict vibratory blade response accurately.<br>Master of Science
24
Reilly, Daniel Oliver. "Inlet Distortion Effects on the Unsteady Aerodynamics of a Transonic Fan Stage." Wright State University / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=wright1482139741887976.
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Horvath, Nathan Rosendo. "Inlet Vortex Formation Under Crosswind Conditions." Digital WPI, 2013. https://digitalcommons.wpi.edu/etd-theses/302.
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A jet engine operating near the ground at low aircraft speeds, high thrust, and subject to a crosswind, can experience a flow separation region on the windward inlet lip and the formation of a vortex that extends from the ground to the engine fan face, known as the inlet vortex. This structure forms from a single point on the ground and is ingested by the engine. Inlet vortices are often observed during engine power-up at the start of the take-off run. They create considerable stagnation pressure losses and flow distortions at the engine fan face, compromising fan efficiency, thrust, and increasing the potential for compressor surge. Inlet vortices have enough suction power to kick up sand and rocks that are then sucked into the engine when an aircraft is operating near the ground and especially over poorly-maintained tarmac. Thus foreign object damage (FOD) becomes a serious threat for an engine under these conditions, and may lead to compressor blade erosion, deteriorating engine performance and reducing service life. The work presented here used ANSYS FLUENT to model a jet engine under crosswind. The 3-D Navier-Stokes equations were solved for compressible, unsteady flow. The mesh generated contained 5.6 million tetrahedral and wedge elements. The goal of this research was to better understand the inlet vortex formation mechanisms by studying its transient formation process, and to provide new information for future development of vortex prevention techniques. This work has shown multiple smaller inlet vortices coexisting on the ground plane during the first 0.9s of the formation process. After about 1s, these vortices are shown to coalesce and form one single inlet vortex, containing the circulation of all the smaller vortices combined. The smaller vortices were weak enough to not present danger of FOD, but once coalesced could lift up a 16cm diameter chunk of tarmac asphalt. The conclusion of this work is a recommendation for the development of a solution to the inlet vortex problem focused on preventing the coalescing of the vortex during its formation, thus eliminating the threat of FOD.
26
Rabe, Angela C. "Effectiveness of a Serpentine Inlet Duct Flow Control Scheme at Design and Off-Design Simulated Flight Conditions." Diss., Virginia Tech, 2003. http://hdl.handle.net/10919/28653.
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An experimental investigation was conducted in a static ground test facility to determine the flow quality of a serpentine inlet duct incorporating active flow control for several simulated flight conditions. The total pressure distortion at the aerodynamic interface plane (AIP) was then used to predict the resulting stability for a compression system. This study was conducted using a model of a compact, low observable, engine inlet duct developed by Lockheed Martin. A flow control technique using air injection through microjets at 1% of the inlet mass flow rate was developed by Lockheed Martin to improve the quality of the flow exiting the inlet duct. Both the inlet duct and the flow control technique were examined at cruise condition and off-design simulated flight conditions (angle of attack and asymmetric distortion). All of the experimental tests were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number of 1.76*105 based on the hydraulic diameter at the inlet throat.
For each of the flight conditions tested, the flow control scheme was found to improve the flow uniformity and reduce the inlet distortion at the AIP. For simulated cruise condition, the total pressure recovery was improved by ~2% with the addition of flow control. For the off-design conditions of angle of attack and asymmetric distortion, the total pressure recovery was improved by 1.5% and 2% respectively. All flight conditions tested showed a reduction in circumferential distortion intensity with flow control. The cruise condition case showed reduced maximum circumferential distortion of 70% with the addition of flow control. A reduction in maximum circumferential distortion of 40% occurred for the angle of attack case with flow control, and 30% for the asymmetric distortion case with flow control.
The inlet total pressure distortion was used to predict the changes in stability margin of a compression system due to design and off-design flight conditions and the improvement of the stability margin with the addition of flow control. A parallel compressor model (DYNTECC) was utilized to predict changes in the stability margin of a representative compression system (NASA Stage 35). Without flow control, all three cases show similar reduced stability margins on the order of 30% of the original stability margin for NASA Stage 35 at 70% corrected rotor speed. With the addition of flow control, the cruise condition tested improved the stability margin to 80% of the original value while the off-design conditions recover to 60% of the original margin. Overall, the flow control has been found to be extremely beneficial in improving the operating range of a compression system for the same inlet duct without flow control.<br>Ph. D.
27
Kozak, Jeffrey D. "3-D flow calculations of a bifurcated 2D supersonic engine inlet at takeoff." Thesis, This resource online, 1995. http://scholar.lib.vt.edu/theses/available/etd-03032009-040729/.
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Chan, Keen Ian 1972. "An assessment of computational procedures for eleven-stage compressor response to inlet distortion." Thesis, Massachusetts Institute of Technology, 2003. http://hdl.handle.net/1721.1/82766.
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Graf, Martin Bowyer. "Investigation of the effect of radial inlet temperature distortion on turbine heat transfer." Thesis, Massachusetts Institute of Technology, 1993. http://hdl.handle.net/1721.1/12359.
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Strang, Eric James. "Influence of unsteady losses and deviations on compression system stability with inlet distortion." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/42511.
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Spakovszky, Zoltán S. (Zoltán Sándor) 1972. "Active control of rotating stall in a transonic compressor stage with inlet distortion." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/50522.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.<br>Includes bibliographical references (p. 101-103).<br>Rotating stall has been stabilized in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor. An annular array of 12 jet-injectors located upstream of the rotor tip was used for forced response testing and to extend the compressor stable operating range. Results for radial and circumferential inlet distortion are reported. First, the effects of radial distortion on the compressor performance and the dynamic behavior were investigated. Control laws were designed using empirical transfer function estimates determined from forced response results. The transfer functions indicated that the compressor dynamics are decoupled with radial inlet distortion, as they are for the case of undistorted inlet flow. Single- input-single-output (SISO) control strategies were therefore used for the radial distortion controller designs. A circumferential total pressure distortion of about one dynamic head and a 120' extent (DC(60) = 0.61) introduced coupling between the harmonics of circumferential pressure perturbations, requiring multi-variable (MIMO) identification and control design techniques. A careful analysis of the coupled pre-stall compressor dynamics revealed a strong first spatial harmonic, dominated by the well known incompressible Moore-Greitzer mode. Constant gain control and more sophisticated MIMO robust control strategies were used for stabilization with circumferential inlet distortion. Steady axisymmetric injection of 4% of the compressor mass flow resulted in a reduction in stalling mass flow of 9.7% relative to the case with radial inlet distortion and no injection. Use of a robust H, controller with unsteady non-axisymmetric injection achieved a further reduction in stalling mass flow of 7.5%, resulting in a total reduction of 17.2%. Steady injection experiments with circumferential inlet distortion resulted in a 6.2% reduction of stalling mass flow. Constant gain feedback, using unsteady asymmetric injection, yielded a further range extension of 9%. Testing of MIMO robust controllers showed only 2% reduction in stalling mass flow. Instead of further tuning the complex MIMO controllers, the same robust H. controller used for radial distortion was tested. This controller achieved a reduction in stalling mass flow of 10.2% relative to steady injection, yielding a total reduction in stalling mass flow of 16.4%.<br>by Zoltán S. Spakovszky.<br>S.M.
32
Ferrar, Anthony Maurice. "Measurement and Uncertainty Analysis of Transonic Fan Response to Total Pressure Inlet Distortion." Diss., Virginia Tech, 2015. http://hdl.handle.net/10919/51747.
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Distortion tolerant fans represent the enabling technology for the successful implementation of highly integrated airframe propulsion system vehicles. This investigation extends the study of fan-distortion interactions to an actual turbofan engine with a total pressure distortion profile representative of a boundary-layer ingesting (BLI) embedded engine. The goal was to make a series of flow measurements that contribute to the overall physical understanding of this complex flow situation.
Proper uncertainty analysis is critical to extracting meaning from the data measured in this study. The important information in the measurements is contained in small differences that lead to large impacts on the fan performance. In some cases, these differences were measured to a useful degree of accuracy, while in others they were not. One important application of the uncertainty analysis techniques developed in this work is the identification of the dominant error sources that resulted in unacceptable uncertainties.
This dissertation presents an experimental study of transonic fan response to inlet total pressure distortion. A Pratt and Whitney JT15D-1 turbofan engine was subjected to a total pressure distortion representative of a boundary layer ingesting serpentine inlet. A 5-hole probe measured the aerodynamic response of the fan rotor in terms of flow angles, total pressure, and static pressure. A thermocouple embedded in the probe measured the rotor outlet total temperature. These measurements enabled the full characterization of the flow condition at each measurement point. The results indicate that a trailing edge separation and reattachment cycle experienced by the blades caused variations in the work input to the flow and resulted in a non-uniform rotor outlet flow profile. The details of the aerodynamic process and several means for improving distortion response are presented in this context.
As a second theme, the modern measurement and uncertainty analysis techniques required to obtain useful information in this situation are developed and explored. Uncertainty analysis is often treated as a less glamorous afterthought in experimental research. However, as technology develops along lines of ever increasing system-level integration, simply suggesting the solution to a single flow situation does not repre- sent closure to the larger problem. In addition to frameworks for developing distortion tolerant fans, frameworks for developing frameworks are required. Uncertainty-drivenexperimental techniques represent the enabling methodology for the discovery and un- derstanding of the subtle phenomena associated with such coupled performance. These considerations are required to extend the usefulness of the results to the overarching issue of integrating the complex performance of individual components into an overall superior system. The experimental methods and uncertainty analysis developed in this study are presented in this context.<br>Ph. D.
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Unrau, Mikkel Andreas. "Analysis of the Effects of Inlet Distortion on Stall Cell Formation in a Transonic Compressor Using CREATE-AV Kestrel." BYU ScholarsArchive, 2018. https://scholarsarchive.byu.edu/etd/7712.
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Accurately predicting fan performance, including bounds of operation, is an important function of any Computational Fluid Dynamics (CFD) package. The presented research uses a CFD code developed as part of the Computational Research and Engineering Acquisition Tools and Environment (CREATE), known as Kestrel, to evaluate a single stage compressor at various operating conditions. Steady-state, single-passage simulations are carried out to validate capabilities recently added to Kestrel. The analysis includes generating speedlines of total pressure ratio and efficiency, as well as radial total temperature and total pressure profiles at two axial locations in the compressor at various operating conditions and fan speeds, and simulation data from the single-passage runs is compared to experimental data. Time-accurate, full annulus simulations are also carried out to capture and analyze the processes leading to stall inception for both uniform and distorted inlet conditions. The distortion profile used contains a 90 degree sector of lower total pressure at the inlet. The observed fan behavior at stall inception is compared to previous research, and it is concluded that the inlet distortion significantly changes the behavior of the part-span stall cells that develop after stall inception. Understanding the physical processes that lead to stall inception allows fan designers to design more robust fans that can safely take advantage of the better performance associated with operating closer to stall.
34
Boller, Shaun M. "One-Dimensional Dynamic Wake Response in an Isolated Rotor due to Inlet Total Pressure Distortion." Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/9587.
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An experimental investigation of the wake of a low-speed axial-flow compressor rotor was conducted with and without the presence of steady inlet total pressure distortions. The steady three-dimensional rotor inlet flow was obtained by a five-hole pneumatic pressure probe, while the one-dimensional rotor exit data were obtained using a piggyback steady/unsteady total pressure probe in non-nulling mode. Both inlet and exit flow conditions were measured in the stationary frame of reference.
Results indicate increases in wake thickness and magnitude of total pressure defect as blade loading increased into the distortion cycle. The wake suction side jet increased in width and magnitude as blade loading increased, which appears to be a response to flow blockage caused by the growing boundary layer on the blades. Based on one-dimensional exit total pressure conditions with respect to the distortion screen, the dynamic response of the intra-blade passage flow does not appear to be a function of blade loading, measurement span, or distortion intensity within the ranges tested. Unsteady one-dimensional rotor exit suction side jet width and magnitude varied a great deal within and outside of the distorted region, and were only moderately correlated to inlet flow conditions. Changes in the unsteady one-dimensional rotor wake width and magnitude were usually in phase with and strongly correlated to changes in the inlet flow conditions.<br>Master of Science
35
Weston, David Bruce. "High Fidelity Time Accurate CFD Analysis of a Multi-stage Turbofan at Various Operating Points in Distorted Inflow." BYU ScholarsArchive, 2014. https://scholarsarchive.byu.edu/etd/5604.
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Inlet distortion is an important consideration in fan performance. Distortion can be caused through flight conditions and airframe-engine interfaces. The focus of this paper is a series of high-fidelity time accurate Computational Fluid Dynamics (CFD) simulations of a multistage fan. These investigate distortion transfer and generation as well as the underlying flow physics of these phenomena under different operating conditions. The simulations are performed on the full annulus of a 3 stage fan. The code used to carry out these simulations is a modified version of OVERFLOW 2.2 developed as part of the Computational Research and Engineering Acquisition Tools and Environment (CREATE) program. Several modifications made to the code are described within this thesis. The inlet boundary condition is specified as a 1/rev total pressure distortion. Simulations at choke, design, and near stall points are analyzed and compared to experimental data. Analysis includes the phase and amplitude of total temperature and pressure distortion through each stage of the fan and blade loading plots. An understanding of the flow physics associated with distorted flows will help designers account for unsteady flow physics at design and off-design operating conditions and build more robust fans with a greater stability margin.
36
Farge, Talib Z. "The effect of tip leakage, backswept blades and inlet distortion on centrifugal compressor flow." Thesis, University of Liverpool, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.304987.
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Defoe, Jeff (Jeffrey James). "Inlet swirl distortion effects on the generation and propagation of fan rotor shock noise." Thesis, Massachusetts Institute of Technology, 2011. http://hdl.handle.net/1721.1/68404.
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Thesis (Ph. D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2011.<br>Cataloged from PDF version of thesis.<br>Includes bibliographical references (p. 195-200).<br>A body-force-based fan model for the prediction of multiple-pure-tone noise generation is developed in this thesis. The model eliminates the need for a full-wheel, three-dimensional unsteady RANS simulation of the fan blade row, allowing Euler calculations to be used to capture the phenomena of interest. The Euler calculations reduce numerical wave dissipation and enable the simultaneous computation of source noise generation and propagation through the engine inlet to the far-field in non-uniform flow. The generated shock Mach numbers are in good agreement with experimental results, with the peak values predicted within 6%. An assessment of the far-field acoustics against experimental data showed agreement of 8 dB on average for the blade-passing tone. In a first-of-its-kind comparison, noise generation and propagation are computed for a fan installed in a conventional inlet and in a boundary-layer-ingesting serpentine inlet for a free-stream Mach number of 0.1. The key effect of boundary layer ingestion is the creation of streamwise vorticity which is ingested into the inlet, resulting in co- and counter-rotating streamwise vortices in the inlet. The fan sound power level increases by 38 dB due to this distortion, while the vortex whose circulation is in the same direction as the fan rotation enhances the sound power attenuation within the inlet duct such that the far-field overall sound pressure levels are increased by only 7 dB on average. The far-field spectra are altered in the following manner due to inlet distortion: (1) tones at up to 3 times the blade-passing frequency are amplified; and (2) tones above one-half of the blade-passing frequency are attenuated and appear to be cut-off. To quantify the effects of serpentine inlet duct geometry on the generation and propagation of multiple-pure-tone noise, a parametric study of inlets is conducted. The conclusions are that (1) the ingestion of streamwise vorticity alters multiple-pure-tone noise more than changes in inlet area ratio or offset ratio do; and (2) changes in the far-field spectra relative to the conventional inlet results are only weakly affected by the duct geometry changes investigated and are instead predominantly caused by flow non-uniformities. A response-surface correlation for the effects of inlet geometry on far-field noise is also developed.<br>by Jeff Defoe.<br>Ph.D.
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Hale, Alan A. "A three-dimensional turbine engine analysis compressor code (TEACC) for steady-state inlet distortion." Diss., This resource online, 1996. http://scholar.lib.vt.edu/theses/available/etd-06062008-154548/.
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Traore, Abdoulaye S. "Mixed Network Interference Management with Multi-Distortion Measures." International Foundation for Telemetering, 2010. http://hdl.handle.net/10150/604294.
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ITC/USA 2010 Conference Proceedings / The Forty-Sixth Annual International Telemetering Conference and Technical Exhibition / October 25-28, 2010 / Town and Country Resort & Convention Center, San Diego, California<br>This paper presents a methodology for the management of interference and spectrum for iNET. It anticipates a need for heavily loaded test environments with Test Articles (TAs) operating over the horizon. In such cases, it is anticipated that fixed and ad hoc networks will be employed, and where spectrum reuse and interference will limit performance. The methodology presented here demonstrates how this can be accomplished in mixed networks.
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Nelson, Michael Allan. "Stereoscopic Particle Image Velocimetry Measurements of Swirl Distortion on a Full-Scale Turbofan Engine Inlet." Thesis, Virginia Tech, 2014. http://hdl.handle.net/10919/64993.
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There is a present need for simulation and measuring the inlet swirl distortion generated by airframe/engine system interactions to identify potential degradation in fan performance and operability in a full-scale, ground testing environment. Efforts are described to address this need by developing and characterizing methods for complex, prescribed distortion patterns. A relevant inlet swirl distortion profile that mimics boundary layer ingesting inlets was generated by a novel new method, dubbed the StreamVane method, and measured in a sub scale tunnel using stereoscopic particle image velocimetry (SPIV) as a precursor for swirl distortion generation and characterization in an operating turbofan research engine. Diagnostic development efforts for the distortion measurements within the research engine paralleled the StreamVane characterization. The system used for research engine PIV measurements is described. Data was obtained in the wake of a total pressure distortion screen for engine conditions at idle and 80% corrected fan speed, and of full-scale StreamVane screen at 50% corrected fan speed. The StreamVane screen was designed to generate a swirl distortion that is representative for hybrid wing body applications and was made of Ultem*9085 using additive manufacturing. Additional improvements to the StreamVane method are also described. Data reduction algorithms are put forth to reduce spurious velocity vectors. Uncertainty estimations specific to the inlet distortion test rig, including bias error due to mechanical vibration, are made. Results indicate that the methods develop may be used to both generate and characterize complex distortion profiles at the aerodynamic interface plane, providing new information about airframe/engine integration.<br>Master of Science
41
Kennedy, Stefan Andrew. "A computational investigation into the effects of lipskin damage on inlet flow distortion in aircraft engines." Thesis, Queen's University Belfast, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.557636.
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Damage to the lipskin of an engine nacelle can have a significant impact on engine performance and safety, especially in combination with other sources of inlet flow distortion. While the degradation of performance in aircraft propulsion systems as a result of inlet distortion has been well-studied, the impact of lipskln damage upon inlet distortion levels is not well understood. Current practice can result in the grounding of aircraft in order to perform repairs if the damage exceeds the specified tolerance in size and the position on the lip, with these tolerances varying from aircraft to aircraft. If these tolerances are overly conservative, aircraft operating times may be adversely affected. The effect of lipskin damage upon the performance of the inlet of the CF34-3A engine has been investigated using the commercial Computational Fluid Dynamics Package ANSYS CFX in a nacelle model which included the fan. The fan was included in order to capture the interaction between the blades and the upstream distortion. The fan was found to have a stabilising effect upon the flow upstream of the fan, reducing the inlet distortion due to a redistribution of the flow. Other sources of inlet distortion have also been studied, with the effects of high angles of attack and crosswinds having been modelled in isolation, and combined with cases of lipskin damage. A typical take-off configuration was primarily used to assess the effect of the damage due to this being the most critical segment of the flight envelope. The resulting distortion levels have been compared to the engine manufacturers limits in order to assess the severity of the damage.
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Frohnapfel, Dustin Joseph. "Methodology Development and Investigation of Turbofan Engine Response to Simultaneous Inlet Total Pressure and Swirl Distortion." Diss., Virginia Tech, 2019. http://hdl.handle.net/10919/88866.
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As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation.
For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility.
A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Computational full-field total pressure recovery and swirl angle profiles matched within approximately 1% and 2.5° (RMSD), respectively. Experimental turbofan engine ground test results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Experimental full-field total pressure recovery and swirl angle profiles matched within approximately 1.25% and 3.0° (RMSD), respectively.
Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased by approximately 5% in the presence of distortion.
While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion.<br>Doctor of Philosophy<br>As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation. For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility. A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational and experimental results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased in the presence of distortion. While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion.
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Ferrar, Anthony Maurice. "Measurements of Flow in Boundary Layer Ingesting Serpentine Inlets." Thesis, Virginia Tech, 2011. http://hdl.handle.net/10919/36408.
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Highly integrated airframe-propulsion systems featuring ingestion of the airframe boundary
layer oer reduced noise, emissions, and fuel consumption. Embedded engine systems are
envisioned which require boundary layer ingesting (BLI) serpentine inlets to provide the
needed air
ow to the engine. These inlets produce distorted
ow proles that can cause
aeromechanical, stability, and performance changes in embedded engines. Proper design
of embedded engine systems requires understanding of the underlying
uid dynamics that
occur within serpentine inlets.
A serpentine inlet was tested in a specially designed wind tunnel that simulated boundary
layer ingestion in a full-scale realistic environment. The measured total pressure proles at
the inlet and exit planes of the duct, and the static pressure distributions along the walls
provided useful data related to the
ow in BLI serpentine inlet systems. A bleed
ow control
system was tested that utilized no more than 2% of the total inlet
ow. Two bleed slots
were employed, one near the rst bend of the S-duct and one near second. The bleed system
successfully reduced inlet distortions by as much as 30%, implying improvements in stall
margin and engine performance.
Analysis of the wake shape entering the S-duct showed that the airframe and inlet duct are
both important components of a wake-ingesting inlet/diusion system. Shape eects and
static pressure distributions determined
ow transport within the serpentine inlet. Flow
separation within the S-duct increased distortion at the engine inlet plane. Discussion
of airframe/inlet/engine compatibility demonstrates that embedded engine systems require
multi-disciplinary collaborative design eorts. An included fundamental analysis provides
performance estimates and design guidelines. The ideal airframe performance improvement
associated with wake-ingestion is estimated.<br>Master of Science
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Eddy, Grant Lee. "Study of Steady-State Wake Characteristics of Variable Angle Wedges." Thesis, Virginia Tech, 2001. http://hdl.handle.net/10919/35205.
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Current methods of creating inlet total pressure distortion for testing in gas turbine engines are only able to simulate steady-state distortion patterns. With modern military aircraft it is becoming necessary to examine the effects of transient inlet distortion on engines. One alternative being evaluated is a splitting airfoil that is essentially a wedge that can be set at different opening angles. An array of such devices would be placed in front of the engine for testing that would be capable of creating steady-state distortion patterns as well as transient distortion patterns by changing the opening angle of the airfoils.
The work here analyzes the steady-state wake characteristics of some of the splitting airfoil concepts. Single-wedge tests were conducted with various opening angles in an attempt to classify the various aspects found in the wake pattern. It was found that the wake has completely different characteristics with larger opening angles. In addition, several different combinations of wedges were also examined to see if single wedge analysis could be applied to arrays of wedges. Analysis was done on combinations of wedges aligned vertically as well as combinations that were done horizontally. It was found that single wedge characteristics change considerably when different wake patterns interact with each other<br>Master of Science
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Schwartz, Jeffrey R. "An Experimental and Analytical Investigation of Dynamic Flow Response of a Fan Rotor with Distorted Inlet Flow." Thesis, Virginia Tech, 1999. http://hdl.handle.net/10919/44314.
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An experimental and analytical investigation was conducted to gain insight and ultimately predict the dynamic flow response of a fan rotor with inlet flow distortion. Rotor exit total pressure circumferential profiles were accurately predicted using frequency response functions derived from experimental rotor response data. Using these predicted profiles, an initial attempt was made at predicting the dynamic (distorted) stage characteristics of the test machine with promising results.
The first step of this research was an experimental investigation to gather unsteady rotor response data. The steady three-dimensional inlet flow of an isolated rotor subjected to inlet distortion was obtained using a five-hole pneumatic prism probe. Exit flow dynamic wake data were obtained using a piggyback steady/unsteady total pressure probe in non-nulling mode. Inlet and exit data were collected for eighteen different combinations of distortion level, operating point, and measurement span.
Frequency response functions were generated and then averaged for each operating regime, span, and distortion intensity, assuming the data to be stationary and ergodic. These 'generalized' FRF's were used to predict the rotor exit total pressure profile. These pressure profiles were then used in an initial attempt to predict the dynamic stage (distorted) characteristics of the test machine. Best predictions resulted when an FRF was used for individual operating regimes, defined with respect to rotor blade mean aerodynamic loading.<br>Master of Science
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Barton, Scott Andrew. "An experimental investigation of the influence of inlet distortion on the fluid borne noise of a centrifugal pump." Thesis, Massachusetts Institute of Technology, 1991. http://hdl.handle.net/1721.1/33251.
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Peterson, Marshall Warren. "Implementations of Fourier Methods in CFD to Analyze Distortion Transfer and Generation Through a Transonic Fan." BYU ScholarsArchive, 2016. https://scholarsarchive.byu.edu/etd/6384.
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Inlet flow distortion is a non-uniform total pressure, total temperature, or swirl (flow angularity) condition at an aircraft engine inlet. Inlet distortion is a critical consideration in modern fan and compressor design. This is especially true as the industry continues to increase the efficiency and operating range of air breathing gas turbine engines. The focus of this paper is to evaluate the Computational Fluid Dynamics (CFD) Harmonic Balance (HB) solver in STAR-CCM+ as a reduced order method for capturing inlet distortion as well as the associated distortion transfer and generation. New methods for quantitatively describing and analyzing distortion transfer and generation are investigated. The geometry used is the rotor 4 fan geometry, consisting of one rotor and one stator. The inlet boundary condition is a 90-degree sector total pressure distortion profile with total pressure and swirl held constant. Multiple HB simulations with varying mode combinations and distortion intensities are analyzed and compared against full annulus Unsteady Reynolds Averaged Navier-Stokes (URANS) simulations. Best practices and recommendations for the implementation of the HB solver are given. The pre-existing Society of Automotive Engineers Aerospace Recommended Practice (SAE-ARP) 1420b descriptors are demonstrated to be inadequate for the purposes of analyzing distortion transfer and generation on a stage-to-stage basis. New implementations of Fourier methods are presented as an alternative to the SAE-ARP 1420b descriptors. These Fourier descriptors are shown to describe distortion transfer and generation to a higher degree of fidelity than the SAE-ARP 1420b descriptors. These new descriptors are demonstrated on the analysis of full annulus URANS and HB simulations. The HB solver is shown to be capable of capturing distortion transfer, generation and performance degradation. Recommendations for the optimal implementation of the HB method are given.
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Hylton, Michael Ronnie. "Assessment of an Innovative Experimental Facility for Testing Diffusing Serpentine Inlets with Large Amounts of Boundary Layer Ingestion." Thesis, Virginia Tech, 2008. http://hdl.handle.net/10919/33969.
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An innovative experimental facility was developed for testing flush-mounted, diffusing serpentine inlets intended for use on blended-wing-body aircraft. The static ground test facility was able to simulate the boundary layer profile expected to be ingested by inlets mounted on the aft sections of these aircraft. It generated Mach numbers ranging from 0.19 to 0.4 and boundary layer thicknesses between 36% and 45%. The circumferential distortions at the aerodynamic interface plane of the serpentine inlet were also calculated, and ranged between 0.0042 for the lowest Mach number, to 0.0098 for the highest Mach number. Reynolds numbers for the tests ranged between 1.2 million and 2.4 million depending on engine speed and Mach number. The results of the experiment were compared to a previous NASA report, and showed close agreement in distortion patterns and pressure losses at a Mach number of 0.25.<br>Master of Science
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List, Michael G. "Numerical Quantification of Interaction Effects in a Closely-Coupled Diffuser-Fan System." University of Cincinnati / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1396530464.
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Novikov, Yaroslav. "Development Of A High-fidelity Transient Aerothermal Model For A Helicopter Turboshaft Engine For Inlet Distortion And Engine Deterioration Simulations." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12614389/index.pdf.
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Presented in this thesis is the development of a high-fidelity aerothermal model for GE T700 turboshaft engine. The model was constructed using thermodynamic relations governing change of flow properties across engine components, and by applying real component maps for the compressor and turbines as well as empirical relations for specific heats. Included in the model were bleed flows, turbine cooling and heat sink effects. Transient dynamics were modeled using inter-component volumes method in which mass imbalance between two engine components was used to calculate the inter-component pressure. This method allowed fast, high-accuracy and iteration-free calculation of engine states. Developed simulation model was successfully validated against previously published simulation results, and was applied in the simulation of inlet distortion and engine deterioration. Former included simulation of steady state and transient hot gas ingestion as well as transient decrease in the inlet total pressure. Engine deterioration simulations were performed for four different cases of component deterioration with parameters defining engine degradation taken from the literature. Real time capability of the model was achieved by applying time scaling of plenum volumes which allowed for larger simulation time steps at very little cost of numerical accuracy. Finally, T700 model was used to develop a generic model by replacing empirical relations for specific heats with temperature and FAR dependent curve fits, and scaling T700 turbine maps. Developed generic aerothermal model was applied to simulate steady state performance of the Lycoming T53 turboshaft engine.