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1
Karki, K. C., V. L. Oechsle, and H. C. Mongia. "A Computational Procedure for Diffuser-Combustor Flow Interaction Analysis." Journal of Engineering for Gas Turbines and Power 114, no. 1 (January 1, 1992): 1–7. http://dx.doi.org/10.1115/1.2906301.
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This paper describes a diffuser-combustor flow interaction analysis procedure for gas turbine combustion systems. The method is based on the solution of the Navier–Stokes equations in a generalized nonorthogonal coordinate system. The turbulence effects are modeled via the standard two-equation (k-ε) model. The method has been applied to a practical gas turbine combustor-diffuser system that includes support struts and fuel nozzles. Results have been presented for a three-dimensional simulation, as well as for a simplified axisymmetric simulation. The flow exhibits significant three-dimensional behavior. The axisymmetric simulation is shown to predict the static pressure recovery and the total pressure losses reasonably well.
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Carrotte, J. F., D. W. Bailey, and C. W. Frodsham. "Detailed Measurements on a Modern Combustor Dump Diffuser System." Journal of Engineering for Gas Turbines and Power 117, no. 4 (October 1, 1995): 678–85. http://dx.doi.org/10.1115/1.2815453.
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An experimental investigation has been carried out to determine the flow characteristics and aerodynamic performance of a modern gas turbine combustor dump diffuser. The system comprised a straight walled prediffuser, of area ratio 1.35, which projected into a dump cavity where the flow divided to pass either into the flame tube or surrounding feed annuli. In addition, a limited amount of air was removed to simulate flow used for turbine cooling. The flame tube was relatively deep, having a radial depth 5.5 times that of the passage height at prediffuser inlet, and incorporated burner feed arms, cowl head porosity, cooling rings, and primary ports. Representative inlet conditions to the diffuser system were generated by a single-stage axial flow compressor. Results are presented for the datum configuration, and for a further three geometries in which the distance between prediffuser exit and the head of the flame tube (i.e., dump gap) was reduced. Relatively high values of stagnation pressure loss were indicated, with further significant increases occurring at smaller dump gaps. These high losses, which suggest a correlation with other published data, are due to the relatively deep flame tube and short diffuser length. Furthermore, the results also focus attention on how the presence of a small degree of diffuser inlet swirl, typical of that which may be found within a gas turbine engine, can result in large swirl angles being generated farther downstream around the flame tube. This is particularly true for flow passing to the inner annulus.
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Agrawal, A. K., J. S. Kapat, and T. T. Yang. "An Experimental/Computational Study of Airflow in the Combustor–Diffuser System of a Gas Turbine for Power Generation." Journal of Engineering for Gas Turbines and Power 120, no. 1 (January 1, 1998): 24–33. http://dx.doi.org/10.1115/1.2818084.
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This paper presents an experimental/computational study of cold flow in the combustor–diffuser system of industrial gas turbines employing can-annular combustors and impingement-cooled transition pieces. The primary objectives were to determine flow interactions between the prediffuser and dump chamber, to evaluate circumferential flow nonuniformities around transition pieces and combustors, and to identify the pressure loss mechanisms. Flow experiments were conducted in an approximately one-third geometric scale, 360-deg annular test model simulating practical details of the prototype including the support struts, transition pieces, impingement sleeves, and can-annular combustors. Wall static pressures and velocity profiles were measured at selected locations in the test model. A three-dimensional computational fluid dynamic analysis employing a multidomain procedure was performed to supplement the flow measurements. The complex geometric features of the test model were included in the analysis. The measured data correlated well with the computations. The results revealed strong interactions between the prediffuser and dump chamber flows. The prediffuser exit flow was distorted, indicating that the uniform exit conditions typically assumed in the diffuser design were violated. The pressure varied circumferentially around the combustor casing and impingement sleeve. The circumferential flow nonuniformities increased toward the inlet of the turbine expander. A venturi effect causing flow to accelerate and decelerate in the dump chamber was also identified. This venturi effect could adversely affect impingement cooling of the transition piece in the prototype. The dump chamber contained several recirculation regions contributing to the losses. Approximately 1.2 dynamic head at the prediffuser inlet was lost in the combustor–diffuser, much of it in the dump chamber where the fluid passed though narrow pathways. A realistic test model and three-dimensional analysis used in this study provided new insight into the flow characteristics of practical combustor–diffuser systems.
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Antas, Stanisław. "Exhaust System for Radial and Axial-Centrifugal Compressor with Pipe Diffuser." International Journal of Turbo & Jet-Engines 36, no. 3 (August 27, 2019): 297–304. http://dx.doi.org/10.1515/tjj-2016-0068.
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Abstract The application of radial and axial-centrifugal compressors in turboprop, turboshaft and turbofan engines may require the construction of small diameters diffuser in order to obtain lower weight and smaller frontal area. Conventional exhaust diffusers typically have large outlet diameters for exit Mach numbers lower than 0,2 and low swirl flow to the combustor, hence the design of channel of the low-diameter diffusers called controlled-contour, fishtail-shaped diffuser or diffusing trumpet is complex. The cross-sectional shape of these channels is varied from circular to oval to elliptic and to rectangular. The paper presents an original method for determining the flow parameters in the channel and at the outlet section of the downstream diffusing trumpet for a pipe diffuser, which constitutes the downstream duct of the radial or axial-centrifugal compressor with the pipe diffuser. It also illustrates a new method for determining the geometrical parameters of the diffuser. Mentioned methods (for conceptual design of a compressor with pipe diffuser) are based on Pythagorean theorem, properties of ellipse, equation of continuity, energy equation, first law of thermodynamics, Euler’s moment of momentum equation, gasdynamics functions and definitions used in theory of turbo-machines.
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Walker, A. Duncan, Paul A. Denman, and James J. McGuirk. "Experimental and Computational Study of Hybrid Diffusers for Gas Turbine Combustors." Journal of Engineering for Gas Turbines and Power 126, no. 4 (October 1, 2004): 717–25. http://dx.doi.org/10.1115/1.1772403.
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The increasing radial depth of modern combustors poses a particularly difficult aerodynamic challenge for the pre-diffuser. Conventional diffuser systems have a finite limit to the diffusion that can be achieved in a given length and it is, therefore, necessary for designers to consider more radical and unconventional diffuser configurations. This paper will report on one such unconventional diffuser; the hybrid diffuser which, under the action of bleed, has been shown to achieve high rates of diffusion in relatively short lengths. However, previous studies have not been conducted under representative conditions and have failed to provide a complete description of the relevant flow mechanisms making optimization difficult. Utilizing an isothermal representation of a modern gas turbine combustor an experimental investigation was undertaken to study the performance of a hybrid diffuser compared to that of a conventional, single-passage, dump diffuser system. The hybrid diffuser achieved a 53% increase in area ratio within the same axial length generating a 13% increase in the pre-diffuser static pressure recovery coefficient which, in turn, produced a 25% reduction in the combustor feed annulus total pressure loss coefficient. A computational investigation was also undertaken in order to investigate the governing flow mechanisms. A detailed examination of the flow field, including an analysis of the terms within the momentum equation, demonstrated that the controlling flow mechanisms were not simply a boundary layer bleed but involve a more complex interaction between the accelerating bleed flow and the diffusing mainstream flow. A greater understanding of these mechanisms enabled a more practical design of hybrid diffuser to be developed that not only simplified the geometry but also improved the quality of the bleed air making it more attractive for use in component cooling.
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Koutmos, P., and J. J. McGuirk. "Numerical Calculations of the Flow in Annular Combustor Dump Diffuser Geometries." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 203, no. 5 (September 1989): 319–31. http://dx.doi.org/10.1243/pime_proc_1989_203_121_02.
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A method for calculating the turbulent isothermal flow in axisymmetric annular dump diffuser geometries is described and appraised. The calculation method is based on the numerical solution of the time-averaged transport equations for momentum, continuity, turbulence kinetic energy and energy dissipation, using a finite difference formulation. A boundary-fitted curvilinear orthogonal grid obtained from a solution of the inverse Laplace equations is used to represent the curved combustor head accurately and reduce numerical diffusion errors due to better alignment of the flow streamlines with the grid lines. Comparison between predicted results and measurements indicates that variations in (a) the overall pressure recovery and (b) the loss coefficient performance of the dump diffuser system, with changes in diffuser design features (for example inner/outer annulus mass flow split or dump gap), can be predicted to within 7 per cent of the inlet dynamic head without adopting a more refined turbulence closure. The method is therefore demonstrated to be a useful design tool for dump diffuser geometries.
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Carrotte, J. F., P. A. Denman, A. P. Wray, and P. Fry. "Detailed Performance Comparison of a Dump and Short Faired Combustor Diffuser System." Journal of Engineering for Gas Turbines and Power 116, no. 3 (July 1, 1994): 517–26. http://dx.doi.org/10.1115/1.2906850.
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A rectangular model simulating four sectors of a combustion chamber was used to compare the performance of a standard dump diffuser, of overall length 180 mm, with that of a faired design 25.5 mm shorter. The performance of each system was assessed in terms of total pressure loss and static pressure recovery between prediffuser inlet and the annuli surrounding the flame tube. Since the program objective was to test design concepts only, no allowance was made for the presence of burner feed arms or flame tube support pins. In addition, tests were performed with relatively low levels of inlet turbulence and no wake mixing effects from upstream compressor blades. Relative to the dump design, the mass weighted total pressure loss to the outer and inner annuli was reduced by 30 and 40 percent, respectively, for the faired diffuser. Measurements around the flame tube head were used to identify regions of high loss within each system and account for the differences in performance. Within a dump diffuser the flow separates at prediffuser exit resulting in a free surface diffusion around the flame tube head and a recirculating flow in the dump cavity. This source of loss is eliminated in the faired system where the flow remains attached to the casings. Furthermore, the faired system exhibited similar velocity magnitudes and gradients around the combustor head despite its shorter length.
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Hendricks, R. C., D. T. Shouse, W. M. Roquemore, D. L. Burrus, B. S. Duncan, R. C. Ryder, A. Brankovic, N. S. Liu, J. R. Gallagher, and J. A. Hendricks. "Experimental and Computational Study of Trapped Vortex Combustor Sector Rig with High-Speed Diffuser Flow." International Journal of Rotating Machinery 7, no. 6 (2001): 375–85. http://dx.doi.org/10.1155/s1023621x0100032x.
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The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL modes of combustion. The present work describes the operational principles of the TVC, and extends diffuser velocities toward choked flow and provides system performance data. Performance data include EINOx results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable compared to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.
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Hubbard, S., and A. P. Dowling. "Acoustic Resonances of an Industrial Gas Turbine Combustion System." Journal of Engineering for Gas Turbines and Power 123, no. 4 (October 1, 2000): 766–73. http://dx.doi.org/10.1115/1.1370975.
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A theory is developed to describe low-frequency acoustic waves in the complicated diffuser/combustor geometry of a typical industrial gas turbine. This is applied to the RB211-DLE geometry to give predictions for the frequencies of the acoustic resonances at a range of operating conditions. The main resonant frequencies are to be found around 605 Hz (associated with the plenum) and around 461 Hz and 823 Hz (associated with the combustion chamber), as well as one at around 22 Hz (a bulk mode associated with the system as a whole). The stabilizing effects of a Helmholtz resonator, which models damping through nonlinear effects, are included, together with effects of coupled pressure waves in the fuel supply system.
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Hestermann, R., S. Kim, A. Ben Khaled, and S. Wittig. "Flow Field and Performance Characteristics of Combustor Diffusers: A Basic Study." Journal of Engineering for Gas Turbines and Power 117, no. 4 (October 1, 1995): 686–94. http://dx.doi.org/10.1115/1.2815454.
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Results of a detailed study concerning the influence of geometric as well as fluid mechanic parameters on the performance of a plane model combustor diffuser in cold flow are presented. For a qualitative insight into the complex flow field inside the prediffuser, the sudden expansion region, and the flow field around the flame tube dome, results of a flow visualization study with the hydrogen bubble method as well as with the ink jet method are presented for different opening angles of the prediffuser and for different flame tube distances. Also, quantitative data from detailed measurements with LDV and conventional pressure probes in a geometrically similar air-driven setup are presented. These data clearly demonstrate the effect of boundary layer thickness as well as the influence of different turbulence levels at the entry of the prediffuser on the performance characteristics of combustor diffusers. The possibility of getting an unseparated flow field inside the prediffuser even at large opening angles by appropriately matching the diffuser’s opening angle and the flame tube distance is demonstrated. Also, for flows with an increased turbulence level at the entrance—all other conditions held constant—an increased opening angle can be realized without experiencing flow separation. The comparison of the experimental data with predictions utilizing a finite-volume-code based on a body-fitted coordinate system for diffusers with an included total opening angle less than 18 deg demonstrates the capability of describing the flow field in combustor diffusers with reasonable accuracy.
11
Walker, A. Duncan, Jon F. Carrotte, and James J. McGuirk. "The Influence of Dump Gap on External Combustor Aerodynamics at High Fuel Injector Flow Rates." Journal of Engineering for Gas Turbines and Power 131, no. 3 (February 11, 2009). http://dx.doi.org/10.1115/1.3028230.
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The increasing demand to reduce fuel burn, hence CO2 emissions, from the gas turbine requires efficient diffusion to reduce the system pressure loss in the combustor. However, interactions between prediffuser and combustor can have a significant effect on diffuser performance. For example, the consequence of increased fuel injector flow at a dump gap set using conventional design guidelines has been shown (Walker, A. D., Carrotte, J. F., and McGuirk, J. J., 2007. “Compressor∕Diffuser∕Combustor Aerodynamic Interactions in Lean Module Combustors,” ASME Turbo Expo 2007—Power for Land Sea and Air, Paper No. GT2007-27872) to introduce a destabilizing interaction between fuel injector and upstream components. The present paper concentrates on examining the effects of increased dump gap. Dump gap ratios of 0.8, 1.2, and 1.6 were employed, with each test utilizing the same inlet guide vane, compressor rotor, integrated outlet guide vane (OGV)∕prediffuser, and dump geometry. The flow fraction of compressor efflux entering the combustor cowl was set to be representative of lean combustors (50–70%). Measurements were made on a fully annular rig using a generic flame tube with metered cowl and inner∕outer annulus flows. The results demonstrate that, with fixed cowl flow, as the dump gap increases, component interactions decrease. At a dump gap ratio of 0.8, the proximity of the flame tube influences the prediffuser providing a beneficial blockage effect. However, if increased to 1.2, this beneficial effect is weakened and the prediffuser flow deteriorates. With further increase to 1.6, the prediffuser shows strong evidence of separation. Hence, at the dump gaps probably required for lean module injectors, it is unlikely the prediffuser will be influenced beneficially by the flame tube blockage; this must be taken into account in the design. Furthermore, with small dump gaps and high cowl flow fraction, the circumferential variation in cowl flow can feed upstream and cause OGV∕rotor forcing. At larger dump gaps, the circumferential variation does not penetrate upstream to the OGV, and the rotor is unaffected. The optimum dump gap and prediffuser design for best overall aerodynamic system performance from rotor through to feed annuli is a compromise between taking maximum advantage of upstream blockage effects and minimizing any 3D upstream forcing.
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Walker, A. Duncan, Bharat Koli, and Peter A. Beecroft. "Influence of Purge Flow Swirl at Exit to the High-Pressure Compressor on OGV/Pre-Diffuser and Combustion System Aerodynamics." Journal of Turbomachinery 141, no. 9 (June 14, 2019). http://dx.doi.org/10.1115/1.4043781.
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As aero gas turbine designs strive for ever greater efficiencies, the trend is for engine overall pressure ratios to rise. Although this provides greater thermal efficiency, it means that cycle temperatures also increase. One potential solution to managing the increasing temperatures is to employ a cooled cooling air system. In such a system, a purge flow into the main gas path downstream of the compressor will be required to prevent hot gas being ingested into the rotor drive cone cavity. However, the main gas path in compressors is aerodynamically sensitive and it is important to understand, and mitigate, the impact such a flow may have on the compressor outlet guide vanes, pre-diffuser, and the downstream combustion system aerodynamics. Initial computational fluid dynamics (CFD) predictions demonstrated the potential of the purge flow to negatively affect the outlet guide vanes and alter the inlet conditions to the combustion system. The purge flow modified the incidence onto the outlet guide vane, at the hub, such that the secondary flows increased in magnitude. An experimental assessment carried out using an existing fully annular, isothermal test facility confirmed the CFD results and importantly demonstrated that the degradation in the combustor inlet flow resulted in an increased combustion system loss. At the proposed purge flow rate, equal to ∼1% of the mainstream flow, these effects were small with the system loss increasing by ∼4%. However, at higher purge flow rates (up to 3%), these effects became notable and the outlet guide vane and pre-diffuser flow degraded significantly with a resultant increase in the combustion system loss of ∼13%. To mitigate these effects, CFD was used to examine the effect of varying the purge flow swirl fraction in order to better align the flow at the hub of the outlet guide vane. With a swirl fraction of 0.65 (x rotor speed), the secondary flows were reduced below that of the datum case (with no purge flow). Experimental data showed good agreement with the predicted flow topology and performance trends but the measured data showed smaller absolute changes. Differences in system loss were measured with savings of around 10% at the turbine feed ports for a mass flow ratio of 1% and a swirl fraction of 0.65.
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Lacombe, Florent, and Yoann Méry. "Mixed Acoustic-Entropy Combustion Instabilities in a Model Aeronautical Combustor: Large Eddy Simulation and Reduced Order Modeling." Journal of Engineering for Gas Turbines and Power 140, no. 3 (October 25, 2017). http://dx.doi.org/10.1115/1.4037960.
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This article focuses on combustion instabilities (CI) driven by entropy fluctuations which is of great importance in practical devices. A simplified geometry is introduced. It keeps the essential features of an aeronautical combustion chamber (swirler, dilution holes, and outlet nozzle), while it is simplified sufficiently to ease the analysis (rectangular vane, one row of holes of the same diameter, no diffuser at the inlet of the chamber, and circular nozzle at the outlet). A large eddy simulation (LES) is carried out on this geometry and the limit cycle of a strong CI involving the convection of an entropy spot is obtained. The behavior of the instability is analyzed using phenomenological description and classical signal analysis. One shows that the system can be better described by considering two reacting zones: a rich mainly premixed flame is located downstream of the swirler and an overall lean diffusion flame is stabilized next to the dilution holes. In a second step, dynamic mode decomposition (DMD) is used to visualize, analyze, and model the complex phasing between different processes affecting the reacting zones. Using these data, a zero-dimensional (0D) modeling of the premixed flame and of the diffusion flame is proposed. These models provide an extended understanding of the combustion process in an aeronautical combustor and could be used or adapted to address mixed acoustic-entropy CI in an acoustic code.
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Lee, Jinwook, Choon S. Tan, Borislav T. Sirakov, Hong-Sik Im, Martin Babak, Denis Tisserant, and Chris Wilkins. "Performance Characterization of Twin-Scroll Turbine Stage for Vehicular Turbocharger Under Unsteady Pulsating Flow Environment." Journal of Engineering for Gas Turbines and Power 139, no. 7 (February 28, 2017). http://dx.doi.org/10.1115/1.4035629.
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Unsteady three-dimensional computations have been implemented on a turbocharger twin-scroll turbine system (volute–turbine wheel–diffuser). The flow unsteadiness in a turbocharger turbine system is essentially driven by a highly pulsating flow from the upstream combustor which causes a pulsating stagnation pressure boundary condition at the inlet to the turbine system. Computed results have been postprocessed and interrogated in depth in order to infer the significance of the induced flow unsteadiness on performance. The induced flow unsteadiness could be deemed important, since the reduced frequency of the turbine system (based on the time scale of the inlet flow fluctuation and the flow through time) is higher than unity. Thus, the computed time-accurate pressure field and the loss generation process have been assessed to establish the causal link to the induced flow unsteadiness in the turbine system. To do this consistently both for the individual subcomponents and the system, a framework of characterizing the operation of the turbine system linked to the fluctuating inlet stagnation pressure is proposed. The framework effectively categorizes the operation of the unsteady turbine system in both spatial and temporal dimensions; such a framework would facilitate determining whether the loss generation process in a subcomponent can be approximated as unsteady (e.g., volute) or as locally quasi-steady (LQS) (e.g., turbine wheel) in response to the unsteady inlet pulsation in the inlet-to-outlet stagnation pressure ratios of the two inlets. The notion that a specific subcomponent can be approximated as locally quasi-steady while the entire turbine system in itself is unsteady is of interest as it suggests a strategy for an appropriate flow modeling and scaling as well as for the turbine system performance improvement. Also, computed results are used to determine situations where the flow effects in a specific subcomponent can be approximated as quasi-one-dimensional; thus, for instance, the flow mechanisms in the volute can reasonably be approximated on an unsteady one-dimensional basis. For a turbine stage with sudden-expansion type diffuser, the framework for integrating subcomponent models into a turbine system is formulated. The effectiveness and generality of the proposed framework are demonstrated by applying it to three distinctly different turbocharger operating conditions. The estimated power from the integrated turbine system model is in good agreement with the full unsteady computational fluid dynamics (CFD) results for all three situations. The formulated framework will be generally applicable for assessing the new design configurations as long as the corresponding high-fidelity steady CFD results are utilized to determine the quasi-steady (or acoustically compact) behavior of each new subcomponent.
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Zaludin, Zairil A. "Designing a Control Failure Survival System for High Speed Transport Aircraft Using Eigenvalue Assignment Method." Jurnal Teknologi, January 20, 2012. http://dx.doi.org/10.11113/jt.v36.560.
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Jika kerosakan berlaku kepada permukaan kawalan penerbangan, tujuan “Sistem Pereka Bentuk Kawalan Penerbangan” ialah untuk membahagi dan menyelaras usaha kawalan antara permukaan-permukaan kawalan yang masih aktif untuk tujuan mengekalkan mutu penerbangan yang diingini. Tugas utama ‘Sistem Pereka Bentuk Kawalan Penerbangan’ adalah untuk menyelaraskan unit kawalan semasa kerosakan berlaku ataupun menukar unit kawalan kepada sistem kawalan yang lebih sesuai untuk tujuan membaiki kerosakan tersebut. Dalam kertas ini, cara yang kedua dipertimbangkan. Reka bentuk “Sistem Keselamatan Kegagalan Kawalan” untuk pesawat hipersonik dibentangkan. Cara tersebut adalah berdasarkan cara penetapan nilai eigen dan teori pengatur kuadratik linear. Terdapat tiga masukan kawalan ke pesawat tersebut. Jika cara yang dibentangkan di dalam kertas ini digunakan, keputusan analisis yang dibentangkan menunjukkan bahawa sistem kawalan penerbangan untuk pesawat ini boleh direka bentuk sehingga kestabilan pesawat tersebut dicapai semula apabila salah satu ataupun gabungan permukaan kawalan gagal berfungsi pada masa yang sama. Didapati juga gerakan tabii pesawat yang mengalami kerosakan dapat dibaik pulih seperti sebelum kerosakan berlaku. Satu contoh disertakan dalam kertas ini menggunakan model matematik pesawat hipersonik. Kata kunci: dinamik penerbangan; penerbangan hipersonik; kawalan optimal; penetapan nilai eigen; Teori LQR In the event of a control surface failure, the purpose of a reconfigurable flight control system is to redistribute and coordinate the control effort among the aircraft’s remaining effective surfaces such that satisfactory flight performance is retained. A major task in control reconfiguration deals with adjusting the controller gains on-line or switching to a different control law to compensate for the failure. In this paper, the former option is considered. The design of a Control Failure Survival System (CFSS) for a hypersonic transport (HST) aircraft is presented. The method is based on eigenvalue assignment which was developed using Linear Quadratic Regulator theory. There are three control inputs available on board the HST; the change in the flaps deflection, the change in the propulsion diffuser area ratio and the change in the total temperature across combustor. Using the method discussed in this paper, the results showed that it was possible to reconfigure the flight control system such that the aircraft stability is regained when either a single or a combination of, control failures occurred simultaneously. In addition, the natural motion characteristics (i.e short period, phugoid and height motion) of the aircraft before the failure occurred are conserved and the transient response of the aircraft state variables after failure was almost the same as before failure occurred. An example is included in this paper using the mathematical model of the longitudinal motion of the HST. Key words: Aircraft dynamics; hypersonic flight; optimal control; eigen value assignment; LQR Theory
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Blunck, David L., Dale T. Shouse, Craig Neuroth, Amy Lynch, Timothy J. Erdmann, David L. Burrus, Joseph Zelina, Daniel Richardson, and Andrew Caswell. "Experimental Studies of Cavity and Core Flow Interactions With Application to Ultra-Compact Combustors." Journal of Engineering for Gas Turbines and Power 136, no. 9 (March 26, 2014). http://dx.doi.org/10.1115/1.4026975.
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Reducing the weight and decreasing pressure losses of aviation gas turbine engines improves the thrust-to-weight ratio and improves efficiency. In ultra-compact combustors (UCC), engine length is reduced and pressure losses are decreased by merging a combustor with adjacent components using a systems engineering approach. High-pressure turbine inlet vanes can be placed in a combustor to form a UCC. In this work, experiments were performed to understand the performance and associated physics within a UCC. Experiments were performed using a combustor operating at pressures in the range of 520–1030 kPa (75–150 psia) and inlet temperature equal to 480–620 K (865 R–1120 R). The primary reaction zone is in a single trapped-vortex cavity where the equivalence ratio was varied from 0.7 to 1.8. Combustion efficiencies and NOx emissions were measured and exit temperature profiles were obtained for various air loadings, cavity equivalence ratios, and configurations with and without representative turbine inlet vanes. A combined diffuser-flameholder (CDF) was used to study the interaction of cavity and core flows. Discrete jets of air immediately above the cavity result in the highest combustion efficiencies. The air jets reinforce the vortex structure within the cavity, as confirmed through coherent structure velocimetry of high-speed images. The combustor exit temperature profile is peaked away from the cavity when a CDF is used. Testing of a CDF with vanes showed that combustion efficiencies greater than 99.5% are possible for 0.8 ≤ Φcavity ≤ 1.8. Temperature profiles at the exit of the UCC with vanes agreed within 10% of the average value. Exit-averaged emission indices of NOx ranged from 3.5 to 6.5 g/kgfuel for all test conditions. Increasing the air loading enabled greater mass flow rates of fuel with equivalent combustion efficiencies. This corresponds to increased vortex strength within the cavity due to the greater momentum of the air driver jets.
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Mohammad, Bassam, San-Mou Jeng, and M. Gurhan Andac. "Influence of the Primary Jets and Fuel Injection on the Aerodynamics of a Prototype Annular Gas Turbine Combustor Sector." Journal of Engineering for Gas Turbines and Power 133, no. 1 (September 24, 2010). http://dx.doi.org/10.1115/1.4002004.
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Transverse dilution jets are widely used in combustion systems. The current research provides a detailed study of the primary jets of a realistic annular combustion chamber sector. The combustor sector comprises an aerodynamic diffuser, inlet cowl, combustion dome, primary dilution jets, secondary dilution jets, and cooling strips to provide convective cooling to the liner. The chamber contracts toward the end to fit the turbine nozzle ring. 2D PIV is employed at an atmospheric pressure drop of 4% (isothermal) to delineate the flow field characteristics. The laser is introduced to the sector through the exit flange. The interaction between the primary jets and the swirling flow as well as the sensitivity of the primary jets to perturbations is discussed. The perturbation study includes: effect of partially blocking the jets, one at a time, the effect of blocking the convective cooling holes, placed underneath the primary jets and shooting perpendicular to it. In addition, the effect of reducing the size of the primary jets as well as off-centering the primary jets is explained. Moreover, PIV is employed to study the flow field with and without fuel injection at four different fuel flow rates. The results show that the flow field is very sensitive to perturbations. The cooling air interacts with the primary jet and influences the flow field although the momentum ratio has a 100:1 order of magnitude. The results also show that the big primary jets dictate the flow field in the primary zone as well as the secondary zone. However, relatively smaller jets mainly influence the primary combustion zone because most of the jet is recirculated back to the CRZ. Also, the jet penetration is reduced with 25% and 11.5% corresponding to a 77% and 62% reduction of the jet’s area, respectively. The study indicates the presence of a critical jet diameter beyond which the dilution jets have minimum impact on the secondary region. The jet off-centering shows significant effect on the flow field though it is in the order of 0.4 mm. The fuel injection is also shown to influence the flow field as well as the primary jets angle. High fuel flow rate is shown to have very strong impact on the flow field and thus results in a strong distortion of both the primary and secondary zones. The results provide useful methods to be used in the flow field structure control. Most of the effects shown are attributed to the difference in jet opposition. Hence, the results are applicable to reacting flow.