Zugl.: Berlin, Freie Univ., Diss., 2008

T he casing treatment technique for the axial fan has never been more significant since its potential applications were recognized in gas turbines, tunnel ventilation and many other industrial applications where the axial fan would benefit from the casing treatment. In the last two decades experimental investigations were carried out at Cranfield University to examine the influence of recess casing treatment on stall margin, operating efficiency and flow field of a low-speed axial flow fan. They showed more than 50% improvement in the stall margin with a negligible loss in the efficiency. However, a little work has been done on the numerical simulation of casing treatments due to its complexities, even though in recent years computational fluid dynamics [CFD] analysis has been very active in the prediction of various phenomena in turbomachinery. This work presents numerical investigation of flow in a single axial-flow fan with and without recess casing treatment. It involves the detailed effect of the recess casing on stall margin improvement as well as its influence on global performance parameters. The project offers a contribution to the understanding of the physical processes occurring when approaching stall and the working mechanism by which recess casing treatments improve stall margin. A Reynolds-Averaged Navier-Stokes CFD code was used for the analysis using steady and unsteady simulations. The numerical investigation of the overall performance, efficiency and work-input characteristics of the fan were found to agree very well with the previously reported experimental results. The effect of casing treatment was investigated using two types of configurations, vaneless and vaned casing. The vaneless casing treatment produced a sizeable stall margin improvement with a measurable loss in both pressure rise and efficiency. The recess was fitted later with vanes and was shown to offer both a further stall margin improvement and an increase in the pressure rise coefficient without any significant drop in efficiency at design conditions. The effect of number of vanes inside the recess was also investigated by doubling and halving the number of vanes originally adopted. The predicted results highlighted the importance of the vane inside the casing. Unsteady simulations for the fan with solid and treated casing were carried out. The solid casing simulated for a single blade passage as well as for the entire fan containing all 27 blades highlighted the flow physics of the tip stall growth process, as a large amount of radial flow injected from the hub at the blade suction side near the trailing edge towards the outer casing and occupy this through a mechanism of radial low momentum flow transport. This transport process is the main contributor to the very large separation observed in the shroud region in addition to the locally induced separation due to high blade loading and tip clearance. Although the examination of the unsteady simulation of the recess treatment cavities does not offer an image of large scale unsteady activity at the flow condition investigated, this is on itself quite significant and enables the drawing of an important conclusion namely that large casing treatments rely primarily on a steady-state flow process. The corollary of this conclusion is of course that a steady-state simulation should then be sufficient to capture the essential features of the recess treatment.

No description available.

Zugl.: Berlin, Freie Universiẗat, Diss., 2008.

Zugl.: Berlin, Freie Universiẗat, Diss., 2008.

Thesis (MScEng (Electrical and Electronic Engineering))--University of Stellenbosch, 2010. / ENGLISH ABSTRACT: This thesis presents the development of a stall prevention flight control subsystem, which can easily be integrated into existing flight control architectures of fixed-wing unmanned aerial vehicles (UAV’s). This research forms an important part of faulttolerant flight control systems and will ensure that the aircraft continues to operate safely within its linear aerodynamic region. The focus of this thesis was the stall detection and prevention problem. After a thorough literature study on the topic of stall, a model based stall prevention control algorithm with feedback from an angle of attack sensor was developed. This algorithm takes into account the slew rate and saturation limits of the aircraft’s servos and is able to predict when the current flight condition will result in stall. The primary concern was stall during wings-level flight and involved the prevention of stall by utilising only the elevator control surface. A model predictive slew rate control algorithm was developed to override and dynamically limit the elevator command to ensure that the angle of attack does not exceed a predefined limit. The stall prevention control system was designed to operate as a switching control scheme, to minimise any restrictions imposed on the existing flight control system. Finally, software in the loop simulations were conducted using a nonlinear aircraft model and realistic sensor noise, to verify the theoretical results obtained during the development of this stall prevention control strategy. A worst-case performance analysis was also conducted to investigate the robustness of the control algorithms against model uncertainties. / AFRIKAANSE OPSOMMING: Hierdie tesis handel oor die ontwikkeling van ’n staak voorkomings-vlugbeheer substelsel wat maklik geïntegreer kan word in bestaande vlugbeheer argitektuur van onbemande vaste-vlerk lugvaartuie. Hierdie tesis vorm ’n belangrike deel van fouttolerante vlugbeheertegnieke en sal verseker dat die vliegtuig slegs binne sy lineêre aerodinamiese werksgebied bly. Die fokus van hierdie tesis is die staak opsporing en voorkomings probleem. Na afloop van ’n deeglike literatuurstudie oor die onderwerp van staak, is ’n model gebaseerde staak voorkomings-beheertegniek ontwikkel, wat terugvoer van ’n invalshoek sensor ontvang. Hierdie algoritme neem die sleur tempo en defleksie limiete van die vliegtuig se servos in ag en is in staat om staak te voorspel. Die primêre oorweging was staak tydens simmetriese vlugte en behels slegs die voorkoming van staak deur gebruik te maak van die hei beheer oppervlak. ’n Model voorspellings sleur tempo beheeralgoritme is ontwikkel om die hei-roer dinamies te beperk sodat die invalshoek nie ’n sekere vooraf bepaalde limiet oorskry nie. Die staak voorkomings beheerstelsel is ontwerp om te funksioneer as ’n skakel beheer skema om die beperkings op die bestaande vlugbeheerstelsel te minimaliseer. Laastens was sagteware-in-die-lus simulasies gebruik om die teoretiese resultate, wat verkry is tydens die ontwikkeling van hierdie staak voorkomings beheer-strategie, te kontroleer. Om die robuusthied van hierdie beheeralgoritmes teen model onsekerhede te ondersoek, is ’n ergste-geval prestasie analise ook uitgevoer.

The leading edge flow structure of the NACA 0012 airfoil is experimentally investigated under dynamic stall conditions (M = 0.1; α = 16.7◦, 22.4◦; Rec = 1× 10^6) using planar particle image velocimetry. The airfoil was dynamically pitched about the 1/4 chord at a reduced frequency, k = 0.1. As expected, on the upstroke the flow remains attached in the leading edge region above the static stall angle, whereas during downstroke, the flow remains separated below the static stall angle. A phase averaging procedure involving triple velocity decomposition in combination with the Hilbert transform enables the entire dynamic stall process to be visualized in phase space, with the added benefit of the complete phase space composed of numerous wing oscillations. The formation and complex evolution of the leading edge vortex is observed. This vortex is seen to grow, interact with surrounding vorticity, detach from the surface, and convect downstream. A statistical analysis coupled with instantaneous realizations results in the modification of the classical dynamic stall conceptual model, specifically related to the dynamics of the leading edge vortex.

In todays highly maneuverable jet aircraft designs, aircraft are required to have a propulsion system that can operate during sudden accelerations and rapid changes in angle-of-attack. Consequently, the compressor of the jet engine occasionally must operate at low-flow rates and rapid changes in inlet conditions. The high angle-of-attack and low-flow regime of compressor operation is often plagued by rotating stall and surge. Rotating stall and surge can result in loss of engine performance, rapid heating of the blades, and severe mechanical stresses. Traditional methods for suppressing rotating stall and surge only partially protect against rotating stall or reduce compressor efficiency. The objective of this research is to design a stall suppression system that will introduce oscillatory blowing into one of the rotor blade (stall suppression blade). This oscillatory blowing method has been tested on a wing section in a wind tunnel and has shown to increase the stall angle-of-attack by several degrees.\cite{gilarranzetal02} This increase in stall angle-of-attack will eliminate stall cells as they form in the compressor. The goal of this research is to design a single stage axial compressor that will incorporate the new oscillatory blowing stall suppression system; moreover, this research will design, build, and test a scaled down version of this suppression system.

No description available.

This research will investigate various swept wing models, designing the mechanism for their pitching motion and control, designing wind tunnel implementation, and performing data measurements and analysis using particle image velocimetry. A NACA0012 section with an aspect ratio of AR = 4, free stream velocity of U∞=34 m/s, and Reynolds Number is Rec=2x105. Swept airfoils of Λ=0°, 15°, and 30° will be pitched sinusoidally between an AoA of 4°and 22°, at a reduced frequency of k=πfc/U∞=0.2. Higher sweep angles developing arch-type vortices interact with wing tip flow and abrupt tip stall is observed. Lower sweep angles possessed defined leading edge vortices persist near the tip after lift has collapsed at mid span. Stall angle was delayed during dynamic motion of the wing as well as the presence of arch and ring type vortices increased with sweep angle and contributed to flow reattachment along the top surface of the wing.