Chen, Chao. / Thesis (M.Phil.)--Chinese University of Hong Kong, 2009. / Includes bibliographical references (leaves 43-46). / Abstract also in Chinese. / Abstract --- p.i / Acknowledgement --- p.iii / Chapter 1 --- Introduction --- p.1 / Chapter 2 --- Preliminaries --- p.7 / Chapter 3 --- The mathematical formulation of the problem and main results --- p.11 / Chapter 4 --- Reformulation of the problem --- p.17 / Chapter 5 --- Proof of the main theorems --- p.23 / Chapter 5.1 --- Proof of uniqueness --- p.23 / Chapter 5.2 --- Proof of non-existence --- p.31 / Chapter 6 --- Work in future --- p.40 / Chapter 7 --- Appendix --- p.41 / Bibliography --- p.43

During aeroelastic simulation, forces and displacements must be interpolated between the non-matching fluid and structural meshes, while the volume fluid mesh must deform as the surface moves. Fluidstructure interpolation is necessary because numerical models for fluids and structures use different solvers, and at the interface these meshes do not match. The problem of mesh motion arises from the fact that the discretised fluid volume must conform to the motion of the surface, which means motion of the surface must be diffused into the volume.

The aerodynamics of bodies in compressible ground effect flowfields from low-subsonic to supersonic Mach numbers have been investigated numerically and experimentally. A study of existing literature indicated that compressible ground effect has been addressed sporadically in various contexts, without being researched in any comprehensive detail. One of the reasons for this is the difficulty involved in performing experiments which accurately simulate the flows in question with regards to ground boundary conditions. To maximise the relevance of the research to appropriate real-world scenarios, multiple bodies were examined within the confines of their own specific flow regimes. These were: an inverted T026 wing in the low-to-medium subsonic regime, a lifting RAE 2822 aerofoil and ONERA M6 wing in the transonic regime, and a NATO military projectile at supersonic Mach numbers. Two primary aims were pursued. Firstly, experimental issues surrounding compressible ground effect flows were addressed. Potential problems were found in the practice of matching incompressible Computational Fluid Dynamics (CFD) simulations to wind tunnel experiments for the inverted wing at low freestream Mach numbers (<0.3), where the inverted wing was found to experience significant compressible effects even at Mach 0.15. The approach of matching full-scale CFD simulations to scale model testing at an identical Reynolds number but higher Mach number was analysed and found to be prone to significant error. An exploration was also conducted of appropriate ways to conduct experimental tests at transonic and supersonic Mach numbers, resulting in the recommendation of a symmetry (image) method as an effective means of approximating a moving ground boundary in a small-scale blowdown wind tunnel. Issues of scale with regards to Reynolds number persisted in the transonic regime, but with careful use of CFD as a complement to experiments, discrepancies were quantified with confidence. The second primary aim was to use CFD to gain a broader understanding of the ways in which density changes in the flowfield affect the aerodynamic performance of the bodies in question, in particular when a shock wave reflects from the ground plane to interact again with the body or its wake. The numerical approach was extensively verified and validated against existing and new experimental data. The lifting aerofoil and wing were investigated over a range of mid-to-high subsonic Mach numbers (1>M???>0.5), ground clearances and angles of incidence. The presence of the ground was found to affect the critical Mach number, and the aerodynamic characteristics of the bodies across all Mach numbers and clearances proved to be highly sensitive to ground proximity, with a step change in any variable often causing a considerable change to the lift, moment and drag coefficients. At the lowest ground clearances in both two and three dimensional studies, the aerodynamic efficiency was generally found to be less than that of unbounded (no ground) flight for shock-dominated flowfields at freestream Mach numbers greater than 0.7. In the fully-supersonic regime, where shocks tend to be steady and oblique, a supersonic spinning NATO projectile travelling at Mach 2.4 was simulated at several ground clearances. The shocks produced by the body reflected from the ground plane and interacted with the far wake, the near wake, and/or the body itself depending on the ground clearance. The influence of these wave reflections on the three-dimensional flowfield, and their resultant effects on the aerodynamic coefficients, was determined. The normal and drag forces acting on the projectile increased in exponential fashion once the reflections impinged on the projectile body again one or more times (at a height/diameter ground clearance h/d<1). The pitching moment of the projectile changed sign as ground clearance was reduced, adding to the complexity of the trajectory which would ensue.

Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1997. / Vita. Abstract. Includes bibliographical references (leaves 185-198). Issued also in computer file.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1993. / Vita. Abstract. Includes bibliographical references (leaves 113-115). Also available via the Internet.

<p>The primary objective of this research was to shed light on the changes in performance observed in a high-speed, centrifugal compressor that occur during the transition from subsonic to transonic operating conditions, using experimental data collected on a research compressor developed by Honeywell Aerospace, as well as results from a numerical model of the compressor.</p> <p> An understanding of the flow behavior in transonic centrifugal compressors is critical as the drive for higher stage pressure ratios while maintaining a compact size results in higher rotational speeds and increased aspect ratios in the inducer of the impeller. Both of these design trends result in higher relative Mach numbers near the impeller leading edge, resulting in the formation of shocks and an increasingly complex flow field. Since it is necessary to maintain high efficiency and adequate surge margin at these conditions—to ensure the compressor is stable across the full operating range—it is important to understand the effects of the transition from subsonic to supersonic flow on performance and stability. Due to the limited availability of research in the open literature regarding transonic centrifugal impellers, especially experimental studies, these behaviors are still not fully understood.</p> <p>Experimental data collected during steady state operation as well as during speed transients, showed a sudden decrease in the variance of the unsteady pressure field throughout the compressor, but most dramatically in the inducer shroud. Analysis of the performance also showed a significant increase in impeller efficiency of approximately 2 points as speed was increased from 80% to 90% of the design speed. Temperature measurements upstream of the impeller leading edge indicated a dramatic reduction in the degree of flow recirculation in the same speed range, indicating the increase in performance is related to a decrease in the blockage near the impeller leading edge. A low pressure region was also observed in the inducer passage, which disappeared upon transition to the transonic operating regime, this coupled with decreased inducer static pressure rise and relative diffusion at lower speeds, strongly indicates that increased loss in the inducer at lower speeds is responsible for the observed performance deficiency during subsonic operation.</p> <p>Analysis of the numerical results revealed that the low pressure region in the inducer may be attributable to the interaction between the inlet shroud boundary layer and the low momentum tip leakage flow in the impeller passage, which at lower speeds, results in the tip leakage flow forming a large recirculation region in the inducer passage. It was also determined that the step change in instability coincides with the inducer shock extending to the shroud and reducing the strength of the interaction between the low momentum regions in the inlet and impeller passage, thereby allowing the tip leakage flow to form into a vortex and preventing the development of the recirculation region in the inducer. </p> <p>This research provides a possible explanation for the observed instability in the compressor, which may allow for further testing of techniques to mitigate the instability caused by the blockage in the inducer, such as casing treatment, bleed, or flow injection into the inducer shroud.</p>

<p>The objective of this work is to provide a computational baseline for modelling the flow physics in the tip region of a transonic fan. A transonic fan was donated by Honeywell Aerospace to the Purdue University High-Speed Compressor Research Laboratory for the purposes of studying casing treatments and inlet distortion under the Office of Naval Research Power and Propulsion Program. The purpose of casing treatment is to extend the stall margin of the fan without being detrimental to fan efficiency. Hence, before an effective casing treatment can be designed, understanding the instabilities that lead to stall or surge and understanding the flow field near the rotor tip at different operating conditions is necessary. </p> <p>The behavior of the flow field was studied at design speed using steady simulations for near stall, peak efficiency, and choke operating conditions. The details of the passage shock, tip leakage vortex, and the shock-vortex interaction were investigated. The passage shock moves forward in the rotor passage as the loading increases, until eventually becoming unstarted near stall. The tip leakage vortex convects from the rotor tip leading edge to the pressure side of the adjacent blade, and its trajectory becomes parallel to the rotor inlet plane as the loading increases. The shock-vortex interaction does not cause the tip leakage vortex to breakdown, although distortion of the shock front and diffusion of the tip leakage vortex is significant near stall.</p> <p>To validate this computational model, steady simulations were used to conduct a grid convergence study. A single passage mesh of 8 million elements is sufficient to capture the flow qualitatively, but a mesh of at least 22 million elements is recommended to lower discretization error if quantitative details are important. A brief comparison of turbulence models is made, and the SST model was found to predict stronger radial flows than the BSL-EARSM and BSL-RSM models. However, the SST model still captures the flow features qualitatively, and the more complex models would be too costly for iterative design simulations.</p> <p>The importance of unsteady effects was also considered for a point near peak efficiency. Near peak efficiency, the effect of shock oscillations near the rotor shroud are small. Compared to steady simulations, the unsteady simulation predicts a slightly stronger horseshoe vortex at the hub and a passage shock closer to the rotor leading edge. The tip leakage vortex trajectory appears to be the same between the steady and unsteady simulations.</p> <p>The modelling decisions made in this research are currently only based on comparison between simulations. This model will be calibrated with experimental data in the future to provide a more accurate view of the flow physics inside this transonic fan.</p>

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. / Master of Science

This investigation, in consideration of the sudden separation increase involved in wing drop, was to determine if the incorporated 2-D airfoil exhibits abnormal shock sensitivity. A comparative airfoil study was used to determine if this particular transonic airfoil is prone to abrupt shock movement, resulting in increased regions of separation. / Master of Science

Striving to improve performance and lower weight of aircraft engines, modern compressor blades become thinner and lighter but higher loaded resulting in an increased vulnerability towards flutter. This trend is further aggravated through blisk designs that diminish structural damping and therewith flutter margin. Modern 3D wide-chord blade designs result in complex structural behaviors that add to the difficulty of correctly predicting flutter occurrence. To counteract above tendencies by driving the physical understanding of flutter and thereby helping to improve aero engine design tools, free flutter as well as forced response will be investigated in the 1.5 stage transonic compressor at TU Darmstadt. Aim of the forced response campaign is to determine the system damping in the stable compressor regime. Hence a novel excitation system capable of dynamically exciting specific rotor blade modes is needed. It is aim of the present work to find a promising concept for such a system. In the present work, the requirements for an excitation system to be used in the TUD compressor are defined with respect to achievable frequency, phase controllability, transferred excitation level, mechanical robustness, integrability and cleanliness. Different excitation system concepts, i.e. oscillating VIGVs, rotating airfoils, tangential and axial air injection are investigated numerically. An evaluation of the results obtained through 2D numerical studies proposes axial air injection as the most favorable concept. / Master of Science