3-hole pressure probes are capable of accurately measuring flow angles in the yaw plane. These probes can be utilized inside a jet engine hot section for diagnostics and flow characterization. Sand and other particulate pose a significant risk to hot section components and measurement devices in gas turbine engines. The objective of this experiment was to develop a better understanding of the sensitivity of experimental 3-hole pressure probe designs to engine realistic sand fouling. In this study, Wedge, Cylindrical, and Trapezoidal probes were exposed to realistic hot section turbine environments of 1050 C at 65-70 m/s. 0-5 micron Arizona Road Dust(ARD) is heated under these conditions and used to foul the yaw probes. The sand deposited on the probe was observed to peel off the probe in thin sheets during ambient cool down. Sand fouling was assessed using a stereoscope and digital camera. Probe calibrations were performed in an ambient temperature, open air, calibration jet to mimic engine cold start conditions at Mach numbers of 0.3 and 0.5. Yaw coefficients were calculated for each probe using probe pressure and jet dynamic pressure readings. These coefficients were used to develop calibration curves for each probe initially, and again for every fouling test. Each probe performed differently, but the trends showed that the sand fouling had little impact on the probe error at Mach 0.3, and a slightly increased effect on the probe error at Mach 0.5. The experiment showed that when flow direction was determined using a true dynamic pressure reading from the jet, the probes were able to accurately measure flow direction even after being significantly sanded, some probes holes being over 50% blocked by sand accumulation. Accelerated erosion testing showed that the trapezoidal yaw probe was by far the most sensitive to sand accumulation, followed by the cylindrical probes, and the least sensitive was the wedge probe. A yaw angle range of interest was chosen to ±10 deg of yaw. The least errors from the Yaw Coefficient, as defined in this report, were found to be in the Trapezoidal and Perpendicular probe configurations. The least error found in the wedge probe. / MS / 3-hole pressure probes are used to measure the speed and direction of air and other fluid flows. These probes can be used inside an active jet engine to measure aspects of the airflow inside the engine during flight. One risk to aircraft engines is sand being ingested into the engine. This can cause significant damage to the engine as well as the hardware inside the engine. The objective of this experiment will be to determine how sand accumulation affects the performance of these probes. The experiment involved sanding the probes in a hot jet, then placing them in front of a room temperature air jet to take measurements. A microscope was used to determine how much sand was on the holes of the probe. Sand was observed to peel off naturally, as the probe cooled from the hot jet. Sand was also noticed to break off during the room temperature jet. The experiment showed that when the Jet pressures was measured from inside the jet, the probes were able to accurately measure flow direction even after being significantly sanded, <50% of the holes being blocked by sand. Of all the probes tested, the Wedge probe performed the best, though a close second was the Trapezoidal probe.

Thesis (MEng.)-- University of Stellenbosch, 1992. 140 leaves on single pages, preliminary pages i-xi and numbered pages 1-113. Includes bibliography. Digitized at 600 dpi grayscale to pdf format (OCR),using an Bizhub 250 Konica Minolta Scanner and at 300 dpi grayscale to pdf format (OCR), using a Hp Scanjet 8250 Scanner. / Thesis (MEng (Mechanical and Mechatronic Engineering))--University of Stellenbosch, 1992 / ENGLISH ABSTRACT: The large variety of turbo-machinery blade root geometries in use in industry prompted the question if a optimum geometry could be found. An optimum blade root was defined as a root with a practical geometry which, when loaded, returns the minimum fillet stress concentration factor. A literature survey on the subject provided guidelines but very little real data to work from. An initial optimization was carried out using a formula developed by Heywood to determine loaded projection fillet stresses. The method was found to produce unsatisfactory results, prompting a photoelastic investigation. This experimental optimization was conducted in two stages. A single tang defined load stage and a single tang in-rotor stage which modeled the practical situation. The defined load stage was undertaken in three phases. The first phase was a preliminary investigation, the second phase was a parameter optimization and the third phase was a geometric optimization based on a material utilization optimization. This material optimization approach produced good results. From these experiments a practical optimum geometry was defined. A mathematical model which predicts the fillet stress concentration factor for a given root geometry is presented. The effect of expanding the single tang optimum to a three tang root was examined. / AFRIKAANSE OPSOMMING: Die groot verskeidenheid lemwortelgeometrieë wat in turbomasjiene gebruik word het die vraag na 'n optimum geometrie laat ontstaan. Vir hierdie ondersoek is 'n optimum geometrie gedefineer as 'n praktiese geometrie wat, as dit belas word, die mimimum vloeistukspanningskonsentrasiefaktor laat ontstaan. 'n Literatuur studie het riglyne aan die navorsing gegee maar het wynig spesifieke en bruikbare data opgelewer. Die eerste optimering is met die Heywood formule, wat vloeistukspannings in belaste projeksies bepaal, aangepak. Die metode het nie bevredigende resultate opgelewer nie. 'n Fotoelastiese ondersoek het die basis vir verdere optimeering gevorm. Hierdie eksperimentele optimering is in twee stappe onderneem. 'n Enkelhaak gedefineerde lasgedeelte en 'n enkelhaak in-rotor gedeelte het die praktiese situasie gemodeleer. Die gedefineerde lasgedeelte is in drie fases opgedeel. Die eerste fase was n voorlopige ondersoek. Die tweede fase was 'n parameter optimering. 'n Geometrie optimering gebasseer op 'n materiaal benuttings minimering het die derde fase uitgemaak. Die materiaal optimerings benadering het goeie resultate opgelewer. Vanuit hierdie eksperimente is 'n optimum praktiese geometrie bepaal. 'n Wiskundige model is ontwikkel, wat die vloeistukspanningskonsentrasiefaktor vir 'n gegewe wortelgeometrie voorspel. Die resultaat van 'n geometriese uitbreiding van die enkelhaaklemwortel na 'n driehaaklemwortel op die spanningsverdeling is ondersoek.

This paper gives an overview of the mathematical and practical tools that can be used in turbo-machinery RANS simulation to realize unstructured mesh adaptation. It first presents the concept of metric and recalls that the hessian of the physical flow properties can become, thanks to small modifications, both a metric and a upper bound of the P1 projection error. The resulting metric is then studied on a simple 2D case. In a second part, the industrial application of this concept is addressed and the tools used to overcome the turbo-machinery specificities are explained. Finally, some 2D and 3D results are presented.

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Due to the requirements for high blade loading, modern turbo‐machine blades operate very close to the stall regime. This can lead to flow separation with periodic shedding of vortices, which could lead to self induced oscillations or stall flutter of the blades. Previous studies on stall flutter have focused on flows at high Reynolds number (Re ~ 106). The Reynolds numbers for fans/propellers of Micro Aerial Vehicles (MAVs), high altitude turbofans and small wind turbines are substantially lower (Re < 105). Aerodynamic characteristics of flows at such low Re is significantly different from those at high Re, due in part to the early separation of the flow and possible formation of laminar separation bubbles (LSB). The present study is targeted towards study of stall flutter in a cascade of blades at low Re. We experimentally study stall flutter of a cascade of symmetric NACA 0012 blades at low Reynolds number (Re ~ 30, 000) through forced sinusoidal pitching of the blades about mean angles of incidences close to stall. The experimental arrangement permits variations of the inter‐blade phase (σ) in addition to the oscillation frequency (f) and amplitude; the inter‐blade phase angle (σ) being the phase difference between the motions of adjacent blades in the cascade. The unsteady moments on the central blade in the cascade are directly measured, and used to calculate the energy transfer from the flow to the blade. This energy transfer is used to predict the propensity of the blades to undergo self‐induced oscillations or stall flutter. Experiments are also conducted on an isolated blade in addition to the cascade. A variety of parameters can influence stall flutter in a cascade, namely the oscillation frequency (f), the mean angle of incidence, and the inter‐blade phase angle (σ). The measurements show that there exists a range of reduced frequencies, k (=πfc/U, c being the chord length of the blade and U being the free stream velocity), where the energy transfer from the flow to the blade is positive, which indicates that the flow can excite the blade. Above and below this range, the energy transfer is negative indicating that blade excitations, if any, will get damped. This range of excitation is found to depend upon the mean angle of incidence, with shifts towards higher values of k as the mean angle of incidence increases. An important parameter for cascades, which is absent in the isolated blade case is the inter‐blade phase angle (σ). An excitation regime is observed only for σ values between ‐450 and 900, with the value of excitation being maximum for σ of 900. Time traces of the measured moment were found to be non‐sinusoidal in the excitation regime, whereas they appear to be sinusoidal in the damping regime. Stall flutter in a cascade has differences when compared with an isolated blade. For the cascade, the maximum value of excitation (positive energy transfer) is found to be an order of magnitude lower compared to the isolated blade case. Further, for similar values of mean incidence angle, the range of excitation is at lower reduced frequencies for a cascade when compared with an isolated blade. A comparison with un‐stalled or classical flutter in a cascade at high Re, shows that the inter‐blade phase angle is a major factor governing flutter in both cases. Some differences are observed as well, which appear to be due to stalled flow and low Re.

This thesis aims to explore the potential for using large eddy simulation (LES) as a predictive tool for gas-turbine compressor flows. Compressors present a significant challenge for the Reynolds Averaged Navier-Stokes (RANS) based CFD methods commonly used in industry. RANS models require extensive calibration to experimental data, and thus cannot be used predictively. This thesis explores how LES can offer a more predictive alternative, by exploring the sensitivity of LES to sources of uncertainty. Specifically, the importance of the numerical scheme, the Sub-Grid Scale (SGS) model, and the correct specification of inflow turbulence is examined. The sensitivity of LES to the numerical scheme is explored using the Taylor-Green vortex test case. The numerical smoothing, controlled by a user defined smoothing constant, is found to be important. To avoid tuning the numerical scheme, a locally adaptive smoothing (LAS) scheme is implemented. But, this is found to perform poorly in a forced isotropic turbulence test case, due to the intermittency of the dispersive error. A novel scheme, the LAS with windowing (LASW) scheme, is thus introduced. The LASW scheme is shown to be more suitable for predictive LES, as it does not require tuning to a known solution. The LASW scheme is used to perform LES on a compressor cascade, and results are found to be in close agreement with direct numerical simulations. Complex transition mechanisms, combining characteristics of both natural and bypass modes, are observed on the pressure surface. These mechanisms are found to be sensitive to numerical smoothing, emphasising the importance of the LASW scheme, which returns only the minimum smoothing required to prevent dispersion. On the suction surface, separation induced transition occurs. The flow here is seen to be relatively insensitive to numerical smoothing and the choice of SGS model, as long as the Smagorinsky-Lilly SGS model is not used. These findings are encouraging, as they show that, with the LASW scheme and a suitable SGS model, LES can be used predictively in compressor flows. In order to be predictive, the accurate specification of inflow conditions was shown to be just as important as the numerics. RANS models are shown to over-predict the extent of the three dimensional separation in the endwall - suction surface corner. LES is used to examine the challenges for RANS in this region. The LES shows that it is important to accurately capture the suction surface transition location, with early transition leading to a larger endwall separation. Large scale aperiodic unsteadiness is also observed in the endwall region. Additionally, turbulent anisotropy in the endwall - suction surface corner is found to be important. Adding a non-linear term to the RANS model leads to turbulent stresses that are in better agreement with the LES. This results in a stronger corner vortex which is thought to delay the corner separation. The addition of a corner fillet reduces the importance of anisotropy, thereby reducing the uncertainty in the RANS prediction.

The aerodynamic integration of an aero-engine intake system with the airframe can pose some notable challenges. This is particularly so for many military air- craft and is likely to become a more pressing issue for both new military systems with highly embedded engines as well as for novel civil aircraft configurations. During the late 1960s with the advent of turbo-fan engines, industry became in- creasingly aware of issues which arise due to inlet total pressure distortion. Since then, inlet-engine compatibility assessments have become a key aspect of any new development. In addition to total temperature and total pressure distortions, flow angularity and the associated swirl distortion are also known to be of notable con- cern. The importance of developing a rigorous methodology to understand the effects of swirl distortion on turbo-machinery has also become one of the major concerns of current design programmes. The goal of this doctoral research was to further the current knowledge on swirl distortion, and its adverse effects on engine performance, focusing on the turbo-machinery components (i.e. fans or compressors). This was achieved by looking into appropriate swirl flow descriptors and by correlating them against the compressor performance parameters (e.g loss in stability pressure ratios). To that end, a number of high-fidelity three-dimensional Computational Fluid Dynamics (CFD) models have been developed using two sets of transonic rotors (i.e. NASA Rotor 67 and 37), and a stator (NASA Stator 67B). For the numerical purpose, a boundary condition methodology for the definition of swirl distortion patterns at the inlet has been developed. Various swirl distortion numerical parametric studies have been performed using the modelled rotor configurations. Two types of swirl distortion pattern were investigated in the research, i.e. the pure bulk swirl and the tightly-wound vortex. Numerical simulations suggested that the vortex core location, polarity, size and strength greatly affect the compressor performance. The bulk swirl simula- tions also showed the dependency on swirl strength and polarity. This empha- sized the importance of quantifying these swirl components in the flow distortion descriptors. For this, a methodology have been developed for the inlet-engine compatibility assessment using different types of flow descriptors. A number of correlations have been proposed for the two types of swirl distortion investigated in the study.

The aerodynamic integration of an aero-engine intake system with the airframe can pose some notable challenges. This is particularly so for many military air- craft and is likely to become a more pressing issue for both new military systems with highly embedded engines as well as for novel civil aircraft configurations. During the late 1960s with the advent of turbo-fan engines, industry became in- creasingly aware of issues which arise due to inlet total pressure distortion. Since then, inlet-engine compatibility assessments have become a key aspect of any new development. In addition to total temperature and total pressure distortions, flow angularity and the associated swirl distortion are also known to be of notable con- cern. The importance of developing a rigorous methodology to understand the effects of swirl distortion on turbo-machinery has also become one of the major concerns of current design programmes. The goal of this doctoral research was to further the current knowledge on swirl distortion, and its adverse effects on engine performance, focusing on the turbo-machinery components (i.e. fans or compressors). This was achieved by looking into appropriate swirl flow descriptors and by correlating them against the compressor performance parameters (e.g loss in stability pressure ratios). To that end, a number of high-fidelity three-dimensional Computational Fluid Dynamics (CFD) models have been developed using two sets of transonic rotors (i.e. NASA Rotor 67 and 37), and a stator (NASA Stator 67B). For the numerical purpose, a boundary condition methodology for the definition of swirl distortion patterns at the inlet has been developed. Various swirl distortion numerical parametric studies have been performed using the modelled rotor configurations. Two types of swirl distortion pattern were investigated in the research, i.e. the pure bulk swirl and the tightly-wound vortex. Numerical simulations suggested that the vortex core location, polarity, size and strength greatly affect the compressor performance. The bulk swirl simula- tions also showed the dependency on swirl strength and polarity. This empha- sized the importance of quantifying these swirl components in the flow distortion descriptors. For this, a methodology have been developed for the inlet-engine compatibility assessment using different types of flow descriptors. A number of correlations have been proposed for the two types of swirl distortion investigated in the study.

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