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Gas turbine companies are always focused on reducing capital costs and increasing overall efficiency. There are numerous advantages in reducing the number of airfoils per stage in the turbine section. While increased airfoil loading offers great advantages like low cost and weight, they also result in increased aerodynamic losses and associated issues. The strength of secondary flows is influenced by the upstream boundary layer thickness as well as the overall flow turning angle through the blade row. Secondary flows result in stagnation pressure loss which accounts for a considerable portion of the total stagnation pressure loss occurring in a turbine passage. A turbine designer strives to minimize these aerodynamic losses through design changes and geometrical effects. Performance of airfoils with varying loading levels and turning angles at transonic flow conditions are investigated in this study. The pressure difference between the pressure side and suction side of an airfoil gives an indication of the loading level of that airfoil. Secondary loss generation and the 3D flow near the endwalls of turbine blades are studied in detail. Detailed aerodynamic loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 axial chord and 1.0 axial chord locations downstream of the trailing edge. Static pressure measurements on the airfoil surface and endwall pressure measurements were carried out in addition to downstream loss measurements. The application of endwall contouring to reduce secondary losses is investigated to try and understand when contouring can be beneficial. A detailed study was conducted on the effectiveness of endwall contouring on two different blades with varying airfoil spacing. Heat transfer experiments on the endwall were also conducted to determine the effect of endwall contouring on surface heat transfer distributions. Heat transfer behavior has significant effect on the cooling flow needs and associated aerodynamic problems of coolant-mainstream mixing. One of the primary objectives of this study is to provide data under transonic conditions that can be used to confirm/refine loss predictions for the effect of various Mach numbers and gas turning. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. A published experimental study on the effect of end wall contouring on such high turning blades at high exit Mach numbers is not available in open literature. Hence, the need to understand the parametric effects of endwall contouring on aerodynamic and heat transfer performance under these conditions. / Ph. D.

Airfoil endwall contouring has been widely studied during the past two decades for the reduction of secondary losses in turbine passages. Although many endwall contouring methods have been suggested by researchers, an analytical tool based on the passage design parameters is still not available for designers. Hence, the best endwall contour shape is usually decided through an optimization study. Moreover, a general guideline for the endwall shape variation can be extrapolated from the existing literature. It has not been validated whether the optimum endwall shape for one passage can be fitted to other similar passage geometry to achieve, least of all a non-optimum but a definite, reduction in losses. Most published studies were conducted at low exit Mach numbers and only recently some studies on the effect of endwall contouring on aerodynamics performance of a turbine passage at high exit Mach numbers have been published. There is, however, no study available in the open literature for a very high turning blade with a transonic design exit Mach number and the effect of endwall contouring on the heat transfer performance of a turbine passage. During the present study, a robust, aerodynamic performance based numerical optimization methodology for turbine endwall contouring has been developed. The methodology is also adaptable to a range of geometry optimization problems in turbomachinery. It is also possible to use the same methodology for multi-objective aero-thermal optimization. The methodology was applied to a high turning transonic turbine blade passage to achieve a geometry based on minimum total pressure loss criterion. The geometry was then compared with two other endwall geometries. The first geometry is based on minimum secondary kinetic energy value instead of minimum total pressure loss criterion. The second geometry is based on a curve combination based geometry generation method found in the literature. A normalized contoured surface topology was extracted from a previous study that has similar blade design parameters. This surface was then fitted to the turbine passage under study in order to investigate the effect of such trend based surface fitting. Aerodynamic response of these geometries has been compared in detail with the baseline case without any endwall contouring. A new non-contoured baseline design and two contoured endwall designs were provided by Siemens Energy, Inc. The pitch length for these designs is about 25% higher than the turbine passage used for the endwall optimization study. The aerodynamic performance of these endwalls was studied through numerical simulations. Heat transfer performance of these endwall geometries was experimentally investigated in the transonic turbine cascade facility at Virginia Tech. One of the contoured geometries was based on optimum aerodynamic loss reduction criterion while the other was based on optimum heat transfer performance criterion. All the three geometries were experimentally tested at design and off-design Mach number conditions. The study revealed that endwall contouring results in significant performance benefit from the heat transfer performance point of view. / Ph. D.

The possibility of misalignments at interfaces would be increased due to individual parts' assembly and external factors during its operation. In actual engine representative conditions, the upstream misalignments have effects on turbines performance through the nozzle guide vane passages. The current experimental aerodynamic investigation over the nozzle guide vane passage was concentrated on the backward-facing step of upstream misalignments. The tests were performed using two types of vane endwall platforms in a 2D linear cascade: flat endwall and axisymmetric converging endwall. The test conditions were a Mach number of 0.85, Re_ex 1.5*10^6 based on exit condition and axial chord, and a high freestream turbulence intensity (16%), at the Virginia tech transonic cascade wind tunnel. The experimental results from the surface flow visualization and the five-hole probe measurements at the vane-passage exit were compared with the two cases with and without the backward-facing step for both types of endwall platforms. As a main source of secondary flow, a horseshoe vortex at stagnation region of the leading edge of the vane directly influences other secondary flows. The intensity of the vortex is associated with boundary layer thickness of inlet flow. In this regard, the upstream backward-facing step as a misalignment induces the separation and attachment of the inlet flow sequentially, and these cause the boundary layer of the inlet flow to reform and become thinner locally. The upstream-step positively affects loss reduction in aerodynamics due to the thinner inlet boundary layer, which attenuates a horseshoe vortex ahead of the vane cascade despite the development of the additional vortices. And converging endwall results in an increase of the effect of the upstream misalignment in aerodynamics, since the inlet boundary layer becomes thinner near the vane's leading edge due to local flow acceleration caused by steep contraction of the converging endwall. These results show good correlation with many previous studies presented herein. / Master of Science

The present study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade consisting of seven blades, the three center blades having a variable tip clearance gap. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects on the near-tip regions at 94% based on engine blade span of the pressure and suction sides. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility with a seven-blade linear cascade. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on engine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 x 105, 9.0 x 105, and 1.1 x 106 based on blade true chord. The tests were performed with a freestream turbulence intensity of 12%. Results at 0.85 exit Mach showed that an increase in the tip gap clearance translates into a 12% increase in the heat transfer coefficients on the blade tip surface. Similarly, at 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 also led to a 24% increase in heat transfer on the tip. High heat transfer was obtained at the leading edge area of the blade tip, and an increase in the tip clearance gap and exit Mach number augmented this leading edge heat transfer. At 94% of engine blade span on the suction side near the tip, a peak in heat transfer was observed in all test cases at an s/C of 0.66 due to the onset of a downstream leakage vortex. At the design condition, this peak represents an increase of a factor of 2.5 from the immediate preceding s/C location. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer as a result of the suction side leakage vortices. / Master of Science

A high-frequency combination probe was used to measure dynamic flow quality in the test section of the NASA Langley 0.3-m Transonic Cryogenic Tunnel. The probe measures fluctuating stagnation (total) temperature and pressure, static pressure, and flow angles in two orthogonal planes. Simultaneous unsteady temperature and pressure measurements were also made in the settling chamber of the tunnel. The data show that the stagnation temperature fluctuations remain constant, and the stagnation pressure fluctuations increase by a factor of two, as the flow accelerates from the settling chamber to the test section. In the test section, the maximum rms value of the normalized fluctuating velocity is 0.7 percent. Correlation coefficients l failed to show vortlcity, entropy, or sound as the dominant mode of turbulence in the tunnel. At certain tunnel operating conditions, periodic disturbances are seen in the data taken in the test section. A possible cause for the disturbances is found to be acoustic coupling of the test section and plenum chamber via the perforated side walls in the tunnel. The experimental data agree well with the acoustic coupling theory. / Master of Science

This thesis presents the results from measurements taken during the transient unstable operation of an axial-flow transonic core-compressor rotor. The measurements were taken to better understand the unstable flow physics of transonic rotors. The rotor, commonly referred to as Rotor 37, was designed by NASA Lewis to be the first stage of an advanced, eight-stage, core-compressor having a high pressure ratio (about 20:1), good efficiency and sufficient stall margin. The rotor was tested without the presence of a stator (or any of the following seven stages) at the NASA Lewis single-stage, high-speed, core-compressor test-rig. The measurements were obtained with a single circumferential, high-response, total pressure and total temperature probe. The measurements were taken immediately after the machine was ’tripped’ into unstable operation by slowly closing the downstream throttle valve. Measurements were obtained at several different span-wise locations and at two different operating speeds. The rotor was shown to exhibit many of the same characteristics typical of low-speed axial-flow machines. Both rotating stall cells and surge cycles were present during unstable operation. The surge cycles present immediately after the inception of unstable operation involved a large-extent single-cell type rotating stall that was present only during the first half of the surge cycles (the second half of these surge cycles involved operation in the stable operating region). However, as the unstable operation progressed (approximately three to five surge cycles later), surge cycles were present that contained a multiple-cell smaller-extent type rotating stall that existed throughout the entire surge cycle with no partial operation in the stable operating region. Thus, compressor system recovery from single-cell large-extent rotating stall (partial operation in stable operating range during the surge cycle) is more probable than recovery from multiple-cell small-extent rotating stall (no operation in stable operating range during the surge cycle). Rotor wheel speed was shown to be an important variable in influencing the form of unstable operation. Surge and rotating stall were shown to be coupled during the unstable operation. Furthermore, the surge/stall coupling was shown to be related more by pressure interactions than by temperature or efficiency interactions. Also, this high hub-tip ratio transonic rotor was shown to exhibit instantaneous stalling across the entire blade span (typical of low-speed, high hub-tip ratio machines). Attempts to fit the data to Greitzer’s one-dimensional lumped-parameter model are presented and the reasons for poor agreement are discussed. / Master of Science

Recent transonic airliner designs have generally converged upon a common cantilever low-wing configuration. It is unlikely that further large strides in performance are possible without a significant departure from the present design paradigm. One such alternative configuration is the strut-braced wing, which uses a strut for wing bending load alleviation, allowing increased aspect ratio and reduced wing thickness to increase the lift to drag ratio. The thinner wing has less transonic wave drag, permitting the wing to unsweep for increased areas of natural laminar flow and further structural weight savings. High aerodynamic efficiency translates into reduced fuel consumption and smaller, quieter, less expensive engines with lower noise pollution. A Multidisciplinary Design Optimization (MDO) approach is essential to understand the full potential of this synergistic configuration due to the strong interdependency of structures, aerodynamics and propulsion. NASA defined a need for a 325-passenger transport capable of flying 7500 nautical miles at Mach 0.85 for a 2010 date of entry into service. Lockheed Martin Aeronautical systems (LMAS), our industry partner, placed great emphasis on realistic constraints, projected technology levels, manufacturing and certification issues. Numerous design challenges specific to the strut-braced wing became apparent through the interactions with LMAS, and modifications had to be made to the Virginia Tech code to reflect these concerns, thus contributing realism to the MDO results. The SBW configuration is 9.2-17.4% lighter, burns 16.2-19.3% less fuel, requires 21.5-31.6% smaller engines and costs 3.8-7.2% less than equivalent cantilever wing aircraft. / Master of Science

Experimental heat transfer data has been collected at engine representative conditions in this work to use in future work to improve computational models. Tests were carried out in a transonic cascade wind tunnel with the data collected using thin film gages. All of the necessary development to use the thin film gages has been completed, including construction of electronics and analysis tools to reduce the data. Gage installation and calibration techniques have been successfully implemented for the current research facility and those procedures have been documented. Heat transfer tests were carried out at engine design speed as well as conditions both above and below design speed. The resulting effect of different Reynolds numbers on heat transfer has been studied and the data collected has been compared with computer predictions, analytical correlations, and data from other published literature to validate the results obtained. Finally, it has been shown that a transient analysis technique can be used to process the data for gages that do not exhibit steady results during the quasi-steady test run. This transient technique resulted in data that agrees well with published literature and analytical correlations. / Master of Science