<p>With the present drive in turbomachine engine developmenttowards thinner and lighter bladings, closer spaced blade rowsand higher aerodynamic loads per blade row and blade, advanceddesign criteria and accurate prediction methods for vibrationalproblems such as forced response become increasingly importantin order to be able to address and avoid fatigue failures ofthe machine early in the design process. The present worksupports both the search for applicable design criteria and thedevelopment of advanced prediction methods for forced responsein transonic turbine stages. It is aimed at a betterunderstanding of the unsteady aerodynamic mechanisms thatgovern forced response in transonic turbine stages and furtherdevelopment of numerical methods for rotor stator interactionpredictions.</p><p>The investigation of the unsteady aerodynamic excitationmechanisms is based on numerical predictions of thethree-dimensional unsteady flow field in representative testturbine stages. It is conducted in three successive steps. Thefirst step is a documentation of the pressure perturbations onthe blade surface and the distortion sources in the bladepassage. This is performed in a phenomenological manner so thatthe observed pressure perturbations are related to thedistortion phenomena that are present in the blade passage. Thesecond step is the definition of applicable measures toquantify the pressure perturbation strength on the bladesurface. In the third step, the pressure perturbations areintegrated along the blade arc to obtain the dynamic bladeforce. The study comprises an investigation of operationvariations and addresses radial forcing variations. With thehelp of this bottom-up approach the basic forcing mechanisms oftransonic turbine stages are established and potential routesto control the aerodynamic forcing are presented.</p><p>For the computation of rotor stator interaction aerodynamicsfor stages with arbitrary pitch ratios a new numerical methodhas been developed, validated and demonstrated on a transonicturbine test stage. The method, which solves the unsteadythree-dimensional Euler equations, is formulated in thefour-dimensional time-space domain and the derivation of themethod is general such that both phase lagged boundaryconditions and moving grids are considered. Time-inclination isutilised to account for unequal pitchwise periodicity bydistributing time co-ordinates at grid nodes such that thephase lagged boundary conditions can be employed. The method isdemonstrated in a comparative study on a transonic turbinestage with a nominal non integer blade count ratio and anadjusted blade count ratio with a scaled rotor geometry. Thepredictions show significant differences in the blade pressureperturbation signal of the second vane passing frequency, whichwould motivate the application of the new method for rotorstator predictions with non-integer blade count ratios.</p>

Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1992. / Vita. Abstract. Includes bibliographical references (leaves 121-127). Also available via the Internet.

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1994. / Vita. Abstract. Includes bibliographical references (leaves 56-58). Also available via the Internet.

No description available.

An experiment was conducted to investigate the aerodynamic losses of two high-pressure steam turbine nozzles (526A, 525B) subjected to a large range of incident angle and exit Mach number. The blades were tested in a 2D transonic windtunnel. The exit Mach number ranged from 0.60 to 1.15 and the incidence was varied from -34o to 35o. Measurements included downstream Pitot probe traverses, upstream total pressure, and endwall static pressures. Flow visualization techniques such as shadowgraph photography and color oil flow visualization were performed to complement the measured data. When the exit Mach number for both nozzles increased from 0.9 to 1.1, the total pressure loss coefficient increased by a factor of 7 as compared to the total pressure losses observed at subsonic condition (M2<0.9). For the range of incidence tested, the effect of flow incidence on the total pressure losses is less pronounced. Based on shadowgraphs taken during the experiment, it's believed that the large increase in losses at transonic conditions is due to strong shock/boundary layer interaction that may lead to flow separation on the blade suction surface. From the measured total pressure coefficients, a modified loss model that accounts for higher losses at transonic conditions was developed. The new model matches the data much better than the existing Kacker-Okapuu model for transonic exit conditions. / Master of Science

No description available.

The influence of freestream turbulence representative of the flow downstream of a modern gas turbine combustor and first stage vane on turbine blade heat transfer has been measured and analytically modeled in a linear, transonic turbine cascade. Measurements were performed on a high turning, transonic turbine blade. The facility is capable of heated flow with inlet total temperature of 120C and inlet total pressure of 10 psig. The Reynolds number based on blade chord and exit conditions (5x106) and the inlet and exit Mach numbers (0.4 and 1.2, respectively) are representative of conditions in a modern gas turbine engine. High intensity, large length-scale freestream turbulence was generated using a passive turbulence-generating grid to simulate the turbulence generated in modern combustors after it has passed through the first stage vane row. The grid produced freestream turbulence with intensity of approximately 10-12% and an integral length scale of 2 cm near the entrance of the cascade passages, which is believed to be representative of the core flow entering a first stage gas turbine rotor blade row. Mean heat transfer results showed an increase in heat transfer coefficient of approximately 8% on the suction surface of the blade, with increases on the pressure surface on the order of two times higher than on the suction surface (approximately 17%). This corresponds to increases in blade surface temperature of 5-10%, which can significantly reduce the life of a turbine blade. The heat transfer data were compared with correlations from published literature with good agreement. Time-resolved surface heat transfer and passage velocity measurements were performed to investigate and quantify the effects of the turbulence on heat transfer and to correlate velocity fluctuations with heat transfer fluctuations. The data demonstrates strong coherence in velocity and heat flux at a frequency correlating with the most energetic eddies in the turbulence flow field (the integral length-scale). An analytical model was developed to predict increases in surface heat transfer due to freestream turbulence based on local measurements of turbulent velocity fluctuations (u'RMS) and length-scale (Lx). The model was shown to predict measured increases in heat flux on both blade surfaces in the current data. The model also successfully predicted the increases in heat transfer measured in other work in the literature, encompassing different geometries (flat plate, cylinder, turbine vane and turbine blade) as well as both laminar and turbulent boundary layers, but demonstrated limitations in predicting early transition and heat transfer in turbulent boundary layers. Model analyses in the frequency domain provided valuable insight into the scales of turbulence that are most effective at increasing surface heat transfer. / Ph. D.

An experimental and numerical investigation was conducted to determine the film cooling effectiveness of a normal slot and angled slot under realistic engine Mach number conditions. Freestream Mach numbers of 0.65 and 1.3 were tested. For the normal slot, hot gas ingestion into the slot was observed at low blowing ratios (M < 0.25). At high blowing ratios (M > 0.6) the cooling film was observed to "lift off" from the surface. For the 30o angled slot, the data was found to collapse using the blowing ratio as a scaling parameter (x/Ms). Results from the current experiment were compared with the subsonic data published to confirm this test procedure. For the angled slot, at the supersonic freestream Mach number, the current experiment shows that at the same x/Ms, the film cooling effectiveness increases by as much as 25% as compared to the subsonic case. The results of the experiment also show that at the same x/Ms, the film cooling effectiveness of the angled slot is considerably higher than that of the normal slot, at both subsonic and supersonic Mach numbers. The flow physics for the slot tests considered here are also described with computational fluid dynamic (CFD) simulations in the subsonic and supersonic regimes. / Master of Science

The relationship between the plenum to freestream total pressure ratio on film cooling performance is experimentally investigated. Measurements of both the heat transfer coefficient and the adiabatic effectiveness were made on the suction side of the center blade in a linear transonic cascade. Entrance and exit Mach numbers were 0.3 and 1.2 respectively. Reynolds number based on chord and exit conditions is 3 x 10⁶. The blade contour is representative of a typical General Electric first stage, high turning, turbine blade. Tunnel freestream conditions were 10 psig total pressure and approximately 80 °C. A chilled air coolant film was supplied to a generic General Electric leading edge showerhead coolant scheme. Pressure ratios were varied from run to run over the ranges of 1.02 to 1.20. The density ratio was near a value of 2. A method to determine both the heat transfer coefficient and film cooling effectiveness from experimental data is outlined. Results show that the heat transfer coefficient is independent of the pressure ratio over these ranges of blowing parameters. Also, there is shown to be a weak reduction of film cooling effectiveness with higher pressure ratios. Results are shown for effectiveness and heat transfer coefficient profiles along the blade. / Master of Science

The effects of using an insert Heat Flux Microsensor (HFM) versus an HFM deposited directly on a turbine blade to measure heat flux in a transonic cascade are investigated. The HFM is a thin-film sensor, 6.35 mm (0.250") in diameter (for an insert gage, including the housing) which measures heat flux and surface temperature. The thermal time response of both gages was modeled using a 1-D, finite difference technique and a 2-D, finite element solver. The transient response of the directly deposited gage was also tested against insert gages using an unsteady shock wave in a bench test setup and using a laser of known output. The effects of physical gage offset from the blade surface were also investigated. The physical offset of an insert HFM near the stagnation point on the suction side of a turbine blade was intentionally varied and the average heat transfer coefficient measured. Turbulence grids were used to study how offset affects the heat transfer coefficient with freestream turbulence added to the flow. The time constant of the directly deposited gage was measured to be 856 ms compared to less than 30 ms for the insert gages. Model results predict less than 20 ms for both gages and rule out the anodization layer (used for electrical isolation of the directly deposited gage from the blade) as the cause for the directly deposited gage's much slower time response. Offsets of ± 0.254 mm (0.010") at the gage location with an estimated boundary layer thickness of 0.10 mm (0.004") produced a higher average heat transfer coefficient than the 0.000" offset case. Using an insert HFM resulted in a higher average heat transfer coefficient than using the directly deposited gage and reduced the effects of freestream turbulence. To accurately measure heat transfer coefficients and the effects of freestream turbulence, the disruption of the flow caused by a gage must be minimized. Depositing a gage directly on the blade minimizes the effects of offset, but the cause of the slow time response must first be resolved if high speed data is to be taken. / Master of Science