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Swirl stabilized combustion is a common technique used in gas turbine engine combustors and is accomplished by introducing swirl into the inlet flow, which enhances mixing and stabilizes the combustion event. Coaxial swirlers introduce the fuel and air axially through concentric inlets and use vanes to impart a tangential component to either the fuel, air, or both flows. The present study conducted a parametric analysis of coaxial swirler design by manufacturing an array of 14 coaxial swirlers scaled for use in low flow, small engine operations which incorporated the same base design but varied the swirl number, Sn, by changing the vane angle between 0 deg and 63 deg, vane count between four and ten, and vane shape between traditional and helical. Each design was experimentally evaluated using nonpremixed air and propane at different flow conditions to correlate swirler design with lean blowout limits, pressure loss, and flame liftoff. Lean blowout was primarily influenced by swirl number, while vane count and shape had significant impact at Sn = 0.8 but little impact at Sn = 1.5. Pressure loss was unchanged below a Sn of 0.6, and unlike lean blowout, Sn had little impact at 0.8 but significant impact at 1.5. Additionally, an initial equation correlating pressure loss with swirler number, vane count, and mass flowrate was developed. Finally, flame liftoff was mainly driven by swirl number, with vane count and shape the next significant design parameters. [ABSTRACT FROM AUTHOR]

Film cooling plays a critical role in providing effective thermal protection to components in modern gas turbine engines. A significant effort has been undertaken over the last 40 years to improve the distribution of coolant and to ensure that the airfoil is protected by this coolant from the hot gases in the freestream. This film, under conditions with high fuel-air ratios, may actually be detrimental to the underlying metal. The presence of unburned fuel from an upstream combustor may interact with this oxygen rich film coolant jet resulting in secondary combustion. The completion of the reactions can increase the gas temperature locally resulting in higher heat transfer to the airfoil directly along the path line of the film coolant jet. This secondary combustion could damage the turbine blade, resulting in costly repair, reduction in turbine life, or even engine failure. However, knowledge of film cooling in a reactive flow is very limited. The current study explores the interaction of cooling flow from typical cooling holes with the exhaust of a fuel-rich well-stirred reactor operating at high temperatures over a flat plate. Surface temperatures, heat flux, and heat transfer coefficients are calculated for a variety of reactor fuel-to-air ratios, cooling hole geometries, and blowing ratios. Emphasis is placed on the difference between a normal cylindrical hole, an inclined cylindrical hole, and a fan-shaped cooling hole. When both air and nitrogen are injected through the cooling holes, the changes in surface temperature can be directly correlated with the presence of the reaction. Photographs of the localized burning are presented to verify the extent and locations of the reaction. [DOI: 10.1115/1.3204616]

It is desirable to accurately predict the heat load on turbine hot section components within the design cycle of the engine. Thus, a set of predictions of the heat flux on the blade outer air seal of a transonic turbine is here validated with time-resolved measurements obtained in a single-stage high-pressure turbine rig. Surface pressure measurements were also obtained along the blade outer air seal, and these are also compared to three-dimensional, Reynolds-averaged Navier-Stokes predictions. A region of very high heat flux was predicted as the pressure side of the blade passed a fixed location on the blade outer air seal, but this was not measured in the experiment. The region of high heat flux was associated both with very high harmonics of the blade-passing event and a discrepancy between predicted and measured time-mean heat-fux levels. Further analysis of the predicted heat flux in light of the experimental technique employed in the test revealed that the elevated heat flux associated with passage of the pressure side might be physical. Improvements in the experimental technique are suggested for future efforts.

A set of experimental data is presented investigating the unsteady aerodynamics associated with a high pressure turbine vane (HPV) and rotor blade (HPB). The data was acquired at the Turbine Research Facility (TRF) of the Air Force Research Laboratory. The TRF is a transient, blowdown facility generating several seconds of experimental data on full scale engine hardware at scaled turbine operating conditions simulating an actual engine environment. The pressure ratio and freestream Reynolds number were varied for this investigation. Surface unsteady pressure measurements on the HPV, total pressure traverse measurements downstream of the vane, and surface unsteady pressure measurements for the rotor blade were obtained. The unsteady content of the HPV surface was generated by the rotor potential field. The first harmonic decayed more rapidly than the second harmonic with a movement upstream causing the second harmonic to be most influential at the vane throat. The blade unsteadiness appears to be caused by a combination of shock, potential field, and vane wake interactions between the vane and rotor blade. The revolution averaged data resulted in higher unsteadiness than a passing ensemble average for both vane and rotor indicating a need to understand each passage for high cycle fatigue (HCF) effects.

Turbine tip and shroud flow and heat transfer are some of the most complex, yet important, issues in turbine design. Most of the work performed to date has been performed in linear cascades and has investigated such items as the effect of tip geometries and turbulence on tip and shroud pressure and heat transfer. There have been very few full annulus or rotating measurements in the literature. Experimental measurements have been made on a single stage high pressure turbine at the US Air Force Turbine Research Facility (TRF) to aid in the understanding of this phenomena. The TRF is a full scale, rotating rig that operates at matched flow conditions to the true turbine environment. Heat flux measurements were acquired with both Pyrex insert strip and button gages, while the pressure measurements were taken with surface-mounted Kulite[R] pressure transducers. This paper presents one of the first full rotating, simultaneous pressure and heat transfer measurements to be taken in the turbine tip shroud region. These measurements provide some of the details needed for accurately quantifying the true flow condition in this complex flow regime. Comparisons between the present data and the existing 2-D cascade data were made. This investigation quantified the effects of Reynolds number, inlet temperature, turbine pressure ratio and inlet flow temperature profiles. This provides a benchmark data set for validation of numerical codes.

The showerhead region of a film-cooled turbine vane in a gas turbine engine involves a complex interaction between mainstream flow and coolant jets. This flow field was studied using three component laser Doppler velocimeter measurements in a simulated turbine vane test facility. Measurements were focused around the stagnation row of holes. Low and high mainstream turbulence conditions were used. The spanwise orientation of the coolant jets, typical for showerhead coolant holes, had a dominating effect. Very high levels of turbulence were generated by the mainstream interaction with the coolant jets. Furthermore, this turbulence was highly anisotropic, with the spanwise component of the turbulent fluctuations being twice as large as the other components. Finally, there was an interaction of the high mainstream turbulence with the coolant injection resulting in increased turbulence levels for the spanwise velocity component, but had little effect on the other velocity components. [DOI: 10.1115/1.1459733]