M. D. Barringer

Flow Field Simulations of a Gas Turbine Combustor

The flow field exiting the combustor in a gas turbine engine is quite complex considering the presence of large dilution jets and complicated cooling schemes for the combustor liner. For the most part, however, there has been a disconnect between the combustor and turbine when simulating the flow field that enters the nozzle guide vanes. To determine the effects of a representative combustor flow field on the nozzle guide vane, a large-scale wind tunnel section has been developed to simulate the flow conditions of a prototypical combustor. This paper presents experimental results of a combustor simulation with no downstream turbine section as a baseline for comparison to the case with a turbine vane. Results indicate that the dilution jets generate turbulence levels of 15-18% at the exit of the combustor with a length scale that closely matches that of the dilution hole diameter. The total pressure exiting the combustor in the near wall region neither resembles a turbulent boundary layer nor is it completely uniform putting both of these commonly made assumptions into question.

Migration of Combustor Exit Profiles Through High Pressure Turbine Vanes

The high pressure turbine stage within gas turbine engines is exposed to combustor exit flows that are nonuniform in both stagnation pressure and temperature. These highly turbulent flows typically enter the first stage vanes with significant spatial gradients near the inner and outer diameter endwalls. These gradients can result in secondary flow development within the vane passage that is different than what classical secondary flow models predict. The heat transfer between the working fluid and the turbine vane surface and endwalls is directly related to the secondary flows. The goal of the current study was to examine the migration of different inlet radial temperature and pressure profiles through the high turbine vane of a modern turbine engine. The tests were performed using an inlet profile generator located in the Turbine Research Facility at the Air Force Research Laboratory. Comparisons of area-averaged radial exit profiles are reported as well as profiles at three vane pitch locations to document the circumferential variation in the profiles. The results show that the shape of the total pressure profile near the endwalls at the inlet of the vane can alter the redistribution of stagnation enthalpy through the airfoil passage significantly. Total pressure loss and exit flow angle variations are also examined for the different inlet profiles.

An Experimental Study of Combustor Exit Profile Shapes on Endwall Heat Transfer in High Pressure Turbine Vanes

The design and development of current and future gas turbine engines for aircraft propulsion have focused on operating the high pressure turbine at increasingly elevated temperatures and pressures. The drive toward thermal operating conditions near theoretical stoichiometric limits as well as increasingly stringent requirements on reducing harmful emissions both equate to the temperature profiles exiting combustors and entering turbines becoming less peaked than in the past. This drive has placed emphasis on determining how different types of inlet temperature and pressure profiles affect the first stage airfoil endwalls. The goal of the current study was to investigate how different radial profiles of temperature and pressure affect the heat transfer along the vane endwall in a high pressure turbine. Testing was performed in the Turbine Research Facility located at the Air Force Research Laboratory using an inlet profile generator. Results indicate that the convection heat transfer coefficients are influenced by both the inlet pressure profile shape and the location along the endwall. The heat transfer driving temperature for inlet profiles that are nonuniform in temperature is also discussed.

Computational Design of a Louver Particle Separator for Gas Turbine Engines

The extreme temperatures in a jet engine require the use of thermal barrier coatings and internal cooling channels to keep the components in the turbine section below their melting temperature. The presence of solid particles in the engine’s gas path can erode thermal coatings and clog cooling channels, thereby reducing part life and engine performance. This study uses computational fluid dynamics to design the geometry of a static, inertial particle separator to remove small particles, such as sand, from the engine flow. The concept for the inertial separator includes the usage of a multiple louver array followed by a particle collector. The results of the study show a louver design can separate particles while not incurring large pressure loss.Copyright

Effects of Combustor Exit Profiles on Vane Aerodynamic Loading and Heat Transfer in a High Pressure Turbine

The flow and thermal fields exiting gas turbine combustors dictate the overall performance of the downstream turbine. The goal of this work was to investigate the effects of engine representative combustor exit profiles on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using a nonreacting turbine inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface was affected by different turbine inlet pressure and temperature profiles at different span locations. The results indicate that the inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer relative to a baseline test with uniform inlet total pressure and total temperature. Near the inner diameter endwall, the baseline heat transfer was reduced 30‐40% over the majority of the vane surface. Near the outer dimeter endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 10‐ 20%, while other profiles resulted in a decrease in the baseline heat transfer by 25‐35%. This study also shows the importance of knowing an accurate prediction of the local flow driving temperature when determining vane surface heat transfer. DOI: 10.1115/1.2950051

Experimental Evaluation of an Inlet Profile Generator for High-Pressure Turbine Tests

Improving the performance and durability of gas turbine aircraft engines depends highly on achieving a better understanding of the flow interactions between the combustor and turbine sections. The flow exiting the combustor is very complex and it is characterized primarily by elevated turbulence and large variations in temperature and pressure. The heat transfer and aerodynamic losses that occur in the turbine passages are driven primarily by these spatial variations. To better understand these effects, the goal of this work is to benchmark an adjustable turbine inlet profile generator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory. The research objective was to experimentally evaluate the performance of the nonreacting simulator that was designed to provide representative combustor exit profiles to the inlet of the TRF turbine test section. This paper discusses the verification testing that was completed to benchmark the performance of the generator. Results are presented in the form of temperature and pressure profiles as well as turbulence intensity and length scale. This study shows how a single combustor geometry can produce significantly different flow and thermal field conditions entering the turbine. Engine designers should place emphasis on obtaining accurate knowledge of the flow distribution within the combustion chamber. Turbine inlet conditions with significantly different profile shapes can result in altered flow physics that can change local aerodynamics and heat transfer.

Microchannels With Manufacturing Roughness Levels

As turbine operating temperatures continue to rise, there is the ever present problem of cooling airfoil components. To accomplish this, cooling methods have morphed into a complex array of specialized internal and external schemes. A majority of cooling methods involve routing cool air from the compressor, through internal cooling channels and out through various film cooling holes around the blade surface. These cooling channels serve as a means to transfer heat from the airfoil surfaces to a cooling flow passing through the inside of the vanes and blades. Internal cooling has evolved from simple smooth-walled channels to ones that include various types of turbulators that serve to increase the convective heat transfer coefficient and heat transfer surface area. Turbulators have been studied in detail over the years, and the most preferred cooling designs involve pedestals, ribs, and various surface roughening effects such as dimples. They serve to increase the turbulent mixing of the flow as it passes through the channel, thereby increasing the convective heat transfer coefficient along the wall surfaces. There are many advantages from a cooling perspective to use microchannels to cool components, provided that the pressure to drive the flow through these channels is available. These microchannels approach the level where manufacturing random roughness and tolerances can play a key role in their heat transfer capabilities 1. Average manufacturing roughness in common casting techniques can approach 1‐10 m 2 before polishing or finishing, and this seemly inconsequential inherent roughness can discernibly affect the microchannels’ heat transfer and pressure drop. This paper provides a summary of the findings of an experimental program focused on examining the heat transfer and pressure drop of microchannels. Accurately quantifying these two parameters will enable better prediction of the engine hardware thermal performance and life. The primary objective of this study was to develop a testing procedure capable of evaluating the heat transfer and pressure drop performance of several microchannels over a range of Reynolds numbers. The performance is reported in the form of an overall Nusselt number and friction factor for each manufacturing roughness level. While the results indicate little dependence of heat transfer and pressure drop augmentation upon Reynolds number, the range studied was in the fully turbulent regime in which compressibility effects were not present.

Effects of Purge Jet Momentum on Sealing Effectiveness

Driven by the need for higher cycle efficiencies, overall pressure ratios for gas turbine engines continue to be pushed higher thereby resulting in increasing gas temperatures. Secondary air, bled from the compressor, is used to cool turbine components and seal the cavities between stages from the hot main gas path. This paper compares a range of purge flows and two different purge hole configurations for introducing the purge flow into the rim cavities. In addition, the mate face gap leakage between vanes is investigated. For this particular study, stationary vanes at engine-relevant Mach and Reynolds numbers were used with a static rim seal and rim cavity to remove rotational effects and isolate gas path effects. Sealing effectiveness measurements, deduced from the use of CO2 as a flow tracer, indicate that the effectiveness levels on the stator and rotor side of the cavity depend on the mass and momentum flux ratios of the purge jets relative to the swirl velocity. For a given purge flow rate, fewer purge holes resulted in better sealing than the case with a larger number of holes. [DOI: 10.1115/1.4034545]

The Design of a Steady Aero Thermal Research Turbine (START) for Studying Secondary Flow Leakages and Airfoil Heat Transfer

This paper describes a new gas turbine research facility that was designed and is currently being constructed with a number of unique features that allow experiments which aim at developing and improving sealing and cooling technologies at a reasonable cost. Experiments will include enginerepresentative rotating turbine hardware in a continuous, steady-state, high-pressure flow environment. The facility includes a 1.5 stage turbine that will simulate the aerodynamic flow and thermal field interactions in the engine between the hot mainstream gas path and the secondary air flows at relevant corrected operating conditions and scaling parameters. Testing in the new facility is planned to begin in 2014 and will take place in two phases. The first phase is focused on understanding the behavior of inner-stage gap flow leakages in the presence of the main gas path and the rotating blade platform. The second phase is focused on developing and testing novel cooling methods for turbine airfoils, platforms, and disks, ultimately leading to an integrated approach for more effective use of the secondary cooling flow. The uniqueness of this facility includes a continuous duration facility with engine-relevant rotational and axial Reynolds and Mach numbers at the blade inlet.

Using a Tracer Gas to Quantify Sealing Effectiveness for Engine Realistic Rim Seals

As overall pressure ratios increase in gas turbine engines, both the main gas path and cooling temperatures increase leading to component durability concerns. At the same time effective use of the secondary air for both cooling and sealing becomes increasingly important in terms of engine efficiency. To fully optimize these competing requirements, experiments at engine-relevant conditions are required to validate new designs and computational tools. A test turbine has been commissioned in the Steady Thermal Aero Research Turbine (START) lab. The test turbine was designed to be a 1.5 stage turbine operating under continuous flow simulating engine-relevant conditions including Reynolds and Mach numbers with hardware true to engine scale. The first phase of research conducted using the test turbine, which was configured for a half-stage (vane only), was to study hot gas ingestion through turbine rim seals. This paper presents a series of facility benchmarks as well as validation experiments conducted to study ingestion using a tracer gas to quantify the performance of rim seals and purge flows. Sensitivity studies included concentration levels and sampling flow rates in flow regimes that ranged from stagnant to compressible depending upon the area of interest. The sensitivity studies included a range of purge and leakage flow conditions for several locations in the rim seal and cavity areas. Results indicate reasonable sampling methods were used to achieve isokinetic sampling conditions. INTRODUCTION Secondary air bled from the compressor is required to cool components in the hot section of a gas turbine engine. These components are exposed to temperatures that can degrade the components and lead to durability concerns. Frequent maintenance is costly and undesirable, so cooling air must be provided to the hot section of the engine. Additionally, the cavities between the rotating and stationary components typically do not feature the advanced cooling technologies seen in main gas path hardware; however, hot gas can be ingested into these cavities leading to high temperatures. Some of the secondary air is provided to the cavities to seal or purge the ingested hot gas. Although some secondary air is required for cooling and sealing in a gas turbine, the excessive use of secondary air negatively impacts the efficiency of the engine. As the bypass ratios of aircraft engines increase the fans get larger accelerating more air and the core mass flow rate decreases. To drive the fan the overall pressure ratio increases, leading to higher turbine inlet and cooling air temperatures. Sealing the turbine rim cavities is increasingly important as engine core diameter decreases. Absolute clearances remain consistent, but relative rim seal clearances increase resulting in significant ingestion. To purge the rotor-stator cavities more secondary flow is required. It is thus critical to quantify the sealing effectiveness of rim seal geometries used in decreasing engine core sizes. Quantifying the sealing effectiveness in actual engines is challenging given the high temperatures and the complexity of the flow. The high temperatures in engines can cause sensors to fail so detailed and reliable measurements in the engine are difficult. Conductive heat transfer in the metal turbine components also confounds effectiveness measurements. Frictional heating in the cavities can increase the air and disk temperatures further complicating effectiveness measurements. Sealing effectiveness is the quantification of how well a rim seal prevents main gas path air from being ingested into the Proceedings of ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition GT2016 June 13 – 17, 2016, Seoul, South Korea