Hyoun-Woo Shin

Boundary Layer Development in Axial Compressors and Turbines: Part 3 of 4— LP Turbines

This is Part Three of a four-part paper. It begins with Section 11.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this part, we present the experimental evidence that we used to construct the composite picture for LP turbines that was given in the discussion in Section 5.0 of Part 1. We present and interpret the data from the surface hot-film gages and the boundary layer surveys for the baseline operating condition. We then show how this picture changes with variations in Reynolds number, airfoil loading, and nozzle-nozzle clocking.

Boundary Layer Development in Axial Compressors and Turbines: Part 1 of 4—Composite Picture

Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading. Part 1 of this paper summarizes all of our experimental findings by using sketches to show how boundary layers develop on compressor and turbine blading. Parts 2 and 3 present the detailed experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data. Readers not interested in experimental detail can go directly from Part 1 to Part 4. For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region, which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level, and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien-Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment. Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.

Boundary Layer Development in Axial Compressors and Turbines: Part 2 of 4—Compressors

This is Part Two of a four-part paper. It begins with Section 6.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this part, we present the experimental evidence used to construct the composite picture for compressors given in the discussion in Section 5.0 of Part 1. We show the data from the surface hot-film gages and the boundary layer surveys, give a thorough interpretation for the baseline operating condition, and then show how this picture changes with variations in Reynolds number, airfoil loading, frequency of occurrence of wakes and wake turbulence intensity. Detailed flow features are described using raw time traces. The use of rods to simulate airfoil wakes is also evaluated.

Boundary layer development in axial compressors and turbines: Part 4 of 4 : Computations and analyses

This is Part Four of a four-part paper. It begins with Section 16.0 and concludes the description of the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery. In this paper, the computational predictions made using several modern boundary layer codes are presented. Both steady codes and an unsteady code were evaluated. The results are compared with time-averaged and unsteady integral parameters measured for the boundary layers. Assessments are made to provide guidance in using the predictive codes to locate transition and predict loss. Conclusions from the computational analyses are then presented.

A Stochastic Model for a Compressor Stability Measure

A stability measure rooted in the unsteady characteristics of the flow field over the compressor rotor has been previously developed. The present work explores the relationship between the stochastic properties of this measure, called the correlation measure, and the compressor stability boundary. A stochastic model has been developed to gauge the impact of the correlation measures stochastic nature on its applicability to compressor stability management. The genesis of this model is in the fundamental properties of a specific stochastic process, one that is created by the threshold crossings of a random process. The model validation utilizes data obtained on three different axial compressor facilities. These include a single-stage low-speed axial compressor, a four-stage low-speed research compressor, and an advanced technology demonstrator high-speed compressor. This paper presents details of the model development and validation, as well as closed loop experimental results to demonstrate correlation measures usefulness in coinpressor stability management.

Boundary Layer Development in Axial Compressors and Turbines: Part 1 of 4 — Composite Picture

Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading.Part 1 of this paper draws a composite picture of boundary layer development in turbomachinery based upon a synthesis of all of our experimental findings for the compressor and turbine. Parts 2 and 3 present the experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data.For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien-Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment.Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.Copyright

Boundary Layer Development in Axial Compressors and Turbines: Part 3 of 4 — LP Turbines

This is Part Three of a four-part paper. It begins with Section 11.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery.In this part, we present the experimental evidence that we used to construct the composite picture for LP turbines that was given in the discussion in Section 5.0 of Part 1. We present and interpret the data from the surface hot-film gauges and the boundary layer surveys for the baseline operating condition. We then show how this picture changes with variations in Reynolds number, airfoil loading and nozzle-nozzle clocking.Copyright

Boundary Layer Development in Axial Compressors and Turbines: Part 2 of 4 — Compressors

This is Part Two of a four-part paper. It begins with Section 6.0 and continues to describe the comprehensive experiments and computational analyses that have led to a detailed picture of boundary layer development on airfoil surfaces in multistage turbomachinery.In this part, we present the experimental evidence used to construct the composite picture for compressors given in the discussion in Section 5.0 of Part 1. We show the data from the surface hot-film gauges and the boundary layer surveys, give a thorough interpretation for the baseline operating condition and then show how this picture changes with variations in Reynolds number, airfoil loading, frequency of occurrence of wakes and wake turbulence intensity. Detailed flow features are described using raw time traces. The use of rods to simulate airfoil wakes is also evaluated.© 1995 ASME

Transonic Fan Tip-Flow Features Revealed by High Frequency Response Over-Tip Pressure Measurements

Data from an array of high frequency response pressure transducers embedded in the casing wall over the tip of a transonic fan are reported. Phase-lock averaging of data from this array was successful in resolving an axial-tangential map of the static pressure rise in the rotor tip, as has been reported by other workers. Phase-lock ensemble RMS processing of the data is shown to be a useful technique that provides insight into the flow physics around the blade tip. Comparison with CFD results allows for more definite identification of features observed in the data. A complex flow field involving the casing wall boundary layer, the blade shock system and the over-tip leakage flow is observed. Differences between CFD data and measurements are explored by way of computational sensitivity studies. Results are reported for a range of throttle settings and speeds.Copyright

Design and Test Results of a Ultra High Loaded Single Stage High Pressure Turbine

This paper describes the engine and rig design, and test results of an ultra-highly loaded single stage high pressure turbine. In service aviation single stage HPTs typically operate at a total-to-total pressure ratio of less than 4.0. At higher pressure ratios or energy extraction the nozzle and blade both have regions of supersonic flow and shock structures which, if not mitigated, can result in a large loss in efficiency both in the turbine itself and due to interaction with the downstream component which may be a turbine center frame or a low pressure turbine. Extending the viability of the single stage HPT to higher pressure ratios is attractive as it enables a compact engine with less weight, and lower initial and maintenance costs as compared to a two stage HPT.The present work was performed as part of the NASA UEET (Ultra-Efficient Engine Technology) program from 2002 through 2005. The goal of the program was to design and rig test a cooled single stage HPT with a pressure ratio of 5.5 with an efficiency at least two points higher than the state of the art. Preliminary design tools and a design of experiments approach were used to design the flow path. Stage loading and through-flow were set at appropriate levels based on prior experience on high pressure ratio single stage turbines. Appropriate choices of blade aspect ratio, count, and reaction were made based on comparison with similar HPT designs. A low shock blading design approach was used to minimize the shock strength in the blade during design iterations. CFD calculations were made to assess performance.The HPT aerodynamics and cooling design was replicated and tested in a high speed rig at design point and off-design conditions. The turbine met or exceeded the expected performance level based on both steady state and radial/circumferential traverse data. High frequency dynamic total pressure measurements were made to understand the presence of unsteadiness that persists in the exhaust of a transonic turbine.© 2013 ASME