David C. Wisler

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.

Effects of rotating inlet distortion on multistage compressor stability

In multispool engines, rotating stall in an upstream compressor will impose a rotating distortion on the downstream compressor, thereby affecting its stability margin. In this paper experiments are described in which this effect was simulated by a rotating screen upstream of several multistage low-speed compressors. The measurements are complemented by, and compared with, a theoretical model of multistage compressor response to speed and direction of rotation of an inlet distortion. For corotating distortions (i.e., distortions rotating in the same direction as rotor rotation), experiments show that the compressors exhibited significant loss in stability margin and that they could be divided into two groups according to their response. The first group exhibited a single peak in stall margin degradation when the distortion speed corresponded to roughly 50 percent of rotor speed. The second group showed two peaks in stall margin degradation corresponding to distortion speeds of approximately 25-35 percent and 70-75 percent of rotor speed. These new results demonstrate that multistage compressors can have more than a single resonant response. Detailed measurements suggest that the two types of behavior are linked to differences between the stall inception processes observed for the two groups of compressors and that a direct connection thus exists between the observed forced response and the unsteady flow phenomena at stall onset. For counterrotational distortions, all the compressors tested showed minimal loss of stability margin. The results imply that counterrotation of the fan and core compressor, or LP and HP compressors, could be a worthwhile design choice. Calculations based on the two-dimensional theoretical model show excellent agreement for the compressors, which had a single peak for stall margin degradation. We take this first-of-a-kind comparison as showing that the model, though simplified, captures the essential fluid dynamic features of the phenomena. Agreement is not good for compressors that had two peaks in the curve of stall margin shift versus distortion rotation speed. The discrepancy is attributed to the three-dimensional and short length scale nature of the stall inception process in these machines; this includes phenomena that have not yet been addressed in any model.

Effects of nonaxisymmetric tip clearance on axial compressor performance and stability

The effects of circumferentially nonuniform tip clearance on axial compressor performance and stability have been investigated experimentally and analytically. A theoretical model for compressor behavior with nonaxisymmetric tip clearance has been developed and used to design a series of first-of-a-kind experiments on a four-stage, low-speed compressor. The experiments and computational results together show clearly the central physical features and controlling parameters of compressor response to nonaxisymmetric tip clearance. It was found that the loss in stall margin was more severe than that estimated based on average clearance. The stall point was, in fact, closer to that obtained with uniform clearance at the maximum clearance level. The circumferential length scale of the tip clearance (and accompanying flow asymmetry) was an important factor in determining the stall margin reduction. For the same average clearance, the loss in peak pressure rise was 50 percent higher for an asymmetry with fundamental wavelength equal to the compressor circumference than with wavelength equal to one-half the circumference. The clearance asymmetry had much less of an effect on peak efficiency; the measured maximum efficiency decrease obtained was less than 0.4 percent compared to the 8 percent decrease in peak pressure rise due to the asymmetric clearance. The efficiency penalty due to nonaxisymmetric tip clearance was thus close to that obtained with a uniform clearance at the circumferentially averaged level. The theoretical model accurately captured the decreases in both steady-state pressure rise and stable operating range which are associated with clearance asymmetry. It also gave a good description of the observed trends of: (i) increasing velocity asymmetry with decreasing compressor flow, and (ii) decreasing effect of clearance asymmetry with decreasing dominant wavelength of the clearance distribution. The time-resolved data showed that the spatial structure of the prestall propagating disturbances in the compressor annulus was well represented and that the stability limiting process could be linked to the unsteady structure of these disturbance modes. The model was also utilized for parametric studies to define how compressor performance and stability are affected by the circumferential distribution of clearance, steady-state compressor pressure-rise characteristic, and system dynamic parameters. Sensitivity to clearance asymmetry was found to fall off strongly with the (asymmetry-related) reduced frequency and to increase with peak pressure rise and increasing curvature of the characteristic near the peak.

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

Study of Boundary Layer Development in a Two-Stage Low-Pressure Turbine

Experimental data from jet-engine tests have indicated that unsteady blade row (wake) interactions and separation can have a significant impact on the efficiency of turbine stages. The effects of these interactions can be intensified in low-pressure turbine stages because of the low Reynolds number operating environment. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Thus, during the last decade a significant amount of effort has been put into determining the effects of transition and turbulence on the performance of low pressure turbine stages. Experimental investigations have been performed, for example, by Hodson et al. and Halstead et al. These investigations have helped identify/clarify the roles that factors such as the Reynolds number, free stream turbulence intensity, pressure gradient and curvature have in the generation of losses. In parallel to the experimental investigations, there have been significant analytical efforts to improve the modeling of transition. Examples of such efforts include the works of Mayle and Gostelow et al. These newer models show promise of providing accurate transition predictions over a wide range of flow conditions, although they have yet to be implemented into the numerical flow analyses used by the turbine design community. Some recent computational investigations of interest include the works of Chernobrovkin and Lakshminarayana and Eulitz and Engel. The focus of the current effort has been to -use a viscous, unsteady quasi-three-dimensional Navier-Stokes analysis to study boundary layer development in a two-stage low-pressure turbine. A two-layer algebraic turbulence model, along with a natural transition model and a bubble transition model, have been used, The geometry used in the simulations has been the subject of extensive experiments. The predicted results have been compared with experimental data, including airfoil loadings and time-averaged/unsteady integral boundary layer quantities.

ENGINEERING - WHAT YOU DON'T NECESSARILY LEARN IN SCHOOL

As young engineers progress in their careers, they begin to understand that there is far more to being an outstanding engineer than they might have thought during their days as an undergraduate. In fact, some of the things they need to know weren’t necessarily learned in school. And this is understandable, given the relatively short time spent in school and the significant differences between the missions of academe and industry/government. This paper focuses on twelve vital aspects in engineering that are usually learned after graduation but can make the difference between success and failure in one’s engineering career. To succeed, engineers must: learn to be business oriented; expect tough, multi-disciplinary problems; learn to work and network in the new multi-cultural and multi-national environment; understand the differences between academe and industry; learn to differentiate all over again; understand the values and culture of their particular company or organization; be open to ideas from everywhere; have unyielding integrity; make their manager a success; support their university and technical society; have fun; and most importantly, manage their careers. Each of these aspects is discussed in detail.

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