Research Papers

J. Turbomach. 2017;139(5):051001-051001-11. doi:10.1115/1.4035103.

The flow field inside a triangular cooling channel for the leading edge of a gas turbine blade has been investigated. The efforts were focused on the investigation of the interaction between effects of rotation, of buoyancy forces, and those induced by turbulence promoters, i.e., perpendicular square ribs placed on both leading and trailing sides of the duct. Particle image velocimetry (PIV) and stereo-PIV measurements have been performed for ReDh = 104, rotation number of 0, 0.2, and 0.6, and buoyancy parameter equal to 0, 0.08, and 0.7. Coriolis secondary flows are detected in the duct cross section, but contrary to the smooth case, they are characterized by a single main vortex and are less affected by an increase of the rotation parameter. Moreover, their main topology is only marginally sensitive to the buoyancy forces. Conversely, the features of the recirculation structure downstream the ribs turned out to be more sensitive to both the buoyancy forces and to the stabilizing/destabilizing effect on the separated shear layer induced by rotation.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051002-051002-10. doi:10.1115/1.4035202.

Turbine blade components in an engine are typically designed with gaps between parts due to manufacturing, assembly, and operational considerations. Coolant is provided to these gaps to limit the ingestion of hot combustion gases. The interaction of the gaps, their leakage flows, and the complex vortical flow at the endwall of a turbine blade can significantly impact endwall heat transfer coefficients and the effectiveness of the leakage flow in providing localized cooling. In particular, a platform gap through the passage, representing the mating interface between adjacent blades in a wheel, has been shown to have a significant effect. Other important turbine blade features present in the engine environment are nonaxisymmetric contouring of the endwall, and an upstream rim seal with a gaspath cavity, which can reduce and increase endwall vortical flow, respectively. To understand the platform gap leakage effect in this environment, measurements of endwall heat transfer, and film cooling effectiveness were performed in a scaled blade cascade with a nonaxisymmetric contour in the passage. A rim seal with a cavity, representing the overlap interface between a stator and rotor, was included upstream of the blades and a nominal purge flowrate of 0.75% of the mainstream was supplied to the rim seal. The results indicated that the endwall heat transfer coefficients increased as the platform gap net leakage increased from 0% to 0.6% of the mainstream flowrate, but net heat flux to the endwall was reduced due to high cooling effectiveness of the leakage flow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051003-051003-10. doi:10.1115/1.4035072.

Airborne particles ingested in aircraft engines deposit on compressor blading and end walls. Aerodynamic surfaces degrade on a microscopic and macroscopic scale. Blade row, compressor, and engine performance deteriorate. Optimization of maintenance scheduling to mitigate these effects requires modeling of the deterioration process. This work provides a deterioration model on blade row level and the experimental validation of this model in a newly designed deposition test rig. When reviewing previously published work, a clear focus on deposition effects in industrial gas turbines becomes evident. The present work focuses on quantifying magnitudes and timescales of deposition effects in aircraft engines and the adaptation of the generalized Kern and Seaton deposition model for application in axial compressor blade rows. The test rig's cascade was designed to be representative of aircraft engine compressor blading. The cascade was exposed to an accelerated deposition process. Reproducible deposition patterns were identified. Results showed an asymptotic progression of blade row performance deterioration. A significant increase in total pressure loss and decrease in static pressure rise were measured. Application of the validated model using existing particle concentration and flight cycle data showed that more than 95% of the performance deterioration due to deposition occurs within the first 1000 flight cycles.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051004-051004-13. doi:10.1115/1.4035160.

Experiments to investigate the effect of target wall curvature on heat transfer and pressure loss from jet array impingement are performed. A jet plate configuration is studied with constant hole diameters and spacings. The geometry of the jet plate has streamwise jet spacings of 5.79 jet diameters, spanwise jet spacings of 4.49 jet diameters, and a jet-to-target plate distance of 3 jet diameters. For the curved case, the radius of the target plate is r/D = 31.57. A flat target wall setup with identical geometric spacing is also tested for direct comparison. Jet spacings were chosen such that validation and comparison can be made with open literature. For all configurations, spent air is drawn out in a single direction, which is tangential to the target plate curvature. Average jet Reynolds numbers ranging from 55,000 to 125,000 are tested. A steady-state measurement technique utilizing temperature-sensitive paint (TSP) is used on the target surface to obtain Nusselt number distributions. Pressure taps placed on the sidewall of the channel are used to evaluate the flow distribution in the impingement channel. Alongside the experimental work, CFD was performed utilizing the v2 − f turbulence model to better understand the relationship between the flow field and the heat transfer on the target surface. The main target of the current study is to quantify the impact of target wall radius and the decay of heat transfer after the impingement section, and to check the open literature correlations. It was found that the target wall curvature did not cause any significant changes in either the flow distribution or the heat transfer level. Comparisons with established correlations show similar level but different trends in heat transfer, potentially caused by differences in L/D. CFD results were able to show agreement in streamwise pitch-averaged Nusselt number levels with experimental results for the curved target plate at higher Re numbers.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051005-051005-12. doi:10.1115/1.4035161.

The performance of a showerhead arrangement of film cooling in the leading edge region of a first-stage nozzle guide vane was experimentally and numerically evaluated. A six-vane linear cascade was tested at an isentropic exit Mach number of Ma2s = 0.42, with a high inlet turbulence intensity level of 9%. The showerhead cooling scheme consists of four staggered rows of cylindrical holes evenly distributed around the stagnation line, angled at 45 deg toward the tip. The blowing ratios tested are BR = 2.0, 3.0, and 4.0. Adiabatic film cooling effectiveness distributions on the vane surface around the leading edge region were measured by means of thermochromic liquid crystals (TLC) technique. Since the experimental contours of adiabatic effectiveness showed that there is no periodicity across the span, the computational fluid dynamics (CFD) calculations were conducted by simulating the whole vane. Within the Reynolds-averaged Navier–Stokes (RANS) framework, the very widely used realizable k–ε (Rke) and the shear stress transport k–ω (SST) turbulence models were chosen for simulating the effect of the BR on the surface distribution of adiabatic effectiveness. The turbulence model which provided the most accurate steady prediction, i.e., Rke, was selected for running detached eddy simulation (DES) at the intermediate value of BR = 3. Fluctuations of the local temperature were computed by DES, due to the vortex structures within the shear layers between the main flow and the coolant jets. Moreover, mixing was enhanced both in the wall-normal and spanwise directions, compared to RANS modeling. DES roughly halved the prediction error of laterally averaged film cooling effectiveness on the suction side of the leading edge. However, neither DES nor RANS provided the expected decay of effectiveness progressing downstream along the pressure side, with 15% overestimation of ηav at s/C = 0.2.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051006-051006-11. doi:10.1115/1.4035252.

Changes in loss generation associated with altering rotor tip blade loading of an embedded rotor–stator compressor stage are assessed with unsteady three-dimensional computations, complemented by control volume analyses. Tip-fore-loaded and tip-aft-loaded rotor blades are designed to provide variation in rotor tip blade loading distributions for determining a compressor design hypothesis that aft-loading a rotor blade tip yields a reduction in loss generation in a stage environment. Aft-loading a rotor blade tip delays the formation of tip leakage flow, resulting in a relatively less mixed-out tip leakage flow at the rotor outlet and a reduction in overall tip leakage mass flow, hence a lower loss generation. However, the attendant changes in tip flow angle distribution are such that there is an overall increase in the flow angle mismatch between tip flow and main flow, leading to higher loss generation. The latter outweighs the former; therefore, rotor passage loss from aft-loading a rotor tip is higher unless a constraint is imposed on tip flow angle distribution so that the associated induced loss is negligible. Tip leakage flow, which is not mixed-out at the rotor outlet, is recovered in the downstream stator. The tip leakage flow recovery process yields a higher benefit for a relatively less mixed-out tip leakage flow in the tip-aft-loaded rotor blades on a time-averaged basis. These characterizing parameters together determine the attendant overall loss associated with rotor tip leakage flow in a compressor stage environment. The revised design hypothesis is thus as follows: A rotor should be tip-aft-loaded and hub-fore-loaded while a stator should be hub-aft-loaded and tip-fore-loaded with tip/hub leakage flow angle distribution such that it results in no additional loss. For the compressor stage being assessed here, an estimated 0.15 points enhancement in stage efficiency is possible from aft-loading rotor tip only.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051007-051007-11. doi:10.1115/1.4035210.

The objective of this work is to study the influence of a pressure side separation bubble on the profile losses and the development of the bubble in the blade passage. For the experimental investigations, the T106 profile is used, with an increased loading due to an enlarged pitch to chord ratio from 0.799 to 0.95 (T106C). The experiments were performed at the high-speed cascade wind tunnel of the Institute of Jet Propulsion at the University of the Federal Armed Forces Munich. The main feature of the wind tunnel is to vary Reynolds and Mach number independently to achieve realistic turbomachinery conditions. The focus of this work is to determine the influence of a pressure side separation on the profile losses and hence the robustness to suction side incidence flow. The cascade is tested at four incidence angles from 0 deg to −22.7 deg to create separation bubbles of different sizes. The influence of the Reynolds number is investigated for a wide range at constant exit Mach number. Therefore, a typical exit Mach number for low pressure turbines in the range of 0.5–0.8 is chosen in order to consider compressible effects. Furthermore, two inlet turbulence levels of about 3% and 7.5% have been considered. The characteristics of the separation bubble are identified by using the profile pressure distributions, whereas wake traverses with a five hole probe are used to determine the influence of the pressure side separation on the profile losses. Further, time-resolved pressure measurements near the trailing edge as well as single hot wire measurements in the blade passage are conducted to investigate the unsteady behavior of the pressure side separation process itself and also its influence on the midspan passage flow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051008-051008-9. doi:10.1115/1.4035075.

Airfoil shapes tailored to specific inflow conditions and loading requirements can offer a significant performance potential over classic airfoil shapes. However, their optimal operating range has to be matched thoroughly to the overall compressor layout. This paper describes methods to organize a large set of optimized airfoils in a database and its application in the throughflow design. Optimized airfoils are structured in five dimensions: inlet Mach number, blade stagger angle, pitch–chord ratio, maximum thickness–chord ratio, and a parameter for aerodynamic loading. In this space, a high number of airfoil geometries are generated by means of numerical optimization. During the optimization of each airfoil, the performance at design and off-design conditions is evaluated with the blade-to-blade flow solver MISES. Together with the airfoil geometry, the database stores automatically calibrated correlations which describe the cascade performance in throughflow calculation. Based on these methods, two subsonic stages of a 4.5-stage transonic research compressor are redesigned. Performance of the baseline and updated geometry is evaluated with 3D CFD. The overall approach offers accurate throughflow design incorporating optimized airfoil shapes and a fast transition from throughflow to 3D CFD design.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051009-051009-13. doi:10.1115/1.4035277.

Experimental results are presented for a double wall cooling arrangement which simulates a portion of a combustor liner of a gas turbine engine. The results are collected using a new experimental facility designed to test full-coverage film cooling and impingement cooling effectiveness using either cross flow, impingement, or a combination of both to supply the film cooling flow. The present experiment primarily deals with cross flow supplied full-coverage film cooling for a sparse film cooling hole array that has not been previously tested. Data are provided for turbulent film cooling, contraction ratio of 1, blowing ratios ranging from 2.7 to 7.5, coolant Reynolds numbers based on film cooling hole diameter of about 5000–20,000, and mainstream temperature step during transient tests of 14 °C. The film cooling hole array consists of a film cooling hole diameter of 6.4 mm with nondimensional streamwise (X/de) and spanwise (Y/de) film cooling hole spacing of 15 and 4, respectively. The film cooling holes are streamwise inclined at an angle of 25 deg with respect to the test plate surface and have adjacent streamwise rows staggered with respect to each other. Data illustrating the effects of blowing ratio on adiabatic film cooling effectiveness and heat transfer coefficient are presented. For the arrangement and conditions considered, heat transfer coefficients generally increase with streamwise development and increase with increasing blowing ratio. The adiabatic film cooling effectiveness is determined from measurements of adiabatic wall temperature, coolant stagnation temperature, and mainstream recovery temperature. The adiabatic wall temperature and the adiabatic film cooling effectiveness generally decrease and increase, respectively, with streamwise position, and generally decrease and increase, respectively, as blowing ratio becomes larger.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(5):051010-051010-9. doi:10.1115/1.4035278.

The measurement of unsteady total temperature is of great interest for the examination of loss mechanisms in turbomachinery with respect to the improvement of the efficiency. Since conventional thermocouples are limited in frequency response, several fast-response total temperature probes have been developed over the past years. To improve the spatial resolution compared to these existing probes and maintaining a high temporal resolution, a new fast-response total temperature probe has been developed at the Institute of Aircraft Propulsion Systems (ILA), Stuttgart, Germany in cooperation with Berns Engineers, Gilching, Germany. The design of the probe allows a sensitive measuring surface below 1 mm2. A detailed insight into the design of the probe, the measurement principle, the calibration process, and an estimation of the measurement uncertainty is given in the present paper. Furthermore, to prove the functionality of the probe, first experimental results of a simple test bed and of area traverses downstream of the first rotor of a two-stage low pressure turbine are presented. It is shown, that the new probe is capable of detecting rotor characteristic effects as well as rotor-stator-interactions. In addition, a hot-spot is investigated downstream of the first rotor of the turbine, and the findings are compared to the effects known from the literature.

Commentary by Dr. Valentin Fuster

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