Research Papers

J. Turbomach. 2019;141(5):051001-051001-9. doi:10.1115/1.4041819.

A passive shock wave control method, using a grooved surface instead of the original smooth surface of a gas turbine nozzle vane to alter a single shock wave into a multiple shock wave structure, is investigated in this paper, so as to gain insight into the flow characteristics of a multiple shock wave system and its variations with various grooved surface geometry parameters. With the combination of numerical and experimental approaches, the shock wave structure and the flow behavior in a linear turbine nozzle channel with different grooved surface configurations were compared and analyzed in details. The numerical and experimental results indicate that the multiple shock wave structure induced by the grooved surface is beneficial for mitigating the intensity of the shock wave, reducing the potential excitation force of the shock wave and decreasing the shock wave loss as well. It was also found that the benefits are related to the geometry of the grooved surface, such as groove width, depth, and number. However, the presence of the grooved surface inevitably causes more viscous boundary layer loss and wake loss, which maybe a bottleneck for general engineering application of such a passive shock wave mitigation method.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051002-051002-11. doi:10.1115/1.4041907.

This paper deals with the numerical and theoretical investigations of the effect of geometrical dimensions and one-dimensional (1D)-design parameters on the impeller pressure slope of a transonic centrifugal compressor stage for industrial process application. A database being generated during the multi-objective and multipoint design process of a high flow coefficient impeller, comprising 545 computational fluid dynamics (CFD) designs is investigated in off-design and design conditions by means of Reynolds-averaged Navier–Stokes (RANS) simulation of an impeller with vaneless diffuser. For high flow coefficients of 0.16 < ϕdes < 0.18, the CFD-setup has been validated against measurement data regarding stage and impeller performance taken from MAN test rig experimental data for a centrifugal compressor stage of similar flow coefficient. This paper aims at answering the question how classical design parameter, such as the impeller blade angle distribution, impeller suction diameter, and camber line length affect the local and total relative diffusion and pressure slope toward impeller stall operation. A second-order analysis of the CFD database is performed by cross-correlating the CFD data with results from impeller two-zone 1D modeling and a rapid loading calculation process by Stanitz and Prian. The statistical covariance of first-order 1D-analysis parameters such as the mixing loss of the impeller secondary flow, the slip factor, impeller flow incidence is analyzed, thereby showing strong correlation with the design and off-design point efficiency and pressure slope. Finally, guide lines are derived in order to achieve either optimized design point efficiency or maximum negative pressure slope characteristics toward impeller stall operation.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051003-051003-9. doi:10.1115/1.4041818.

A rig, simulating two adjacent cooling cavities on the trailing side of an airfoil, made up of two trapezoidal channels is tested. Eleven crossover holes on the partition wall between the two channels create the jets. Two exit flow arrangements are investigated—(a) jets, after interaction with the target surface, are turned toward the target channel exit axially and (b) jets exit from a row of racetrack-shaped slots along the target channel. Flow measurements are reported for individual holes and heat transfer coefficients on the eleven target walls downstream the jets are measured using liquid crystals under steady-state conditions. Smooth as well as ribbed target surfaces with four rib angles are tested. Correlations are developed for mass flow rate through each crossover hole, varying the number of crossover holes. Heat transfer coefficient variations along the target channel are reported for a range of 5000–50,000 local jet Reynolds numbers. Major conclusions are: (1) Correlations are developed to successfully predict the air flow rate through each crossover hole for partition walls with six to eleven crossover holes, based on the pressure drop across the holes, (2) impingement heat transfer coefficient correlates well with local jet Reynolds number for both exit flow arrangements, and (3) case of target channel flow exiting from the channel end, at higher jet Reynolds numbers, produce higher heat transfer coefficients than those in the case of flow exiting through a row of slots along the target channel opposite to the crossover holes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051004-051004-9. doi:10.1115/1.4042208.

In low-pressure turbines (LPT) at design point, around 60–70% of losses are generated in the blade boundary layers far from end walls, while the remaining 30–40% is controlled by the interaction of the blade profile with the end-wall boundary layer. Increasing attention is devoted to these flow regions in industrial design processes. This paper discusses the end-wall flow characteristics of the T106 profile with parallel end walls at realistic LPT conditions, as described in the experimental setup of Duden, A., and Fottner, L., 1997, “Influence of Taper, Reynolds Number and Mach Number on the Secondary Flow Field of a Highly Loaded Turbine Cascade,” Proc. Inst. Mech. Eng., Part A, 211(4), pp.309–320. Calculations are carried out by both Reynolds-averaged Navier–Stokes (RANS), due to its continuing role as the design verification workhorse, and highly resolved large eddy simulation (LES). Part II of this paper focuses on the loss generation associated with the secondary end-wall vortices. Entropy generation and the consequent stagnation pressure losses are analyzed following the aerodynamic investigation carried out in the companion paper (GT2018-76233). The ability of classical turbulence models generally used in RANS to discern the loss contributions of the different vortical structures is discussed in detail and the attainable degree of accuracy is scrutinized with the help of LES and the available test data. The purpose is to identify the flow features that require further modeling efforts in order to improve RANS/unsteady RANS (URANS) approaches and make them able to support the design of the next generation of LPTs.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051005-051005-9. doi:10.1115/1.4041923.

In the present paper, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through four rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle in the range ±20 deg. The aero-thermal characterization of vane platform was obtained through five-hole probe and end wall adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. A relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of −20 deg.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051006-051006-9. doi:10.1115/1.4041909.

The role of gas-to-wall temperature ratio on bypass transition along a highly loaded turbine guide vane is investigated through time-resolved heat flux measurements for different flow conditions. The tests are conducted in the von Karman Institute (VKI) isentropic compression tube (CT-2) facility. High response thin films mounted on the vane (VKI LS89 airfoil) coupled with analog circuits are used for the heat flux measurements. The first detectable wall heat flux fluctuations denote the onset of transition which is also evaluated by both the heat transfer coefficient and the intermittency factor distributions along the suction side. The exit Mach number is kept constant during each test by means of a downstream sonic throat while the gas-to-wall temperature ratio is varied from 1.14 to 1.51 by changing the inlet gas temperature. Four test cases are considered for different exit Mach numbers (0.52 and 0.86) and freestream turbulence intensity levels (0.8% and 5.3%) while the isentropic exit chord Reynolds number is maintained at 106. In the present test campaign, the length and the evolution of the phenomenon indicate a measurable effect of the gas-to-wall temperature ratio for test cases with mild pressure gradients.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051007-051007-10. doi:10.1115/1.4042423.

As engine development continues to advance toward increased efficiency and reduced fuel consumption, efficient use of compressor bypass cooling flow becomes increasingly important. In particular, optimal use of compressor bypass flow yields an overall reduction of harmful emissions. Cooling flows used for cavity sealing between stages are critical to the engine and must be maintained to prevent damaging ingestion from the hot gas path. To assess cavity seals, the present study utilizes a one-stage turbine with true-scale engine hardware operated at engine-representative rotational Reynolds number and Mach number. Past experiments have made use of part-span (PS) rather than full-span (FS) blades to reduce flow rate requirements for the test rig; however, such decisions raise questions about potential influences of the blade span on sealing effectiveness measurements in the rim cavity. For this study, a tracer gas facilitates sealing effectiveness measurements in the rim cavity to compare data collected with FS engine airfoils and simplified, PS airfoils. The results from this study show sealing effectiveness does not scale as a function of relative purge flow with respect to main gas path flow rate when airfoil span is changed. However, scaling the sealing effectiveness for differing spans can be achieved if the fully purged flow rate is known. Results also suggest reductions of purge flow may have a relatively small loss of seal performance if the design is already near a fully purged condition. Rotor tip clearance is shown to have no effect on measured sealing effectiveness.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051008-051008-10. doi:10.1115/1.4041810.

The film cooling effectiveness distribution and its uniformity downstream of a row of film cooling holes on a flat plate are investigated by pressure sensitive paint (PSP) under different density ratios. Several hole geometries are studied, including streamwise cylindrical holes, compound-angled cylindrical holes, streamwise fan-shape holes, compound-angled fan-shape holes, and double-jet film-cooling (DJFC) holes. All of them have an inclination angle (θ) of 35 deg. The compound angle (β) is 45 deg. The fan-shape holes have a 10 deg expansion in the spanwise direction. For a fair comparison, the pitch is kept as 4d for the cylindrical and the fan-shape holes, and 8d for the DJFC holes. The uniformity of effectiveness distribution is described by a new parameter (Lateral-Uniformity, LU) defined in this paper. The effects of density ratios (DR = 1.0, 1.5 and 2.5) on the film-cooling effectiveness and its uniformity are focused. Differences among geometries and effects of blowing ratios (M = 0.5, 1.0, 1.5, and 2.0) are also considered. The results show that at higher density ratios, the lateral spread of the discrete-hole geometries (i.e., the cylindrical and the fan-shape holes) is enhanced, while the DJFC holes is more advantageous in film-cooling effectiveness. Mostly, a higher lateral-uniformity is obtained at DR = 2.5 due to better coolant coverage and enhanced lateral spread, but the effects of the density ratio on the lateral-uniformity are not monotonic in some cases. Utilizing the compound angle configuration leads to an increased lateral-uniformity due to a stronger spanwise motion of the jet. Generally, with a higher blowing ratio, the lateral-uniformity of the discrete-hole geometries decreases due to narrower traces, while that of the DJFC holes increases due to a stronger spanwise movement.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051009-051009-12. doi:10.1115/1.4041817.

A test facility for aeroeolastic investigations has been set up at the chair of Aero Engines at the Technische Universität Berlin. The test rig provides data for tool and code validation, and is used for basic aeroelastic experiments. It is a low-speed wind tunnel, which allows free and controlled flutter testing. The test section contains a linear cascade with eleven compressor blades. Nine of them are elastically suspended. The paper presents a detailed description of the test facility results to evaluate the overall flow quality alongside an aeroelastic model to predict the flutter velocity and critical interblade phase angles (IBPAs). Furthermore, chordwise pressure distributions, measured with traveling wave (TW) mode experimental tests, are presented. These measurements have been carried out for a wide range of IBPAs and have been compared to numerical results. Hot-wire anemometry has been applied to examine the inlet flow for several Mach numbers and Reynolds numbers. The results show small turbulence intensities. The blade surface pressure distribution and the flow field of the blade's suction and pressure sides have been obtained by oil flow visualization.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051010-051010-9. doi:10.1115/1.4041807.

The present study deals with the application of the transient thermochromic liquid crystal (TLC) technique in a flow network of intersecting circular passages as a potential internal turbine component cooling geometry. The investigated network consists of six circular passages with a diameter d = 20 mm that intersect coplanar at an angle θ = 40 deg, the innermost in three, the outermost in one intersection level. Two additional nonintersecting passages serve as references. Such a flow network entails specific characteristics associated with the transient TLC method that have to be accounted for in the evaluation process: the strongly curved surfaces, the mixing and mass flow redistribution at each intersection point, and the resulting gradients between the wall and passage centerline temperatures. All this impedes the choice of a representative fluid reference temperature, which results in deviations using established evaluation methods. An alternative evaluation approach is introduced, which is supported by computational results obtained from steady-state three-dimensional (3D) Reynolds-averaged Navier–Stokes equations (RANS) simulations using the shear-stress transport (SST) turbulence model. The presented analysis uncouples local heat transfer (HT) coefficients from actually measured local temperatures but uses the time information of the thermocouples (TC) instead that represents the fluid temperature step change and evolution along the passages. This experimental time information is transferred to the steady-state numerical bulk temperatures, which are finally used as local references to evaluate the transient TLC experiments. As effective local mass flow rates in the passage sections are considered, the approach eventually allows for a conclusion whether HT is locally enhanced due to higher mass flow rates or the intersection effects.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051011-051011-9. doi:10.1115/1.4041808.

In aerodynamic design, accurate and robust surrogate models are important to accelerate computationally expensive computational fluid dynamics (CFD)-based optimization. In this paper, a machine learning framework is presented to speed-up the design optimization of a highly loaded transonic compressor rotor. The approach is threefold: (1) dynamic selection and self-tuning among several surrogate models; (2) classification to anticipate failure of the performance evaluation; and (3) adaptive selection of new candidates to perform CFD evaluation for updating the surrogate, which facilitates design space exploration and reduces surrogate uncertainty. The framework is demonstrated with a multipoint optimization of the transonic NASA rotor 37, yielding increased compressor efficiency in less than 48 h on 100 central processing unit cores. The optimized rotor geometry features precompression that relocates and attenuates the shock, without the stability penalty or undesired reacceleration usually observed in the literature.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051012-051012-9. doi:10.1115/1.4041820.

A nonlinear full-wheel time-domain simulation of a two-stage low pressure turbine is presented, analyzed, and compared with the available experimental data. Recent improvements to the computational fluid dynamics (CFD) solver TRACE that lead to significantly reduced wall-clock times for such large scale simulations are described in brief. Since the configuration is characterized by significant unsteady turbulence and transition effects, it is well suited for the validation and benchmarking of frequency-domain methods. Transition, flow separation and wall pressure fluctuations on the stator blades of the second stage are analyzed in detail. A strong azimuthal π-periodicity is observed, manifesting in a significantly varying stability of the midspan trailing edge flow with a quasi-steady closed separation bubble on certain blades and highly dynamic partially open separation bubbles with recurring transition and turbulent reattachment on other blades. The energy spectrum of fluctuating wall quantities in that regime shows a high bandwidth and considerable disharmonic content, which is challenging for frequency-domain-based simulation methods.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(5):051013-051013-9. doi:10.1115/1.4041935.

The measurement accuracy of the temperature/pressure probe mounted at the leading edge of a turbine/compressor blade is crucial for estimating the fuel consumption of a turbo-fan engine. Apart from the measurement error itself, the probe also introduces extra losses. This again would compromise the measurement accuracy of the overall engine efficiency. This paper utilizes high-fidelity numerical analysis to understand the complex flow around the probe and quantify the loss sources due to the interaction between the blade and its instrumentation. With the inclusion of leading-edge probes, three-dimensional flow phenomena develop, with some flow features acting in a similar manner to a jet in cross flow. The separated flow formed at the leading edge of the probe blocks a large area of the probe bleedhole, which is one of the reasons why the probe accuracy can be sensitive to Mach and Reynolds numbers. The addition of 4% free-stream turbulence is shown to have a marginal impact on the jet trajectory originated from the probe bleedhole. However, a slight reduction is observed in the size of the separation bubble formed at the leading edge of the probe, preceding the two bleedhole exits. The free-stream turbulence also has a significant impact on the size of the separation bubble near the trailing edge of the blade. With the addition of the free-stream turbulence, the loss observed within the trailing edge wake is reduced. More than 50% of the losses at the cascade exit are generated by the leading-edge probe. A breakdown of the dissipation terms from the mean flow kinetic energy equation demonstrates that the Reynolds stresses are the key terms in dissipating the counter-rotating vortex pairs with the viscous stresses responsible for the boundary layer.

Commentary by Dr. Valentin Fuster

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