Research Papers

J. Turbomach. 2019;141(6):061001-061001-10. doi:10.1115/1.4041900.

A further investigation of an acoustic theory-based stall-warning approach is presented in this paper, which contains the basis of this approach and an application on a low-speed compressor (LSC) with a stabilization system. In the present work, this stall-warning approach is first explained through a numerical simulation in which the periodicity of pressure signals is analyzed, and then application experiments of this approach are actualized on a LSC with a stall precursor-suppressed (SPS) casing treatment (CT) as a stabilization system. For this stall-warning approach, a parameter named Rc is calculated through pressure signals of compressor to evaluate the periodicity of pressure signal, and statistical estimates are implemented on Rc so that the probabilities for Rc less than a threshold Rcth can be used as a criterion for stall warning. The numerical and experimental results both show that the signal resolution is determined by the sensor position, the prestall signal amplitude decreases rapidly with the increase of the distance between sensor and blades. Results also show that the probability increases significantly when the operating point is nearing the stall boundary. And at the same operating point, the probability value of Rc will decrease when the SPS CT is engaged. Through this stall-warning approach, a stabilization system based on SPS CT can be activated when the stall margin needs to be extended.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061002-061002-17. doi:10.1115/1.4042214.

This study addresses flutter that can occur in compressors when operating at high relative incidence. Studies are performed on a subsonic annular compressor cascade containing a sector of blades that can be subjected to controlled torsional oscillation. Measurements taken on the centrally located blade are used to study the unsteady surface pressures developed. Three large mean incidences are considered to characterize the aeroelastic performance. Aerodynamic damping is calculated for each test condition and its variation due to interblade phase angle (IBPA), reduced frequency, and incidence is studied. The source of stability or instability is traced to the nature of unsteady pressures. When the incidence is below the static-stall limit, an increasing incidence is found to exhibit aeroelastically more stable performance, whereas beyond the limit, the stability worsens. For the latter, the amount of improvement in stability by increasing reduced frequency is less compared to the former and its variation with IBPA is not as regular owing to the associated large uncertainty. The nonlinearity effects were found to be relatively higher for this case, especially from the aft region of the suction surface. It is also found that the phase of the local fluctuating pressure and its location on the chord has a decisive influence on the aerodynamic damping and its trends with respect to various parameters are discussed. The results are expected to be useful in the assessing aerodynamic damping trends in relation to the pressure phase variations in specific regions along the chord.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061003-061003-10. doi:10.1115/1.4041908.

Flow structures in lattice cooling channels are investigated experimentally by measuring three-dimensional (3D) velocity components over entire duct. The lattice cooling structure is formed by crossing two sets of parallel inclined ribs. Heat transfer is enhanced when coolant flows through the narrow subchannels between the ribs. According to the past literature, longitudinal vortex structures are formed inside the subchannels due to interactions between crossing flows. In this study, 3D velocity field measurement is performed using magnetic resonance imaging (MRI) scanner to clarify the flow mechanism. The rib inclination angle is varied from 30 to 60 deg. Reynolds number is set at approximately 8000 based on the whole duct inlet hydraulic diameter and bulk velocity. Working fluid is 0.015 mol/L copper sulfate aqueous solution. Measured results show that coolants in the upper and lower subchannels interact not only at the both ends of the duct, but also at diamond-shaped openings formed by opposite subchannels. The exchange of momentum between the upper and lower subchannels occurs at the openings, leading to sustained longitudinal vortex in each subchannel as mentioned in the literature. When the ribs are arranged with obtuse angle, a large vortex spreads across the contact surface, while the vortex structure independently stays in each subchannel for acute rib angle. The measured velocity fields are compared with numerically-simulated ones using a Reynolds-averaged Navier-Stokes (RANS) solver. Overall flow pattern is captured, but flow interaction between the upper and lower subchannels is underestimated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061004-061004-10. doi:10.1115/1.4041821.

Present paper developed a series of additively manufactured parallel cooling channels with streamwise wall jets inside, which could be suitable for double-wall and near-wall cooling configurations in gas turbine hot section components. The tested coupons consisted of parallel channels, each channel further divided into small chambers using several spanwise separation walls. Height of these walls was kept less than channel height, thus forming a slot with one of the end walls. Coolant entered from one side of channel and formed streamwise wall jet while crossing through the slot over to the downstream chamber. The test coupons were additively manufactured by selective laser sintering (SLS) technique using Inconel 718 alloy. Steady-state heat transfer experiments with constant wall temperature boundary condition were performed to analyze effect of pitch between subsequent slots and blockage ratio (ratio of separation wall height to channel height) on heat transfer. The channel Reynolds number ranged from 1800 to 5000. Numerical simulations were performed using ANSYS cfx solver with SST k–ω turbulence model to obtain detailed understanding of existing flow field. Experimental results showed heat transfer enhancement of up to 6.5 times that of a smooth channel for the highest blockage ratio of 0.75. Numerical results revealed complex flow field which consisted of wall jets along with impingement, separation, and recirculation zones in each chamber. For all configurations, gain in heat transfer was accompanied with high pressure drops. However, coupled with the high heat transfer, this design could lead to potential coolant savings.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061005-061005-14. doi:10.1115/1.4041867.

The research presented in this paper strives to exploit the benefits of near-wall measurement capabilities using hotwire anemometry and flowfield measurement capabilities using particle image velocimetry (PIV) for analysis of the injection of a staggered array of film cooling jets into a turbulent cross-flow. It also serves to give insight into the turbulence generation, jet structure, and flow physics pertaining to film cooling for various flow conditions. Such information and analysis will be applied to both cylindrical and diffuser shaped holes, to further understand the impacts manifesting from hole geometry. Spatially resolved PIV measurements were taken at the array centerline of the holes and detailed temporally resolved hotwire velocity and turbulence measurements were taken at the trailing edge of each row of jets in the array centerline corresponding to the PIV measurement plane. Flowfields of jets emanating from eight staggered rows, of both cylindrical and diffuser shaped holes inclined at 20 deg to the main-flow, are studied over blowing ratios in the range of 0.3–1.5. To allow for deeper interpretation, companion local adiabatic film cooling effectiveness results will also be presented for the geometric test specimen from related in-house work. Results show “rising” shear layers for lower blowing ratios, inferring boundary layer growth for low blowing ratio cases. Detachment of film cooling jets is seen from a concavity shift in the urms line plots at the trailing edge of film cooling holes. Former rows of jets are observed to disrupt the approaching boundary layer and enhance the spreading and propagation of subsequent downstream jets. Behavior of the film boundary layer in the near-field region directly following the first row of injection, as compared to the near-field behavior after the final row of injection (recovery region), is also measured and discussed. The impact of the hole geometry on the resulting film boundary layer, as in this case of cylindrical verses diffuser shaped holes, is ascertained in the form of mean axial velocity, turbulence level (urms), and length scales profiles.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061006-061006-10. doi:10.1115/1.4041866.

Discrete hole film cooling is widely employed to protect turbine blades and vanes from hot combustion gases entering the high-pressure turbine stage. Accurate prediction of the heat transfer near film cooling holes is critical, and high-fidelity experimental data sets are needed for validation of new computational models. Relatively few studies have examined the effects of periodic main flow unsteadiness resulting from the interaction of turbine blades and vanes, with a particular lack of data for shaped hole configurations. Periodic unsteadiness was generated in the main flow over a laidback, fan-shaped cooling hole at a Strouhal number (St = fD/U) of 0.014 by an airfoil oscillating in pitch. Magnetic resonance imaging (MRI) with water as the working fluid was used to obtain full-field, phase-resolved velocity and scalar concentration data. Operating conditions consisted of a hole Reynolds number of 2900, channel Reynolds number of 25, 000, and blowing ratio of unity. Both mean and phase-resolved data are compared to the previous measurements for the same hole geometry with steady main flow. Under unsteady freestream conditions, the flow separation pattern inside the hole was observed to change from an asymmetric separation bubble to two symmetric bubbles. The periodic unsteadiness was characterized by alternating periods of slow main flow, which allowed the coolant to penetrate into the freestream along the centerplane, and fast, hole-impinging main flow, which deflected coolant toward the laidback wall and caused ejection of coolant from the hole away from the centerplane. Mean adiabatic surface effectiveness was reduced up to 23% inside the hole, while mean laterally averaged effectiveness outside the hole fell 28–36% over the entire measurement domain. A brief comparison to a round jet with and without unsteadiness is included; for the round jet, no disturbance was observed inside the hole, and some fluctuations directed coolant toward the wall, which increased mean film cooling effectiveness. The combined velocity and concentration data for both cases are suitable for quantitative validation of computational fluid dynamics predictions for film cooling flows with periodic freestream unsteadiness.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061007-061007-10. doi:10.1115/1.4041961.

Optimal turbine blade tip designs have the potential to enhance aerodynamic performance while reducing the thermal loads on one of the most vulnerable parts of the gas turbine. This paper describes a novel strategy to perform a multi-objective optimization of the tip geometry of a cooled turbine blade. The parameterization strategy generates arbitrary rim shapes around the coolant holes on the blade tip. The tip geometry performance is assessed using steady Reynolds-averaged Navier–Stokes simulations with the k–ω shear stress transport (SST) model for the turbulence closure. The fluid domain is discretized with hexahedral elements, and the entire optimization is performed using identical mesh characteristics in all simulations. This is done to ensure an adequate comparison among all investigated designs. Isothermal walls were imposed at engine-representative levels to compute the convective heat flux for each case. The optimization objectives were a reduction in heat load and an increase in turbine row efficiency. The multi-objective optimization is performed using a differential evolution strategy. Improvements were achieved in both the aerodynamic efficiency and heat load reduction, relative to a conventional squealer tip arrangement. Furthermore, this work demonstrates that the inclusion of over-tip coolant flows impacts the over-tip flow field, and that the rim–coolant interaction can be used to create a synergistic performance enhancement.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061008-061008-10. doi:10.1115/1.4042023.

Large eddy simulation (LES) is used to explore the boundary layer transition mechanisms in two rectilinear compressor cascades. To reduce numerical dissipation, a novel locally adaptive smoothing (LAS) scheme is added to an unstructured finite volume solver. The performance of a number of subgrid scale (SGS) models is explored. With the first cascade, numerical results at two different freestream turbulence intensities (Ti's), 3.25% and 10%, are compared. At both Ti's, time-averaged skin-friction and pressure coefficient distributions agree well with previous direct numerical simulations (DNS). At Ti = 3.25%, separation-induced transition occurs on the suction surface, while it is bypassed on the pressure surface. The pressure surface transition is dominated by modes originating from the convection of Tollmien–Schlichting waves by Klebanoff streaks. However, they do not resemble a classical bypass transition. Instead, they display characteristics of the “overlap” and “inner” transition modes observed in the previous DNS. At Ti = 10%, classical bypass transition occurs, with Klebanoff streaks incepting turbulent spots. With the second cascade, the influence of unsteady wakes on transition is examined. Wake-amplified Klebanoff streaks were found to instigate turbulent spots, which periodically shorten the suction surface separation bubble. The celerity line corresponding to 70% of the free-stream velocity, which is associated with the convection speed of the amplified Klebanoff streaks, was found to be important.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061009-061009-11. doi:10.1115/1.4042305.

Improvements in stage isentropic efficiency and reductions in total pressure loss are sought in a 1.5 stage axial turbine. This is representative of power generation equipment used in thermal power cycles, which delivers about 80% of the 20 × 1012 kWh world-wide electricity. Component-level improvements are therefore timely and important toward achieving carbon dioxide global emission targets. Secondary flow loss reduction is sought by applying a nonaxisymmetric endwall design to the turbine stator hub. A guide groove directs the pressure side branch of the horseshoe vortex away from the airfoil suction side, using a parametric endwall hub surface, which is defined as to obtain first-order smooth boundary connections to the remainder of the passage geometry. This delays the onset of the passage vortex and reduces its associated loss. The Automatic Process and Optimization Workbench (apow) generates a Kriging surrogate model from a set of Reynolds-averaged Navier–Stokes simulations, which is used to optimize the hub surface. The three-dimensional steady Reynolds-averaged Navier–Stokes model with an axisymmetric hub is validated against reference experimental measurements from the Rheinisch-Westfälische Technische Hochschule (RWTH) Aachen. Comparative computational fluid dynamics (CFD) predictions with an optimized nonaxisymmetric hub show a decrease in the total pressure loss coefficient and an increase in the isentropic stage efficiency at and off design conditions.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(6):061010-061010-10. doi:10.1115/1.4042250.

This paper describes the role of tip leakage flow in creating the leading edge separation necessary for the onset of spike-type compressor rotating stall. A series of unsteady multipassage simulations, supported by experimental data, are used to define and illustrate the two competing mechanisms that cause the high incidence responsible for this separation: blockage from a casing-suction-surface corner separation and forward spillage of the tip leakage jet. The axial momentum flux in the tip leakage flow determines which mechanism dominates. At zero tip clearance, corner separation blockage dominates. As clearance is increased, the leakage flow reduces blockage, moving the stall flow coefficient to lower flow, i.e., giving a larger unstalled flow range. Increased clearance, however, means increased leakage jet momentum and contribution to leakage jet spillage. There is thus a clearance above which jet spillage dominates in creating incidence, so the stall flow coefficient increases and flow range decreases with clearance. As a consequence, there is a clearance for maximum flow range; for the two rotors in this study, the value was approximately 0.5% chord. The chordwise distribution of the leakage axial momentum is also important in determining stall onset. Shifting the distribution toward the trailing edge increases flow range for a leakage jet dominated geometry and reduces flow range for a corner separation dominated geometry. Guidelines are developed for flow range enhancement through control of tip leakage flow axial momentum magnitude and distribution. An example is given of how this might be achieved.

Commentary by Dr. Valentin Fuster

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