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Research Papers

J. Turbomach. 2019;141(8):081001-081001-24. doi:10.1115/1.4042652.

Mixed flow turbines (MFTs) offer potential benefits for turbocharged engines when considering off-design performance and engine transient behavior. Although the performance and use of MFTs are described in the literature, little is published on the combined impact of the cone angle and the inlet blade angle, which are the defining features of such turbines. Numerical simulations were completed using a computational fluid dynamics (CFD) model that was validated against experimental measurements for a baseline geometry. The mechanical impact of the design changes was also analyzed. Based on the results of the numerical study, two rotors of different blade angle and cone angle were selected and manufactured. These rotors were tested using the Queen's University Belfast (QUB) low-temperature turbine test rig, which allowed for accurate and wide-range mapping of the turbine performance to low values of the velocity ratio. The performance results from these additional rotors were used to further validate the numerical findings. The numerical model was used to understand the underlying physical reasons for the measured performance differences through detailed consideration of the flow field at the rotor inlet and to document how the loss mechanisms and secondary flow structures developed with varying rotor inlet geometry. It was observed that large inlet blade cone angles resulted in strong separation and flow blockage near the hub at off-design conditions, which greatly reduced efficiency. However, the significant rotor inertia benefits achieved with the large blade cone angles were shown to compensate for the efficiency penalties and could be expected to deliver improved transient performance in downsized automotive engine applications.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081002-081002-8. doi:10.1115/1.4042892.

The increasing demands on jet engines require progressive thermodynamic process parameters, which typically lead to higher aerothermal loadings and accordingly to designs with high complexity. State-of-the-art high-pressure turbine (HPT) nozzle guide vane (NGV) design involves vane profiles with three-dimensional features including a high amount of film cooling and profiled endwalls (PEWs). Typically, the specific mass flow, also called capacity, which governs the engine's operation, is set by the HPT NGV. Hence, geometric variations due to manufacturing scatter of the HPT NGV's passage can affect relevant aerodynamic quantities and the entire engine behavior. Within the traditional deterministic design approach, the influences of those geometric variations are covered by conservative assumptions and engineering experience. This paper addresses the consideration of variability due to the manufacturing of HPT NGVs through probabilistic CFD investigations. To establish a statistical database, 80 HPT NGVs are digitized with a high precision optical 3D scanning system to record the outer geometry. The vane profiles are parametrized by a section-based approach. For this purpose, traditional profile theory is combined with a novel method that enables the description of NGV profile variability taking the particular leading edge (LE) shape into account. Furthermore, the geometric variability of PEWs is incorporated by means of principle component analysis (PCA). On this basis, a probabilistic system assessment including a sensitivity analysis in terms of capacity and total pressure loss coefficient is realized. Sampling-based methods are applied to conduct a variety of 3D CFD simulations for a typical population of profile and endwall geometries. This probabilistic investigation using realistic input parameter distributions and correlations contributes to a robust NGV design in terms of relevant aerodynamic quantities.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081003-081003-11. doi:10.1115/1.4042891.

Shock control bumps can help to delay and weaken shocks, reducing loss generation and shock-induced separation and delaying stall inception for transonic turbomachinery components. The use of shock control bumps on turbomachinery blades is investigated here for the first time using 3D analysis. The aerodynamic optimization of a modern research fan blade and a highly loaded compressor blade is carried out using shock control bumps to improve their performance. Both the efficiency and stall margin of transonic fan and compressor blades may be increased through the addition of shock control bumps to the geometry. It is shown how shock-induced separation can be delayed and reduced for both cases. A significant efficiency improvement is shown for the compressor blade across its characteristic, and the stall margin of the fan blade is increased by designing bumps that reduce shock-induced separation near to stall. Adjoint surface sensitivities are used to highlight the critical regions of the blade geometries, and it is shown how adding bumps in these regions improves blade performance. Finally, the performance of the optimized geometries at conditions away from where they are designed is analyzed in detail.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081004-081004-13. doi:10.1115/1.4042728.

The present work combines experimental measurements and unsteady, three-dimensional computational fluid dynamics predictions to gain further insight into the complex flow-field within an automotive turbocharger centrifugal compressor. Flow separation from the suction surface of the main impeller blades first occurs in the mid-flow range, resulting in local flow reversal near the periphery, with the severity increasing with decreasing flow rate. This flow reversal improves leading-edge incidence over the remainder of the annulus, due to (a) reduction of cross-sectional area of forward flow, which increases the axial velocity, and (b) prewhirl in the direction of impeller rotation, as a portion of the tangential velocity of the reversed flow is maintained when it mixes with the core flow and transitions to the forward direction. As the compressor operating point enters the region where the slope of the constant speed compressor characteristic (pressure ratio versus mass flow rate) becomes positive, rotating stall cells appear near the shroud side diffuser wall. The angular propagation speed of the diffuser rotating stall cells is approximately 20% of the shaft speed, generating pressure fluctuations near 20% and 50% of the shaft frequency, which were also experimentally observed. For the present compressor and rotational speed, flow losses associated with diffuser rotating stall are likely the key contributor to increasing the slope of the constant speed compressor performance curve to a positive value, promoting the conditions required for surge instabilities. The present mild surge predictions agree well with the measurements, reproducing the amplitude and period of compressor outlet pressure fluctuations.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081005-081005-12. doi:10.1115/1.4042990.

Secondary flow limits the aerodynamic loading level of turbomachinery. Vortex generators (VGs) offer the potential to attenuate secondary flow when implemented at the endwall of the blade passage. Customary design usually relies on computational fluid dynamics (CFD); however, VG geometry modeling and mesh generation are challenging. This paper presents an efficient method for designing the optimal VG layout. In this approach, first, a mathematical model (BAYC) is introduced to replace the actual VGs; hence, simulation can be carried out without detailed VG gridding. Second, an optimization procedure with response surface methods is employed to determine the optimal VG layout. To illustrate the proposed method, compressor cascades with one and three VGs are used as the test cases. The results demonstrate that the optimal VG layout may effectively weaken the secondary flow and can decrease the aerodynamic loss by 15–25% in almost all incidence angle ranges, particularly at positive incidence angles. Flow mechanism analysis indicates that VGs can enhance the boundary layer kinetic energy, thereby elevating the capability to withstand adverse pressure gradients.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081006-081006-11. doi:10.1115/1.4043071.

Recent advances in gas turbine film cooling technology such as round film cooling holes embedded in craters or trenches, and shaped film cooling holes are of interest due to a marked improvement in the effectiveness of film cooling jets. Typically, shaped film cooling holes have higher manufacturing cost, while film cooling holes embedded in craters/trenches etched in thermal barrier coatings (TBC) are seen as a cost-effective alternative. In a recent numerical study Kalghatgi and Acharya (2015, “Improved Film Cooling Effectiveness With a Round Film Cooling Hole Embedded in Contoured Crater,” ASME J. Turbomach., 137(10), p. 101006) reported a novel crater shape to generate anti-counter rotating vortex pair (CRVP) beneath the film cooling jet and showed a significant improvement in film cooling performance. In the present paper, a comprehensive study of flow dynamics is presented to gain insight into the unsteady flow physics of film cooling jet issued from a round hole embedded in the contoured crater. As a baseline case, a round film cooling hole with a 35 deg inclined short delivery tube (l/D = 1.75) is used as from a previous study with freestream Reynolds number based on jet diameter set to ReD = 16,000 and density ratio of coolant to freestream fluid of ρjo = 2.0. These flow conditions are used for the cases of film cooling jet embedded in contoured crater. The results are presented for two crater depths: (i) shallow crater with 0.2D depth and (ii) deep crater with 0.75D depth. First- and second-order flow statistics are presented for all the cases, including the experimental data for baseline case from the literature. Time-averaged and instantaneous flow structures are visualized to reveal the mechanisms of anti-CRVP and attenuating CRVP. The dynamics of flow structures studied using single-point spectral analysis in the shear layer and modal analysis of three-dimensional flow field shows a loss of coherency and increased time scales of shear layer structures as the crater depth is increased, primarily due to attenuating of CRVP in the downstream vicinity of the crater. The modal analysis confirmed reduced magnitude of temperature fluctuations (hot spots) on the cooling wall compared with baseline round film cooling hole. Finally, a 2–5% additional pressure loss due to the crater is reported over the existing ≈7% loss in pressure for baseline case.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081007-081007-11. doi:10.1115/1.4043120.

Gas turbines engine designers are leaning toward aircraft engine architectures due to their footprint, weight, and performance advantages. Such engines need some modifications to both the combustion system, to comply with emission limits, and turbine rotational speed. Aeroderivative engines maintain the same legacy aircraft engine architecture and replace the fan and booster with a higher speed compressor booster driven by a single-stage intermediate turbine. A multistage free power turbine (FPT) sits on a separate shaft to drive compressors for liquefied natural gas (LNG) applications or generators. The intermediate-power turbine (IPT) design is important for the engine performance as it drives the booster compressor and sets the inlet boundary conditions to the downstream power turbine. This paper describes the experience of Baker Hughes, a GE company (BHGE) in the design of the intermediate turbine that sits in between a GE legacy aircraft engine core exhaust and the downstream power turbine. This paper focuses on the flow path of the turbine center frame (TCF)/intermediate turbine and the associated design, as well as on the 3D steady and unsteady computational fluid dynamics (CFD)-assisted design of the IPT stage to control secondary flows in presence of through flow curvature induced by the upstream TCF.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081008-081008-9. doi:10.1115/1.4043139.

Variable pitch fans are of interest for future low-pressure ratio fan systems since they provide improved operability relative to fixed pitch fans. If they can also be re-pitched such that they generate sufficient reverse thrust they could eliminate the engine drag and weight penalty associated with bypass duct thrust reversers. This paper sets out to understand the details of the 3D fan stage flow field in reverse thrust operation. This study uses the Advanced Ducted Propulsor variable pitch fan test case, which has a design fan pressure ratio of 1.29. Comparison with spanwise probe measurements show that the computational approach is valid for examining the variation of loss and work in the rotor in forward thrust. The method is then extended to a reverse thrust configuration using an extended domain and appropriate boundary conditions. Computations, run at two rotor stagger settings, show that the spanwise variation in relative flow angle onto the rotor aligns poorly to the rotor inlet metal angle. This leads to two dominant rotor loss sources: one at the tip associated with positive incidence and the second caused by negative incidence at lower span fractions. The second loss is reduced by opening the rotor stagger setting, and the first increases with rotor suction surface Mach number. The higher mass flow at more open rotor settings provide higher gross thrust, up to 49% of the forward take-off value, but is limited by the increased loss at high speed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081009-081009-8. doi:10.1115/1.4043217.

Impinging jet arrays are typically used to cool several gas turbine parts. Some examples of such applications can be found in the internal cooling of high-pressure turbine airfoils or in the turbine blade tip clearances control of aero-engines. The effect of the wall-to-jets temperature ratio (TR) on heat transfer is generally neglected by the correlations available in the open literature. In the present contribution, the impact of the temperature ratio on the heat transfer for a real engine active clearance control system is analyzed by means of validated computational fluid dynamics (CFD) computations. At different jets Reynolds number and considering several impingement array arrangements, a wide range of target wall-to-jets temperature ratio is accounted for. Computational results prove that both local and averaged Nusselt numbers reduce with increasing. An in-depth analysis of the numerical data shows that the last mentioned evidence is motivated by both the heat transfer incurring between the spent coolant flow and the fresh jets and the variation of gas properties with temperature through the boundary layer. A scaling procedure, based on the TR power law, was proposed to estimate the Nusselt number at different wall temperature levels necessary to correct available open-literature correlations, typically developed with small temperature differences, for real engine applications.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081010-081010-11. doi:10.1115/1.4043252.

This paper presents a novel tool for the shape optimization of turbomachinery blade profiles operating with fluids in non-ideal thermodynamic conditions and in complex flow configurations. In novel energy conversion systems, such as organic Rankine cycles or supercritical CO2 cycles, the non-conventional turbomachinery layout as well as the complex thermodynamics of the working fluid complicate significantly the blade aerodynamic design. For such applications, the design of turbomachinery may considerably benefit from the use of systematic optimization methods, especially in combination with high-fidelity computational fluid dynamics (CFD), as it is shown in this paper. The proposed technique is implemented in the shape-optimization package FORMA (Fluid-dynamic OptimizeR for turbo-Machinery Aerofoils) developed in-house at the Politecnico di Milano. FORMA is constructed as a combination of a generalized geometrical parametrization technique based on B-splines, a CFD solver featuring turbulence models and arbitrary equations of state, and multiple surrogate-based evolutionary strategies based on either trust-region or training methods. The application to the re-design of a supersonic turbine nozzle shows the capabilities of applying a high-fidelity optimization, consisting of a 50% reduction in the cascade loss coefficient and in an increased flow uniformity at the inlet of the subsequent rotor. Two alternative surrogate-based evolutionary strategies and different fitness functions are tested and discussed, including nonlinear constraints within the design process. The optimization study reveals relevant insights into the design of supersonic turbine nozzles as well on the performance, reliability, and potential of the proposed design technique.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(8):081011-081011-12. doi:10.1115/1.4043219.

The optimum aspect ratio at which maximum efficiency occurs is relatively low, typically between 1 and 1.5. At these aspect ratios, inaccuracies inherently exist in the decomposition of the flow field into freestream and endwall components due to the absence of a discernible freestream. In this paper, a unique approach is taken: a “linear repeating stage” concept is used in conjunction with a novel way of defining the freestream flow. Through this approach, physically accurate decomposition of the flow field for aspect ratios as low as ∼0.5 can be achieved. This ability to accurately decompose the flow leads to several key findings. First, the endwall flow is found to be dependent on static pressure rise coefficient and endwall geometry, but independent of the aspect ratio. Second, the commonly accepted relationship that endwall loss coefficient varies inversely with the aspect ratio is shown to be physically inaccurate. Instead, a new term, which the authors refer to as the “effective aspect ratio,” should replace the term “aspect ratio.” Moreover, not doing so can result in efficiency errors of ∼0.6% at low aspect ratios. Finally, there exists a low aspect ratio limit below which the two endwall flows interact causing a large separation to occur along the span. From these findings, a low-order model is developed to model the effect of varying aspect ratio on compressor performance. The last section of the paper uses this low-order model and a simple analytical model to show that to a first order, the optimum aspect ratio is just a function of the loss generated by the endwalls at zero clearance and the rate of change in profile loss due to blade thickness. This means that once the endwall configuration has been selected, i.e., cantilever or shroud, the blade thickness sets the optimum aspect ratio.

Commentary by Dr. Valentin Fuster

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