Research Papers

J. Turbomach. 2017;140(1):011001-011001-15. doi:10.1115/1.4038022.

This paper presents an innovative stability analysis and design approach for time-domain impedance boundary conditions to simulate noise propagation and radiation from a lined turbomachinery duct in the presence of a mean flow. A control-oriented model is developed for the stability analysis of the impedance boundary condition by using generalized function at the lining surface. The mean flow effect and sound propagation are considered in the model as well. Then, the numerical stability issue is analyzed by using the Bode plots before stabilized accordingly by employing the phase lead compensator method, which results in a rational transfer function. Finally, the corresponding time-domain implementation is achieved by using the so-called controllable canonical form rather than an inconvenient convolution operation. The performance of the current proposed approach is first validated in an in-duct propagation case by comparing to analytical solutions obtained by employing the Wiener–Hopf method and then demonstrated in a couple of duct acoustic problems with representative turbomachinery setups. The innovative cross-disciplinary nature of the current proposed approach can shed light on impedance problems and is very useful to time-domain acoustic simulations for turbomachinery applications.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011002-011002-8. doi:10.1115/1.4037820.

The integrated combustor vane concept for power generation gas turbines with can combustors has been shown to have significant benefits compared to conventional nozzle guide vanes (NGV). Aerodynamic loss, heat transfer levels, and cooling requirements are reduced while stage efficiency is improved by approximately 1.5% (for a no-swirl scenario). Engine realistic combustor flow with swirl, however, leads to increased turning nonuniformity downstream of the integrated vanes. This paper thus illustrates the altered integrated vane stage performance caused by inlet swirl. The study shows a distinct performance penalty for the integrated vane rotor as a result of increased rotor incidence and the rotor's interaction with the residual swirl core. The stage efficiency advantage of the integrated combustor vane concept compared to the conventional design is thus reduced to 0.7%. It is furthermore illustrated how integrated vane profiling is suitable to reduce the turning variation across the span downstream of the vane, further improve stage efficiency (in this case by 0.23%) and thus mitigate the distinct impact of inlet swirl on integrated vane stage performance.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011003-011003-10. doi:10.1115/1.4037997.

The effect of feeding shaped film cooling holes with an internal crossflow is not well understood. Previous studies have shown that internal crossflow reduces film cooling effectiveness from axial shaped holes, but little is known about the mechanisms governing this effect. It was recently shown that the crossflow-to-mainstream velocity ratio is important, but only a few of these crossflow velocity ratios have been studied. This effect is of concern because gas turbine blades typically feature internal passages that feed film cooling holes in this manner. In this study, film cooling effectiveness was measured for a single row of axial shaped cooling holes fed by an internal crossflow with crossflow-to-mainstream velocity ratio varying from 0.2 to 0.6 and jet-to-mainstream velocity ratios varying from 0.3 to 1.7. Experiments were conducted in a low speed flat plate facility at coolant-to-mainstream density ratios of 1.2 and 1.8. It was found that film cooling effectiveness was highly sensitive to crossflow velocity at higher injection rates while it was much less sensitive at lower injection rates. Analysis of the jet shape and lateral spreading found that certain jet characteristic parameters scale well with the crossflow-to-coolant jet velocity ratio, demonstrating that the crossflow effect is governed by how coolant enters the film cooling holes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011004-011004-10. doi:10.1115/1.4037998.

In gas turbine engines, film cooling holes are commonly fed with an internal crossflow, the magnitude of which has been shown to have a notable effect on film cooling effectiveness. In Part I of this study, as well as in a few previous studies, the magnitude of internal crossflow velocity was shown to have a substantial effect on film cooling effectiveness of axial shaped holes. There is, however, almost no data available in the literature that shows how internal crossflow affects compound angle shaped film cooling holes. In Part II, film cooling effectiveness, heat transfer coefficient augmentation, and discharge coefficients were measured for a single row of compound angle shaped film cooling holes fed by internal crossflow flowing both in-line and counter to the spanwise direction of coolant injection. The crossflow-to-mainstream velocity ratio was varied from 0.2 to 0.6 and the injection velocity ratio was varied from 0.2 to 1.7. It was found that increasing the magnitude of the crossflow velocity generally caused degradation of the film cooling effectiveness, especially for in-line crossflow. An analysis of jet characteristic parameters demonstrated the importance of crossflow effects relative to the effect of varying the film cooling injection rate. Heat transfer coefficient augmentation was found to be primarily dependent on injection rate, although for in-line crossflow, increasing crossflow velocity significantly increased augmentation for certain conditions.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011005-011005-13. doi:10.1115/1.4038120.

Impeller recirculation is a loss which has long been considered in one-dimensional (1D) modeling; however, the full extent of its impact on stage performance has not been analyzed. Recirculation has traditionally been considered purely as a parasitic (or external) loss, i.e., one which absorbs work but does not contribute to total pressure rise across the stage. Having extensively analyzed the impact of recirculation on the impeller exit flow field, it was possible to show that it has far-reaching consequences beyond that of increasing total temperature. The overall aim of this package of work is to apply a much more physical treatment to the impact of impeller exit recirculation (and the aerodynamic blockage associated with it) and hence realize an improvement in the 1D stage performance prediction of a number of turbocharger centrifugal compressors. The factors influencing the presence and extent of this recirculation are numerous, requiring detailed investigations to successfully understand its sources and to characterize its extent. A combination of validated three-dimensional computational fluid dynamics (CFD) data and gas stand test data for six automotive turbocharger compressor stages was employed to achieve this aim. In order to capture the variation of the blockage presented to the flow with both geometry and operating condition, an approach involving the impeller outlet to inlet area ratio and a novel flow coefficient term were employed. The resulting data permitted the generation of a single equation to represent the impeller exit blockage levels for the entire operating map of all the six compressor stages under investigation. With an understanding of the extent of the region of recirculating flow realized and the key drivers leading to its creation identified, it was necessary to comprehend how the resulting blockage influenced compressor performance. Consideration was given to the impact on impeller work input through modification of the impeller exit velocity triangle, incorporating the introduction of the concept of an “aerodynamic meanline” to account for the reduction in the size of the active flow region due to the presence of blockage. The sensitivity of the stage to this change was then related back to the level of backsweep applied to the impeller. As a result of this analysis, the improvement in the 1D performance prediction of the six compressor stages is demonstrated. In addition, a number of design recommendations are presented to ensure that the detrimental effects associated with the presence of impeller exit recirculation can be minimized.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011006-011006-11. doi:10.1115/1.4038150.

The combined effects of upstream purge flow, slashface leakage flow, and discrete hole film cooling on turbine blade platform film cooling effectiveness were studied using the pressure sensitive paint (PSP) technique. As a continued study, discrete cylindrical holes were replaced by laidback fan-shaped (10-10-5) holes, which generally provide better film coverages on the endwall. Experiments were done in a five-blade linear cascade. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. A wide range of parameters was evaluated in this study. The coolant-to-mainstream mass flow ratio (MFR) was varied from 0.5%, 0.75%, to 1% for the upstream purge flow. For the platform film cooling holes and slashface gap, average blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1 (close to low-temperature experiments) to 1.5 (intermediate DR) and 2 (close to engine conditions) were also examined. Purge flow swirl effect was studied particularly at a typical swirl ratio (SR) of 0.6. Area-averaged film cooling effectiveness results were compared between cylindrical and fan-shaped holes. The results indicate that the fan-shaped holes provide superior film coverage than cylindrical holes for platform film cooling especially at higher blowing ratios and momentum flux ratios.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011007-011007-10. doi:10.1115/1.4038179.

The demand for higher efficiency is ever present in the gas turbine field and can be achieved through many different approaches. While additively manufactured parts have only recently been introduced into the hot section of a gas turbine engine, the manufacturing technology shows promise for more widespread implementation since the process allows a designer to push the limits on capabilities of traditional machining and potentially impact turbine efficiencies. Pin fins are conventionally used in turbine airfoils to remove heat from locations in which high thermal and mechanical stresses are present. This study employs the benefits of additive manufacturing to make uniquely shaped pin fins, with the goal of increased performance over conventional cylindrical pin fin arrays. Triangular, star, and spherical shaped pin fins placed in microchannel test coupons were manufactured using direct metal laser sintering (DMLS). These coupons were experimentally investigated for pressure loss and heat transfer at a range of Reynolds numbers. Spacing, number of pin fins in the array, and pin fin geometry were variables that changed pressure loss and heat transfer in this study. Results indicate that the additively manufactured triangles and cylinders outperform conventional pin fin arrays, while stars and dimpled spheres did not.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011008-011008-12. doi:10.1115/1.4038181.

The stable operating range of a centrifugal compressor stage of an engine turbocharger is limited at low mass flow rates by aerodynamic instabilities which can lead to the onset of rotating stall or surge. There have been many techniques employed to increase the stable operating range of centrifugal compressor stages. The literature demonstrates that there are various possibilities for adding special treatments to the nominal diffuser vane geometry, or including injection or bleed flows to modify the diffuser flow field in order to influence diffuser stability. One such treatment is the porous throat diffuser (PTD). Although the benefits of this technique have been proven in the existing literature, a comprehensive understanding of how this technique operates is not yet available. This paper uses experimental measurements from a high pressure ratio (PR) compressor stage to acquire a sound understanding of the flow features within the vaned diffuser which affect the stability of the overall compression system and investigate the stabilizing mechanism of the porous throat diffuser. The nonuniform circumferential pressure imposed by the asymmetric volute is experimentally and numerically examined to understand if this provides a preferential location for stall inception in the diffuser. The following hypothesis is confirmed: linking of the diffuser throats via the side cavity equalizes the diffuser throat pressure, thus creating a more homogeneous circumferential pressure distribution, which delays stall inception to lower flow rates. The results of the porous throat diffuser configuration are compared to a standard vaned diffuser compressor stage in terms of overall compressor performance parameters, circumferential pressure nonuniformity at various locations through the compressor stage and diffuser subcomponent analysis. The diffuser inlet region was found to be the element most influenced by the porous throat diffuser, and the stability limit is mainly governed by this element.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011009-011009-11. doi:10.1115/1.4038182.

As additive manufacturing (AM) technologies utilizing metal powders continue to mature, the usage of AM parts in gas turbine engines will increase. Current metal AM technologies produce parts with substantial surface roughness that can only be removed from external surfaces and internal surfaces that are accessible for smoothing. Difficulties arise in making smooth the surfaces of small internal channels, which means the augmentation of pressure loss and heat transfer due to roughness must be accounted for in the design. As gas turbine manufacturers have only recently adopted metal AM technologies, much remains to be examined before the full impacts of applying AM to turbine parts are understood. Although discrete film cooling holes have been extensively studied for decades, this objective of this study was to understand how the roughness of film cooling holes made using AM can affect the overall cooling effectiveness. Coupons made from a high temperature nickel alloy with engine-scale film holes were tested in a rig designed to simulate engine relevant conditions. Two different hole sizes and two different build directions were examined at various blowing ratios. Results showed that the effectiveness is dependent on the build direction and the relative size of the hole. It was also discovered that commercially available AM processes could not reliably produce small holes with predictable behavior.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;140(1):011010-011010-12. doi:10.1115/1.4038176.

The detailed design of the inducer of a high pressure ratio transonic radial compressor impeller with a design inlet tip relative Mach number of 1.4 is considered. Numerical analysis has been used to compare a datum impeller with ruled inducer design with a number of different free-form design concepts, generated following the same aerodynamic design philosophy. The datum stage and one with a free-form inducer, referred to as “barrelled forward swept,” with forward swept leading edge near the tip and increased chord at midspan, have been manufactured and tested. The tests were performed with the same stationary components, including the casing, vaned diffuser, and the volute. The design with a barrelled forward sweep of the inducer allows the designer more control of the strength and position of the passage shock at the inlet while meeting mechanical constraints. Interestingly, the performance is also enhanced at off-design points at lower tip-speeds. The measurements show that the stage tested with the swept impeller achieves higher efficiency of between 0.5% and 1.6% compared to the datum design, depending on the operating speed. The computational fluid dynamics (CFD) simulations are used to further study the flow at part speeds, in order to explain the causes of the observed performance differences at off design conditions.

Commentary by Dr. Valentin Fuster

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