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Research Papers: Gas Turbines: Aircraft Engine

J. Eng. Gas Turbines Power. 2016;139(5):051201-051201-10. doi:10.1115/1.4034964.

Porous media model computational fluid dynamics (CFD) is a valuable approach allowing an entire heat exchanger system, including the interactions with its associated installation ducts, to be studied at an affordable computational effort. Previous work of this kind has concentrated on developing the heat transfer and pressure loss characteristics of the porous medium model. Experimental validation has mainly been based on the measurements at the far field from the porous media exit. Detailed near field data are rare. In this paper, the fluid dynamics characteristics of a tubular heat exchanger concept developed for aero-engine intercooling by the authors are presented. Based on a rapid prototype manufactured design, the detailed flow field in the intercooler system is recorded by particle image velocimetry (PIV) and pressure measurements. First, the computational capability of the porous media to predict the flow distribution within the tubular heat transfer units was confirmed. Second, the measurements confirm that the flow topology within the associated ducts can be described well by porous media CFD modeling. More importantly, the aerodynamic characteristics of a number of critical intercooler design choices have been confirmed, namely, an attached flow in the high velocity regions of the in-flow, particularly in the critical region close to the intersection and the in-flow guide vane, a well-distributed flow in the two tube stacks, and an attached flow in the cross-over duct.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):051202-051202-10. doi:10.1115/1.4034986.

The paper presents a thorough analysis of the historical data and results acquired over a period of two years through an on-line real-time monitoring system installed at a combined heat and power (CHP) plant. For gas turbine health and performance assessment, a gas path analysis tool based on the adaptive modeling method is integrated into the system. An engine adapted model built through a semi-automated method is part of a procedure which includes a steam/water cycle simulation module and an economic module used for power plant performance and economic assessment. The adaptive modeling diagnostic method allowed for accurate health assessment during base and part load operation identifying and quantifying compressor recoverable deterioration and the root cause of an engine performance shift. Next, the performance and economic assessment procedure was applied for quantifying the economic benefit accrued by implementing daily on-line washing and for evaluating the financial gains if the off-line washings time intervals are optimized based on actual engine performance deterioration rates. The results demonstrate that this approach allows continuous health and performance monitoring at full and part load operation enhancing decision making capabilities and adding to the information that can be acquired through traditional analysis methods based on heat balance and base load correction curves.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Coal, Biomass, and Alternative Fuels

J. Eng. Gas Turbines Power. 2016;139(5):051401-051401-7. doi:10.1115/1.4034942.

The demand for more environmentally friendly and economic power production has led to an increasing interest to utilize alternative fuels. In the past, several investigations focusing on the effect of low-calorific fuels on the combustion process and steady-state performance have been published. However, it is also important to consider the transient behavior of the gas turbine when operating on nonconventional fuels. The alternative fuels contain very often a large amount of dilutants resulting in a low energy density. Therefore, a higher fuel flow rate is required, which can impact the dynamic behavior of the gas turbine. This paper will present an investigation of the transient behavior of the all-radial OP16 gas turbine. The OP16 is an industrial gas turbine rated at 1.9 MW, which has the capability to burn a wide range of fuels including ultra-low-calorific gaseous fuels. The transient behavior is simulated using the commercial software GSP including the recently added thermal network modeling functionality. The steady-state and transient performance model is thoroughly validated using real engine test data. The developed model is used to simulate and analyze the physical behavior of the gas turbine when performing load sheds. From the simulations, it is found that the energy density of the fuel has a noticeable effect on the rotor over-speed and must be considered when designing the fuel control.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Combustion, Fuels, and Emissions

J. Eng. Gas Turbines Power. 2016;139(5):051501-051501-7. doi:10.1115/1.4034966.

Spark-assisted compression ignition (SACI) offers more practical combustion phasing control and a lower pressure rise rate than homogeneous charge compression ignition (HCCI) combustion and improved thermal efficiency and lower NOx emissions than spark ignition (SI) combustion. Any practical passenger car engine, including one that uses SACI in part of its operating range, must be robust to changes in ambient conditions. This study investigates the effects of ambient temperature and humidity on stoichiometric SACI combustion and emissions. It is shown that at the medium speed and load SACI test point selected for this study, increasing ambient air temperature from 20 °C to 41 °C advances combustion phasing, increases maximum pressure rise rate, causes a larger fraction of the charge to be consumed by auto-ignition (and a smaller fraction by flame propagation), and increases NOx. Increasing ambient humidity from 32% to 60% retards combustion phasing, reduces maximum pressure rise rate, increases coefficient of variation (COV) of indicated mean effective pressure (IMEP), reduces NOx, and increases brake-specific fuel consumption (BSFC). These results show that successful implementation of SACI combustion in real-world driving requires a control strategy that compensates for changes in ambient temperature and humidity.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):051502-051502-8. doi:10.1115/1.4034969.

Two pyrometric tools for measuring soot temperature response in fuel-rich flames under unsteady inlet airflow conditions are developed. High-speed pyrometry using a high-speed color camera is used in producing soot temperature distributions, with its results compared with those of global soot temperature response measured using a multiwavelength pyrometer. For the former, the pixel red, green, and blue (RGB) values pertaining to respective bandwidths of red, green, and blue filters are used to calculate temperature and for the latter, the emission from whole flame at 660 nm, 730 nm, and 800 nm is used to measure temperature. The combustor, running on jet-A fuel, achieves unsteady inlet airflow using a siren running at frequencies of 150 and 250 Hz and with modulation levels (root mean square (RMS)) 20–50% of mean velocity. Spatiotemporal response of flame temperature measured by the high-speed camera is presented by phase-averaged with average subtracted images and by fast Fourier transform (FFT) at the modulation frequencies of inlet velocity. Simultaneous measurement of combustor inlet air velocity and flame soot temperature using the multiwavelength pyrometer is used in calculating the flame transfer function (FTF) of flame temperature response to unsteady inlet airflow. The results of global temperature and temperature fluctuation from the three-color pyrometer show qualitative agreement with the local temperature response measured by the high-speed camera. Over the range of operating conditions employed, the overall flame temperature fluctuation increases linearly with respect to the inlet velocity fluctuation. The two-dimensional map of flame temperature under unsteady combustion determined using a high-speed digital color camera shows that the local temperature fluctuation during unsteady combustion occurs over relatively small region of flame and its level is greater (∼10% to 20%) than that of overall temperature fluctuation (∼1%).

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):051503-051503-10. doi:10.1115/1.4035143.

The current work focuses on the large eddy simulation (LES) of combustion instability in a laboratory-scale swirl burner. Air and fuel are injected at ambient conditions. Heat conduction from the combustion chamber to the plenums results in a preheating of the air and fuel flows above ambient conditions. The paper compares two computations: In the first computation, the temperature of the injected reactants is 300 K (equivalent to the experiment) and the combustor walls are treated as adiabatic. The frequency of the unstable mode (≈ 635 Hz) deviates significantly from the measured frequency (≈ 750 Hz). In the second computation, the preheating effect observed in the experiment and the heat losses at the combustion chamber walls are taken into account. The frequency (≈ 725 Hz) of the unstable mode agrees well with the experiment. These results illustrate the importance of accounting for heat transfer/losses when applying LES for the prediction of combustion instabilities. Uncertainties caused by unsuitable modeling strategies when using computational fluid dynamics for the prediction of combustion instabilities can lead to an improper design of passive control methods (such as Helmholtz resonators) as these are often only effective in a limited frequency range.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):051504-051504-13. doi:10.1115/1.4035228.

To manage the increasing turbine temperatures of future gas turbines a cooled cooling air system has been proposed. In such a system some of the compressor efflux is diverted for additional cooling in a heat exchanger (HX) located in the bypass duct. The cooled air must then be returned, across the main gas path, to the engine core for use in component cooling. One option is do this within the combustor module and two methods are examined in the current paper; via simple transfer pipes within the dump region or via radial struts in the prediffuser. This paper presents an experimental investigation to examine the aerodynamic impact these have on the combustion system external aerodynamics. This included the use of a fully annular, isothermal test facility incorporating a bespoke 1.5 stage axial compressor, engine representative outlet guide vanes (OGVs), prediffuser, and combustor geometry. Area traverses of a miniature five-hole probe were conducted at various locations within the combustion system providing information on both flow uniformity and total pressure loss. The results show that, compared to a datum configuration, the addition of transfer pipes had minimal aerodynamic impact in terms of flow structure, distribution, and total pressure loss. However, the inclusion of prediffuser struts had a notable impact increasing the prediffuser loss by a third and consequently the overall system loss by an unacceptable 40%. Inclusion of a hybrid prediffuser with the cooled cooling air (CCA) bleed located on the prediffuser outer wall enabled an increase of the prediffuser area ratio with the result that the system loss could be returned to that of the datum level.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):051505-051505-10. doi:10.1115/1.4035207.

The present article reports a numerical analysis of instability coupled by a spinning mode in an annular combustor. This corresponds to experiments carried out on the MICCA test facility equipped with 16 matrix burners. Each burner response is represented by means of a global experimental flame describing function (FDF). A harmonic balance nonlinear stability analysis is carried out by combining the FDF with a Helmholtz solver to determine the system dynamics trajectories in a frequency-growth rate plane. The influence of the distribution of the volumetric heat release corresponding to each burner is investigated in a first stage. Even though each of the 16 burners is compact with respect to the transverse mode wavelength, and the 16 flames occupy the same volume, this distribution of heat release is not compact in the azimuthal direction and simulations reveal an influence of this volumetric distribution on frequencies and growth rates. This study emphasizes the importance of providing a suitable description of the flame zone geometrical extension and correspondingly an adequate representation of the level of heat release rate fluctuation per unit volume. It is found that these two items can be deduced from a knowledge of the heat release distribution under steady-state operating conditions. Once the distribution of the heat release fluctuations is unequivocally defined, limit cycle simulations are performed. For the conditions explored, simulations retrieve the spinning nature of the self-sustained mode that was identified in the experiments both in the plenum and in the combustion chamber.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Controls, Diagnostics, and Instrumentation

J. Eng. Gas Turbines Power. 2017;139(5):051601-051601-10. doi:10.1115/1.4034943.

Gas turbines overall efficiency enhancement requires further increasing of the firing temperature and decreasing of cooling flow usage. Multihole (or effusion, or full-coverage) film cooling is widely used for hot gas path components cooling in modern gas turbines. The present study focused on the adiabatic film effectiveness measurement of a round multihole flat-plate coupon. The measurements were conducted in a subsonic open-loop wind tunnel with a generic setup to cover different running conditions. The test conditions were characterized by a constant main flow Mach number of 0.1 with constant gas temperature. Adiabatic film effectiveness was measured by pressure-sensitive paint (PSP) through mass transfer analogy. CO2 was used as the coolant to reach the density ratio of 1.5. Rig computational fluid dynamics (CFD) simulation was conducted to evaluate the impact of inlet boundary layer on testing. Experimental data cover blowing ratios (BRs) at 0.4, 0.6, 0.8, 1.0, and 2.0. Both 2D maps and lateral average profiles clearly indicated that the film effectiveness increases with increasing BR for BR < 0.8 and decreases with increasing BR for BR > 0.8. This observation agreed with coolant jet behavior of single film row, i.e., attached, detached then reattached, and fully detached. PSP data quality was then discussed in detail for validating large eddy simulation.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Heat Transfer

J. Eng. Gas Turbines Power. 2017;139(5):051901-051901-9. doi:10.1115/1.4035144.

Previous studies have indicated some differences between steady computational fluid dynamics (CFD) predictions of flow in a rotor–stator disk cavity with rotating bolts compared to measurements. Recently, time-dependent CFD simulations have revealed the unsteadiness present in the flow and have given improved agreement with measurements. In this paper, unsteady Reynolds averaged Navier–Stokes (URANS) 360 deg model CFD calculations of a rotor–stator cavity with rotor bolts were performed in order to better understand the flow and heat transfer within a disk cavity previously studied experimentally by other workers. It is shown that the rotating bolts generate unsteadiness due to wake shedding which creates time-dependent flow patterns within the cavity. At low throughflow conditions, the unsteady flow significantly increases the average disk temperature. A systematic parametric study is presented giving insight into the influence of number of bolts, mass flow rate, cavity gap ratio, and the bolts-to-shroud gap ratio on the time-dependent flow within the cavity.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Manufacturing, Materials, and Metallurgy

J. Eng. Gas Turbines Power. 2017;139(5):052101-052101-6. doi:10.1115/1.4035159.

Nozzle guide vanes (NGV) of gas turbine engines are the first components to withstand the impingement of hot combustion gas and therefore often suffer thermal fatigue failures in service. A lifting analysis is performed for the NGV of a gas turbine engine using the integrated creep–fatigue theory (ICFT). With the constitutive formulation of inelastic strain in terms of mechanism-strain components such as rate-independent plasticity, dislocation glide-plus-climb, and grain boundary sliding (GBS), the dominant deformation mechanisms at the critical locations are thus identified quantitatively with the corresponding mechanism-strain component. The material selection scenarios are discussed with regards to damage accumulated during take-off and cruise. The interplay of those deformation mechanisms in the failure process is elucidated such that an “optimum” material selection solution may be achieved.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Oil and Gas Applications

J. Eng. Gas Turbines Power. 2016;139(5):052401-052401-13. doi:10.1115/1.4034968.

Solid particle ingestion is one of the principal degradation mechanisms in the compressor and turbine sections of gas turbines. In particular, in industrial applications, the microparticles not captured by the air filtration system can cause deposits on blading and, consequently, result in a decrease in the compressor performance. This paper presents three-dimensional numerical simulations of the microparticle ingestion (0.15–1.50 μm) in a transonic axial compressor stage, carried out by means of a commercial computational fluid dynamic code. Particles of this size can follow the main air flow with relatively little slip, while being impacted by the flow turbulence. It is of great interest to the industry to determine which zones of the compressor blades are impacted by these small particles. Particle trajectory simulations use a stochastic Lagrangian tracking method that solves the equations of motion separately from the continuous phase. A particular computational strategy is adopted in order to take into account the presence of two subsequent annular cascades (rotor and stator) in the case of particle ingestion. The proposed strategy allows the evaluation of particle deposition in an axial compressor stage, thanks to its capability of accounting for rotor/stator interaction. NASA Stage 37 is used as a case study for the numerical investigation. The compressor stage numerical model and the discrete phase model are set up and validated against the experimental and numerical data available in the literature. The blade zones affected by the particle impact and the kinematic characteristics of the impact of micrometric and submicrometric particles with the blade surface are shown. Both blade zones affected by the particle impact and deposition are analyzed. The particle deposition is established by using the quantity called sticking probability, adopted from the literature. The sticking probability links the kinematic characteristics of particle impact on the blade with fouling phenomenon. The results show that microparticles tend to follow the flow by impacting at full span with a higher impact concentration on the pressure side of rotor blade and stator vane. Both the rotor blade and stator vane suction side are only affected by the impact of smaller particles (up to 1 μm). Particular fluid dynamic phenomena, such as separation, shock waves, and tip leakage vortex, strongly influence the impact location of the particles. The kinematic analysis shows a high tendency of particle adhesion on the suction side of the rotor blade, especially for particles with a diameter equal to 0.15 μm.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Structures and Dynamics

J. Eng. Gas Turbines Power. 2016;139(5):052501-052501-12. doi:10.1115/1.4034920.

In this study, experimental and analytical analyses of the vibration stability of a 225 kW class turbo blower with a hybrid foil–magnetic bearing (HFMB) were performed. First, critical speed and unbalance vibration responses were examined as part of the rotordynamic research. Its shaft diameter was 71.5 mm, its total length was 693 mm, and the weight of the rotor was 17.8 kg. The air foil bearing (AFB) utilized was 50 mm long and had a 0.7 aspect ratio. In the experiments conducted, excessive vibration and rotor motion instability occurred in the range 12,000–15,000 rpm, which resulted from insufficient dynamic pressure caused by the length of the foil bearing being too short. Consequently, as the rotor speed increased, excessive rotor motion attributable to aerodynamic and bearing instability became evident. This study therefore focused on improving rotordynamic performance by rectifying rigid mode unstable vibration at low speed, 20,000 rpm, and asynchronous vibration due to aerodynamic instability by using HFMB with vibration control. The experimental results obtained were compared for each bearing type (AFB and HFMB) to improve the performance of the vibration in the low-speed region. The experimental results show that the HFMB technology results in superior vibration stability for unbalance vibration and aerodynamic instability in the range 12,000–15,000 rpm (200–250 Hz). The remarkable vibration reduction achieved from vibration control of the HFMB–rotor system shows that oil-free turbomachinery can achieve excellent performance.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):052502-052502-10. doi:10.1115/1.4034965.

Hole-pattern or honeycomb seals have been commonly used for many years in the Oil & Gas industry as damper seals for turbomachinery. The main motivation has been to introduce additional damping to improve the shaft rotordynamic stability operating under high-pressure conditions. Experience has shown that the dynamic and even static characteristics of those seals are very sensitive to the operating clearance profile as well as the installation tolerances. Rotordynamic stability is related not only to the seal effective damping but to the effective stiffness as well. In fact, for this kind of seal, the effective stiffness can be high enough to alter the rotor system's natural frequency. The seal stiffness is strictly related to the tapering contour: if the clearance profile changes from divergent to convergent, the effective stiffness may change from a strong negative to a strong positive magnitude, thus avoiding the rotor natural frequency drop as it is detrimental for the stability. Unfortunately, the effective damping is reduced at the same time but this effect can be mitigated using proper devices to keep the preswirl low or even negative (e.g., swirl brakes and shunt holes). This paper presents the results from an extended test campaign performed in a high-speed rotor test rig equipped with active magnetic bearings (AMBs) working under high pressure (14 krpm, 200 bar gas inlet pressure), with the aim to validate the rotordynamic characteristics of a negative preswirl, convergent honeycomb seal and demonstrate its ability to effectively act as a gas bearing as well as a seal. The test plan included variations of inlet pressure, differential pressure (given the same inlet pressure), as well as rotational speed in order to fully validate the seal behavior. This kind of test was performed in a “dynamic mode” that is to say exciting the spinning test rotor through a pair of AMBs along linear orbits. Additionally, the impact of the seal to rotor static eccentricity and the seal to rotor angular misalignment were both experimentally investigated and compared to relevant computational fluid dynamics (CFD) simulations. This kind of test was performed in a “static mode,” that is to say imposing through the AMBs the required eccentricity/angular misalignment and then measuring the forces needed to keep the rotor in the original position. Dynamic mode test was also performed in order to check the impact of the seal static eccentricity on its dynamic behavior. Finally, the test results were compared with predictions from a state of the art bulk-flow code in order to check the predictability level for future design applications.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):052503-052503-8. doi:10.1115/1.4034967.

The usual ways to measure the aerodynamic forcing function are complex and expensive. The aim of this work is to evaluate the forces acting on the blades using a relatively simpler experimental methodology based on a time-resolved pressure measurement at the rotor discharge. Upstream of the rotor, a steady three holes probe (3HP) has been used. The postprocessing procedures are described in detail, including the application of a phase-locked average and of an extension algorithm with phase-lag. The algorithm for the computation of the force components is presented, along with the underlying assumptions. In order to interpret the results, a preliminary description of the flowfield, both upstream and downstream of the rotor, is provided. This gives an insight of the most relevant features that affect the computation of the forces. Finally, the analysis of the results is presented. These are first described and then compared with overall section-average results (torque-sensor), and with the results from 3D unsteady simulations (integral of pressure over the blade surface) in order to assess the accuracy of the method. Both the experimental and the numerical results are also compared for two different operating conditions with increasing stage load.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2016;139(5):052504-052504-12. doi:10.1115/1.4035142.

In this paper, a novel and efficient modal analysis method is raised to work on blisk structures based on mixed-dimension finite element model (MDFEM). The blade and the disk are modeled separately. The blade model is figured by 3D solid elements considering its complex configuration and its degrees-of-freedom (DOFs) are condensed by dynamic substructural method. Meanwhile, the disk is structured by 2D axisymmetric element developed specially in this paper. The DOFs of entire blisk are tremendously reduced by this modeling approach. The key idea of this method is derivation of displacement compatibility to different dimensional models. Mechanical energy equivalence and summation further contribute to the model synthesis and modal analysis of blade and disk. This method has been successfully applied on the modal analysis of blisk structures in turbine, which reveals its effectiveness and proves that this method reduces the computational time expenses while maintaining the precision performances of full 3D model. Though there is limitation that structure should have proper coverage of blades, this method is still feasible for most blisks in engineering practice.

Topics: Disks , Blades , Displacement
Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052505-052505-10. doi:10.1115/1.4035176.

The recent move toward subsea oil and gas production brings about a requirement to locate process equipment in deepwater installations. Furthermore, there is a drive toward omitting well stream separation functionality, as this adds complexity and cost to the subsea installation. This in turn leads to technical challenges for the subsea installed pumps and compressors that are now required to handle multiphase flow of varying gas to liquid ratios. This highlights the necessity for a strong research focus on multiphase flow impact on rotordynamic properties and thereby operational stability of the subsea installed rotating machinery. It is well known that careful design of turbomachinery seals, such as interstage and balance piston seals, is pivotal for the performance of pumps and compressors. Consequently, the ability to predict the complex interaction between fluid dynamics and rotordynamics within these seals is key. Numerical tools offering predictive capabilities for turbomachinery seals in multiphase flow are currently being developed and refined, however the lack of experimental data for multiphase seals renders benchmarking and validation impossible. To this end, the Technical University of Denmark and Lloyd's Register Consulting are currently establishing a purpose built state of the art multiphase seal test facility, which is divided into three modules. Module I consists of a full scale active magnetic bearing (AMB) based rotordynamic test bench. The internally designed custom AMBs are equipped with an embedded Hall sensor system enabling high-precision noncontact seal force quantification. Module II is a fully automatized calibration facility for the Hall sensor based force quantification system. Module III consists of the test seal housing assembly. This paper provides details on the design of the novel test facility and the calibration of the Hall sensor system employed to measure AMB forces. Calibration and validation results are presented, along with an uncertainty analysis on the force quantification capabilities.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052506-052506-9. doi:10.1115/1.4035175.

Recent studies have demonstrated that the aerothermal characteristics of turbine rotor blade tip under a transonic condition are qualitatively different from those under a low-speed subsonic condition. The cooling injection adds further complexity to the over-tip-leakage (OTL) transonic flow behavior and aerothermal performance, particularly for commonly studied shroudless tip configurations such as a squealer tip. However there has been no published experimental study of a cooled transonic squealer. The present study investigates the effect of cooling injection on a transonic squealer through a closely combined experimental and CFD effort. Part I of this two-part paper presents the first of the kind tip cooling experimental data obtained in a transonic linear cascade environment (exit Mach number 0.95). Transient thermal measurements are carried out for an uncooled squealer tip and six cooling configurations with different locations and numbers of discrete holes. High-resolution distributions of heat transfer coefficient and cooling effectiveness are obtained. ansysFluent is employed to perform numerical simulations for all the experimental cases. The mesh and turbulence modeling dependence is first evaluated before further computational studies are carried out. Both the experimental and computational results consistently illustrate strong interactions between the OTL flow and cooling injection. When the cooling injection (even with a relatively small amount) is introduced, distinctive series of stripes in surface heat transfer coefficient are observed with an opposite trend in the chordwise variations on the squealer cavity floor and on the suction surface rim. Both experimental and CFD results have also consistently shown interesting signatures of the strong OTL flow–cooling interactions in terms of the net heat flux reduction distribution in areas seemingly unreachable by the coolant. Further examinations and analyses of the related flow physics and underlining vortical flow structures will be presented in Part II.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052507-052507-9. doi:10.1115/1.4035200.

A basic attribute for turbine blade film cooling is that coolant injected should be largely passively convected by the local base flow. However, the effective working of the conventional wisdom may be compromised when the cooling injection strongly interacts with the base flow. Rotor blade tip of a transonic high-pressure (HP) turbine is one of such challenging regions for which basic understanding of the relevant aerothermal behavior as a basis for effective heat transfer/cooling design is lacking. The need to increase our understanding and predictability for high-speed transonic blade tip has been underlined by some recent findings that tip heat transfer characteristics in a transonic flow are qualitatively different from those at a low speed. Although there have been extensive studies previously on squealer blade tip cooling, there have been no published experimental studies under a transonic flow condition. The present study investigates the effect of cooling injection on a transonic squealer tip through a closely combined experimental and computational fluid dynamics (CFD) effort. The experimental and computational results as presented in Part I have consistently revealed some distinctive aerothermal signatures of the strong coolant-base flow interactions. In this paper, as Part II, detailed analyses using the validated CFD solutions are conducted to identify, analyze, and understand the causal links between the aerothermal signatures and the driving flow structures and physical mechanisms. It is shown that the interactions between the coolant injection and the base over-tip leakage (OTL) flow in the squealer tip region are much stronger in the frontal subsonic region than the rear transonic region. The dominant vortical flow structure is a counter-rotating vortex pair (CRVP) associated with each discrete cooling injection. High HTC stripes on the cavity floor are directly linked to the impingement heat transfer augmentation associated with one leg of the CRVP, which is considerably enhanced by the near-floor fluid movement driven by the overall pressure gradient along the camber line (CAM). The strength of the coolant-base flow interaction as signified by the augmented values of the HTC stripes is seen to correlate to the interplay and balance between the OTL flow and the CRVP structure. As such, for the frontal subsonic part of the cavity, there is a prevailing spanwise inward flow initiated by the CRVP, which has profoundly changed the local base flow, leading to high HTC stripes on the cavity floor. On the other hand, for the rear high speed part, the high inertia of the OTL flow dominates; thus, the vortical flow disturbances associated with the CRVP are largely passively convected, leaving clear signatures on the top surface of the suction surface rim. A further interesting side effect of the strong interaction in the frontal subsonic region is that there is considerable net heat flux reduction (NHFR) in an area seemingly unreachable by the injected coolant. The present results have confirmed that this is due to the large reduction in the local HTC as a consequence of the upstream propagated impact of the strong coolant-base flow interactions.

Commentary by Dr. Valentin Fuster

Research Papers: Gas Turbines: Turbomachinery

J. Eng. Gas Turbines Power. 2016;139(5):052601-052601-9. doi:10.1115/1.4034689.

It is well known that compressor surge imposes a significant limit on the flow range of a turbocharged internal combustion engine. The centrifugal compressor is commonly placed upstream of the inlet manifold, and hence, it is exposed to the intermittent flow regime of the inlet valves. Surge phenomena have been well studied over the past decades, and there still remains limited information with regard to the unsteady impact caused by the inlet valves. This study presents an experimental evaluation of such a situation. Engine representative pulses are created by a downstream system comprising a large volume, two rotating valves, a throttle valve, and the corresponding pipe network. Different pulsation levels are characterized by means of their frequency and the corresponding amplitude at the compressor inlet. The stability limit of the system under study is evaluated with reference to the parameter B proposed by Greitzer (1976, “Surge and Rotating Stall in Axial Flow Compressors—Part II: Experimental Results and Comparison With Theory,” ASME J. Eng. Power, 98(2), pp. 199–211; 1976, “Surge and Rotating Stall in Axial Flow Compressors—Part I: Theoretical Compression System Model,” ASME J. Eng. Power, 98(2), pp. 190–198). B describes the dynamics of the compression system in terms of volume, area, equivalent length, and compressor tip speed as well as the Helmholtz frequency of the system. For a given compressor, as B goes beyond a critical value, the system will exhibit surge as the result of the flow instability progression. The reduced frequency analysis shows that the scroll diffuser operates in an unsteady regime, while the impeller is nearly quasi-steady. In the vicinity of the surge point, under a pulsating flow, the instantaneous operation of the compressor showed significant excursions into the unstable side of the surge line. Furthermore, it has been found that the presence of a volume in the system has the greatest effect on the surge margin of the compressor under the unsteady conditions.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052602-052602-8. doi:10.1115/1.4035206.

This paper reports the internal performance evaluation of S-duct diffusers with different entrance aspect ratios as part of a parametric investigation of a generic S-duct inlet. The generic S-duct diffusers studied had a rectangular entrance (aspect ratios of 1.5 and 2.0) transitioning S-duct diffuser in high-subsonic (Mach number > 0.8) flow. The test section was manufactured using rapid prototyping to facilitate the parametric investigation of the geometry. Streamwise static pressure and exit-plane total pressure were measured in a test-rig using surface pressure taps and a five-probe rotating rake, respectively. The baseline and a variant were simulated through computational fluid dynamics (CFD). The investigation indicated the presence of streamwise and circumferential pressure gradients leading to a three-dimensional flow in the S-duct diffuser and to distortion at the exit plane. The static pressure recovery increased for the diffuser with the higher aspect ratio. Total pressure losses and circumferential and radial distortions at the exit plane were higher than that of the podded nacelle type of inlet. An increase in the total pressure recovery was observed for the increase in the aspect ratio for the baseline area ratio (1.57) S-ducts, but without a clear trend for the other area ratio (1.8) ducts. The work represents the development of a database on the performance of a particular type of generic inlet. This database will be useful for predicting the performance of aero-engines and air vehicles in high-subsonic flight.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052603-052603-9. doi:10.1115/1.4035158.

This work aims at investigating the impact of axial gap variation on aerodynamic performance of a high-pressure steam turbine stage. Numerical and experimental campaigns were conducted on a 1.5-stage of a reaction steam turbine. This low speed test rig was designed and operated in different operating conditions. Two different configurations were studied in which blades axial gap was varied in a range from 40% to 95% of the blade axial chord. Numerical analyses were carried out by means of three-dimensional, viscous, unsteady simulations, adopting measured inlet/outlet boundary conditions. Two sets of measurements were performed: steady measurements, from one hand, for global performance estimation of the whole turbine, such as efficiency, mass flow, and stage work; steady and unsteady measurements, on the other hand, were performed downstream of rotor row, in order to characterize the flow structures in this region. The fidelity of computational setup was proven by comparing numerical results to measurements. Main performance curves and spanwise distributions have shown a good agreement in terms of both shape of curves/distributions and absolute values. Moreover, the comparison of two-dimensional maps downstream of rotor row has shown similar structures of the flow field. Finally, a comprehensive study of the axial gap effect on stage aerodynamic performance was carried out for four blade spacings (10%, 25%, 40%, and 95% of S1 axial chord) and five aspect ratios (1.0, 1.6, 3, 4, and 5). The results pointed out how unsteady interaction between blade rows affects stage operation, in terms of pressure and flow angle distributions, as well as of secondary flows development. The combined effect of these aspects in determining the stage efficiency is investigated and discussed in detail.

Commentary by Dr. Valentin Fuster
J. Eng. Gas Turbines Power. 2017;139(5):052604-052604-9. doi:10.1115/1.4035208.

This paper describes an innovative, three-day, turbomachinery research project for Japanese and British high-school students. The project is structured using modern teaching theories that encourage student curiosity and creativity. The experience develops teamwork and communication and helps to break down the cultural and linguistic barriers between students from different countries and backgrounds. The approach provides a framework for other hands-on research projects that aim to inspire young students to undertake a career in engineering. The project is part of the Clifton Scientific Trust's annual UK–Japan Young Scientist Workshop Programme. This work focuses on compressor design for jet engines and gas turbines. It includes lectures introducing students to turbomachinery concepts, a computational design study of a compressor blade section, experimental tests with a low-speed cascade, and tutorials in data analysis and aerodynamic theory. The project also makes use of 3D printing technology, so that students go through the full engineering design process, from theory, through design, to practical experimental testing. Alongside the academic aims, students learn what it is like to study engineering at university, discover how to work effectively in a multinational team, and experience a real engineering problem. Despite a lack of background in fluid dynamics and the limited time available, the lab work and end-of-project presentation show how far young students can be stretched when they are motivated by an interesting problem.

Commentary by Dr. Valentin Fuster

Research Papers: Internal Combustion Engines

J. Eng. Gas Turbines Power. 2017;139(5):052801-052801-7. doi:10.1115/1.4035227.

Valid and quick phase synchronization is essential to diesel engine operation, with impacts on the starting time and emissions. This paper presents an innovative wheel shape of camshaft with four long teeth and four short teeth, which plays a fundamental role in the proposed phase synchronization strategies. A dynamic validity-check algorithm of camshaft and crankshaft sensor signals is developed as a premise. Specifically, the segment pattern matching method is used in the rapid phase synchronization strategy that is applicable given valid camshaft and crankshaft signals. A combination of trial injections and the pattern matching method is applied when either the camshaft or crankshaft sensor signal is missing in the limp home mode. Bench test results showed that phase synchronization could be realized within 180 crankshaft angles, and that the phase synchronization strategies in the limp home mode were feasible.

Commentary by Dr. Valentin Fuster

Technical Brief

J. Eng. Gas Turbines Power. 2016;139(5):054501-054501-4. doi:10.1115/1.4035049.

Three-dimensional solid element models often with a great number of degrees-of-freedom (DOFs) are now widely used for rotor dynamic analysis. While without reduction, it will cost considerable calculating resources and time to solve the equations of motion, especially when Monte Carlo simulation (MCS) is needed for stochastic analysis. To improve the analysis efficiency, the DOFs are partly reduced to modal spaces, and the stochastic results (critical speeds or unbalance response) are expanded to polynomial spaces. First, a reduced rotor model is got by component mode synthesis (CMS), and the stochastic results are expanded by polynomial chaos basis with unknown coefficients. Then, the reduced rotor model is used to calculate the sample results to obtain the coefficients. At last, the expressions of the result by polynomial chaos basis are used as surrogate models for MCS. An aero-engine rotor system with uncertain parameters is analyzed. The accuracy of the method is validated by direct MCS, and the high efficiency makes it possible for stochastic dynamic analysis of complex engine rotor systems modeled by 3D solid element.

Commentary by Dr. Valentin Fuster

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