bluuebunny/crossref_metadata_2025 huggingface dataset:
Near and Far Field Noise Decay From a Quadcopter Propeller With and Without a Leading Edge Notch
Authors: Van Treuren Kenneth , Wisniewski Charles , Cinnamon Emily
Published: August 2018

Electric propulsion is being considered for a wide range of airframes from large commercial transports to the small Unmanned Aerial Systems (UASs). These electric systems, especially for small fixed wing UASs and quadcopters, need to be both efficient and quiet if they are to operate in an urban/populated environment or used in an Intelligence, Surveillance, and Reconnaissance (ISR) scenario.

A propeller test facility was developed to record propeller performance and sound generation in the near field behind UAS propellers. The question of defining near and far field noise was studied by characterizing sound decay with distance from a UAS propeller. Defining near and far field noise is a subject that is not addressed well in the literature. Far field noise generally follows the 1/r decay rate and near field does not. Behind the propeller there are other flow field interactions that also change the decay rate, which this study illustrates. The data presented in this paper shows the difficulty in measuring sound around a UAS propeller and begins to resolve this topic. Previous UAS propeller design work by the authors resulted in propellers that were quieter in the near field and at the same time more efficient. Their studies showed RPM and tip vortex formation both contribute significantly to propeller sound generation. Disrupting the tip vortex formation should decrease the noise being generated. The current work extends these initial findings and examines the noise generation of a stock quadcopter propeller from a DJI Phantom 2 platform. One inch aft of the plane of rotation, this propeller, a 9.4 × 5.0, has a peak sound pressure level (SPL) of approximately 118 dBA under normal static operation producing 0.7 lbf of thrust at approximately 5900 RPM. Modifications were made to four stock propellers by cutting a notch perpendicular to the leading edge of the propeller at the 0.75 r/R and 0.87 r/R locations. The notches were of different depths and widths. Of the modifications, three of the configurations did not noticeably decrease the sound. However; the final configuration reduced the peak near field SPL to 111 dBA, a 6% reduction in dBA over the stock configuration corresponding to a greater than 50% reduction in sound generation. Smoke visualization confirms that a notch located at 0.87 r/R effectively disrupts the tip vortex formation, causing the tip vortices to dissipate much earlier than the stock propeller without the notch. Examining the noise frequency spectrums associated with both the stock and the modified propeller also confirm that the notch changes the magnitude and frequency distribution of the sound being generated.
Axial Force Balance of Supercritical CO2 Radial Inflow Turbine Impeller Through Backface Cavity Design
Authors: Ma Can , Qiu Zhiqiang , Gou Jinlan , Wu Jun , Zhao Zhenxing , Wang Wei
Published: August 2018

The supercritical CO2-based power cycle is very promising for its potentially higher efficiency and compactness compared to steam-based power cycle. Turbine is the critical component in the supercritical CO2-based cycle which delivers the power. Compared to the gas turbine or steam turbine of similar power output, the size of the supercritical CO2 radial turbine is much smaller and the axial force on the impeller is much larger. The load on the thrust bearing could be too heavy for long-term safe operation. Therefore, it is necessary to balance the axial force on the impeller through aerodynamic design to reduce the load on the thrust bearing. The impeller backface design with radial pump-out vanes proves to be an effective design to reduce the axial force on the impeller of radial turbomachinery, which is widely used in the pump industry. This work investigates the impeller backface cavity flow of a supercritical CO2 radial turbine and the application of the pump-out vanes to the impeller through computational fluid dynamics simulations. Design variations of the pump-out vane are presented and their performance variations are discussed from the view of viscous compressible fluid, instead of the commonly assumed inviscid incompressible fluid in the pump industry.
Optimization Analysis of Combined Heat and Power Plant of Multistage Gas Turbine for Marine Applications
Authors: Wang Zhitao , Lei Haoda , Li Yi-Guang , Li Shuying , Wang Weitian
Published: August 2018

Nowadays, the rising demand for energy and serious environmental pollution become the motive to improve the energy structure, saving energy and optimize energy utilization. Based on a gas turbine, a marine multistage gas turbine combined heat and power (CHP) structure is proposed. The CHP system includes the top gas turbine Brayton cycle, the intermediate water Rankine cycle (WRC) and the bottom organic Rankine cycle (ORC). According to the method of screening organic Rankine cycle refrigerant to select the appropriate organic working fluids, and their physical characteristics are described. Based on the modular modelling method, the 3-stage CHP system is established. In order to more effectively absorb low temperature waste heat, three different kinds of 3-stage CHP structures were designed to recover the heat in the exhaust gas from the heat recover steam generator (HRSG). The thermodynamic model of the combined heat and power system of marine multistage gas turbine was used to simulate the performance of three different types of 3-stage CHP structures, the optimal 3-stage CHP structure was selected by comparing and analyzing the simulation results. Based on the simulation results of the design point, it is found that the introduction of the optimal 3-stage CHP structure can increase the power output by about 8.5% and improve the cycle thermal efficiency by about 4.32% compared with a conventional 2-stage CHP cycle where only gas turbine topping cycle and water Rankine bottoming cycle are included.
The Use of Blended Blade and End Wall in Compressor Cascade: Optimization Design and Flow Mechanism
Authors: Li Jiabin , Ji Lucheng , Yi Weilin
Published: August 2018

Nowadays, the flow field at the compressor is more and more complex with the increasing of the aerodynamic loading. The complex flow in the endwall regions is thus key to aerodynamic blockage, loss production, and finally its performance deterioration. The design of Blended Blade and End Wall (BBEW) contouring technology had been proved to be useful in delaying, reducing, and eliminating the corner separation in the compressor. The BBEW technology can adjust the dihedral angle between the suction and the endwall in 30% of the spanwise easily, which is different with the fillet. However, the design of the BBEW always relies on the experiences of the designers, and the effective design results cannot be the optimal result. This paper presents an optimization design method for the BBEW technology, and analyses the flow mechanism of the BBEW design.

Firstly, the parameters for the BBEW design is simplified as two, one is the maximum blended width, the other is the axial position of the maximum blended width. The optimal result can be obtained through the response surface method.

Secondly, based on the optimization method, this paper make an optimization BBEW design at the suction side of a NACA65 linear compressor cascade with the turning angle 42 degrees. The numerical results show that the optimal BBEW design can eliminate the boundary layer separation at the corner intersection region, and reduce the suction side separation. When the incidence angle is 0 degrees, the BBEW technology can reduce the total pressure loss coefficient by 5%, and reduce the aerodynamic blockage coefficient by 14%. The aerodynamic performance of the cascade shows a more obvious improvement with the BBEW design at a larger incidence. The total pressure loss coefficient of the cascade is reduced by 20% at 15 degrees incidence.

The numerical study shows that the design with the BBEW can control the axial development of the dihedral angle between the suction side and the endwall, which can eliminate the boundary layer separation at the corner intersection region. What’s more, the BBEW technology can produce a pressure gradient at the axial position of the maximum blended width, and value of the pressure gradient in proportion to the maximum blended width. This pressure gradient enhance the kinetic energy of the low energy fluid at the endwall region, which is consist of the secondary cross flow, thus elevating the capability to withstand the adverse pressure gradient, and improve the suction side separation around the trailing edge.
Investigation on the Leading Edge Film Cooling of Counter-Inclined Cylindrical and Laid-Back Holes With and Without Impingement: Part II — Heat Transfer Coefficient
Authors: Wang Rui-dong , Liu Cun-liang , Liu Hai-yong , Zhu Hui-ren , Guo Qi-ling , Gao Chao
Published: August 2018

Heat transfer of the counter-inclined cylindrical and laid-back holes with and without impingement on the turbine vane leading edge model are investigated in this paper. To obtain the film cooling effectiveness and heat transfer coefficient, transient temperature measurement technique on complete surface based on double thermochromic liquid crystals is used in this research. A semi-cylinder model is used to model the vane leading edge which is arranged with two rows of holes. Four test models are measured under four blowing ratios including cylindrical film holes with and without impingement tube structure, laid-back film holes with and without impingement tube structure. This is the second part of a two-part paper, the first part paper GT2018-76061 focuses on film cooling effectiveness and this study will focus on heat transfer. Contours of surface heat transfer coefficient and laterally averaged result are presented in this paper. The result shows that the heat transfer coefficient on the surface of the leading edge is enhanced with the increase of blowing ratio for same structure. The shape of the high heat transfer coefficient region gradually inclines to span-wise direction as the blowing ratio increases. Heat transfer coefficient in the region where the jet core flows through is relatively lower, while in the jet edge region the heat transfer coefficient is relatively higher. Compared with cylindrical hole, laid-back holes give higher heat transfer coefficient. Meanwhile, the introduction of impingement also makes heat transfer coefficient higher compared with cross flow air intake. It is found that the heat transfer of the combination of laid-back hole and impingement tube can be very high under large blowing ratio which should get attention in the design process.
An Experimental Investigation of Full-Coverage Film Cooling Characteristics of a Turbine Guide Vane
Authors: Wu Jin , Zhang Li , Cheng Li-jian , Jiang Ru , Fu Zhong-yi , Zhu Hui-ren
Published: August 2018

This paper researches on the effects of Reynolds number and mass flow ratio on the film cooling characteristics at high turbulence intensity (Tu = 15%). The experiment adopted an actual three-dimensional twisted vane and presents the film cooling characteristics on full-coverage film surface in a two-passage, linear cascade. The cooling effectiveness and heat transfer coefficient of the vane’s whole surface were obtained by using transient liquid crystal measurement technique. The transient liquid crystal is SPN/R35C1W, whose bandwidth is 2°C. There are fifteen rows of film cooling holes which have different diameter, injection angle and yaw angle. The secondary flow was supplied by two cavities. The front cavity supplied the secondary flow to thirteen rows of film cooling holes that were arranged in the suction surface, the leading edge and the front half of the pressure surface. The rear cavity supplied the secondary flow to the rear half of pressure surface which included two rows of film cooling holes. The investigated parameters are Reynolds number of 1 × 105, 1.3 × 105 and 1.6 × 105 and the mass flow ratio of MFR = 5.5%∼12.5% (6 cases). The data recorded in the experiment was analyzed with MATLAB.

Results show that the combined effects of mass flow ratio and channel vortex are the maintain reasons that influence the distribution of cooling effectiveness in the contour. Increasing the mass flow ratio can improve the film cooling effectiveness on leading edge and pressure surface, while that presents complex rule on suction surface. Increasing the Reynolds number can improve the heat transfer coefficient at the same mass flow ratio. When increasing the mass flow ratio, the heat transfer coefficient increases on leading edge and pressure surface at Re = 1.6 × 105. However, the decreases at film hole outlet region on the suction side, and not obviously changes at the film hole downstream region.
Effect of Buoyancy and Density Ratio on Heat Transfer in a Smooth Cooling Channel of a Gas Turbine Blade
Authors: Saravani Mandana S. , Beyhaghi Saman , Amano Ryoichi S.
Published: August 2018

The present work investigates the effects of buoyancy and density ratio on the thermal performance of a rotating two-pass square channel. The U-bend configuration with smooth walls is selected for this study. The channel has a square cross-section with a hydraulic diameter of 5.08 cm (2 inches). The lengths of the first and second passes are 514 mm and 460 mm, respectively. The turbulent flow enters the channel with Reynolds numbers of up to 34,000. The rotational speed varies from 0 to 600 rpm with the rotational numbers up to 0.75. For this study, two approaches are considered for tracking the buoyancy effect on heat transfer. In the first case, the density ratio is set constant, and the rotational speed is varied. In the second case, the density ratio is changed in the stationary case, and the effect of density ratio is discussed. The range of Buoyancy number along the channel is 0–6. The objective is to investigate the impact of Buoyancy forces on a broader range of rotation number (0–0.75) and Buoyancy number scales (0–6), and their combined effects on heat transfer coefficient for a channel with aspect ratio of 1:1. Several computational fluid dynamics (CFD) simulation are carried out for this study, and some of the results are validated against experimental data.
Impact Test for the Leading Edge of CMC Vane Based on Actual Aircraft Engine Field Data
Authors: Obuchi Kenro , Watanabe Fumiaki , Kuroki Hiroshi , Yagi Hiroyuki , Arai Kazuyoshi
Published: August 2018

Ceramic matrix composites (CMCs) have lower density and a higher service temperature limit than nickel based alloys which have been used for turbine components of aircraft engines. These properties of CMCs have the potential to reduce the weight of turbine components and improve turbine thermal efficiency with a higher turbine inlet temperature (TIT). One of the technical issues of the CMC turbine vane is a relatively lower impact resistance than nickel based alloy turbine vanes. There are various previous works about impact resistance of CMCs, but there is little work that assumed actual engine conditions.

The objective of this work was to verify the resistance of SiC/SiC CMC turbine vane to the impact phenomena that occur in the actual aircraft engine. The field damage survey was conducted on actual metal turbine vanes of commercial engines overhauled in IHI. The survey made it clear that the typical damage was less-than-0.127-mm-dent at the leading edge. In addition, the dropped weight impact test using the actual turbine airfoil which is made from a nickel based alloy was conducted at ambient temperature. The amount of energy required to make the dent of a certain size that was observed in actual metal turbine vanes was estimated. Then, the dropped weight impact test using the CMC test piece with a leading edge shape was conducted at the impact energy estimated by the metal turbine airfoil. The results showed that the failure mode of the CMC test piece was local damage with dents of a certain size and not a catastrophic failure mode.

From this work, the damage to be assumed on CMC vane in actual aircraft engines was identified. As a future task, the effect of the damage to the fatigue capability of CMC turbine vanes needs to be investigated.
Large Eddy Simulation of a Two-Phase Staged Swirling Burner Using an Euler-Lagrange Approach: Validation of the Injection Strategy
Authors: Cunha Caldeira Mesquita Léo , Vié Aymeric , Ducruix Sébastien
Published: August 2018

A two-staged swirling burner is numerically simulated using Large Eddy Simulation (LES). This combustor uses two types of injection: a multipoint system that consists in 10 holes in a crossflow configuration, and a pilot system that uses a pressure-swirl atomizer. The relation between the rate of fuel injected from each injection system was found to be related with flame shape transition and hysteresis phenomena[4]. Also, the pilot spray was found to have a major role on these transitons, so it is of paramount importance to correctly reproduce its behavior on the numerical modeling, if one is interested in simulating these flame bifurcations. To describe the spray, a point-droplet approximation is used in a Lagrangian framework with the FIM-UR model [1], that has already proven its accuracy for several configurations. However, in this application it fails to reproduce the droplet size distribution, especially in the Central Recirculation Zone (CRZ), as it uses an arbitrary expression to impose the spray opening limits (which are not input parameters). In the present work, the input parameters of the FIM-UR model are modified to enable the recovery of the right droplet size distribution and improve the description of the liquid velocity field, resulting in a better numerical representation of the experimental results, essential for further studies.
Numerical Investigation of Diffuser Rotating Stall and System Instability in a Centrifugal Compressor With Vaned Diffuser
Authors: Zhao Yang , Xi Guang , Zhao Jiayi
Published: August 2018

The operating range of a centrifugal compressor is often limited by the occurrence of the flow instability, such as diffuser rotating stall or system surge. In the paper, the unsteady numerical simulations are performed on a low-speed centrifugal compressor to investigate the characteristic of the rotating stall in the vaned diffuser. And also, the developed model of lumped parameter is used to predict the system instability.

The flow field in the diffuser is firstly investigated at near stall condition. It is found that the leading-edge vortex and the secondary flow induce the hub-corner separation at the suction side of the diffuser blade. When the mass flow rate is reduced gradually, the fore part of the volute turns to act as a diffuser from a nozzle. Under the influence of the asymmetry induced by the volute, the hub-corner separation firstly develops into rotating stall in the passage with the lowest mass flow rate when at critical stall point. And then the diffuser rotating stall propagates along the circumferential direction at about 7% of the impeller speed.

And also, the model of lumped parameter considering the effect of rotating stall is developed to analyze the system instability of mild surge. The predicted vibration frequency is within 5.8% of the measurement and the predicted transient process in mild surge matches well with the measurement. With different volume of the compressed air, the transient compressor characteristic tends to be stabilized or oscillates in a cycle along the counter-clockwise with different magnitude.
Effects of a Diverging Cup on Swirl Number, Flow Pattern and Topology of Premixed Flames
Authors: Degeneve A. , Jourdaine P. , Mirat C. , Caudal J. , Vicquelin R. , Schuller T.
Published: August 2018

Impact of the diverging cup angle of a swirling injector on the flow pattern and stabilization of technically premixed flames is investigated both theoretically and experimentally with the help of laser diagnostics. Recirculation enhancement with a lower position of the internal recirculation zone and a flame leading edge protruding further upstream in the swirled flow are observed as the injector nozzle cup angle is increased. A theoretical analysis is carried out to examine if this could be explained by changes of the swirl level as the diffuser cup angle is varied. It is shown that pressure effects need in this case to be taken into account in the swirl number definition and expressions for its variation through a diffuser are derived. They indicate that changes of the swirl level including or not the pressure contribution to the axial momentum flux cannot explain the changes observed of the flow and flame patterns in the experiments. The swirl number without the pressure term, designated as pressure-less swirl, is then determined experimentally for a set of diffusers with increasing quarl angles under non-reacting conditions and the values found corroborate the predictions. It is finally shown that the decline of axial velocity and the rise of adverse axial pressure gradient, both due to the cross section area change through the diffuser cup, are the dominant effects that control the leading edge position of the internal recirculation zone of the swirled flow. This in turn is used to develop a model for the change of this position as the quarl angle varies that shows good agreement with experiments.
Application of Kriging Metamodels to the Automated Start Value Generation for Gas Turbine Performance Simulations
Authors: Schmeink Jens , Becker Richard-Gregor
Published: August 2018

This paper presents a process for automated start value generation for gas turbine performance simulations using Kriging metamodels. The metamodels are trained on a small number of pre-selected operating points in the flight envelope. Predictions of the trained metamodels are used as initialization parameters for subsequent performance simulations of arbitrary operating points in order to increase robustness and computational speed of the numerical process. Different approaches for the selection of the training points are evaluated. A comparison to the classical approach of table-based initialization is carried out to highlight the advantages and disadvantages of the new methodology.

Furthermore, the inclusion of supplementary operating points into the training sample of the metamodels is analyzed. Depending on certain criteria, such as the difference of prediction and simulation result, operating points are included into the sample and a retraining of the metamodels is performed. Simulations using the retrained models as guess value generators are compared to the previous Kriging approach. The advantages and disadvantages of the retraining approach are discussed.
Numerical Study on Analogy Principle of Overall Cooling Effectiveness in Engine and Laboratory Condition
Authors: Xie Gang , Liu Cun-liang , Ye Lin , Wang Rui
Published: August 2018

The overall cooling effectiveness, which represents the distribution of dimensionless temperature on gas turbines surface, is an important parameter for conjugate heat transfer analysis of gas turbines. Generally, it is difficult to measure the overall cooling effectiveness in engine condition. However, the overall cooling effectiveness can be measured in the laboratory by matching the appropriate parameters to those of the actual turbine blade. Thus, it is important to evaluate the key parameters of matching methods. In this paper, the effects of adiabatic film effectiveness and Biot number on the overall cooling effectiveness were investigated with an impingement/effusion model by numerical simulation, in which 3-D steady RANS approach with the k–ω SST turbulence model were used. The tested plate had 8 cylinder hole rows of 30 degree inclined angle, and the internal cooling employed staggered array jet impingements. The matching performance was evaluated by comparing the results in both typical engine condition and laboratory condition. The analogy principles were discussed in detail, the results showed that the overall cooling effectiveness can be matched by using different matching principles in different lab condition. The theoretical analysis was verified by numerical results. The distribution and values of overall cooling effectiveness can be matched well between engine condition and lab condition by matching both temperature ratio, mainstream side Biot number and blowing ratio. If the temperature ratio is mismatched, the momentum flux ratio will be an important parameter for overall cooling effectiveness. Matching momentum flux ratio will reduce the difference of the adiabatic cooling effectiveness and heat transfer ratio between engine condition and laboratory condition.
Detailed Analysis of Boundary Layer Control by Fluidic Oscillators on a Highly Loaded Profile
Authors: Kurz Julia , Niehuis Reinhard
Published: August 2018

One application method of active flow control is the exploitation of the interaction between transition and flow separation on a profile. As turbulent flows are able to withstand higher adverse pressure gradients the enforcement of the transition process can be utilized to prevent or to reduce flow separation. This paper focuses on gaining a better understanding of high frequency active flow control (AFC) by fluidic oscillators and its influence on the transition process for a separated boundary layer. Flow control is applied on a highly loaded turbine exit case (TEC) profile which was in particular designed for this application. The profile is investigated in the high-speed cascade wind tunnel at the Bundeswehr University Munich. Significant loss reduction by AFC could be observed by total pressure loss determination in the low Reynolds number regime. In order to gain a better understanding of development of the suction side boundary layer, several boundary layer profiles are determined by hot-wire measurements at six axial positions on the suction side of the profile. Differences between the boundary layer development and the extent of the separation can be detected. Furthermore, a stability analysis of the boundary layer upstream of separation is conducted and compared to the measured frequency spectra.
Design Characteristics of an Aerodynamic Foil Bearing With Adaptable Bore Clearance
Authors: Sadri Hossein , Schlums Henning , Sinapius Michael
Published: August 2018

Aerodynamic foil bearings are suitable to support light, high-speed rotors under extreme operating conditions such as very low or very high temperatures, e.g. in cooling turbines, small gas turbines or exhaust gas turbochargers. The required bearing load capacity is generated by an aerodynamic pressure build-up in the corresponding lubrication gap. Due to the high dependence of the bearing performance on the bore geometry, the rotordynamic behavior (e.g. bearing stability) and static properties (e.g. load capacity) as a function of radial clearance and hydrodynamic preload are one of the main points of interest in recent studies. The outcome of both the experimental and the numerical investigations show the advantages and disadvantages of the various configurations of the bearing bore in different operating conditions. These observations lead to the basic idea of an adaptive air foil bearing (AAFB) in which, depending on the operating conditions, the bearing bore contour is changed by means of piezoelectric actuators applied to the compliant supporting shell. Similar to other shape morphing approaches, optimization with regard to various components of the mechanism is the next step in the design process after targeting the design pattern. This paper concentrates on an AAFB as an efficient approach to actively shape the contour of the bore clearance in a 3-pad bearing. Numerous FEM analyses of a functional model for an AAFB in addition to the experimental efforts reveal the main concerns of the design. Finally, the result of this study is a working graph for the AAFB under various loading conditions while operating with different input voltages of the actuators.
Performance of a Semi-Closed Oxy-Fuel Combustion Combined Cycle (SCOC-CC) With an Air Separation Unit (ASU)
Authors: Sammak Majed , Thern Marcus , Genrup Magnus
Published: August 2018

The objective of this paper is to evaluate the performance of a semi-closed oxy-fuel combustion combined cycle (SCOC-CC) and its power penalties. The power penalties are associated with CO2 compression and high-pressure oxygen production in the air separation unit (ASU). The paper discusses three different methods for high pressure oxygen (O2) production. Method 1 is producing O2 directly at high pressure by compressing the air before the air separation takes place. Method 2 is producing O2 at low pressure and then compressing the separated O2 to the desired pressure with a compressor. Method 3 is alike the second method, except that the separated liquid O2 is pressurized with a liquid oxygen pump to the desired pressure.

The studied SCOC-CC is a dual-pressure level steam cycle due to its comparable efficiency with three pressure level steam cycle and less complexity.

The SCOC-CC, ASU and CO2 compression train are modeled with the commercial heat and mass balance software IPSEpro. The paper analyzed the SCOC-CC performance at different combustion outlet temperatures and pressure ratios. The combustion outlet temperature (COT) varied from 1200 °C to 1550 °C and the pressure ratio varied from 25 to 45.

The study is concerned with mid-sized SCOC-CC with a net power output 100 MW. The calculations were performed at the selected design point which was at 1400°C and pressure ratio at 37. The calculated power consumption of the O2 separation at a purity of 95 % was 719 kJ/kgO2. The power consumption for pressurizing the separated O2 (method 2) was 345 kJ/kgO2 whereas it was 4.4 kJ/kgO2 for pumping liquid O2 to the required pressure (method 3). The calculated power consumption for pressurizing and pumping the CO2-enriched stream was 323 kJ/kgCO2.

The SCOC-CC gross efficiency was 57.6 %. The SCOC-CC net efficiency at method 2 for air separation was 46.7 %. The gross efficiency was reduced by 9 % due to ASU and other 2 % due to CO2 compression. The SCOC-CC net efficiency at method 3 of the air separation was 49.6 %. The ASU reduced the gross efficiency by 6 % and additional 2 % by CO2 compression. Using method 3 for air separation gave a 3 % gain in cycle efficiency.
Influence of Turbocharger Turbine Blade Geometry on Vibratory Blade Stresses
Authors: Naik Pavan , Lehmayr Bernhard , Homeier Stefan , Klaus Michael , Vogt Damian M.
Published: August 2018

In this paper, a method to influence the vibratory blade stresses of mixed flow turbocharger turbine blade by varying the local blade thickness in spanwise direction is presented. Such variations have an influence on both the static and the vibratory stresses and therefore can be used for optimizing components with respect to High-Cycle Fatigue (HCF) tolerance. Two typical cyclic loadings that are of concern to turbocharger manufacturers have been taken into account. These loadings arise from the centrifugal forces and from blade vibrations.

The objective of optimization in this study is to minimize combined effects of centrifugal and vibratory stresses on turbine blade HCF and moment of inertia. Here, the conventional turbine blade design with trapezoidal thickness profile is taken as baseline design. The thicknesses are varied at four span-wise equally spaced planes and three stream-wise planes to observe their effects on static and vibratory stresses. The summation of both the stresses is referred to as combined stress. In order to ensure comparability among the studied design variants, a generic and constant excitation order dependent pressure field is used at a specific location on blade.

The results show that the locations of static and vibratory stresses, and hence the magnitude of the combined stresses, can be influenced by varying the blade thicknesses while maintaining the same eigenfrequencies. By shifting the maximum vibratory stresses farther away from the maximum static stresses, the combined stresses can be reduced considerably, which leads to improved HCF tolerance.
Effect of Solidity on Non-Axisymmetric Endwall Contouring Performance in Compressor Linear Cascades
Authors: Xiwu Liu , Donghai Jin , Xingmin Gui , Xiaoheng Liu , Hanwen Guo
Published: August 2018

This paper presents both the computational and experimental results to assess the effectiveness of non-axisymmetric endwall contouring in linear cascades under different solidities. Endwalls were designed by geometric scaling of a prior optimized endwall. The results show that the total pressure loss can be reduced by the contoured endwall (CEW) under different solidities. The mechanism of the improvement of CEW is that the adverse pressure gradient (APG) has been reduced mainly through the groove configuration near the leading edge of the suction surface. Besides, the cross-passage pressure gradient (CPG) has also been reduced, which has the benefits to further suppress the corner separation. Moreover, there is an optimum range of the solidity for the CEW. For a lower solidity, the performance of the CEW at +7 degree incidence presents a rapid deterioration, due to the risk of flow separation near the mid-span, for a higher solidity, the profile loss can be more dominant.
Modeling Deposition in Turbine Cooling Passages With Temperature Dependent Adhesion and Mesh Morphing
Authors: Bowen Christopher P. , Libertowski Nathan D. , Mortazavi Mehdi , Bons Jeffrey P.
Published: August 2018

The role of temperature on deposition in gas turbine internal cooling geometries is investigated. Single impingement cones are developed by an oversized (6 mm) impinging jet over a range of temperatures and flow velocities using 0–5 μm ARD. Cone size was found to increase with increasing temperature and decrease with increasing velocity. Capture efficiency and cone angle effects are presented, and packing factor data is used as a metric to determine if the contact area (Acont) for adhesion explains the trends seen with temperature. It is systematically demonstrated that the surface free energy (γ) is likely a first order function of temperature in internal deposition for the range of temperatures investigated. Candidate physical mechanisms that may cause increased adhesive force at elevated temperatures are identified. Temperature dependent adhesion is added to the OSU Deposition Model which is then used with a simplified morphing approach to match temperature induced blockage patterns in a vane leading edge cooling experiment. This process is improved upon using a full mesh morphing routine and matching two of the experimental deposition cones at varied flow temperatures. The added fidelity that mesh morphing affords is demonstrated.
Circumferential Mode Analysis of Axial Compressor Tip Flow Using Fourier Transform and Proper Orthogonal Decomposition
Authors: Yao Dan , Tian Jie , Wu Yadong , Ouyang Hua
Published: August 2018

Rotating instability (RI) of a single-stage axial compressor was studied by both numerical and experimental methods. A circumferential mode decomposition method based on spatial Fourier transform was used to analyze the circumferential pressure distribution of the tip flow. Circumferential mode characteristics were captured both on blade passing frequency (BPF) and rotational instability frequency (RIF) under several flow conditions. The characteristic spectrum of RI with broadband hump existed in a large range of flow conditions. Both frequency range and dominant circumferential mode number decreased with flow rate, while circumferential angular velocity of RI increased at the same time. On the other hand, a proper orthogonal decomposition (POD) method was applied to obtain the mode component of tip flow. The feature of tip flow was analyzed with the help of POD mode vector and mode amplitude. The influence of the decrease on the spatial monitor points in POD method was analyzed using CFD data to analysis the potential error from experimental results. It is expected to deeply understand the mechanism of the rotating instability and rotor-stator interaction phenomenon by spatial FT and POD methods in this study.
Measurement of Unsteady Pressure Field in a Turbocharger Compressor Using Pressure-Sensitive Paint
Authors: Kitamura Takaaki , Kameda Masaharu , Watanabe Wataru , Horimoto Kazutaka , Akimoto Kenta , Akahori Akihito
Published: August 2018

It is necessary to quantitatively resolve the flow field generated in compressors at various operating conditions including the surge. In this study, fast-responding pressure-sensitive-paint (PSP) and temperature-sensitive-paint (TSP) was applied to measurement of steady and unsteady pressure and temperature fields in a turbocharger compressor. The surface pressure of a compressor impeller and the diffuser downstream of the impeller was measured under the rotating speed of 133,000 rpm with various mass flow rate of the inlet. Blue LED strobes were used as the excitation light source for PSP/TSP. The luminescence from PSP/TSP were captured by a high-speed camera. These instruments were controlled using the signal synchronized with a fixed rotation angle of the impeller, which was detected using a gap sensor installed near the impeller. The pressure was calculated from the time-series of the PSP signal. The output of PSP/TSP was compared with the signal from semiconductor pressure transducers and the thermal image taken by an infrared camera. For the analysis of the diffuser, the power spectra of the pressure field were obtained using the fast Fourier transform of the time-series data of the pressure fluctuation. The pressure field on the impeller was successfully obtained with the aid of adequate temperature compensation by TSP. It was confirmed that the time-averaged pressure on the impeller blades increased as the inlet mass flow rate was decreased. Large-amplitude points were distributed on the extension line of the impeller blades in the diffuser when the surge occurs.
Investigation on Film Cooling Performance of the Compound Hole and Y-Shaped Hole Configurations With the Cross-Flow Coolant Channel
Authors: Li Lin , Liu Cun-liang , Liu Hai-yong , Zhu Hui-ren , Luo Jian-xia
Published: August 2018

Film cooling performance of the compound hole and the Y-shaped hole configurations with the cross-flow coolant channel is investigated experimentally and numerically in this paper. The Reynolds numbers of coolant flow are fixed as 50000 and 100000 respectively. The film cooling effectiveness and heat transfer coefficient were measured by the transient liquid crystal measurement technique under three blowing ratios of 0.5, 1.0 and 2.0 respectively. And flow resistance measurements were also performed to obtain the discharge coefficient of the two film hole configurations. Numerical simulations with Reynolds-averaged Navier-Stokes (RANS) method were performed to explain the experiment results. The results show that the distribution feature of film cooling effectiveness for the compound hole and Y-shaped hole configurations is different. The film cooling effectiveness of the Y-shaped hole configuration is higher than that of the compound hole, and the film spanwise coverage is larger than that of the compound hole under all cases. For two film hole configurations, the heat transfer coefficient increases with the increase of the cross-flow Reynolds number and the blowing ratio. The heat transfer coefficient of the compound hole and Y-shaped hole configurations is close to each other under small cross-flow Reynolds number. However, under large cross-flow Reynolds number, the compound hole configuration has much higher heat transfer coefficient. The discharge coefficient increases gradually with the rising blowing ratio, then tend to a fixed value. Under the condition of the small cross-flow Reynolds number, the discharge coefficient of two film hole configurations is high. The discharge coefficient of the Y-shaped hole configuration is a little higher than that of the compound hole configuration under the condition of the large blowing ratio.
Experimental Investigation of Ignition and LBO Characteristics of SPP Injector: The Effect of Pilot Stage Air Split Ratio
Authors: Yang Jinhu , Liu Cunxi , Wu Haowei , Liu Fuqiang , Mu Yong , Hu Chunyan , Xu Gang
Published: August 2018

The influence of PASR (Pilot stage Air Split Ratio) on the ignition and LBO (Lean Blow Out) performances is experimentally investigated for an SPP (Stratified Partially Premixed) injector in this paper. The pilot stage of the SPP injector comprises two axial air swirlers as well as an air blast prefilm atomizer for pilot fuel preparation. It is believed that the variation of the air split ratio between the outer swirler and the inner swirler of the pilot stage will transform the flow structure and fuel distribution of the local flame anchoring zone, and consequently improves or deteriorates the stability of the pilot flame. The ignition and LBO characteristics were measured for PASR = 8:2, 7:3 and 6:4, and several inexplicable but interesting results are observed. In order to make out the underlying reasons for the differences of the obtained ignition and LBO data, the velocity field and spray concentration at the meridian plane were acquired experimentally with the help of optical diagnostics at isothermal conditions. It it concluded that two dominant mechanisms of flame stability exist depending on the range of the injector pressure drop (Δ Psw/P3t). At low pressure drop of the injector, the flame stability is mainly affected by the fuel distribution, however, the flow structure will play a more important role at high Δ Psw/P3t in that it can transform the local flow structures around the pilot flame root. The inherent correlations between the combustion stability and the flow structure as well as the fuel distribution are disscussed and conclusions are drawn for this research work in the end of this paper.
Investigation on Creep Characteristic of Nickel-Base Single Crystal Alloy Double Wall Cooling Specimen
Authors: Li Lei , Wan Huan , Gao Wenjing , Tong Fujuan , Sun Shouyi
Published: August 2018

At present, advanced gas turbine are designed to work at increasingly higher turbine inlet temperatures (TIT) to obtain higher thermal efficiency and power. TIT of modern aircraft engine has exceeded temperature limit of high-temperature alloys application over 600C°. So turbine blade has to be cooled in order to work at an acceptable temperature environment. Double wall cooling technology with higher cooling efficiency which is developed by the RR and GE companies has become development trends of blade cooling. In this work, a flat double wall cooling structure is introduced. On basis of creep damage model considering interaction of cavitation and degradation, the creep behaviors of double wall cooling structure are investigated. Research shows that there are stress interferences among holes and pin fin. The higher stress region locates around holes. The lower stress region locates at pin fin. The distance of impingement hole and pin fin in the analysis model affect creep behavior seriously.
Twin-Web Turbine Discs: Part 1 — Design and Analysis of Their Efficiency
Authors: Vasilyev Boris , Salnikov Anton , Semenov Artem , Magerramova Liubov
Published: August 2018

Part 1 of this work discusses the numerical results of the performance efficiency investigation conducted for twin-web (hollow) discs in terms of stress and mass reductions.

Three typical turbine discs and one small-sized turbine blisk are analysed using parameterised finite-element models and the modern methods of multi-criteria optimisation. The discs and blisk have different overall sizes and load conditions (rotation speed [rpm]), thermal state, blade load, etc.). The design of disc variants with a split in the hub is also considered. This option will help to simplify the manufacturing technology and improve design efficiency.

An analysis of the results shows that a twin-web turbine disc design permits increases in turbine disc strength as well as decreases in its mass. The extent of these advantages depends on the overall size of the disc and its rpm: a large-sized disc with a relatively low rpm shows a 4% mass decrease and a 15% increase in cyclic life; a medium-sized disc with an average rpm has an 18% mass decrease and a doubled cyclic life; and a small-sized disc with high rpm has a 30% mass decrease and a 2.8 times longer cyclic life. Additionally, a small-sized blisk with a high rpm shows a 60% mass and a doubled cyclic life.
Large Eddy Simulations of Separated Boundary Layer With Pressure Gradient and Heat Transfer
Authors: Wu Yifei , Zhang Weihao , Zou Zhengping , Chen Jiang
Published: August 2018

Separated boundary layers of the low pressure turbine blade suction surface under wall heat transfer and pressure gradient conditions are investigated using large eddy simulations (LES) in this paper. The study constructed a converging-diverging channel with a flat plate as the bottom wall, and the pressure distribution of the bottom wall is similar to that of a high lift low pressure turbine blade suction surface. The boundary layer was investigated under different heat transfer boundary conditions of the bottom wall (i.e., the adiabatic wall and the isothermal wall with the wall temperature being 0.8 times of the inflow temperature). The time-averaged flow parameters and the separation bubble characteristics were analyzed and discussed. The evolution of coherent structure diagrams of the boundary layer was also obtained to study the evolution process of the vortex. The results show that the cooled isothermal wall condition can significantly suppress the separation bubble and reduce the frequency of the large scale spanwise vortex roll-up. Under the two wall heat transfer conditions, the scale of the near wall small scale spanwise vortex is similar as well as the scale of the large scale spanwise wortex. The location of the vortex are also approximate under the two wall heat transfer conditions, but the position of the large scale spanwise vortex shedding from separated laminar boundary layers moves upstream under the cooled isothermal wall condition, and the transition process is more rapid than that of the adiabatic wall condition.
An Experimental and Numerical Study of Tip-Leakage Flows in an Idealized Turbine Tip Gap at High Mach Numbers
Authors: Passmann Maximilian , aus der Wiesche Stefan , Joos Franz
Published: August 2018

This paper presents results of a detailed investigation of turbine tip-leakage flows at high Mach numbers. The experimental work was carried out using a small blow-down wind tunnel. An idealized blade test section was used to study blade tip-clearance effects in transonic conditions. Unshrouded blade tips are considered and different tip gap heights are investigated. A high blade exit Mach number of Me = 2 was selected deliberately. While conventional transonic turbine stages generally operate at lower supersonic exit Mach numbers, the conditions are representative for ORC turbines. Both experimental and numerical results are presented in this contribution. The results indicate, that tip leakage flow under transonic conditions leads to a complex three-dimensional flow field. A strong interaction between tip gap vortex and trailing edge shocks was observed, that also had a profound effect on the base region. While no final statement on losses could be made in the present configuration, the results indicate a weakened shock system.