bluuebunny/crossref_metadata_2025_split huggingface dataset:
Design Optimization of Splitter Blades for Rocket Engine Turbopump
Authors: Torre Francesco , Konno Shinichi , Lettieri Claudio , Pini Matteo , Kawata Yutaka 
Published: August 2018

In this paper we present and validate a shape optimization framework for the design of splitter blades that extends the operative range under cavitation while maintaining the wetted performance of rocket engine turbopumps. For a target turbopump application, the optimization framework allows for independent changes to the blade angle distributions across the span and to the pitchwise position of the splitter blades while preserving the thickness distributions.
The optimization is conducted with a surrogate-based gradient method. The geometry is optimized at a fixed cavitation number corresponding to a 5% head coefficient dropoff, while constraints are imposed on the wet pump performance. It is found that this approach, coupled with the optimal design points distribution provided by the Design of Experiment method, reduces the computational cost of the optimization process by minimizing the number of multiphase calculations.
The numerical results suggest that the optimized splitter blades successfully increase the pump operative range by 2.2% and increase the head coefficient by 5.3% compared to the baseline case with non-optimized splitters. These results are corroborated by experiments conducted in a closed-loop water test facility. Several pump geometries are tested through rapid prototyping using additive manufacturing. The experimental data validate the optimization framework, demonstrating a 4.7% increase of pump operative range and a 7.6% increase in head coefficient. The calculations are used to gain insight in the physical mechanisms for the performance improvement. The analysis of the results indicates that the improved performance is due to the optimized position and shape of the splitter blades which increase the pump slip factor.
Power to Ice: A Novel Approach to Stabilize Non-Programmable Renewable by Means of Gas Turbine IACC
Authors: Ancona M. A. , Bianchi M. , Branchini L. , Brilloni A. , De Pascale A. , Melino F. , Peretto A. 
Published: August 2018

Large penetration of non-programmable energy sources, such as wind, is a challenging issue for grid operators and quick ramping fossil fuel generators. Indeed, the variability and fluctuation of renewables (i.e. rapid change in generation over relatively short time periods) is increasing the need for regulating power. In this context and according to Authors knowledge, the innovative approach proposed in this study is to use a gas turbine equipped with continuous cooling in order to stabilize the power output generation of renewable generators. The gas turbine power boost, obtained thanks to inlet air cooling, will compensate for renewable generators underproduction. On the opposite, in case of renewable generators power output surplus, the excess of power is driven to a compressor chiller device for cold water storage, then used to chill gas turbine inlet air whenever needed.
A detailed performance evaluation on the proposed system is carried out showing the influence of most important parameters (i.e. additional pressure losses introduced with the cooling device, gas turbine models and assumed derating coefficients, etc.).
A feasibility analysis of the investigated system is presented for several case studies in case of wind generators (two different gas turbine models, different wind farm nameplate capacity, different power output set points), investigating the volume of tanks necessary to stabilize wind fluctuations.
The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments
Authors: Holgate Nicholas E. , Ireland Peter T. , Romero Eduardo 
Published: August 2018

Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface.
Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.
Determination of Aerodynamic Damping at High Reduced Frequencies
Authors: Pan Minghao , Petrie-Repar Paul , Mårtensson Hans , Sun Tianrui , Gezork Tobias 
Published: August 2018

In turbomachines, forced response of blades is blade vibrations due to external aerodynamic excitations and it can lead to blade failures which can have fatal or severe economic consequences. The estimation of the level of vibration due to forced response is dependent on the determination of aerodynamic damping. The most critical cases for forced response occur at high reduced frequencies. This paper investigates the determination of aerodynamic damping at high reduced frequencies. The aerodynamic damping was calculated by a linearized Navier-Stokes flow solver with exact 3D non-reflecting boundary conditions.
The method was validated using Standard Configuration 8, a two-dimensional flat plate. Good agreement with the reference data at reduced frequency 2.0 was achieved and grid converged solutions with reduced frequency up to 16.0 were obtained. It was concluded that at least 20 cells per wavelength is required. A 3D profile was also investigated: an aeroelastic turbine rig (AETR) which is a subsonic turbine case. In the AETR case, the first bending mode with reduced frequency 2.0 was studied. The 3D acoustic modes were calculated at the far-fields and the propagating amplitude was plotted as a function of circumferential mode index and radial order. This plot identified six acoustic resonance points which included two points corresponding to the first radial modes. The aerodynamic damping as a function of nodal diameter was also calculated and plotted. There were six distinct peaks which occurred in the damping curve and these peaks correspond to the six resonance points. This demonstrates for the first time that acoustic resonances due to higher order radial acoustic modes can affect the aerodynamic damping at high reduced frequencies.
Experimental Evaluations of the Relative Contributions to Overall Effectiveness in Turbine Blade Leading Edge Cooling
Authors: Bryant Carol E. , Wiese Connor J. , Rutledge James L. , Polanka Marc D. 
Published: August 2018

Gas turbine hot gas path components are protected through a combination of internal cooling and external film cooling. The coolant typically travels through internal passageways, which may involve impingement on the internal surface of a turbine component, before being ejected as film cooling. Internal cooling effects have been studied in facilities that allow measurement of heat transfer coefficients within models of the internal cooling paths, with large heat transfer coefficients generally desirable. External film cooling is typically evaluated through measurements of the adiabatic effectiveness and its effect on the external heat transfer coefficient.
Efforts aimed at improving cooling are often focused on either only the internal cooling or the film cooling; however, the common coolant flow means the internal and external cooling schemes are linked and the coolant holes themselves provide another convective path for heat transfer to the coolant. Recently, measurements of overall cooling effectiveness using matched Biot number turbine component models allow evaluation of the nondimensional wall temperature achieved for the fully cooled component. However, the relative contributions of internal cooling, external cooling, and convection within the film cooling holes is not well understood.
Large scale, matched Biot number experiments, complemented by CFD simulations, were performed on a fully film cooled cylindrical leading edge model to evaluate the effects of various alterations in the cooling design on the overall effectiveness. The relative influence of film cooling and cooling within the holes was evaluated by selectively disabling individual holes and quantifying how the overall effectiveness changed. Several internal impingement cooling schemes in addition to a baseline case without impingement cooling were also tested. In general, impingement cooling is shown to have a negligible influence on the overall effectiveness in the showerhead region. This indicates that the cost and pressure drop penalties for implementing impingement cooling may not be compensated by an increase in thermal performance. Instead, the internal cooling provided by convection within the holes themselves was shown, along with external film cooling, to be a dominant contribution to the overall cooling effectiveness. Indeed, the numerous holes within the showerhead region impede the ability of internal surface cooling schemes to influence the outside surface temperature. The results of this research may allow improved focus of future efforts on the forms of cooling with the greatest potential to improve cooling performance.
Automated Condition Evaluation of Hot-Gas Path Components of Jet Engines Through Exhaust Jet Analysis
Authors: Hartmann Ulrich , Seume Joerg R. 
Published: August 2018

This paper determines the influence of different defective components in the hot-gas path (HGP) of a civil aircraft engine on the density distribution in the exhaust. The intention is to automate the identification of defective components inside the HGP through an analysis of the density distribution in the exhaust jet. The defects include an increased radial gap of the blades in the high-pressure turbine (HPT), and a reduction of the film cooling air mass flow in the first stage of the HPT. In addition, several combinations of both defects are simulated. In the present paper the exhaust density distributions are generated numerically using CFD simulations of the HGP. The density distribution in the exhaust jet is reconstructed with synthetic Background-Oriented Schlieren (BOS) measurements and automatically analyzed. The methodology for the automated defect detection consists of two algorithms, a Support Vector Machine (SVM) algorithm to automatically classify each measurement into a corresponding defect or reference class and an outlier detection algorithm to detect variations from the reference state — without assignment.
It is shown that BOS provides a sufficient reconstruction quality to automatically detect defective HGP components with a SVM algorithm. It is possible to automatically detect both defects, even when they occur at the same time. For this purpose, different features were calculated to isolate the influence of each defect on the density distribution. The outlier detection algorithm allows for an automated detection of variations in the density distribution compared to the reference state without any previous knowledge of the influence of the defects on the density distributions during the training procedure. With this algorithm it is possible to detect unknown or new defects which have not been observed or regarded yet. These results strengthen the hypothesis, that an automated detection of defects in jet engines prior to the disassembly is possible.
Experimental Validation of Forced Response Methods in a Multi-Stage Axial Turbine
Authors: Hauptmann Thomas , Meinzer Christopher E. , Seume Joerg R. 
Published: August 2018

Depending on the in service condition of jet engines, turbine blades may have to be replaced, refurbished, or repaired in the course of an engine overhaul. Thus, significant changes of the turbine blade geometry can be introduced due to regeneration and overhaul processes. Such geometric variances can affect the aerodynamic and aeroelastic behavior of turbine blades. One goal in the development of the regeneration process is to estimate the aerodynamic excitation of turbine blades depending on these geometric variances caused during the regeneration. Therefore, this study presents an experimentally validated comparison of two methods for the prediction of forced response in a multistage axial turbine. Two unidirectional fluid structure interaction (FSI) methods, a time-linearized and a time-accurate with a subsequent linear harmonic analysis, are employed and the results validated against experimental data. The results show that the vibration amplitude of the time-linearized method is in good agreement with the experimental data and, also requires lower computational time than the time-accurate FSI.
Based on this result, the time-linearized method is used to perform a sensitivity study of the tip clearance size of the last rotor blade row of the five stage axial turbine. The results show that an increasing tip clearances size causes an up to 1.35 higher vibration amplitude compared to the reference case, due to increased forcing and decreased damping work.
Aerodynamic Optimization of a Two-Part Variable Inlet Guide Vane in a Highly Loaded Low Pressure Compressor
Authors: Hemmert-Pottmann Stefan , Gouézou William , Nicke Eberhard 
Published: August 2018

Continuous reduction of fuel consumption for a wide range of operating conditions leads to a high efficiency demand for all engine parts of modern jet engines and especially the compressor. To meet these requirements a two-part Variable Inlet Guide Vane (VIGV), composed of a fixed strut and a variable flap, can be used. Besides the aerodynamic aspects, the VIGV strut is a substantial part for the structural integrity of the compressor. The aerodynamic design optimization of such a VIGV, located upstream of the first rotor of a 2.5 stage low pressure compressor, under the conditions of three different operating points is presented in this paper.
In a previous study the shape of the axial gap between strut and flap was optimized without changing the envelope of both parts [1]. The new design tool SplitBlade, developed at the DLR, enables the creation of an axial gap and has been integrated in the design process of the in-house optimization tool AutoOpti. The target of the optimization was to decrease the total pressure loss coefficients for all three operating points.
The design optimization presented in this paper is more complex by allowing the VIGV blade geometry to change. The basic dimensions of the VIGV such as the axial chord and the maximum profile thickness are still frozen. In total, 88 parameters are free to change in the optimization process. Additionally to the main target of loss reduction, the circumferential outflow angles are restricted to maintain the deflection of the blade and therewith the required rotor inflow conditions to ensure the operability of the entire compressor in the whole working range.
The final result is a two-part VIGV with an axial gap, which is optimized in terms of total pressure losses in three operating points. Compared to a reference geometry without an axial gap, the losses are almost equal at nominal speed, and about one to two percentage points higher in the two part speed operating points.
Development of a Criterion for a Robust Identification of Diffuser Rotating Stall Onset in Industrial Centrifugal Compressors
Authors: Bianchini Alessandro , Andreini Giulia , Ferrara Giovanni , Ferrari Lorenzo , Rubino Dante Tommaso 
Published: August 2018

Recent studies showed that a prompt detection of the stall inception, connected with a specific model to predict its associated aerodynamic force, could provide room for an extension of the left margin of the operating curve of high-pressure centrifugal compressors.
In industrial machines working in the field, however, robust procedures to detect and identify the phenomenon are still missing, i.e. the operating curve is almost ever cut preliminary by the manufacturer by a proper safety margin; moreover, no agreement is found in the literature about a well-defined threshold to define the onset of the stall. In particular, in some cases the intensity of the arising subsynchronous frequency is compared to the revolution frequency, while in many other ones it is compared to the blade passage frequency. A large experience in experimental stall analyses collected by the authors revealed that in some cases unexpected spikes could make this direct comparison not reliable for a robust automatic detection.
To this end, a new criterion was developed based on an integral analysis of the area subtended to the entire subsynchronous spectrum of the dynamic pressure signal of probes positioned just outside the impeller exit. A dimensionless parameter was then defined to account for the spectrum area increase in proximity to stall inception. This new parameter enabled the definition of a reference threshold to highlight the arising of stall conditions, whose validity and increased robustness was here verified based on a set of experimental analyses of different types of full-stage test cases of industrial centrifugal compressors at the test rig.
Full-Annulus URANS Study of Inlet Hot-Streak Transportation in a Four-Stage Gas Turbine
Authors: Chi Zhongran , Liu Haiqing , Zang Shusheng , Pan Chengxiong , Zhang Meibao 
Published: August 2018

Nonuniformity of combustor-exit-flow temperature, also named hot-streak, could strongly increase the complexity in unsteady flow and heat transfer of gas turbine cascades. Detailed investigation into the hot-streak transportation in multi-stage gas turbine should provide a valuable reference to both fault diagnosis of combustors based on the distribution of turbine exhaust temperature.
In this study, transportation of hot-streaks inside a four-stage gas turbine was numerically investigated using three-dimensional (3D), Unsteady Reynolds Averaged Navier-Stokes equations (URANS) Computational Fluid Dynamics (CFD) study with simulation domain including eight full-annulus cascades. The two-dimensional (2D) temperature fields at the rotor-stator interfaces and turbine exit was acquired by the full-annulus URANS CFD, and then discussed through comparing the circumferential and radial distributions and also the ranges of time-dependent fluctuations. The time-average temperature distributions were analyzed by Discrete Fourier Transform (DFT) method in circumferential direction to characterize the distribution.
The URANS CFD results show that the hot-streaks could travel through the entire turbine and result in nonuniformity of turbine exhaust temperature. The rotor-stator interaction could change the 2D shape of hot-streaks and intensity the time-dependent fluctuation of flow temperature. DFT results show that the typical frequencies of circumferential temperature distribution are coinciding with the number of hot-streaks and the blade number of upstream/ downstream cascades. The DFT results also indicated that the hot-streaks migrate 35°∼45° in the circumferential direction, and the maximal relative amplitude of hot-streaks is reduced from 28% to 5%. The redistribution of hot/cold fluid in stator cascade could affect the mixing of hot-streaks depending on clocking position.
Experimental and Numerical Investigation on Gas Turbine Package Scale Model
Authors: Lucherini Gabriele , Minotti Stefano , Ragni Giacomo , Bologna Francesco 
Published: August 2018

Gas turbine packages require a ventilation system in order to keep temperatures under acceptable limits and to dilute any hazardous accumulation of gas due to unexpected leakages. As part of the design phase, a detailed CFD (Computational Fluid Dynamics) analysis is performed on the complete system to assess the fluid dynamic behavior of the flow in terms of flow path, temperature distributions and velocity field.
In this work, as additional approach, a detailed experimental 3D assessment of an entire aero-derivative gas turbine (GT) package was performed by creating a scale model (1:8) of the real configuration. The original package can be as much as 60 m3 in volume where details from pipes to valves can create severe flow distortions and the 3D CFD study might not necessarily include all details up to such level during the design phase. The scale model, built using sintered plastic material through rapid prototyping, was used for a test campaign reproducing the operation of the ventilation system, copying the dynamic similarity of the real scale. The model was equipped with a set of instruments to acquire measurements of pressure and velocity in several locations and at different flow rates. A significant benefit of using a scale model with transparent plexiglass for the external structure of the enclosure and ventilation ducts walls, was that it allowed to carry out a smoke test. This has been done by injecting a visible gas from several locations allowing the visualization of the streak lines of the local flow field. The aim of this approach was to find a fast and reliable way to investigate in detail complex phenomena such as gas leakage dilution and local flow distribution.
A good agreement between experimental and computational data was found confirming that the CFD studies currently performed during the standard design phase are accurate and reliable enough to provide a proper prediction of the flow field inside the entire package even when a high level of details is included.
Characterizing the Influence of Cyclic Re-Priming on the Prediction of Long Term Creep Damage for Gas Turbine Components
Authors: Green Richard , Douglas Jonathan , Moffat Andrew , Bellows Richard 
Published: August 2018

Industrial Gas Turbines have traditionally been designed for base load operation. However, modern applications require flexibility as well as availability. These requirements lead to components in the hot section experiencing both cyclic and hold time loading, at high temperatures. These loading profiles inevitably lead to damage from both fatigue, due to engine cycles, and creep, due to dwells at load, as well as the interaction of these two damage mechanisms over the duration of the service interval. It is these interactions which can lead to higher average creep rates and more damage than expected.
This paper presents the results of a study into the influence of cyclic loading on the average creep rate for a proprietary single crystal nickel based alloy, which is based on the chemistry of INCONEL® 792, a relatively high chromium, gamma/gamma-prime strengthened superalloy. Creep tests have been conducted isothermally with reload cycles of varying duration (or dwell) that result in an unexpected reduction in creep life and an increase in overall creep rate when compared to the continuous isothermal creep tests performed at constant stress without reloading. A hypothesis is presented which attributes the increase in overall creep rate to the influence of a recovery potential stress. The dwell period of each reload cycle is critical to calculating the recovery potential and subsequent creep rate. Results show that tests with relatively short dwell periods exhibited lower lives and higher creep rates than tests with fewer cycles and longer dwell periods. The implication of these findings are significant when considering actual operation, where variations in engine cycles and dwell periods, could influence the accumulation of creep damage. Therefore, an initial approach is presented which accounts for cycles as part of the creep damage calculation.