bluuebunny/crossref_metadata_2025 huggingface dataset:
Prediction of CO and NOx Pollutants in a Stratified Bluff Body Burner
Authors: Gruhlke Pascal , Proch Fabian , Kempf Andreas M. , Illana Mahiques Enric , Dederichs Stefan , Beck Christian
Published: August 2017

The major exhaust gas pollutants from heavy duty gas turbine engines are CO and NOx. The difficulty of predicting the concentration of these combustion products originates from their wide range of chemical time scales. In this paper, a combustion model that includes the prediction of the carbon monoxide and nitric oxide emissions is tested. Large eddy simulations (LES) are performed using a compressible code (OpenFOAM). A modified flamelet generated manifolds (FGM) approach is applied with a thickened flame approach (ATF) to resolve the flame on the numerical grid, with a flame sensor to ensure that the flame is only thickened in the flame region. For the prediction of the CO and NOx emissions, pollutant species transport equations and a second, CO based, progress variable are introduced for the flame burnout zone to account for slow chemistry effects. For the validation of the models, the Cambridge burner of Sweeney and Hochgreb [1, 2] is employed, as both carbon monoxide and nitric oxide [3] data is available.
A Practical Approach From the MEE Toward Hybrid Propulsion
Authors: Morioka Noriko , Oyori Hitoshi , Asako Tomoaki , Takahashi Katsuyuki , Ando Takumi
Published: August 2017

The MEA (More Electric Aircraft)/AEA (All Electric Aircraft) and MEE (More Electric Engine) is representative of system evolution, contributing to improved energy management of the entire aircraft system and reducing fuel consumption of aircraft. Looking at future aviation, it is necessary to have further evolution beyond the MEA/AEA and MEE to achieve further improvement of aircraft efficiency and “greener aviation”. The electrified propulsion system, which uses electrical energy as propulsive power, would be a possible option, and various kinds of configurations have been studied and developed. Small one- or two-seater airplanes have already successfully flown with the all-electrified propulsion; however, there are a lot of perspectives for the system configuration for larger-sized commercial aircraft. It is considered that a hybrid propulsion system, in which the E3M (Engine Embedded Electric Machine) provides assisting power to the engine shafts, would be one of the probable configurations. The authors are currently developing the E3M as an LP (Low Pressure) spool-driven generator, and the approach is considered to be practical because E3M technology can be commonly applicable for both large electrical power generation and hybrid propulsion. In this paper, several advanced propulsion system configurations are compared, and perspectives of the hybrid propulsion for future commercial aircraft are described. Also technical challenges for the MEE/E3M, continuing toward future hybrid propulsion, are outlined and further R&D steps are considered.
Interaction Between Precessing Vortex Core and Thermoacoustic Coupling in a Lab-Scale Lean Premixed Gas Turbine Combustor: Numerical Simulation Studies
Authors: Wang Zhenlin , Li Xiangsheng , Feng Zhenping
Published: August 2017

To evaluate the complex dynamic phenomena occurring during combustion, Large Eddy Simulations of the swirl-stabilized flame of PRECCINSTA burner were carried out. The existence of a precessing vortex core in the initial stage of unsteady combustion was proved, both in the cold flow and initial stage of the combustion. The PVC flow structure finally disappeared and became a flapping toroidal flowfield structure with the increase in the amplitude of pressure oscillation when combustion instabilities occurred. The LES results indicate the interactions between the PVC and thermoacoustic coupling and the underlying mechanism of the phenomenon. This study evaluated the effect of PVC on the thermoacoustic coupling in the initial stage of unsteady combustion and elucidated the mechanism for the disappearance of PVC under large-scale pressure oscillation. This has a great significance in the practical industrial applications of controlling combustion instabilities.
An Experimental Investigation on the Overall Cooling Performances of Two Turbine End-Wall Structures
Authors: Wang Wei , Pu Jian , Yuan Rui-ming , Wang Jian-hua , Luan Yong-xian , Kang Bin-peng
Published: August 2017

This paper presents an experimental investigation on the overall cooling performances of two endwall specimens, one is a simple film cooling, the other is a laminated structure with pin-fins and impingement holes, but has the same cover board of the simple film cooling endwall. The two specimens are made of stainless steel with a real size of a gas turbine endwall. The experiments were carried out in a large temperature ratio (TR) of mainstream to cooling air (TR = 2.47), which is close to real operation conditions of modern gas turbines. The surface temperatures were measured by an infrared thermal imaging system (ITIS) at three blowing ratios (BRs = 0.63, 1.02, 2.04). The overall effectiveness of the two specimens was analyzed and compared. Through the analysis and comparisons, some interesting phenomena are discovered: 1) In general, as like as the previous studies of endwall cooling, the laminated endwall specimen can provide a higher cooling effectiveness averaged over the entire area than the simple film cooling specimen; 2) However, if we want to get more exact designs, the combination of the two structures may be better. Because of at low BR = 0.63, the simple endwall specimen can locally demonstrate a better cooling effect in an elliptic region of the endwall; 3) At high BR = 2.04, the design strategy of the simple endwall specimen can also locally be used in a triangle region. The exact designs can not only increase overall cooling effect, but also reduce the weight of whole endwalls.
Experimental Investigations of the Aerodynamics of Highly Loaded Tandem Vanes in a High-Speed Stator Cascade
Authors: Heinrich Alexander , Tiedemann Christine , Peitsch Dieter
Published: August 2017

Tandem blade configurations are said to being able to outperform conventional single blades in terms of loss behaviour and flow turning. The work presented here is supported by experimental investigations which were conducted at a 2D linear stator cascade at the Chair for Aero Engines at the Technische Universität Berlin. Two different tandem blade configurations with a different load split are examined and compared against a conventional single reference blade.

Multi-coloured oil flow visualisation complemented by wake flow measurements show the development of the secondary flow structures and enable a thorough understanding of the impact of varying tangential displacement for tandem blades. The results show that tandem blades can reduce the total pressure losses, especially in the midspan region, which also results in smaller overall values. In addition, the incidence angle variation has shown that the working range of the tandem vanes exceeds the originally designed working range of the single reference blade. The tandem configurations still allow flow guidance at stall condition and prove the capability of the concept.
Validation of a Numerical Model for Predicting Stalled Flows in a Low-Speed Fan
Authors: Lee Kuen-Bae , Wilson Mark , Vahdati Mehdi
Published: August 2017

This paper investigates the rotating stall for a low-speed/low-pressure ratio fan blade. The type of stall studied in this paper is similar to the multi cell, part span rotating stall, which can occur on the front stages of the core compressors. Although this type of rotating stall may not result in surge or massive loss of power, it can cause excessive vibration and noise.

It was found that the standard Spalart-Allmaras (SA) model implemented in the CFD solver used in this work predicts premature stall, which is in line with the observation of other researchers who use the SA model. Therefore, to improve the prediction of the stall boundary, the standard SA model was modified by scaling the source term in the model based on the local pressure gradient and the velocity helicity of the flow. Furthermore, a generalized wall function valid for non-zero wall pressure gradient was implemented to improve the accuracy of boundary conditions at the wall. The turbulence modelling part of this work aims to produce a turbulence model which can be used to model the flows near the stall boundary for the transonic fan blades on relatively coarse grids of around 300k points per passage. Initially, two fan blades with different design and operating speeds were used to optimize the new parameters in the modified turbulence model. The optimization was based on improving the correlation between measured and numerical radial profiles of the pressure ratio. Thereafter, steady computations were performed for two other blades (with the same parameters) and the predictions were compared with the experimental data for all the four fan blades. Numerical results showed a significant improvement over those obtained with the original SA model, when compared against the measured data.

In the second part, the modified turbulence model was used to study the flow near the stalled region. Three-dimensional, whole assembly unsteady simulations were performed for a modern low speed fan rig for which extensive measured data were available. Stall simulations were conducted at 80% speed, which contained a part with positive slope on the measured constant speed characteristic. The results indicated that by using an unsteady whole assembly approach and the modified turbulence model, it is possible to predict the flow for all the points on the measured constant speed characteristic (including those on the positive slope part).
Different Effects of Cantilevered and Shrouded Stators on Axial Compressor Performance
Authors: Si Xiayi , Teng Jinfang , Qiang Xiaoqing , Feng Jinzhang
Published: August 2017

Numerical simulations with the steady 3D RANS were performed on the rear stage of a modern high pressure compressor. The labyrinth seal cavity model of the shrouded stator was simplified according to the actual stator structure, which the seal cavity gap is 1% of blade height. Several typical configurations (shrouded stator, idealized stator and cantilevered stators) were designed and carried out, and cantilevered stators contained no gap, small gap (CS1%), design gap (CS2.5%) and large gap (CS4%/CS5%). The results indicate due to the effect of leakage flow from 1% span seal cavity gap, the total pressure loss of SS is larger than IS, while IS instead of SS in the process of the compressor design, the stall margin will be enlarged nearly 6% numerically. At the design point, when the hub gap is 3.5% span clearance CS has the same loss with IS, and when the hub gap is 4.5% span clearance CS has almost the same loss with SS. Among all operation range, the total pressure loss of S1 increases with the increase of the hub clearance. When the hub gap is 0 (CS0), there is no leakage flow and the loss is the least. At the design point, comparing with SS, the total pressure loss coefficient of CS0 decreases 18.34%, CS2.5% decreases 8.46% and IS decreases 6.45%. It means if the cantilevered stator with 2.5% span hub clearance were adopted in the HPC, the performance would be better than the shrouded stator. However, because of the matching condition, the rotor that follows after cantilevered stator should be redesigned according to blade loading and inlet flow angle changed. The performance of cantilevered stator is impacted of various hub clearance, the loss below 25% span increases significantly with hub clearance, the maximum value of outlet flow angle deviation is 2.3 degree. The stator hub peak loading is shifted upstream toward the leading edge when hub clearance size is increased. The total pressure loss coefficient and pressure coefficient at different axial position had the function relation. When the hub clearance increases, the position of double leakage flow start backwards, in the rear part of stator the secondary flow becomes stronger leading to more mixing loss and lower total pressure.
Design and Testing for Ceramic Matrix Composite Turbine Vane
Authors: Watanabe Fumiaki , Nakamura Takeshi , Mizokami Yousuke
Published: August 2017

Ceramic matrix composite (CMC) have higher temperature capability and lower density than nickel based alloys which have been used for hot section components of gas turbine engines. These properties are expected to bring many benefits, such as higher turbine inlet temperature (TIT), reduction of cooling air, and reduction of weight, when it is used as the material for hot section components of gas turbine engine.

The authors have been developing CMC turbine vane for aircraft engines. In this paper, the authors present the summary of design, manufacturing, and testing, which were conducted from 2010 to 2012.

The purpose of this work was to verify that the SiC-SiC CMC which IHI has developed has the applicability to aircraft turbine vanes. The concept was planned for CMC hollow turbine vanes, in which the airfoil and the platform are fabricated in CVI process. As the demonstration of this concept, the first stage turbine vane was designed with CMC for IHI IM270 that is the 2MW-class small industrial gas turbine engine.

Bending rig test was conducted at room temperature in order to check the structural feasibility of the airfoil-platform joint. The outer platform of vane was fixed in the same way with the engine parts, and the load simulating the aerodynamic force was applied at the airfoil portion. The fracture load was higher than the load which the vanes receive in the actual engine.

Burner rig test was conducted in order to check the durability against thermal cycle. A CMC vane was set between dummy metal vanes, and cyclically heated by gas burner. The maximum airfoil surface temperature was set to 1200 degree C, and the maximum temperature difference between airfoil and platform was about 700 degree C. The minimum airfoil temperature at the interval of heating was about 300 degree C. The time of one thermal cycle was 6 minutes that consisted of 3 minute heating and 3 minute natural cooling. The test was conducted for 1,000 cycles. In post-test inspection there was no defect like a crack.

Engine test for CMC vanes was conducted using IHI IM270. The four CMC vanes were mounted into the first stage turbine nozzle assembly in place of the normal metal vanes. The test was conducted for 400 hours. The inlet temperature of CMC vanes were measured by thermocouples installed at the leading edge, and the measured temperature was about 1050 degree C at the steady state.

From this work, the applicability of the design concept for the CMC vane to actual engine was verified in which airfoil-platform are fabricated in CVI process.
Detailed Numerical Study of the Main Sources of Loss and Flow Behavior in Low Pressure Steam Turbine Exhaust Hoods
Authors: Munyoki Dickson , Schatz Markus , Vogt Damian M.
Published: August 2017

The performance of the axial-radial diffuser downstream of the last low-pressure steam turbine stages and the losses occurring subsequently within the exhaust hood directly influences the overall efficiency of a steam power plant. It is estimated that an improvement of the pressure recovery in the diffuser and exhaust hood by 10% translates into 1% of last stage efficiency [11].

While the design of axial-radial diffusers has been the object of quite many studies, the flow phenomena occurring within the exhaust hood have not received much attention in recent years. However, major losses occur due to dissipation within vortices and inability of the hood to properly diffuse the flow. Flow turning from radial to downward flow towards the condenser, especially at the upper part of the hood is essentially the main cause for this.

This paper presents a detailed analysis of the losses within the exhaust hood flow for two operating conditions based on numerical results. In order to identify the underlying mechanisms and the locations where dissipation mainly occurs, an approach was followed, whereby the diffuser inflow is divided into different sectors and pressure recovery, dissipation and finally residual kinetic energy of the flow originating from these sectors is calculated at different locations within the hood.

Based on this method, the flow from the topmost sectors at the diffuser inlet is found to cause the highest dissipation for both investigated cases. Upon hitting the exhaust hood walls, the flow on the upper part of the diffuser is deflected, forming complex vortices which are stretching into the condenser and interacting with flow originating from other sectors, thereby causing further swirling and generating additional losses.

The detailed study of the flow behavior in the exhaust hood and the associated dissipation presents an opportunity for future investigations of efficient geometrical features to be introduced within the hood to improve the flow and hence the overall pressure recovery coefficient.
Numerical Investigation of Stator Clocking Effects on the Downstream Stator in a 1.5-Stage Axial Turbine
Authors: Zhu Yalu , Luo Jiaqi , Liu Feng
Published: August 2017

The unsteady nature of the flow in multi-stage turbomachines makes it possible for designers to improve the aerodynamic performance by changing the relative pitchwise positions of consecutive stators/rotors. An in-house three-dimensional multi-stage flow solver with a volume-weighted interpolation method on the interface between adjacent blade rows is employed to compute the unsteady flow in a 1.5-stage axial turbine. The variation of entropy at the inlet of the second stator and the pressure fluctuation over the second stator blade are firstly presented to illustrate the unsteady wake characteristics and the development of pressure fluctuations in the second stator passages. Then the variations of overall aerodynamic parameters, wake characteristics and pressure fluctuations at 50% span position are further investigated for five different stator clocking configurations. The results are presented in detail, demonstrating that at 50% span position the adiabatic efficiency of the stage increases for the configurations with the upstream stator wake cores passing through the downstream stator passages near the suction side. For the fluctuation produced by the impingement of upstream low entropy on the stator blade, the amplitude is much greater for the clocking configuration with higher efficiency. However, for the fluctuation produced by the collision of two other fluctuations, the amplitude is larger for the clocking configuration with lower efficiency.
Development and Validation of an On-Wing Engine Thrust Measurement System
Authors: Bauer Marc , Friedrichs Jens , Wulff Detlev , Werner-Spatz Christian
Published: August 2017

Maintenance on aircraft engines is usually performed on an on-condition basis. Monitoring the engine condition during operation is an important prerequisite to provide efficient maintenance. Engine Condition Monitoring (ECM) has thus become a standard procedure during operation. One of the most important parameters, the engine thrust, is not directly measured, however, and can therefore not be monitored, which makes it difficult to distinguish whether deteriorating trends e.g. in fuel comsumption must be attributed to the engine (e.g. due to thermodynamic wear) or to the aircraft (e.g. due to increased drag). Being able to make this distinction would improve troubleshooting and maintenance planning and thus help to reduce the cost of ownership of an aircraft. As part of the research project APOSEM (Advanced Prediction of Severity effects on Engine Maintenance), Lufthansa Technik (LHT) and the Institute of Jet Propulsion and Turbomachinery of Technische Universität Braunschweig develop a method for direct measurement of engine thrust during the operation. In this paper, the design process of the On-Wing (OW) Measurement System is presented, including the validation in labratory tests, the mechanical and thermal calibration as well as the final ground test during an engine test run at LHT test cell and the work on the flight test certification.
Optimization of Hybrid-Linked Jet Impingement Cooling Channels Based on Response Surface Methodology and Genetic Algorithm
Authors: Yang Li , Min Zheng , Parbat Sarwesh Narayan , Chyu Minking K.
Published: August 2017

In recent years, development of new manufacturing technologies like additive manufacturing has made it possible to make complex cooling structures to improve the efficiency of jet impingement. Present paper considers hybrid-linked jet impingement cooling channels which involve both parallel linked jets and serial linked jets. Systematic analysis was conducted with the aid of Computational Fluid Dynamics and Response Surface Methodology, focusing on the influence of topology on performance. An optimization platform was established with aid of the regressed database and the Genetic Algorithm. Of particular interest is the influence of optimization strategies on results. Results obtained indicates that the topology number developed in this study works well with the Response Surface Methodology. Topology can be considered to be a new degree of freedom of jet impingement design. Among the tested topologies, serial linked jet impingement has significantly higher heat transfer and pressure drop than the traditional parallel linked jet impingement. In the first optimization strategy, mass flow rate was used as the objective function while heat transfer and pressure drop were constrained. Optimized results under this strategy show consistent parameters and purely serial linked topology for all cases, due to the high cooling efficiency of serial linked jets. In the second optimization strategy, pressure drop was minimized while heat transfer and mass flow rate were constrained. Contrast with the first strategy, optimal results of this strategy have different topologies under different constraint conditions, which is caused by the complex influence of geometric parameters on pressure drop. Such results indicate the capability of hybrid-linked jet impingement to fit a wide various requirement by changing topology.
Transient Modeling of a Supercritical CO2 Power Cycle and Comparison With Test Data
Authors: Avadhanula Vamshi K. , Held Timothy J.
Published: August 2017

A transient model of a commercially-available 7.3MWe sCO2 power cycle was developed using an implicit 1-D Navier Stokes solver. The power cycle underwent extensive factory testing, which validated the component and system performance, and the design and performance of the control system. The model structure simulates the as-tested configuration of the power cycle, with major components consisting of water-cooled heat rejection heat exchanger, turbine-driven compressor, recuperator, primary heat exchanger, power turbine, gearbox and generator. Subsystem models are developed to validate individual component (compressor, drive turbine, power turbine, heat exchangers etc.) model performance against design data. The component models are then assembled into the full system model. For the system transient simulation, the input parameters (or boundary conditions) are taken from the test data. More than eight hours of test data is used in the present transient simulation. The simulation results, including thermodynamic state points, component performance and system performance, are compared against the corresponding test data.
Geometric Modeling and Analysis for Gooseneck II: 2D Simplified Model for Quick Assessment
Authors: Tu Qiuye , Deng Rutan , Zhang Dongdong , Sun Xingjian
Published: August 2017

In the previous studies, the proposed method for gooseneck geometric modeling employed two polynomials to construct the inner-wall and area distribution curves. The inflection point location served as the variable to control the inner-wall polynomial curve, and the peak point location and peak value to control the area distribution polynomial curve. In the effort to be quickly located, the control variables were provided with more geometric meaning. 3D numerical simulations indicated that there existed a total pressure recovery island for given solution area of the three control variables. Consequently, the relationships between the geometric parameters and the total pressure loss were set up.

This paper focused on the 2D simplifications to quickly address the control variables for the total pressure island. The studies were conducted in three aspects. First, the simplified model took the constructional blocking effects of struts into account. The baseline of the 2D simplified modeling was set at 30% spanwise near the hub through comparisons of different settings. Therefore, 70% blocking area compensated to outer-wall and 30% to inner-wall along the normal direction of the baseline. The 2D simulation results indicated that the static pressure distribution was consistent with the 3D results, but waves exited at the end walls of both leading and trailing edges due to the geometric changes. Second, the simplification considered the blocking effects of the wake. The wake was converted to boundary layer thickness, and moreover, compensation to the end wall was similar with the constructional blocking of struts. The simulation results revealed that wake blocking had very small impacts to the simplification, although the peak values of static pressure slightly increased at the end wall. Third, smoothing treatments were done for both inner-wall and outer-wall after the above compensating transformations. The results showed that smoothing treatments were very necessary and improved the waves located at end wall on the static pressure distribution which was nearly the same with 3D results. After all the simplifying treatments above, the final 2D results had almost the same total pressure loss distributions with the 3D results, and could save at least 40% calculation time as a quick assessment used to search the reasonable geometric solution areas of inflection point location and peak point location for minimum total pressure loss of the gooseneck.
Low Reynolds Number Response of High Efficiency, Intermediate Pressure Compressor Profiles
Authors: Lázaro Benigno J. , González Ezequiel , Cadrecha David , Antoranz Antonio , Parra Jorge
Published: August 2017

Current trends on intermediate pressure, axial compressors designs for aeroengine applications demand to extend their operation envelope into low Reynolds number regimes, of the order of 105 based on the real chord and inlet velocity (Re). In this range, a very limited open experimental database on profile performance can be found. Furthermore, in order to propose high efficiency designs for this regime, it is critical to determine and to understand the profile behaviour with respect to different operating parameters. In this work, a detailed experimental study of a proposed high efficiency, intermediate pressure compressor aerofoil has been carried out, both for design and off-design flow incidences, in the range 1.5 · 105 < Re < 3.5 · 105,. The experimental facility is a low-speed linear cascade where different boundary suction strategies have been implemented to optimize the flow periodicity and to minimize pressure gradient perturbations induced by end-wall secondary flow development, in an effort to ensure high quality, 2D passage flow evolution both at design and significant off-design incidences. High resolution total pressure loss and LDV traverses performed at different streamwise locations have been carried out to describe the flow evolution. The characterizations performed at close to nominal incidence give a profile loss dependence on the Reynolds number that exhibits two clearly differentiated ranges, with the lower one exhibiting a higher profile loss dependence on the Reynolds number. At large off-design incidences, the profile loss coefficient practically becomes independent of the Reynolds number, rapidly increasing as the incidence is increased. In both cases physical arguments and scaling laws based on the experimental evidence are proposed to explain the profile behavior. RANS and URANS based CDF simulations have been also conducted, showing their ability and limitations to capture the experimentally observed aerofoil behavior.
High Speed Imaging of Forced Ignition Kernels in Non-Uniform Jet Fuel/Air Mixtures
Authors: Wei Sheng , Sforzo Brandon , Seitzman Jerry
Published: August 2017

This paper describes experimental measurements of forced ignition of prevaporized liquid fuels in a well-controlled facility that incorporates non-uniform flow conditions similar to those of gas turbine engine combustors. The goal here is to elucidate the processes by which the initially unfueled kernel evolves into a self-sustained flame. Three fuels are examined: a conventional Jet-A and two synthesized fuels that are used to explore fuel composition effects. A commercial, high-energy recessed cavity discharge igniter located at the test section wall ejects kernels at 15 Hz into a preheated, striated crossflow. Next to the igniter wall is an unfueled air flow; above this is a premixed, prevaporized, fuel-air flow, with a matched velocity and an equivalence ratio near 0.75. The fuels are prevaporized in order to isolate chemical effects. Differences in early ignition kernel development are explored using three, synchronized, high-speed imaging diagnostics: schlieren, emission/chemiluminescence, and OH planar laser-induced fluorescence (PLIF). The schlieren images reveal rapid entrainment of crossflow fluid into the kernel. The PLIF and emission images suggest chemical reactions between the hot kernel and the entrained fuel-air mixture start within tens of microseconds after the kernel begins entraining fuel, with some heat release possibly occurring. Initially, dilution cooling of the kernel appears to outweigh whatever heat release occurs; so whether the kernel leads to successful ignition or not, the reaction rate and the spatial extent of the reacting region decrease significantly with time. During a successful ignition event, small regions of the reacting kernel survive this dilution and are able to transition into a self-sustained flame after ∼1–2 ms. The low aromatic/low cetane number fuel, which also has the lowest ignition probability, takes much longer for the reaction zone to grow after the initial decay. The high aromatic, more easily ignited fuel, shows the largest reaction region at early times.
Using Dynamic Simulation to Evaluate Attemperator Operation in a Natural Gas Combined Cycle With Duct Burners in the Heat Recovery Steam Generator
Authors: Liese Eric , Zitney Stephen E.
Published: August 2017

A generic training simulator of a natural gas combined cycle was modified to match operations at a real plant. The objective was to use the simulator to analyze cycling operations of the plant. Initial operation of the simulator revealed the potential for saturation conditions in the final high pressure superheater as the attemperator tried to control temperature at the superheater outlet during gas turbine loading and unloading. Subsequent plant operational data confirmed simulation results. Multiple simulations were performed during loading and unloading of the gas turbine to determine operational strategies that prevented saturation and increased the approach to saturation temperature. The solutions included changes to the attemperator temperature control setpoints and strategic control of the steam turbine inlet pressure control valve.
Thermal Properties for the Simulation of Direct-Fired sCO2 Combustor
Authors: Manikantachari K. R. V. , Martin Scott , Bobren-Diaz Jose O. , Vasu Subith
Published: August 2017

The direct-fired supercritical CO2 (sCO2) cycle is currently considered as a zero-emission power generation concept. It is of interest to know how to optimize various components of this cycle using computational tools, however, a comprehensive effort on this area is currently lacking. In this work, the behavior of thermal properties of sCO2 combustion at various reaction stages has been investigated by coupling real gas CHEMKIN (CHEMKIN-RG) with an in-house Premixed Conditional Moment Closure (PCMC) code and the high pressure Aramco-2.0 kinetic mechanism. Also, the necessary fundamental information for sCO2 combustion modelling is reviewed.

The Soave-Redlich-Kwong equation of state (SRK EOS) is identified as the most accurate EOS to predict the thermal states at all turbulence levels. Also, an empirical model for the compression factor Z is proposed for sCO2 combustors, which is a function of mixture inlet conditions and the reaction progress variable. This empirical model is validated between the operating conditions 250–300 bar, inlet temperatures of 800–1200 K and within the current designed inlet mole fractions and the accuracy is estimated to be less than 0.5% different from the exact relation. For sCO2 operating conditions the compression factor Z always decreases as the reaction progresses and this leads to the static pressure loss between inlet and exit of the sCO2 combustor.

Further, a review of high pressure viscosity and thermal conductivity models of mixtures and pure-components are presented from the literature and suggestions are made for their adoptability in sCO2 combustor simulations. The thermal properties such as specific heats, speed of sound, pressure exponent and isothermal compressibility are accurately quantified.
Thermoacoustic Damping Rate Determination From Combustion Noise Using Bayesian Statistics
Authors: Stadlmair Nicolai V. , Hummel Tobias , Sattelmayer Thomas
Published: August 2017

In this paper, we present a method to determine the quantitative stability level of a lean-premixed combustor from dynamic pressure data. Specifically, we make use of the autocorrelation function of the dynamic pressure signal acquired in a combustor where a turbulent flame acts as a thermoacoustic driver. In the proposed approach, the unfiltered pressure signal including several modes is analyzed by an algorithm based on Bayesian statistics. For this purpose, a Gibbs sampler is used to calculate parameters like damping rates and eigenfrequencies in the form of probability density functions (PDF) by a Markov-Chain Monte-Carlo (MCMC) method. The method provides a robust solution algorithm for fitting problems without requiring initial values. A further advantage lies in the nature of the statistical approach since the results can be assessed regarding its quality by means of the PDF and its standard deviation for each of the obtained parameters. First, a simulation of a stochastically forced van-der-Pol oscillator with pre-set input values is carried out to demonstrate accuracy and robustness of the method. In this context it is shown that, despite of a large amount of uncorrelated background noise, the identified damping rates are in a good agreement with the simulated parameters. Secondly, this technique is applied to measured pressure data. By doing so, the combustor is initially operated under stable conditions before the thermal power is gradually increased by adjusting the fuel mass flow rate until a limit-cycle oscillation is established. It is found that the obtained damping rates are qualitatively in line with the amplitude levels observed during operation of the combustor.
Impact of Water Injection on Thermoacoustic Modes in a Lean Premixed Combustor Under Atmospheric Conditions
Authors: Stadlmair Nicolai V. , Mohammadzadeh Keleshtery Payam , Zahn Max , Sattelmayer Thomas
Published: August 2017

The goal of the study presented in this paper is to analyze how the stability margin of a premixed combustor changes when liquid water is injected. Within the scope of this work, experimental results are presented and serve as a validation basis for a numerical model of the combustor. Experiments comprise qualitative as well as quantitative analyses by means of damping rate measurements under varying water-to-fuel ratios. For this purpose, a natural gas fueled low-NOX burner which is equipped with a coaxially placed water injection system is operated in a single burner test rig under atmospheric conditions. Water-to-fuel ratios investigated in this study range between 0 and 2 whereby either the adiabatic flame temperature or the equivalence ratio is held constant. With technical applications in mind, the former case is chosen to represent the conditions occurring in a gas turbine combustor when the thermal output is increased at a constant flame temperature level. A recently developed approach is applied to extract the net damping rates of the occurring modes from dynamic pressure measurements. More specifically, the autocorrelation function of the combustion noise signal is evaluated by an algorithm based on Bayesian statistics to ensure for robust extraction of the complex eigenfrequencies. Beyond that, the dynamic system response is measured for various operating conditions with and without water injection by means of Scattering Matrices (SM) using the Multi-Microphone Method (MMM). Measured SM’s are deployed in a numerical model of the combustor and acoustic conservation equations are solved in frequency space using a Finite Element (FEM) approach. The results are, on the one hand, evaluated to analyze occurring thermoacoustic modes by comparing its spatial distribution with experiments. On the other hand, eigenvalue studies are carried out and experimentally identified eigenfrequencies and damping rates are used to benchmark the results of the numerical analyses.
Development and Testing of a Gas Turbine Engine Combustion Air Inlet Filtration System for the USMC Amphibious Combat Vehicle
Authors: Gastopoulos Thomai , Lawton Joseph
Published: August 2017

The Auxiliary Ships and New Acquisition Support Branch (Code 425) of the Naval Surface Warfare Center, Philadelphia Division conducted a study to assist the Marine Corps Systems Command in assessing the feasibility of using a gas turbine engine as a propulsion system on future United States Marine Corps Amphibious Combat Vehicles (ACV). The study was focused on developing and testing a gas turbine intake solution for the ACV that can remove saltwater from the intake airstream of a notional 3,000 horsepower ACV engine.

Code 425 developed a two-part solution for the intake of the ACV. The first part of the solution is an intake shroud designed to elevate the intake to protect the engine from deck water wash. The second part of the solution is the Combustion Air Separation System (CASS), a gas turbine intake filtration system designed to remove marine contaminants that enter the intake. Code 425 tested a CASS prototype for its efficiency at removing saltwater spray and bulk water up to 10 gallons per minute. Test results showed that the CASS met each requirement and that an ACV intake system incorporating both the intake shroud and the CASS should protect the gas turbine engine from saltwater ingestion.
Investigation of the Coupling Mechanism Between Bent Pipes and Volute on the Stall Inception at the Centrifugal Compressor Inlet
Authors: Yang Ce , Wang Yingjun , Lao Dazhong , Zhang Hanzhi , Qi Mingxu , Tong Ding
Published: August 2017

The rotating stall of a centrifugal compressor not only deteriorates its efficiency but also impacts the blade fatigue failure. The inlet total pressure distortion is generated by a 90° bent pipe placed upstream from the inlet. The volute causes the circumferential non-uniform static pressure distribution of the impeller outlet, and the impeller is under the inlet distortion and the non-uniform outlet distribution condition. Current research pays little attention to the stall inception location and its formation process under the coupling interaction between the bent pipe and volute. In this paper, two installation angles of the inlet bent pipe were compared concerning the stall inception process, including the 115° (M1) and the 295° (M2) models. The circumferential angle between the volute tongue and the elbow axial plane approaches a blade passage width in model M1, and model M2 has the opposite installation angle to M1. The model M1 inlet low total pressure region caused by the bent pipe, and the outlet high static pressure region induced by the volute tongue together affect the same impeller passage at the 115° location causing the leading edge spillover. The coupling effect of the Model M1 accelerates the process of stall. However, the low total pressure region for the model M2 is located at the circumferential 295° point, and the high static outlet pressure affects the 115° impellers, resulting in a different stall inception location and process compared to the model M1. The leading edge spillover first occurs at the 295° location because of inlet distortion, and the second spillover appears at the 115° location due to the reversed propagation of the outlet high pressure region induced by the volute tongue. Compared to model M1, the stall formation of model M2 is relatively slow. Meanwhile, because of the recirculation flow, the static temperature rises sharply and the axial velocity drops significantly at the spillover region during the stall process. These results indicate that the coupling interaction between the low total inlet pressure and the high outlet static pressure jointly determines the stall inception location and its process at the centrifugal compressor inlet.
LES Combustion Model With Stretch and Heat Loss Effects for Prediction of Premix Flame Characteristics and Dynamics
Authors: Tay-Wo-Chong Luis , Scarpato Alessandro , Polifke Wolfgang
Published: August 2017

The present paper extends an approach to include effects of stretch and heat losses into turbulent combustion models from the RANS framework to the LES framework. This approach has shown the potential to improve the prediction of flame stabilization by considering these combined effects. The model is based on the calculation of the consumption speed of laminar premixed flames influenced by variations in strain and heat loss in asymmetric counterflow configurations. The consumption speed depending on strain and heat loss is introduced into a turbulent combustion model based on a progress variable approach. Large Eddy Simulations of a fully-premixed axial swirl burner with and without the influence of stretch and heat loss effects are carried out and validated against flow field and OH* chemiluminescence measurements for different power ratings and equivalence ratios. Flame dynamics are also investigated by extracting the Flame Transfer Function of the fully-premixed axial swirl burner with System Identification methods. Good agreement on the flow field, flame characteristics and dynamics between experiment and simulation was obtained with the inclusion of stretch and heat loss effects into the combustion model. Results show the importance of including these effects into turbulence combustion models for the design of premix burners for gas turbine combustors.
Methodology to Evaluate Turbocharger Turbine Performance at High Blade to Jet Speed Ratio Under Quasi Adiabatic Conditions
Authors: Serrano José Ramón , García-Cuevas Luis Miguel , Inhestern Lukas Benjamin , Mai Holger , Rinaldi Andrea , Miguel-Sanchez Adrian
Published: August 2017

Current key technologies to meet the future emission standards for internal combustion engines are downsizing, down-speeding, and advanced charging concepts. While turbocharging already combines high specific rated engine power with low fuel consumption, there is still potential for optimization to achieve prospective demands for fuel efficiency with low emissions. Using engine exhaust energy, the turbine underlies pulsating flow conditions from high towards zero mass flow at almost constant blade speed. The average turbine efficiency is then affected for the high blade to jet speed ratio conditions, which is very important at low engine loads during urban driving conditions. Since turbocharger performance is very sensitive for the overall engine efficiency, a very accurate measurement of the characteristic maps is desired to evaluate the thermodynamic behavior of the turbocharger and to ensure best possible matching. This paper presents a methodology to extend the turbine performance at low expansion ratio and to characterize the adiabatic efficiency in a wide operating range. This enables measuring turbines on a hot-gas test bench at very high blade to jet speed ratio and very low turbine flow to develop, improve, and validate reliable turbocharger models that can be used for full engine simulations. The industrial applicability has been proven from very low turbine power up to negative turbine power output simply based on using inlet guide vanes (IGV) upstream of the compressor. By generating a swirl in the compressor wheel rotating direction and pressurizing the inlet air, the compressor can be run as a turbine. Thus, the compressor provides power to the shaft and the turbine can be driven with very low flow power. The test campaign has been realized under quasi-adiabatic conditions to limit the heat transfer. While measuring at three different oil temperatures, the impact of remaining internal heat transfer has been taken into account. A turbocharger heat transfer model has also been used to correct residual heat flows from the obtained data set for all oil temperatures.
Thermal Stability Measurement of Alternative Jet Fuels Using Ellipsometry
Authors: Nash Leigh , Vasu Subith
Published: August 2017

Thermal stability is an important characteristic of alternative fuels that must be evaluated before they can be used in aviation engines. Thermal stability refers to the degree to which a fuel breaks down when it is heated prior to combustion. This characteristic is of great importance to the effectiveness of the fuel as a coolant and to the engine’s combustion performance. The thermal stability of Sasol IPK, a synthetic alternative to Jet-A, with varying levels of naphthalene has been studied on aluminum and stainless steel substrates at 300 to 400 °C. This was conducted using a spectroscopic ellipsometer to measure the thickness of deposits left on the heated substrates. Ellipsometry is an optical technique that measures the changes in a light beam’s polarization and intensity after it reflects from a thin film to determine the film’s physical and optical properties. It was observed that, as would be expected, increasing the temperature increased the deposit thickness for a constant concentration of naphthalene on both substrates. The repeatability of these measurements was verified using multiple trials at identical test conditions. Lastly, the effect of increasing the naphthalene concentration at a constant temperature was found to also increase the deposit thickness.
Life Extension of Power Turbine Disks Exposed to In-Service Corrosion Damage
Authors: Dua Dipankar , Khajavi Mohammad , Hill Matthew
Published: August 2017

Siemens Energy has a large fleet of aero derivative gas turbines installed in oil & gas and power generation applications globally. Over the past several years, several incidents of disk corrosion damage have been observed during the power turbine overhauls especially for the units fielded in marine environment. Most of the corrosion attack has been observed at disk firtree (blade attachment features), gas washed disk faces and torque transmission features. In all of these cases the mechanism for corrosion has been identified as Low Temperature Hot Corrosion (LTHC) also known as Type-II sulphidation.

Siemens has explored a number of potential operating and/or design options e.g. utilization of more corrosion resistant alloys and coatings to mitigate the effects of the corrosive attack on these disks. However, many of these proposals are either intermediate or long term solutions and are addressed more towards new power turbine disks rather than those that are already in service.

In the short term, Siemens technical solution has focused on development of proprietary repair processes to effectively remove the corrosion damage and extend the operational life of in-service disks as well as salvage some of the previously scraped disks.

This Paper discusses the methodology and findings from the development and characterization of the repair of the power turbine disks when exposed to corrosion and other operational damage.