bluuebunny/crossref_metadata_2025 huggingface dataset:
An Experimental Study of Mist/Air Film Cooling With Fan-Shaped Holes on an Extended Flat Plate—Part 1: Heat Transfer
Authors: Ragab Reda , Wang Ting
Published: August 2017

Motivated by the need to further improve film cooling in terms of both cooling effectiveness and coolant coverage area, the mist/air film cooling scheme is investigated through experiments using fan-shaped holes over an extended downstream length in this study. Both an existing wind tunnel and test facility, used in previous work, have been retrofitted. The first modification was extending the length of the flat plate test section to cover longer distances downstream of the injection holes, up to X/D = 100, in order to investigate whether mist cooling can be harnessed farther downstream where single-phase film cooling is not effective. X represents the axial distance downstream of the cooling hole of diameter D. The second modification was to incorporate fan-shaped (diffusion) holes which are proven to have a higher film cooling efficiency, than cylindrical holes. The objective is to investigate whether mist can further enhance the film cooling performance of the already highly effective fan-shaped holes. A phase Doppler particle analyzer (PDPA) system is employed to measure the droplet size, velocity, and turbulence information. An infrared camera and thermocouples are both used for temperature measurements. Part I is focused on the heat transfer result on the wall. The results show that, at low blowing ratios when the film is attached to the surface, the enhancement of the mist film cooling effectiveness, compared to the air-only case, on the centerline of the hole ranges from 40% in the near hole region to over 170% at X/D = 100. Due to the diffusive nature of the fan-shaped hole, the laterally averaged enhancement is on par with that on the centerline. The significant enhancement over the extended downstream distance from X/D = 40–100 is attributed to the evaporation time needed to evaporate all of the droplets. Each droplet acts as a cooling sink and flies over a distance before it completely vaporizes. This “distributed cooling” characteristic allows the water droplets to extend the cooling effects farther downstream from the injection location. At higher blowing ratios, when the cooling film is lifted off from the surface, the cooling enhancement drops below 40%. Although the enhancement in the near hole region X/D < 40 is about 20% lower than that achieved by using the cylindrical holes, the magnitudes of the mist adiabatic film cooling effectiveness using fan-shaped holes are still much higher than those of the cylindrical holes. Part II of this study is focused on analyzing the two-phase droplet multiphase flow behavior to explain the fundamental physics involved in the mist film cooling.
Two-Side Laser Processing Method for Producing High Aspect Ratio Microholes
Authors: Nasrollahi Vahid , Penchev Pavel , Dimov Stefan , Korner Lars , Leach Richard , Kim Kyunghan
Published: August 2017

Laser microprocessing is a very attractive option for a growing number of industrial applications due to its intrinsic characteristics, such as high flexibility and process control and also capabilities for noncontact processing of a wide range of materials. However, there are some constrains that limit the applications of this technology, i.e., taper angles on sidewalls, edge quality, geometrical accuracy, and achievable aspect ratios of produced structures. To address these process limitations, a new method for two-side laser processing is proposed in this research. The method is described with a special focus on key enabling technologies for achieving high accuracy and repeatability in two-side laser drilling. The pilot implementation of the proposed processing configuration and technologies is discussed together with an in situ, on-machine inspection procedure to verify the achievable positional and geometrical accuracy. It is demonstrated that alignment accuracy better than 10 μm is achievable using this pilot two-side laser processing platform. In addition, the morphology of holes with circular and square cross sections produced with one-side laser drilling and the proposed method was compared in regard to achievable aspect ratios and holes' dimensional and geometrical accuracy and thus to make conclusions about its capabilities.
Research on Active Control Strategy of Gas Turbine Secondary Air System in Different Ambient Temperature Conditions
Authors: Ji Jingjin , Huang Danping , Sun Bo , Peng Shuhong , Pan Chengxiong
Published: August 2017

Conventional gas turbine secondary air system in early stage typically uses a fixed throttle unit between supply side which is compressor bleeding point and demand side which is turbine blade. The cooling air mass flow strongly depends on the extraction pressure characteristics of compressor. Optimal amount of cooling air is supplied only in design point in this way. The cooling air mass flow would be either too much or too less in off design condition.

Recently, heavy duty gas turbine manufacturers introduced an active control method for secondary air system. The main strategy is to adjust the cooling air valve set point as a function of gas turbine load percentage in order to adjust cooling air pressure ratio and cooling air mass flow as well. With this active control strategy, cooling mass flow is separated from compressor extraction pressure characteristics, and it can provide a better way to deal with combustion contaminant issues. But it is still a problem that there is no dependence relationship between cooling air valve set point and operating ambient temperature in that strategy. That is to say, the cooling air pressure ratio is constant while varying ambient temperature at base load.

In order to quantitatively analyze this phenomena, a 1-dimensional integrated gas turbine thermodynamic analysis method is first applied to obtain the extraction pressure characteristics of compressor for all bleeding points. In the meantime, the optimal cooling air mass flow for turbine blades in different operating conditions is evaluated by a 0-dimensional heat transfer assessment method. A 1-dimensional fluid network analysis method is then employed to calculate the cooling air mass flow variation characteristics for 2 typical throttle configurations between compressor bleeding points and turbine blades, the first one is setting a fixed throttle unit, and the second one is setting constant cooling air pressure ratio by a cooling air control valve. Quantitative calculation results show that the cooling air supply will not always meet the optimal requirements at different ambient temperature conditions with neither of the 2 configurations.

This paper further optimized the active control strategy. With the optimized strategy, cooling air supply not only no longer depends on extraction characteristics of compressor, but also could be actively adjusted according to the optimal requirements of turbine blades at different ambient temperature conditions. Performance evaluation results show that the optimized active control strategy could enhance the overall efficiency without exceeding maximum allowable metal temperature of turbine blades.
System Eigenvalue Identification of Mistuned Bladed Disks Using Least-Squares Complex Frequency-Domain Method
Authors: Huang Yuan , Dimitriadis Grigorios , Kielb Robert E. , Li Jing
Published: August 2017

This paper presents the results from a research effort on eigenvalue identification of mistuned bladed rotor systems using the Least-Squares Complex Frequency-Domain (LSCF) modal parameter estimator. The LSCF models the frequency response function (FRF) obtained from a vibration test using a matrix-fraction description and obtains the coefficients of the common denominator polynomial by minimizing the least squares error of the fit between the FRF and the model. System frequency and damping information is obtained from the roots of the denominator; a stabilization diagram is used to separate physical from mathematical poles. The LSCF estimator is known for its good performance when separating closely spaced modes, but few quantitative analyses have focused on the sensitivity of the identification with respect to mode concentration. In this study, the LSCF estimator is applied on both computational and experimental forced responses of an embedded compressor rotor in a three-stage axial research compressor. The LSCF estimator is first applied to computational FRF data obtained from a mistuned first-torsion (1T) forced response prediction using FMM (Fundamental Mistuning Model) and is shown to be able to identify the eigenvalues with high accuracy. Then the first chordwise bending (1CWB) computational FRF data is considered with varied mode concentration by varying the mistuning standard deviation. These cases are analyzed using LSCF and a sensitivity algorithm is developed to evaluate the influence of the mode spacing on eigenvalue identification. Finally, the experimental FRF data from this rotor blisk is analyzed using the LSCF estimator. For the dominant modes, the identified frequency and damping values compare well with the computational values.
Analysis of the Effect of Multi-Row and Multi-Passage Aerodynamic Interaction on the Forced Response Variation in a Compressor Configuration: Part 1 — Aerodynamic Excitation
Authors: Schoenenborn Harald
Published: August 2017

The aeroelastic prediction of blade forcing is still a very important topic in turbomachinery design. Usually, the wake from an upstream airfoil and the potential field from a downstream airfoil are considered as the main disturbances. In recent years, it became evident that in addition to those two mechanisms Tyler-Sofrin modes, also called scattered or spinning modes, may have a significant impact on blade forcing.

In Schrape et al. [9] it was found that in multi-row configurations not only the next, but also the next but one blade row is very important as it may create a large circumferential forcing variation which is fixed in the rotating frame of reference.

In the present paper a study of these effects is performed on the basis of a quasi 3D multi-row and multi-passage compressor configuration. For the analysis a harmonic balancing code, which was developed by DLR Cologne, is used for various setups and the results are compared to full-annulus unsteady calculations. It is shown that the effect of the circumferentially different blade excitation is mainly contributed by the Tyler-Sofrin modes and not to blade-to-blade variation in the steady flow field.

The influence of various clocking positions, coupling schemes and number of harmonics onto the forcing is investigated. It is also shown that along a speed-line in the compressor map the blade-to-blade forcing variation may change significantly.

In addition, multi-row flutter calculations are performed, showing the influence of the upstream and downstream blade row onto aerodynamic damping.

The effect of these forcing variations onto random mistuning effects is investigated in the second part of the paper.
Impact of Wake Dispersion on Axial Compressor Performance
Authors: Hah Chunill
Published: August 2017

Rotor wake dispersion in a low-speed, one and half stage axial compressor is investigated in detail with a Large Eddy Simulation (LES). The primary focus is to quantify the total pressure recovery due to wake stretching and the total pressure loss from the rotor wake interaction with the stator blade boundary layer. The relative magnitude of the aerodynamic loss due to these two effects is examined at several radial locations. The spacing between the rotor and the stator was varied from 29% to 112% of the rotor axial chord at the mid span to investigate the effects of rotor wake decay before entering the stator passage. The current analysis indicates that the efficiency through the compressor stage is increased about 0.5% when the spacing between the rotor and the stator is decreased from 112% to 29% of the rotor axial chord at mid-span. 22% of the efficiency gain from the narrower axial gap is due to the wake recovery and 63% is due to the stronger unsteady pressure field at the exit of the rotor due to stage interaction. Total pressure loss/recovery across the stator varies significantly in the radial direction for the current compressor, which has a much lower aspect ratio. The total pressure recovery due to wake stretching is larger than the total pressure loss due to the unsteady boundary layer development on the stator blade from 20% to 35% of the span from the hub for 29% spacing and from 35% to 55% of the span for 112% spacing. Above 50% of the span, rotor tip clearance flow affects wake dispersion and the overall wake recovery is less than expected.
Vibration Response Analysis of Mistuned Bladed Disk With Under-Platform Damper: Effect of Variation of Contact Condition on Vibration Characteristics
Authors: Kaneko Yasutomo
Published: August 2017

Blades with a friction damper have been used in a steam turbine and a gas turbine to improve the blade reliability. In particular, for a gas turbine blade of the upstream stage, under-platform dampers have been widely used, where the damper pieces with various geometries are inserted into the platforms of the adjacent blades. The damper piece is designed so that its surface contacts the platform surface uniformly. However, the contact conditions of the damper piece (in other words, the equivalent stiffness and the damping caused by the damper piece) may change appreciably blade by blade because of the likes of manufacturing tolerance, blade deformation in operation, and wear of the damper piece. Therefore, it is essential to consider the mistuning effect caused by the variation of the contact condition of the damper piece in evaluating the vibration response of the bladed disk with the under-platform damper. In this study, a mistuned bladed disk with under-platform dampers is represented by the equivalent spring-mass model. Frequency response analysis and random response analysis are carried out using the direct method and Monte Carlo simulation. Carrying out an extensive parametric study, the effect of the variation of the contact condition caused by the damper piece on the vibration response of the bladed disk is clarified.
Flow Interactions Between Shrouded Power Turbine and Nonaxisymmetric Exhaust Volute for Marine Gas Turbines
Authors: Gao Jie , Lin Feng , Niu Xiying , Zheng Qun , Yue Guoqiang , Fu Weiliang , Wei Ming
Published: August 2017

The marine gas turbine exhaust volute is an important component that connects a power turbine and an exhaust system, and it is of great importance to the overall performance of the gas turbine. Gases exhausted from the power turbine are expanded and deflected 90 degrees in the exhaust volute, and then discharge radially into the exhaust system. The flows in the power turbine and the nonaxisymmetric exhaust volute are closely coupled and inherently unsteady. The flow interactions between the power turbine and the exhaust volute have a significant influence on the shrouded rotor blade aerodynamic forces. However, the interactions have not been taken into account properly in current power turbine design approaches.

The present study aims to investigate the flow interactions between the last stage of a shrouded power turbine and the nonaxisymmetric exhaust volute with struts. Special attention is given to the coupled aerodynamics and pressure response studies. This work was carried out using coupled computational fluid dynamics (CFD) simulations with the computational domain including a stator vane, 76 shrouded rotor blades, 9 struts and an exhaust volute. Three-dimensional (3D) unsteady and steady Reynolds-averaged Navier-Stokes (RANS) solutions in conjunction with a Spalart-Allmaras turbulence model are utilized to investigate the aerodynamic characteristics of shrouded rotors and an exhaust volute using a commercial CFD software ANSYS Fluent 14.0. The asymmetric flow fields are analyzed in detail; as are the unsteady pressures on the shrouded rotor blade. In addition, the unsteady total pressures at the volute outlet is also analyzed without consideration of the upstream turbine effects.

Results show that the flows in the nonaxisymmetric exhaust volute are inherently unsteady; for the studied turbine-exhaust configuration the nonaxisymmetric back-pressure induced by the downstream volute leads to the local flow varying for each shrouded blade and low frequency fluctuations in the blade force. Detailed results from this investigation are presented and discussed in this paper.
Measurement and Prediction of Centrifugal Compressor Axial Forces During Surge: Part 1 — Surge Force Measurements
Authors: Brun Klaus , Simons Sarah , Kurz Rainer , Pinelli Michele , Morini Mirko , Munari Enrico
Published: August 2017

Centrifugal compressor impellers and shafts are subject to severe fluctuating axial and radial forces when operating in surge. These forces can cause severe damage to the close clearance components of a centrifugal compressor such as the thrust and radial bearings, inter-stage and dry gas seals, and balance piston. Being able to accurately quantify the cyclic surge forces on the close clearance components of the compressor allows the user to determine whether an accidental surge event, or emergency shutdown (ESD) transient, has caused damage requiring inspection, repair, or part replacement. For the test, a 700 Hp (∼520 kW) industrial air centrifugal compressor was operated in surge at speeds ranging from 7,000 to 13,000 rpm and pressure ratios from 1.2 to 1.8. The axial surge forces were directly measured using axial load cells on the thrust bearings. Suction and discharge pressures, proximity probe axial shaft position, flows, and temperatures were also measured. Time domain and frequency plots of axial vibration and dynamic pulsations showed the impact of the operating conditions on surge force amplitudes and frequencies. A surge severity coefficient was also derived as a simple screening tool to evaluate the magnitude of potential damage to a compressor during surge.
Model Based Prediction of Off-Design Operation Condition NOx Emission From DLE Gas Turbine Combustors
Authors: Lauer Martin , Färber Jens , Reiß Frank , El Masalme Jaman
Published: August 2017

In this paper a semi-empirical method is presented to predict the NOX emission of a pilot stabilized, technically premixed gas turbine combustor for arbitrary combustor inlet conditions. The method is based on measurement data for a reference operation cycle and ISO ambient condition. A model based correction function is used to account for the change of NOX emission, which results from combustor inlet condition and/or operation cycle deviating from reference data.

The proposed method assumes that a pilot stabilized, technically premixed flame can be simplistically modeled as two separate combustion regimes: An assumed perfectly premixed regime, and an assumed diffusion flame regime. The emission of the combustor is then modeled as a superposition of the emissions from both individual combustion regimes. The emissions from the diffusion flame regime are the dominant contribution, although only a comparably small share of fuel is burnt in this regime.

The emission of the perfectly premixed combustion regime can be modeled as a sole function of flame temperature. On the other hand, the emissions from the diffusion regime are strongly affected by combustor inlet pressure and temperature, which are in turn a function of the gas turbine operation cycle and inlet conditions. To simplify matters, only the change in NOX emission from the diffusion regime, which results from deviation of combustor inlet pressure and/or temperature from values expected for a reference operation cycle and ISO ambient condition, is modeled. NOX emission data for this reference operation cycle is necessary experimental input for the method.

The proposed method can — with minimum experimental input — predict the NOX emission of a pilot stabilized, technically premixed combustor with good accuracy for arbitrary engine operation cycles and inlet conditions. It is aimed at providing reliable NOX emission estimates without the need for time intensive CFD or expensive experimental investigations. Typical applications are the assessment of a gas turbine’s NOX potential for feasibility studies of new operation or engine concepts with a proven combustor design, and for NOX estimates for out-of-experience ambient or site conditions for existing engines.
Blade Tip Leakage Loss Reduction by Means of Passive Tip Injection: Linear Cascade Wind Tunnel Results
Authors: Rejek Jonas , aus der Wiesche Stefan , Willinger Reinhard
Published: August 2017

In the open literature, an innovative concept for turbine blade tip leakage loss reduction by means of passive tip injection was recently proposed. The present paper presents experimental results obtained for an unshrouded turbine blade corresponding to a 50 % reaction stage. The experiments were performed in a low-speed linear cascade wind tunnel facility with air as working fluid. The effect of passive tip injection on the resulting loss was investigated by detailed five-hole-probe measurements. Cascades with three different tip gap heights and blades with and without passive injection were considered. Special attention was spent to the actual upstream conditions. The detailed flow field measurements showed that at the blade tip exit the leakage flow merged with the main flow and rolled up to a tip leakage vortex. The linear cascade wind tunnel results indicated a slight reduction of the resulting total pressure loss coefficient due to the passive tip injection. The observed tip leakage loss reduction was well comparable with the predictions of simplified analytical model.
Verification of Single Digit Emission Performance of a 24 MW Gas Turbine: SGT-600 3rd Generation DLE
Authors: Manrique Carrera Arturo , Geipel Philipp , Larsson Anders , Magnusson Rikard
Published: August 2017

The SGT-600 3rd generation DLE is a 24 MW industrial gas turbine which was recently upgraded from the SGT-600 2nd generation DLE. The upgraded SGT-600 is an addition to the existing low emissions gas turbine portfolio within Siemens. The objectives of the engine upgrade focused on increasing the lifetime of the components, lowering emissions and improving liquid fuel operation. In order to accomplish these objectives the combustion system was fully replaced, an improved gas fuel distribution was implemented and the first stage of the turbine was replaced. Furthermore, the liquid fuel injection system was enhanced in terms of fuel droplet distribution. Both gas and liquid fuel operation were confirmed in Siemens industrial gas turbine test facility located in Finspong, Sweden.

The upgraded combustion system originates from the SGT-700 gas turbine, a 33MW class engine, which consists of 18 “3rd generation DLE” burners that replaced the original “2nd generation DLE” ones that are normally incorporated in the SGT-600 DLE gas turbines. The first stage of the turbine has also been improved and less air is needed for cooling purposes. Moreover, an additional update on the compressor air bleed control was implemented. The liquid fuel burner hardware was optimized with new fuel injection lances manufactured by electron erosion techniques in order to improve the droplet distribution.

The SGT-600 3rd generation DLE is capable to operate in single digit conditions both in NOx and CO from 100% to 70% load and it is possible to extend single digit operation on NOx down to 50% load. CO emissions are below 80ppm at 50% load which is in compliance with current European regulation. Concerning liquid fuel operation the SGT-600 3rd generation DLE does not use water injection for emission control and is capable to reach below 58ppm NOx at full load while keeping CO below 40ppm down to 75% load.

The present work describes the main modifications performed during the engine upgrade and the results of the performed engine tests. Finally, it should be noted that the SGT-600 3rd generation DLE is an excellent example of the development/upgrade efforts within Siemens AG. The commercialization process of the SGT-600 3rd generation DLE has been initiated and three engines are already in commercial operation up to date, and two units will be installed in the near future.
High Temperature Brush Seal Development
Authors: Kirk Tracey , Bowsher Aaron , Crudgington Peter
Published: August 2017

The industry bristle material of choice for brush seals has been the cobalt-based alloy Haynes®25 (also known as L605) for over 30 years. Haynes®25 has excellent oxidation resistance and wear properties in brush seal applications up to temperature of 620°C [1148°F]. Above this temperature creep resistance becomes undesirable for brush seal bristles and has lead to alternative sealing solutions to be implemented in these turbine locations. Nickel-based alloys have been explored as an alternative for Haynes®25 but have been shown to gall badly and wear quickly in comparison.

As increases in turbine performance have resulted in an increase in operating temperatures, it has lead to a need to find a bristle material that performs at temperatures above the limitation of Haynes®25. Initial experimental data has been obtained for a new cobalt-based alloy that shows potential for use as a bristle material at temperatures above 620°C [1148°F]. Further experimental results also indicate that the material appears to have better wear characteristics than Haynes®25 and may prove to be a feasible alternative in some cases.

This paper outlines a material selection process for brush seals, along with development of the alloy for use within brush seals and details of the comparative testing carried out at Cross Manufacturing Company.
ASME PTC-10 and Heat Capacity Relations for Polytropic and Isentropic Compression Process of Real Gas
Authors: Taher Matt
Published: August 2017

ASME PTC-10 (reaffirmed 2009) serves as an internationally recognized standard factory acceptance and field performance testing for centrifugal compressors. It provides a test procedure to determine the thermodynamic performance of centrifugal compressors for gases conforming to ideal gas laws and for real gases. ASME PTC-10 defines ideal gases as those, which fall within the limits of table 3.3. The ratio of heat capacities is one of the parameters used to determine the limits of departure from ideal gas in table 3.3. However, ASME PTC-10 does not clearly define whether to use the ideal gas or a real gas method to calculate the ratio of heat capacities.

The relationship Ĉp – Ĉv = R, is valid for ideal gases, but not real gases. The validity of Ĉp – Ĉv = R is examined across a typical range of pressures and temperatures and a composition applicable to the natural gas industry.

Isentropic processes of ideal gases are accurately described with a simple relationship with the ratio of heat capacities. However, for real gases, that relationship is not valid and a more complex relationship is required for similar accuracy.

Thermodynamic relationships used in calculating isentropic and polytropic exponents are summarized. Limitations for real and ideal gas calculation methods are described. The deviations of real gas isentropic and polytropic volume and temperature exponents from ideal gas calculation methods are presented.
Component Matching of Centrifugal Compressors for Turbocharger Application
Authors: Chen Hua
Published: August 2017

Matching of various components (impeller exducer to impeller inducer, and vaneless diffuser and volute to the impeller) in a centrifugal compressor is critical for stage performance, but this is often neglected during compressor design and selection. This paper studies the importance of flow area matching for stage efficiency and how this matching process can be performed rapidly following a few simple principles. Methods for achieving optimum efficiency under different compressor operating conditions and size constraint are proposed and compared with experimental results. The purpose of this work is to draw attention to the key aspects of the matching, and provide easy-to-use design guidelines for engineers.
Turbine Blade Platform Film Cooling With Fan-Shaped Holes Under Simulated Swirl Purge Flow and Slashface Leakage Conditions
Authors: Chen Andrew F. , Shiau Chao-Cheng , Han Je-Chin
Published: August 2017

The combined effects of upstream purge flow, slashface leakage flow, and discrete hole film cooling on turbine blade platform film cooling effectiveness were studied using the pressure sensitive paint (PSP) technique. Detailed adiabatic film cooling effectiveness distributions on the platform were obtained and analyzed. As a continued study, discrete cylindrical holes [1] were replaced by laidback fan-shaped (10-10-5) holes which generally provide better film coverages on the endwall. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. A wide range of parameters were evaluated in this study. The coolant-to-mainstream mass flow ratio (MFR) was varied from 0.5%, 0.75%, to 1% for the upstream purge flow. For the platform film cooling holes and slashface gap, average blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1 (close to low-temperature experiments) to 1.5 (intermediate DR) and 2 (close to engine conditions) were also examined. Purge flow swirl effect was studied particularly at a typical swirl ratio of 0.6. The results provide the gas turbine engine community a better insight into various parametric effects on turbine blade platform film cooling with fan-shaped holes when the upstream swirl purge flow and slashface leakage flow were presented. Area-averaged film cooling effectiveness results were compared between cylindrical and fan-shaped holes under various parametric conditions. The results indicate that the fan-shaped holes provide superior film coverage than cylindrical holes for platform film cooling especially at higher blowing ratios and momentum flux ratios.
Development of Improvements to “Controlled Flow” Technology for Large Steam Turbines
Authors: Haller B. R. , Singh G. , Millington P. R. , Mund F. C. , Vernon K.
Published: August 2017

The aim of this paper is to present advances in the blading design for large steam turbines — ‘Controlled Flow’ technology. The purpose of the design is to improve the turbine efficiency in a cost neutral manner, adding value for the customer. Controlled Flow is a 3D design philosophy which passes more flow through the efficient middle sections of the blade, and less flow through the comparatively inefficient regions near the endwalls. It has been used for the Impulse Technology Blading ITB guide blades. The current improvement builds on the previous successful Controlled Flow design but incorporates the following new features:

- Ultra High Lift for the mean section (at significantly reduced axial width)

- Ultra High Back Surface Deflection for the mean section

- forward leading edge sweep.

The new guide delivers the same radial distribution of absolute fluid exit angle to the runner as the previous design.

Confirmatory model turbine tests demonstrated that the new guide delivered a stage efficiency improvement of 0.35%, above an already very high datum level.

The endwall sections of the guide are kept the same which maintains the mechanical strength of the diaphragm (same stress and deflection). Therefore, the new design can easily ‘slot-in’ and replace the previous design.

The following will be described in detail:

- History of the designs/background and design philosophy

- Flow physics

- Stage optimization and performance prediction

- Probabilistic analysis and robustness of the design

- Confirmatory model turbine testing and validation (comparison with design predictions).
Transonic Turbine Vane Suction Side Film Cooling With Showerhead Effect Using PSP Measurement Technique
Authors: Shiau Chao-Cheng , Chowdhury Nafiz H. K. , Han Je-Chin , Mirzamoghadam Alexander V. , Riahi Ardeshir
Published: August 2017

This work focuses on the parametric experimental study of film cooling effectiveness on the suction side of a scaled turbine vane under transonic flow condition. The experiments were performed in a five-vane annular sector cascade blowdown facility. The controlled exit Mach numbers were 0.7, 0.9, and 1.1, from high subsonic to transonic conditions. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of coolant-to-mainstream density ratios, ranging from 1.0, 1.5 to 2.0. Three row-averaged coolant-to-mainstream blowing ratios in the range 0.7, 1.0, and 1.6 are studied. The test vane includes three rows of radial-angle cylindrical holes around the leading edge and two rows of compound-angle shaped holes on the suction side. All the cooling holes are active in order to study the resultant film cooling on suction side as well as from leading edge. Pressure sensitive paint (PSP) technique was used to obtain the film cooling effectiveness distributions from suction side holes and the contribution from leading edge showerhead holes. This work shows the effects of coolant-to-mainstream blowing ratio, density ratio, and exit Mach number on the film cooling effectiveness as well as its interaction with a potential shock wave. The results indicate that when the cooling holes are located in a critical region on the vane suction surface, the parametric effect on film cooling performance will significantly deviate from the common trend for a typical hole geometry.
Turbine Vane Endwall Film Cooling Study From Axial-Row Configuration With Simulated Upstream Leakage Flow
Authors: Chowdhury Nafiz H. K. , Shiau Chao-Cheng , Han Je-Chin , Zhang Luzeng , Moon Hee-Koo
Published: August 2017

Endwall film cooling can be greatly improved if the leakage coolant flow from the upstream gap between the combustor and vane endwall is effectively engaged. In this study, such a full coverage film cooling design, called axial-row configuration, is considered and the performance is studied by measuring the film cooling effectiveness distribution using PSP technique. Experiments were performed in a blow-down wind tunnel cascade facility at the isentropic exit Mach number of 0.5 corresponding to inlet Reynolds number of 3.8 × 105, based on the axial chord. Passive turbulence grid was used to generate freestream turbulence level about 19 % with a length scale of 1.7 cm. The results are presented as two-dimensional film cooling effectiveness distributions on the endwall surface with pitchwise averaged distributions in the axial direction. The focus of this study is evaluating the effect of coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR) on a particular endwall cooling design. Increasing coolant amount for the upstream leakage exhibited increased local adiabatic cooling effectiveness levels with relatively uniform coverage area. However, the passage cooling was not improved at highest coolant MFR = 1.5% rather indicated an optimum value of MFR = 1% based on better coolant coverage on the endwall surface. For density ratio effect, results indicated the best performance at DR = 1.5.
On the Force Coefficients of a Flooded, Open Ends Short Length Squeeze Film Damper: From Theory to Practice (and Back)
Authors: San Andrés Luis , Den Sean , Jeung Sung-Hwa
Published: August 2017

Gas turbine aircraft engine manufacturers push for simple squeeze film damper (SFD) designs, short in length, yet able to provide enough damping to ameliorate rotor vibrations. SFDs employ orifices to feed lubricant directly into the film land or into a deep groove. The holes, acting as pressure sources (or sinks), both disrupt the film land continuity and reduce the generation of squeeze film dynamic pressure. Overly simple predictive formulations disregard the feedholes and deliver damping (C) and inertia (M) force coefficients not in agreement with experimental findings. Presently, to bridge the gap between simple theory and practice, the paper presents measurements of the dynamic forced response of an idealized SFD that disposes of the feedholes altogether. The short-length SFD, whose diameter D = 125 mm, has one end submerged (flooded) within a lubricant bath and the other end exposed to ambient. ISO VG 2 lubricant flows by gravity through the film land of length L = 25.4 mm and clearance c = 0.122 mm. From dynamic load tests over excitation frequency range 10–250 Hz, experimental damping coefficients (CXX, CYY) from the flooded damper agree well with predictions from the classical open ends model with a full film for small amplitude whirl motions (r/c << 1), centered and off-centered. Air ingestion inevitably occurs for large amplitude motions (r/c > 0.4) thus exacerbating the difference between predictions and tests results. For reference, identical tests were conducted with a practical SFD supplied with lubricant (Pin = 0.4 bar) via three orifice feedholes, 120° apart at the film land mid plane. A comparison of test results shows as expected, that for small amplitude (r/c ∼ 0.05) orbits, the flooded damper generates on average 30% more damping than the practical configuration as the latter’s feedholes distort the generation of pressure. For large amplitude motions (r/c > 0.4), however, the flooded damper provides slightly lesser damping and inertia coefficients than the SFD with feedholes whose pressurized lubricant delivery alleviates air ingestion in the film land. The often invoked open ends SFD classical model is not accurate for the practical engineered design of an apparently simple mechanical element.
Aeroelastic Analysis of NREL Wind Turbine
Authors: Lu Yaozhi , Zhao Fanzhou , Salles Loic , Vahdati Mehdi
Published: August 2017

The current development of wind turbines is moving toward larger and more flexible units, which can make them prone to fatigue damage induced by aeroelastic vibrations. The estimation of the total life of the composite components in a wind turbine requires the knowledge of both low and high cycle fatigue (LCF and HCF) data. The first aim of this study is to produce a validated numerical model, which can be used for aeroelastic analysis of wind turbines and is capable of estimating the LCF and HCF loads on the blade. The second aim of this work is to use the validated numerical model to assess the effects of extreme environmental conditions (such as high wind speeds) and rotor over-speed on low and high cycle fatigue.

Numerical modelling of this project is carried out using the Computational Fluid Dynamics (CFD) & aeroelasticity code AU3D, which is written at Imperial College and developed over many years with the support from Rolls-Royce. This code has been validated extensively for unsteady aerodynamic and aeroelastic analysis of high-speed flows in gas turbines, yet, has not been used for low-speed flows around wind turbine blades. Therefore, in the first place the capability of this code for predicting steady and unsteady flows over wind turbines is studied. The test case used for this purpose is the Phase VI wind turbine from the National Renewable Energy Laboratory (NREL), which has extensive steady, unsteady and mechanical measured data.

From the aerodynamic viewpoint of this study, AU3D results correlated well with the measured data for both steady and unsteady flow variables, which indicated that the code is capable of calculating the correct flow at low speeds for wind turbines.

The aeroelastic results showed that increase in crosswind and shaft speed would result in an increase of unsteady loading on the blade which could decrease the lifespan of a wind turbine due to HCF. Shaft overspeed leads to significant increase in steady loading which affects the LCF behaviour. Moreover, the introduction of crosswind could result in significant dynamic vibration due to forced response at resonance.
CFD Leakage Predictions of Unworn and Worn Labyrinth Seals With and Without Tooth Axial Offset
Authors: Chougule Hasham H. , Mirzamoghadam Alexander
Published: August 2017

The objective of this study is to develop a Computational Fluid Dynamics (CFD) based methodology for analyzing and predicting leakage of worn or rub-intended labyrinth seals during operation. The simulations include intended tooth axial offset and numerical modeling of the flow field. The purpose is to predict total leakage through the seal when an axial tooth offset is provided after the intended/unintended rub.

Results indicate that as expected, the leakage for the in-line worn land case (i.e. tooth under rub) is higher compared to unworn. Furthermore, the intended rotor/teeth forward axial offset/shift with respect to the rubbed land reduces the seal leakage. The overall leakage of a rubbed seal with axial tooth offset is observed to be considerably reduced, and it can become even less than a small clearance seal designed not to rub. The reduced leakage during steady state is due to a targeted smaller running gap because of tooth offset under the intended/worn land groove shape, higher blockages, higher turbulence and flow deflection as compared to worn seal model without axial tooth offset.
The Wing Foil: A Novel Compliant Radial Foil Bearing Design
Authors: Swanson Erik E. , O’Meara P. Shawn
Published: August 2017

To meet the challenging demands for high performance, affordable compliant foil bearings, a novel compliant support element has been developed. This recently patented, novel support element uses a multidimensional array of multiple, formed, cantilever “wing foil” tabs. The wing foil bearing has all the features required to achieve state of the art performance (Gen III for radial bearings). This paper describes two radial foil beings using the wing foil and the unique design features. Test data for a 31.75 mm diameter bearing operating in air and in steam up to 42 krpm are presented to demonstrate the performance of this bearing. It is shown to have low subsynchronous vibration and reasonable damping through rigid shaft critical speeds.
An Implicit Off-Design Deviation Angle Correlation of Axial Flow Compressor Blade Elements
Authors: Dong-run Wu , Jin-fang Teng , Xiao-qing Qiang , Jin-zhang Feng
Published: August 2017

This paper applies a new analytical/empirical method to formulate the off-design deviation angle correlation of axial flow compressor blade elements. An implicit function of deviation angle is used to map off-design deviation curves into linear correlations (minimum linear correlation coefficient R = 0.959 in this paper). Solution of the coefficients in the correlation is given through the study of classical theories and statistical analysis of the experimental data. The off-design deviation angle can be calculated numerically. The approach requires only knowledge of the blade element geometry. The comparison among 2 classical correlations and the new correlation proposed in this paper shows the new correlation has minimum error over the entire range of incidence angle while classical correlations show high reliability only in a limited range.

Experimental data in this paper is collected from NASA’s open technical reports. Rotors and stators are studied together. Considering there is significant deviation angle variation along spanwise direction, only data at 50% span is studied, if possible.

The error among experimental data, statistical regressions of the experimental data, and numerical results based on the new correlation is discussed. It has to be noted that the influence of the flow condition other than incidence angle is only being discussed but with less break through.
Modal Analyses of an Axial Turbine Blisk With Intentional Mistuning
Authors: Beirow Bernd , Figaschewsky Felix , Kühhorn Arnold , Bornhorn Alfons
Published: August 2017

The potential of intentional mistuning to reduce the maximum forced response is analyzed within the development of an axial turbine blisk for ship diesel engine turbocharger applications. The basic idea of the approach is to provide an increased aerodynamic damping level for particular engine order excitations and mode shapes without any significant distortions of the aerodynamic performance. The mistuning pattern intended to yield a mitigation of the forced response is derived from an optimization study applying genetic algorithms. Two blisk prototypes have been manufactured a first one with and another one without employing intentional mistuning. Hence, the differences regarding the real mistuning and other modal properties can be experimentally determined and evaluated as well. In addition, the experimental data basis allows for updating structural models which are well suited to compute the forced response under operational conditions. In this way, the real benefit achieved with the application of intentional mistuning is demonstrated.
Modeling of a Turbofan Engine With Ice Crystal Ingestion in the NASA Propulsion System Laboratory
Authors: Veres Joseph P. , Jorgenson Philip C. E. , Jones Scott M. , Nili Samaun
Published: August 2017

The main focus of this study is to apply a computational tool for the flow analysis of the turbine engine that has been tested with ice crystal ingestion in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The PSL has been used to test a highly instrumented Honeywell ALF502R-5 (LF11) turbofan engine at simulated altitude operating conditions. Test data analysis with an engine cycle code and a compressor flow code was conducted to determine the values of key icing parameters that can indicate the risk of ice accretion, which can lead to engine rollback (un-commanded loss of engine thrust). The full engine aerothermodynamic performance was modeled with the Honeywell Customer Deck specifically created for the ALF502R-5 engine. The mean-line compressor flow analysis code, which includes a code that models the state of the ice crystal, was used to model the air flow through the fan-core and low pressure compressor. The results of the compressor flow analyses included calculations of the ice-water flow rate to air flow rate ratio (IWAR), the local static wet bulb temperature, and the particle melt ratio throughout the flow field. It was found that the assumed particle size had a large effect on the particle melt ratio, and on the local wet bulb temperature. In this study the particle size was varied parametrically to produce a non-zero calculated melt ratio in the exit guide vane (EGV) region of the low pressure compressor (LPC) for the data points that experienced a growth of blockage, and resulted in an engine called rollback (CRB). At data points where the engine experienced a CRB having the lowest wet bulb temperature of 492 R at the EGV trailing edge, the smallest particle size that produced a non-zero melt ratio (between 3%–4%) was on the order of 1μm. The particle size was varied from 1μm – 9.5μm to achieve the target melt ratio. For data points that did not experience a CRB which had static wet bulb temperatures in the EGV region below 492 R, a non-zero melt ratio could not be achieved even with a 1μm ice particle size. The highest value of static wet bulb temperature for data points that experienced engine CRB was 498 R with a particle size of 9.5μm. Based on this study of the LF11 engine test data, the range of static wet bulb temperature at the EGV exit for engine CRB was in the narrow range of 492 R – 498 R, while the minimum value of IWAR was 0.002. The rate of blockage growth due to ice accretion and boundary layer growth was estimated by scaling from a known blockage growth rate that was determined in a previous study. These results obtained from the LF11 engine analysis formed the basis of a unique icing wedge which defines a region of ice accretion risk that are being applied to other turbofan engines in order to predict the risk of ice accretion at various altitudes and operating conditions.