bluuebunny/crossref_metadata_2025 huggingface dataset:
Testing of Synthesized Aromatic Kerosene (SAK) Aviation Fuel Blends at Simulated Altitudes
Authors: Canteenwalla Pervez , Davison Craig R. , Chishty Wajid A. , Ginestra Cynthia , Dally Brice
Published: September 2016

A number of pathways for producing new alternative fuels are emerging. One such fuel is Hydrodeoxygenated Synthesized Aromatic Kerosene (HDO-SAK) which is composed of approximately 95% mono-aromatic compounds. This fuel is intended as a blending component to increase the aromatic content of other synthetic fuels that do not contain aromatic compounds and can be used to create either semi-synthetic or fully-synthetic jet fuel for use in aviation.

This paper presents the results of engine tests using the HDO-SAK blended with Synthesized Paraffinic Kerosene from Hydroprocessed Esters and Fatty Acids (HEFA-SPK) to create a fully-synthetic fuel with aromatic content that matched that of a comparison conventional jet fuel. Both the HDO-SAK blended fuel and conventional Jet A were tested in a Microturbo TRS-18 turbojet engine at a range of engine conditions and simulated altitudes up to 8,600 m in the National Research Council Canada (NRC) Research Altitude Test Facility (RATFac). This paper details the engine performance results from this testing which represents the first time this particular alternative fuel has been tested in an engine.
DNS-Based Analysis of RANS Predictions of a Low-Pressure Turbine Cascade
Authors: Müller Christoph , Baier Ralf-D. , Seume Joerg R. , Herbst Florian
Published: September 2016

The state of the art design of turbomachinery components is based on RANS solutions. RANS solvers model the effects of turbulence and boundary layer transition and therefore allow for a rapid prediction of the aerodynamic behavior. The only drawback is that modeling errors are introduced to the solution. Researchers and CFD developers are working on reducing these errors by improved model calibrations which are based on experimental data. These experiments do not typically, however, offer detailed insight into three-dimensional flow fields and the evolution of model quantities in an actual machine. This can be achieved through a direct step-by-step comparison of model quantities between RANS and DNS. In the present work, the experimentally obtained model correlations are be recomputed based on DNS of the same turbine profile simulated by RANS. The actual local values are compared to the modeled RANS results, providing information about the source of model deficits. The focus is on the transition process on the blade suction side and on evaluating the development of turbulent flow structures in the blades wake. It is shown that the source of disagreement between RANS and DNS can be traced back to three major deficiencies that should be the focus of further model improvements.
Numerical Study on Aero-Thermal Performance of HP Turbine Endwalls Under Influence of Hot Streak and High Mainstream Turbulence
Authors: Wang Zhiduo , Zhang Wenhao , Liu Zhaofang , Zhang Chen , Feng Zhenping
Published: September 2016

In this paper, unsteady RANS simulations were performed at two hot streak (HS) circumferential positions with inlet turbulence intensity of 5% and 20%. The interacted HS and high mainstream turbulence effects on endwall heat transfer characteristics of a high-pressure (HP) turbine were discussed by analyzing the flow structures and presenting the endwall adiabatic wall temperature, heat transfer coefficient (HTC) and heat flux distributions. The results indicate that both the wall temperature and HTC increase with the turbulence intensity at most stator endwall regions. In addition, the increase of wall temperature plays a greater role than HTC of influencing the wall heat flux. However, higher turbulence intensity decreases the intensity of the stator passage horse-shoe vortex, also the corresponding region HTC and heat flux are reduced. In rotor passage, the variation of HS circumferential position would alter the hub and casing endwall temperature, however, the discrepancy is weakened at higher turbulence. The elevated HS attenuation at higher turbulence results in temperature augmentation at the leading edge of rotor hub and casing endwalls, while temperature decrease after 50% axial chord, thus obtains more uniform temperature distributions on the endwalls. However, the rotor endwall HTC is only augmented significantly at the leading edge on hub endwall, and pressure side and downstream of trailing edge on casing endwall. Variation of HTC and adiabatic wall temperature jointly determines the rotor hub and casing endwall heat flux, and the temperature variation has dominant effects in the most regions. In general, the variation of adiabatic wall temperature and HTC should be considered simultaneously when analyzing the turbine endwall heat transfer characteristics.
Experimental Characterization of Rotor Convective Heat Transfer Coefficients in the Vicinity of a Leaf Seal
Authors: Pelegrin-Garcia Juan D. , Gillespie David R. H. , Pekris Michael J. , Ganin Leonid , Franceschini Gervas
Published: September 2016

Leaf seals are filament seals for use at static to rotating interfaces in the engine secondary air system. They offer reduced leakage rates and better off-design performance over conventional labyrinth seals. If compared with advanced brush seals, leaf seals are more compliant due to their lower stiffness and can withstand higher axial pressure differences. Although leaf seals can exhibit hydrodynamic air-riding, this is not always the case and seal-rotor contact can occur. As a result, friction between the leaf tips and the rotor causes heat generation and wear. To predict the diameter of the rotating shaft and the seal life, the shaft and seal interface temperature needs to be estimated. In the steady state this is determined by the ratio of convective heat transfer through the seal to that through the shaft. To that end, the convective heat transfer characteristics of the flow over the shaft around the seal is required to build accurate thermal models. In this paper the convective heat transfer coefficient (htc) distribution in the close vicinity of a typical leaf seal is investigated in a new test facility. The experimental setup and test method are described in detail and accuracy considerations are included. The methodology employed to derive htc is explained with reference to an analogous CFD model. The importance of the choice of an appropriate driving gas temperature is demonstrated. Experimental htc maps are presented for a blow-down seal geometry operating over a range of engine representative pressure ratios. Insight is gained into the flow field characteristics and heat transfer around the seal.
Investigations on Direct and Hybrid Sound Predictions
Authors: Kendall-Torry Christofer , Danner Florian
Published: September 2016

As a result of dissipative and dispersive properties of numerical methods the accuracy of direct noise prediction degrades with increasing distance from the source. Hybrid approaches are hence applied to predict the acoustic farfield, which rely on an integration of disturbance quantities extracted from the nearfield flow solutions. In order to elaborate the influence of the integration limits on sound prediction, different approaches of the Ffowcs Williams and Hawkings method are systematically applied to an advanced counter rotating propfan configuration within the present study. The solutions of various permeable and impermeable integration surfaces are analysed with respect to nearfield sound radiation and compared against direct sound predictions from the compressible Reynolds-averaged Navier-Stokes solutions which likewise serve as input for the extrapolation routine. Due to the flexibility of the routine, source terms and zones can be selectively excluded from the surface integration, allowing a systematic identification of the origin of dissimilar sound prediction. Subsequent farfield analyses are used to conclude on the propagation and persistency of differences identified in the nearfield predictions.
ALGOR Gas Turbine Performance Modular Tool: One Dimensional Turbine Module Validation
Authors: Piola Stefano , Canepa Roberto , Silingardi Andrea , Cecchi Stefano , Carcasci Carlo , Bolgè Edoardo
Published: September 2016

One dimensional codes play a key role in gas turbine performance simulation: once they are calibrated they can give reliable results within very short computational time if compared to two or three dimensional analysis. Thanks to their ability to quickly evaluate flow, pressure and temperature along the energy conversion from fluid to shaft or reverse, one dimensional tools fit the requirements of modular-structured program for the simulation of complete gas turbine. In ASEN experience, ALGOR heat and mass balance software is used as a platform for system integration between each disciplines by means of a modular structure in which a large number of modules, chosen from the available library, are freely connected allowing to potentially analyze any gas turbine engine configuration. ALGOR code provides advanced cycle calculation capabilities for example in case that cooling and secondary air system layout modification have to be considered in design process. In these situations, a turbine map-based approach is hardly applicable, while a one dimensional aerodynamic row by row simulation can provide a suitable method for off-design turbine behavior prediction. In ASEN practice, ALGOR turbine module is calibrated at design point on one dimensional data provided by turbine designers and is then adopted for the engine configuration optimization or off-design performance evaluation. This paper presents the validation of the off-design performance prediction given by the ALGOR embedded 1D turbine model comparing calculated results with experimental ones. The warm air full scale test rig investigated within the GE-NASA “Energy Efficient Engine” program for the aerodynamic evaluation of a two stages high pressure turbine has been chosen as validation case. It includes both experimental performance maps varying turbine operating conditions such as speed and pressure ratio extending to the sub-idle and starting region and an analysis of cooling flow variation effect on turbine performance. Literature available loss and exit flow angle correlations are implemented and compared to experimental data. The results given by each of them are analyzed to appreciate their accuracy in evaluating efficiency and flow variations. In addition the paper shows the ability of the 1D turbine module to consider secondary air system modification impact on performance comparing calculated results to experimental ones.

Literature correlations tuning on proprietary experimental results could further improve the tool performance for the off-design evaluation of ASEN turbine geometries.
Multi-Stage Coupled Forced Response of Aircraft Engine Compressor and Turbine Bladed Discs
Authors: Rzadkowski Romuald , Maurin Artur
Published: September 2016

Considered here is the effect of multistage coupling on the dynamics of an aircraft engine rotor with eight mistuned bladed discs on a drum-disc shaft during foreign object ingestion (FOI). In the dynamic model, each disc had a different number of rotor blades. Free and forced vibrations were examined using finite element models of single rotating blades, bladed discs and an entire rotor with bladed discs. Calculations of the mode shapes of flexible mistuned bladed disc-shaft assemblies took into account simultaneous excitations of the first and second stages of the compressor and the turbine bladed disc with 0EO, 1EO and 2EO. The thus obtained maximal stress values of all of the rotor blades were carefully examined and compared with a tuned system to discover resonance conditions and coupling effects. Our investigation has shown that mistuning changes the stress distribution in individual rotor blades and the level of maximum stress increases in relation to single or multiple bladed disc excitation.
ALGOR Gas Turbine Performance Modular Tool: Map Based Modules Development
Authors: Canepa Roberto , Piola Stefano , Pirotta Marco , Silingardi Andrea , Bonzani Federico , Carcasci Carlo
Published: September 2016

Commercially available or in-house developed performance tools, mostly based on heat and mass balance, are nowadays widespread among Universities, consulting companies and utilities. Generally these software are based on main gas turbine measurable information and, yet accurate on global performance estimation, are limited in the level of insight on component performance they can provide and also in the range of analysis, generally limited to engine possible operating points. On the other hand, the tools adopted by OEMs generally differ for components (compressor, turbine and combustor) and secondary air system details. In ASEN experience ALGOR heat and mass balance software is used as a platform for system integrations between each disciplines by means of a modular structure in which a large number of modules, chosen from the available library, are freely connected allowing to potentially analyze any gas turbine engine configuration.

This paper describes the structure and the implementation of latest ALGOR updates, developed by ASEN and University of Florence, aimed at creating new map based modules for compressor and secondary air system. With this approach, component performance coming from field data can be continuously adopted to refine the reliability of calculation. Furthermore, nonlinear phenomena occurring in stationary and rotating cooling passages can be evaluated only with devoted calculation tools, which output can be conveniently translated in flow functions maps. Thanks to ALGOR modular structure, with these newly available (map-interpolating) modules, additional levels of analysis are allowed, ranging from “cycle deck” map-matching level, to mixed modelling in which map based modules are linked with 1D mean line analysis modules. Moreover their use can be also foreseen in ASEN conceptual design approach in which just map reference points are adjusted to reflect expected technological leaps required by engine upgrade.
Numerical and Experimental Investigation of Return Channel Vane Aerodynamics With 2D and 3D Vanes
Authors: Hildebrandt A. , Schilling F.
Published: September 2016

The present paper deals with the numerical and experimental investigation of the effect of return channel dimensions of a centrifugal compressor stage on the aerodynamic performance. Three different return channel stages were investigated, two stages comprising 3D (three-dimensional) return channel blades and one stage comprising (2D) two-dimensional RCH (Return Channel) vanes. The analysis was performed regarding both the investigation of overall performance (stage efficiency, RCH total pressure loss coefficient) and detailed flow field performance. For detailed experimental flow field investigation at the stage exit, six circumferentially traversed three-hole probes were positioned downstream the return channel exit in order to get two-dimensional flow field information. Additionally, static pressure wall measurements were taken at the hub and shroud pressure and suction side of the 2D and 3D return channel blades. The return channel system overall performance was calculated by measurements of the circumferentially averaged 1D flow field downstream the diffuser exit and downstream the stage exit. Dependent on the type of return channel blade, the numerical and experimental results show a significant effect on the flow field overall and detail performance. In general, satisfactory agreement between CFD-prediction and test-rig measurements was achieved regarding overall and flow field performance. In comparison with the measurements, the CFD calculated stage performance (efficiency and pressure rise coefficient) of all 3D-RCH stages was slightly over-predicted. Very good agreement between CFD and measurement results was found for the static pressure distribution on the RCH wall surfaces while small CFD-deviations occur in the measured flow angle at the stage exit, dependent on the turbulence model selected.
Potential of Micro Turbine Based Propulsion Systems for Civil UAVs: A Case Study
Authors: Marcellan A. , Visser W. P. J. , Colonna P.
Published: September 2016

There is a high potential for civil applications of Unmanned Aerial Vehicles (UAV) in areas such as goods transport, telecommunication, remote monitoring and sensing, surveillance, search and rescue, and disaster management. Developments in areas such as telecommunication, control and information technology offer opportunities for long range remotely or automatically piloted missions. This requires efficient and light-weight small propulsion systems.

The potential of turboprop propulsion for civil UAVs using micro turbine technology has been explored and compared with existing concepts, such as piston engine driven propellers. Different propulsion concepts have been analyzed and the application areas where advanced turboprops would be superior to other systems such as reciprocating engines and electric motors, identified. However, turboprop engines of the small power capacity required for the aircraft concepts and missions considered are not currently available with competitive performance.

A conceptual design study of a micro turboprop engine has been performed by downscaling an existing reference engine. Scale effects on efficiency have been taken into account, as well as effects of technological progress. Engine cycle optimization has been carried out and the effects of turbine inlet temperature, compressor pressure ratio, engine size, and component efficiency have been investigated. An aerodynamic and flight performance model of a baseline UAV has been developed in order to predict mission performance. This model has been coupled to a turboprop model to evaluate system performance with different engine configurations for the selected mission.

The outcome of the study provides information about the technological improvements in terms of cycle efficiency required to make the micro-turboprop a competitive solution. The Propulsion and Power group of Delft University of Technology will pursue these R&D goals in an attempt to contribute to the development of civil UAV technology.
Effusion Cooling System Optimization for Modern Lean Burn Combustor
Authors: Andreini Antonio , Becchi Riccardo , Facchini Bruno , Mazzei Lorenzo , Picchi Alessio , Peschiulli Antonio
Published: September 2016

Stricter legislation limits concerning NOx emissions are leading main aero-engine manufacturers to update the architecture of the combustors towards the implementation of lean burn combustion concept. Cooling air availability for the thermal management of combustor liners is significantly reduced, demanding even more effective liner cooling schemes. The state-of-the-art of liner cooling technology is represented by effusion cooling, consisting in a very efficient cooling strategy based on multi-perforated liners, where metal temperature is lowered by the combined protective effect of coolant film and heat removal inside the holes. The present research study aims at deepening the knowledge of effusion systems, exploiting the results of a thorough experimental campaign carried out in two different planar test rigs, equipped with a complete liner cooling scheme composed by slot injection and effusion array.

The film cooling protection was analysed using PSP (Pressure Sensitive Paint) technique, while the effect of cooling injection and extraction from the annulus on heat transfer distribution were studied by means of TLC (Thermochromic Liquid Crystals) thermography. Thermal measurements were supported by flow field investigation with standard 2D PIV (Particle Image Velocimetry) in order to highlight the typical velocity distributions generated by a realistic lean injector. These detailed experimental data were exploited in a 1D thermal flow-network solver that allows to better assess the main cooling mechanisms characterising the proposed cooling system. Moreover, an optimized cooling configuration with enhanced back-side convective cooling was proposed and compared with the standard configuration in terms of metal temperature and cooling consumption.
Experimental Investigation of Turbine Stage Flow Field and Performance at Varying Cavity Purge Rates and Operating Speeds
Authors: Dahlqvist Johan , Fridh Jens
Published: September 2016

The aspect of hub cavity purge has been investigated in a high-pressure axial low-reaction turbine stage. The cavity purge is an important part of the secondary air system, used to isolate the hot main annulus flow from cavities below the hub level. A full-scale cold-flow experimental rig featuring a rotating stage was used in the investigation, quantifying main annulus flow field impact with respect to purge flow rate as it was injected upstream of the rotor.

Five operating speeds were investigated of which three with respect to purge flow, namely a high loading case, the peak efficiency, and a high speed case. At each of these operating speeds, the amount of purge flow was varied across a very wide range of ejection rates.

Observing the effect of the purge rate on measurement plane averaged parameters, a minor outlet swirl decrease is seen with increasing purge flow for each of the operating speeds while the Mach number is constant. The prominent effect due to purge is seen in the efficiency, showing a similar linear sensitivity to purge for the investigated speeds. An attempt is made to predict the efficiency loss with control volume analysis and entropy production.

While spatial average values of swirl and Mach number are essentially unaffected by purge injection, important spanwise variations are observed and highlighted. The secondary flow structure is strengthened in the hub region, leading to a generally increased over-turning and lowered flow velocity. Meanwhile, the added volume flow through the rotor leads to higher outlet flow velocities visible in the tip region, and an associated decreased turning. A radial efficiency distribution is utilized, showing increased impact with increasing rotor speed.
Analysis of a LPT Rotor Blade for a Geared Engine: Part I — Aero-Mechanical Design and Validation
Authors: Giovannini Matteo , Rubechini Filippo , Marconcini Michele , Arnone Andrea , Bertini Francesco
Published: September 2016

The rotational speed of low pressure turbines (LPT) for geared turbofan applications is significantly increased looking for potential benefit in performance, weight and overall dimensions. As a drawback, the high speed LPT are characterized by critical mechanical constraints due to the large centrifugal stresses in conjunction with the use of lightweight materials. The present activity was carried out in the framework of the Clean Sky European research project ITURB (Optimal High-Lift Turbine Blade Aero-Mechanical Design), aimed at designing and validating a turbine blade for a geared open rotor engine. This two-part paper presents the redesign and the analysis of an optimized rotor blade starting from a baseline configuration, representative of a state-of-the-art LPT rotor. In the redesign activity high standard of performance was required in conjunction with tight mechanical and geometrical constraints. The design strategy was based on an effective multi-objective optimization strategy. The aerodynamic performance was evaluated by means of 3D steady multi-row viscous computations using a two-equation k-ω turbulence model. At the same time, the mechanical integrity checks were mainly based on the evaluation of the maximum rotor tensile stress due to centrifugal forces. A simplified and very fast tool was developed in order to compute the centrifugal stress. Finally a response-surface approach based on neural-networks (ANNs) was adopted for the design space exploration. The design was validated by means of a comprehensive experimental campaign carried out in a low-speed turbine single-stage facility. A comparison between the numerical and experimental results is presented in terms of the main rotor performance for a fixed Reynolds number while varying the rotor incidence angle. Unsteady numerical analysis of both the baseline and the optimized blade were carried out by using a multi-equation, transition-sensitive, turbulence model and considering the boundary conditions measured on the test rig.
Unsteady Heat Transfer and Pressure Measurements on the Airfoils of a Rotating Transonic Turbine With Multiple Cooling Configurations
Authors: Nickol Jeremy B. , Mathison Randall M. , Dunn Michael G. , Liu Jong S. , Malak Malak F.
Published: September 2016

Measurements are presented for a high-pressure transonic turbine stage operating at design-corrected conditions with forward and aft purge flow and blade film cooling in a short-duration blow-down facility. Four different film-cooling configurations are investigated: simple cylindrical-shaped holes, diffusing fan-shaped holes, an advanced-shaped hole, and uncooled blades. A rainbow turbine approach is used so each of the four blade types comprise a wedge of the overall bladed disk and are investigated simultaneously at identical speed and vane exit conditions. Double-sided Kapton heat-flux gauges are installed at midspan on all three film-cooled blade types, and single-sided Pyrex heat-flux gauges are installed on the uncooled blades. Kulite pressure transducers are installed at midspan on cooled blades with round and fan-shaped cooling holes. Experimental results are presented both as time-averaged values and as time-accurate encoder-averages. In addition, the results of a steady RANS CFD computation are compared to the time-averaged data.

The computational and experimental results show that the cooled blades reduce heat transfer into the blade significantly from the uncooled case, but the overall differences in heat transfer among the three cooling configurations is small. This challenges previous conclusions for simplified geometries that show shaped cooling holes outperforming cylindrical holes by a great margin. It suggests that the more complicated flow physics associated with an airfoil operating in an engine-representative environment reduces the effectiveness of the shaped cooling holes. The experimental results appear to show a small benefit to the advanced cooling holes, but this is on the order of the variation caused by changes in the alignment of heat-flux gauges with cooling holes.
Study on the Effect of Inlet Geometry on the Noise of an Axial Fan, With Involvement of the Phased Array Microphone Technique
Authors: Benedek Tamás , Vad János
Published: September 2016

The paper presents a comparative case study in which a free-inlet, free-outlet industrial ventilating fan has been equipped with various inlet geometries. The original short-tapered entry has been replaced by a standardized bellmouth entry, resulting in remarkable noise reduction. The experimentation presented herein is adaptable to industrial onsite diagnostics. The upstream-radiated broadband noise associated with rotating sources has been mapped in a spatially resolved manner, by means of a Phased Array Microphone system and a Rotating Source Identifier beamforming algorithm. The acoustic measurements have been supplemented with aerodynamic measurements on the inlet velocity profile, and with Computational Fluid Dynamics studies. The acoustic data have been processed for enabling their evaluation in association with the aerodynamic operation of the elemental rotor cascades in a two-dimensional approach, and also for their interpretation in relationship to three-dimensional flow phenomena such as tip leakage flow. For this purpose, the acoustic data have been presented in the form of circumferentially-averaged noise profiles along the blade span, as well as noise source imaging maps. The studies reveal the following acoustic benefits of reconfiguring the original short-tapered entry to the bellmouth entry. A peripheral separation zone is characteristic for the short-tapered entry, provoking double-leakage tip clearance flow, being the dominant source of noise at higher frequencies. Such a peripheral separation zone is suppressed by the bellmouth inlet, and thus, the double-leakage flow and the related noise is eliminated. Farther away from the tip, along the dominant portion of the span, the moderation of endwall blockage due to suppressing the peripheral separation zone has led to the reduction of the rotor inlet velocity, thus moderating the noise associated with the suction side boundary layer developing on the blades.
An Investigation of Coolant Within Serpentine Passages of a High-Pressure Axial Gas Turbine Blade
Authors: Nickol Jeremy B. , Mathison Randall M. , Dunn Michael G. , Liu Jong S. , Malak Malak F.
Published: September 2016

Cooling flow behavior is investigated within the multiple serpentine passages with turbulators on the leading and trailing walls of an axial gas turbine blade operating at design-corrected conditions with accurate external flow conditions. Pressure and temperature measurements at midspan within the passages are obtained using miniature butt-welded thermocouples and miniature Kulite pressure transducers. These measurements, as well as airfoil surface pressure field data from a full CFD simulation, are used as boundary conditions for a model that provides quantitative values of film-cooling blowing ratio for each film cooling hole on the blade. The model accounts for the continuously changing cross-sectional area and shape of the channels, frictional pressure loss, convective heat transfer from the solid portion of the blade, massflow reduction as coolant bleeds out through film-cooling or impingement holes, compressibility effects, and the effects of blade rotation.

The results of the model provide detailed coolant ejection information for a film-cooled rotating turbine airfoil operating at design-corrected conditions, and also accounts for the highly variable freestream conditions on the airfoil. While these values are commonly known for simpler experimental geometries, they have previously either been unknown or estimated crudely for full-stage experiments of this nature. The better-quantified cooling parameters provide a bridge for better comparison with the wealth of film-cooling work already reported for simplified geometries. The calculation also shows the significant range in blowing ratio that can arise even among a single row of cooling holes associated with one of the turbulated passages, due to significant changes in both coolant, and local freestream massfluxes.
Operational Experience of the SSS (Synchro-Self-Shifting) Clutch for Naval Marine Applications
Authors: Hendry Morgan L. , Bellamy Nicholas
Published: September 2016

The Synchro-Self-Shifting (SSS) overrunning clutch is well known, particularly in the Naval Marine field. This paper reviews the clutch operating principle, then outlines some of the service experience since 1962, particularly in naval main propulsion drives beginning with CODOG, CODAG, COGOG and COGAG plant, and then the experience with more recent applications such as combined electric motor propulsion with either gas turbines or diesel engines and hybrid electric plants.

Extra features are then described such as a lock-out control as is usually necessary for turbine applications to permit turbine testing, e.g., when in harbor; also a lock-in control as is essential when the clutch has to transmit power in both directions of rotation. Various clutch mounting arrangements will be presented with respective advantages. The paper concludes with information regarding reliability during many years of service experience.
Gas Turbine Combustor Rig Development and Initial Observations at Cold and Reacting Flow Conditions
Authors: Gomez-Ramirez David , Kedukodi Sandeep , Gadiraju Siddhartha , Ekkad Srinath V. , Moon Hee-Koo , Kim Yong , Srinivasan Ram
Published: September 2016

The present paper describes the first phase of the design and development of a realistic, high-pressure, full-scale research gas turbine combustor at Virginia Tech. The final test rig will be capable of operating at inlet temperatures of 650 K, pressures up to 9.28 Bar (120 psig), maximum air inlet flow rates of 1.27 kg/s (2.8 lbm/s), and allow for variations in the geometry of the combustor model. The first phase consists of a low-pressure (atmospheric) optical combustor for heat transfer and flow-field measurements at isothermal and reacting conditions. The combustor model is equipped with an industrial low emission fuel injector from Solar Turbines Incorporated, used in their land based gas turbine Taurus-60. The primary objective of the developed rig is to provide additional insight into the heat transfer processes that occur within gas turbine combustors, primarily the convective component, which has not been characterized. A future phase of the test rig development will incorporate a pressure vessel that will allow for the operation of the combustor simulator at higher pressures. In the present publication, the design methodology and considerations, as well as the challenges encountered during the design of the first phase of the simulator are briefly discussed. An overview is given on the design of the instrumentation and process piping surrounding the test rig, including ASME codes followed as well as the instrumentation and equipment selected. A detailed description of the test section design is given, highlighting the design for high temperature operation.

As an example of the capabilities of the rig, representative measurements are presented. Characterization of the isothermal flow field using planar Particle Image Velocimetry (PIV) at a Reynolds number of 50 000 was performed and compared with flame imaging data at the same inlet conditions operating at an equivalence ratio of 0.7. The data suggests that the flame location follows the maximum turbulent kinetic energy as measured in the isothermal field. Representative data from the computational efforts are also presented and compared with the experimental measurements. Future work will expand on both reacting and isothermal PIV and heat transfer measurements, as well as computational validations.
Entrance Aspect Ratio Effect on S-Duct Inlet Performance at High-Subsonic Flow
Authors: Asghar Asad , Stowe Robert A. , Allan William D. E. , Alexander Derrick
Published: September 2016

This paper reports the internal performance evaluation of S-duct diffusers with different entrance aspect ratios as part of an ongoing parametric investigation of a generic S-duct inlet. The generic S-duct diffusers were a rectangular-entrance (aspect ratio 1.5 and 2.0) transitioning S-duct diffuser in high subsonic (Mach number > 0.8) flow. The test section was manufactured using rapid prototyping for facilitating the parametric investigation of the geometry. Streamwise static pressure and exit-plane total pressure were measured in a test-rig using surface pressure taps and a 5-probe rotating rake, respectively and the baseline and a variant was simulated through computational fluid dynamics. The investigation indicated the presence of streamwise and circumferential pressure gradients leading to a three dimensional flow in the S-duct diffuser and distortion at the exit plane. The static pressure recovery increased for the diffuser with higher aspect ratio. Total pressure losses and circumferential and radial distortions at the exit plane were higher than that of the podded nacelle type of inlet. The increase in the total pressure recovery was observed for the increase in the aspect ratio for the baseline area ratio (1.57) S-ducts, but without a clear trend for the other area ratio (1.8) ducts. The work represents the beginning of the development of a database for the performance of a particular type of generic inlet. This database will be useful for predicting the performance of aero-engines and air vehicles in high subsonic flight.
Comparison of Different Numerical Approaches for Determination of Compressible Fluid Flow in Narrow Gaps
Authors: Kefalas Alexander , Benra Friedrich-Karl , Brillert Dieter , Dohmen Hans Josef
Published: September 2016

The prediction of fluid flow through a narrow gap is a characteristic problem in fluid mechanics. In today’s turbomachinery, several applications in bearing and sealing technology are based on the phenomena of two surfaces being separated by a thin fluid film of only a few micrometers. The common method for analyzing the non-contact application’s performance usually applies the lubrication theory based on the Reynolds equation. This two-dimensional model is based on the assumption of a laminar viscous flow field, isothermal conditions and it takes aerostatic as well as aerodynamic effects into account. In cases of a complex geometry and challenging flow conditions this approach has its limitations. The usage of commercial state of the art computational fluid dynamics (CFD) software allows the circumvention of these restrictions. As a matter of fact CFD simulations take up high effort in terms of preparation and calculation time. The present contribution compares the numerical approaches with regard to the application’s performance accuracy and calculation effort, using the example of a dry gas seal. To extend the Reynolds equation’s applicability to a wide variety of geometries, a method for implementing the topography design with high fidelity is depicted. The numerical methods are performed for various dry gas seal designs at different operating conditions and are compared with reference data.
Prediction of Turbo Air Classifier Cut Size Based on Particle Trajectory
Authors: Yu Yuan , Saadat Mehdi , Untaroiu Alexandrina , Thomas Benjamin R. , Wood Houston G.
Published: September 2016

As a mainstream dynamic dry classifer, the turbo air classifier is widely used in powder preparation industries for its adjustable cut size, controllable product granularity and high classification performance. As an important indicator for evaluating the classification performance of a turbo air classifier, cut size is often predicted in advance to evaluate classification effect so that the operation parameters can be adjusted suitably according to the production requirement. There are two common ways to obtain cut size of turbo air classifiers. One is based on a theoretical formula; another is based on an experimentally derived formula. There are a few problems with the aforementioned ways of predicting cut size. The theoretical analysis often has some large deviations from the actual values. Analysis based on empirical formula can obtain an accurate predicted cut size at the cost of a large number of training samples. In this paper, a new strategy is introduced to determine the cut size based on numerical simulation of gas-solid two-phase flow in the turbo air classifier using ANSYS® CFX, Release 15.0. The three-dimensional Reynolds-averaged Navier-Stokes equations along with the k-ε turbulence model are adopted to describe the gas flow, and the Lagrangian particle tracking technique is used to calculate the particle trajectory. According to its definition, cut size can be obtained by means of analyzing the particle trajectory. The effects of rotor cage rotational speed and particle density on cut size are also obtained based on analysis of the change of cut size. The simulation results are validated against the experimental data. Numerical simulation provides a new way to obtain the cut size of a turbo air classifier and serves as a method to regulate the operating parameters for classification. It also provides a reference method to study the cut size of various types of classifier.
Preliminary Design and Performance of Counter Rotating Turbines for Open Rotors: Part I — 1-D Methodology
Authors: Bellocq Pablo , Garmendia Inaki , Legrand Jordane , Sethi Vishal
Published: September 2016

Direct Drive Open Rotors (DDORs) have the potential to significantly reduce fuel consumption and emissions relative to conventional turbofans. However, this engine architecture presents many design and operational challenges both at engine and aircraft level. At preliminary design stages, a broad design space exploration is required to identify potential optimum design regions and to understand the main trade offs of this novel engine architecture. These assessments may also aid the development process when compromises need to be performed as a consequence of design, operational or regulatory constraints.

Design space exploration assessments are done with 0-D or 1-D models for computational purposes. These simplified 0-D and 1-D models have to capture the impact of the independent variation of the main design and control variables of the engine. Historically, it appears that for preliminary design studies of DDORS, Counter Rotating Turbines (CRTs) have been modeled as conventional turbines and therefore it was not possible to assess the impact of the variation of the number of stages (Nb) and rotational speed of the propellers. Additionally, no preliminary design methodology for CRTs was found in the public domain.

Part I of this two-part publication proposes a 1-D preliminary design methodology for DDOR CRTs. It allows an independent definition of the Nb, rotational speeds of both parts of the CRT, inlet flow conditions, inlet and outlet annulus geometry as well as power extraction. It includes criteria and procedures to calculate: power extraction in each stage, gas path geometry, blade metal angles, flow conditions at each turbine plane and overall CRT efficiency. The feasible torque ratios of a CRT are discussed in this paper. A form factor for the CRT velocity triangles is defined (similar to stage reaction on conventional turbines) and its impact on performance and blade design is discussed. A method for calculating the off-design performance of a CRT is also described in Part I.

In Part II, a 0-D design point (DP) efficiency calculation for CRTs is proposed as well as a case study of a DDOR for a 160 PAX aircraft. In the case study, three main aspects are investigated: A) the design and performance of a 20 stage CRT for the DDOR application; B) the impact of the control of the propellers on cruise specific fuel consumption, C) the impact of the design rotational speeds and Nb of the CRT on its DP efficiency, engine fuel consumption and engine weight.
An Outlook for Radical Aero Engine Intercooler Concepts
Authors: Petit Olivier , Xisto Carlos , Zhao Xin , Grönstedt Tomas
Published: September 2016

A state of the art turbofan engine has an overall efficiency of about 40%, typically composed of a 50% thermal and an 80% propulsive efficiency. Previous studies have estimated that intercooling may improve fuel burn on such an engine with a 3–5% reduction depending on mission length. The intercooled engine benefits stem firstly from a higher Overall Pressure Ratio (OPR) and secondly from a reduced cooling flow need. Both aspects relate to the reduced compressor exit temperature achieved by the intercooler action. A critical aspect of making the intercooler work efficiently is the use of a variable intercooler exhaust nozzle. This allows reducing the heat extracted from the core in cruise operation as well as reducing the irreversibility generated on the intercooler external surface which arises from bypass flow pressure losses. In this respect the improvements, higher OPR and lower cooling flow need, are achieved indirectly and not by directly improving the underlying thermal efficiency.

This paper discusses direct methods to further improve the efficiency of intercooled turbofan engines, either by reducing irreversibility generated in the heat exchanger or by using the rejected heat from the intercooler to generate useful power to the aircraft. The performance improvements by using the nacelle wetted surface to replace the conventional intercooler surface is first estimated. The net fuel burn benefit is estimated at 1.6%. As a second option a fuel cooled intercooler configuration, operated during the climb phase, is evaluated providing a net fuel burn reduction of 1.3%.

A novel concept that uses the rejected heat to generate additional useful power is then proposed. A secondary cycle able to convert rejected intercooler heat to useful thrust is used to evaluate three possible scenarios. The two first cases investigate the impact of the heat transfer rate on the SFC reduction. As a final consideration the geared intercooled engine cycle is re-optimized to maximize the benefits of the proposed heat recovery system. The maximum SFC improvement for the three cycles is established to 2%, 3.7% and 3%.
Preliminary Design and Performance of Counter Rotating Turbines for Open Rotors: Part II — 0-D Methodology and Case Study for a 160 PAX Aircraft
Authors: Bellocq Pablo , Garmendia Inaki , Legrand Jordane , Sethi Vishal
Published: September 2016

Direct Drive Open Rotors (DDORs) have the potential to significantly reduce fuel consumption and emissions relative to conventional turbofans. However, this engine architecture presents many design and operational challenges both at engine and aircraft level. At preliminary design stages, a broad design space exploration is required to identify potential optimum design regions and to understand the main trade offs of this novel engine architecture. These assessments may also aid the development process when compromises need to be performed as a consequence of design, operational or regulatory constraints.

Design space exploration assessments are done with 0-D or 1-D models for computational purposes. These simplified 0-D and 1-D models have to capture the impact of the independent variation of the main design and control variables of the engine. Historically, it appears that for preliminary design studies of DDORs, Counter Rotating Turbines (CRTs) have been modelled as conventional turbines and therefore it was not possible to assess the impact of the variation of the number of stages (Nb) of the CRT and rotational speed of the propellers. Additionally, no preliminary design methodology for CRTs was found in the public domain.

Part I of this two-part publication proposes a 1-D preliminary design methodology for DDOR CRTs which allows an independent definition of both parts of the CRT. A method for calculating the off-design performance of a known CRT design is also described.

In Part II, a 0-D design point efficiency calculation for CRTs is proposed and verified with the 1-D methods. The 1-D and 0-D CRT models were used in an engine control and design space exploration case study of a DDOR with a 4.26m diameter an 10% clipped propeller for a 160 PAX aircraft. For this application:

• the design and performance of a 20 stage CRT rotating at 860 rpm (both drums) obtained with the 1-D methods is presented.

• differently from geared open rotors, negligible cruise fuel savings can be achieved by an advanced propeller control.

• for rotational speeds between 750 and 880 rpm (relatively low speeds for reduced noise), 22 and 20 stages CRTs are required.

• engine weight can be kept constant for different design rotational speeds by using the minimum required Nb.

• for any target engine weight, TOC and cruise SFC are reduced by reducing the rotational speeds and increasing Nb (also favourable for reducing CRP noise). However additional CRT stages increase engine drag, mechanical complexity and cost.