bluuebunny/crossref_metadata_2025_split 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.
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.
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.
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.
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.
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.
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.