bluuebunny/crossref_metadata_2025_split huggingface dataset:
Developing Statistically Defensible Propulsion System Test and Evaluation Techniques
Authors: Kidman David , Stevens Craig , Moulder Christopher , Kitto William , Brownlow James , Remund Todd 
Published: July 2013

Acquisition of military hardware typically proceeds from design, development, production, and finally to operational use and support. Prior to the full-rate production, both developmental and operational test and evaluation (DT&E, OT&E) must occur to ensure that the system meets military requirements. The United States Air Force (USAF) is continually looking for ways to improve its test and evaluation techniques. Since 1997, Air Combat Command (ACC) has been successfully using Design of Experiments (DOE) to construct and analyze operational test efforts. This paper highlights recent efforts to pursue statistically defensible test techniques to aid developmental test efforts.
Defensible testing is a statistical approach similar to DOE by emphasizing the need for better test planning by:
• insistence on understanding the system under test;
• requiring clear and achievable test objectives;
• ensuring system performance is measurable;
• requiring that instrumentation accuracy and uncertainty propagation are well understood;
• and requiring confidence, power, and performance thresholds.
This paper highlights the Air Force Flight Test Center’s (AFFTC) first steps to improve aircraft propulsion system test and evaluation (T&E) through the implementation of statistically defensible test techniques. Background on the AF acquisition process, the Air Force vision for defensible testing, and an aircraft propulsion T&E case study are presented.
Experimental Study of Blockerless Cascade Thrust Reverser
Authors: Shi Jingwei , Wang Zhanxue , Liu Chunyang , Gong Hao 
Published: July 2013

Modern Aero-engine increasingly requires higher propulsive efficiency, lower weight and less fuel consumption, which motivates researchers all over the world to explore novel technology to achieve these goals. Blockerless cascade thrust reverser is one of such innovative technology, using high pressure secondary flow injecting into fan bypass to fulfill reverse thrust instead of traditional blocker door.
The efficiency of thrust reverser is a critical index of the technology, which is a function of aerodynamic and geometric parameters, including fan pressure ratio, secondary injection pressure ratio, injection angle, and injection position. The present study aims to investigate the effects of these parameters on reverse performance through basic demonstration experiments. Firstly, a simple experimental model was designed, which would be conducted on cascade experimental facility. Then, experiments were performed at static conditions, with fixed fluidic injection slot, variable injection angles (30°, 45°, 60°, 90°) at positions of Xj = 38.77mm, 58.77mm, 78.77mm, 98.77mm, 118.77mm, with the scheduled fan pressure ratios (FPR from 1.3 to 1.7) and secondary to ambient pressure ratios (SPR from 2.0 to 6.0). The results indicate that there exits an optimum injection location and injection angle for maximum thrust efficiency, and performance of blockerless cascade thrust reverser varies with diverse fan pressure ratio and secondary injection pressure ratio.
Ethanol and Distillate Blends: A Thermodynamic Approach to Miscibility Issues: Part 3 — Generalization to Other Alcohols (Methanol, Isopropanol and 1-Butanol)
Authors: Privat Romain , Jaubert Jean-Noël , Molière Michel 
Published: July 2013

In the framework of a multistep program devoted to the ternary gasoil/alcohol/water system, the authors investigated the miscibility of anhydrous and hydrated ethanol qualities with four classes of industrial gasoil having different compositions and densities. To that end, they considered a pseudo binary system made by the various hydrocarbon species on one hand and the alcohol/water sub-system on another hand. Using the UNIQUAC thermodynamic theory and the Group Contribution approach, the team computed the Minimum Miscibility Temperature (“MMT”) for a series of the gasoil/ethanol/water system having water concentrations in ethanol comprised between 0 and 10%. The TMM is the temperature above which the various components of the system form a sole phase. This work is summarized in two papers already published (Part 1: GT 2010-22126; Part 2: GT2011-45896).
In the continuity of this prior work and considering the potential interest of alternative alcohols as “gasoil extenders”, the team has generalized this approach to selected C1-C4 alcohols: methanol, isopropanol (or 2-propanol) and n-butanol (or 1-butanol). While methanol is an interesting “energy vector” of coal and biomass via the CTL and BTL processes, isopropanol is a widespread commodity produced by the classical petrochemistry and 1-butanol is a promising biofuel candidate of the second, “lingo-cellulosic” generation.
This third part of the project shows that the introduction of these alternative alcohols and their respective interactions with water lead to considerable changes in the liquid-liquid equilibria and important shifts of the MMTs, trends that were difficult to anticipate beforehand.
Influence of Microjet Condition on Characteristics of Supersonic Jet Noise and Flow Field
Authors: Okada Ryuichi , Watanabe Toshinori , Uzawa Seiji , Himeno Takehiro , Oishi Tsutomu 
Published: July 2013

Jet noise reduction is essential for environmentally-friendly civil transport. Since jet noise becomes very intense in the case of supersonic aircraft, noise reduction is crucial topic for the realization of next-generation supersonic transport.
In the present study, experimental investigations were performed to clarify the effect of microjet injection on supersonic jet noise and flow field. The experiments were focused on supersonic jet with Mach number up to 1.47, which was generated from a rectangular nozzle with high aspect ratio.
Far-field acoustic measurements were conducted for widely ranged microjet conditions to understand the influence of the condition on characteristics of supersonic jet noise and flow field. For understanding the unsteady behavior of the flow field and the relation with noise reduction, flow field visualization was performed with schlieren technique using a high-speed camera.
Impact of Fuel Flexibility Needs on a Selected GT Performance in IGCC Application
Authors: Majoumerd Mohammad Mansouri , Breuhaus Peter , Smrekar Jure , Assadi Mohsen , Basilicata Carmine , Mazzoni Stefano , Chennaoui Leila , Cerri Giovanni 
Published: July 2013

As part of a European Union (EU) funded H2-IGCC project, a baseline IGCC power plant was established; this was presented at the ASME Turbo Expo 2011 (GT2011-45701). The current paper focuses on a detailed investigation of the impact of using various fuels considering different operating conditions on the gas turbine performance, and the identification of technical solutions for the realization of the targeted fuel flexibility.
Using a lumped model, based on real engine data, compressor and turbine maps of the targeted engine were generated and implemented into the detailed GT model made in the commercial heat and mass balance program, IPSEpro. The implementation was done in terms of look-up tables. The impact of fuel change on the gas turbine island has been investigated and reported in this paper. Calculation results show that for the given boundary conditions, the surge margin of the compressor was slightly reduced when natural gas was replaced by hydrogen-rich syngas. The use of cleaned syngas instead of hydrogen-rich syngas resulted in a considerable reduction of the surge margin and elevation of the turbine outlet temperature (TOT) at design point conditions, when keeping the turbine inlet temperature (TIT) and compressor inlet mass flow unchanged. To maintain the TOT and improve the surge margin, when operating the engine with cleaned syngas, a combination of adjustment of variable inlet guide vanes (VIGV) and reduced TIT was considered. A parameter study was carried out to provide better understanding of the current limitations of the engine and to identify possible modifications to improve fuel flexibility.
Application of Endoscopic OH*-Chemiluminescence Measurements at a Full-Scale High-Pressure Gas Turbine Combustion Test Rig
Authors: Witzel Benjamin , Heinze Johannes , Kock Boris F. , Stockhausen Guido , Streb Holger , van Kampen Jaap , Schulz Christof , Willert Chris , Fleing Christian 
Published: July 2013

Single burner combustion tests play a key role in the Siemens gas turbine combustion system development process. The main scope of these tests is to assess the performance of combustor design variants in terms emissions or combustion stability at gas turbine relevant operation conditions. Both emissions and combustion stability strongly depend on the flame front and flame position. A pragmatic approach to investigate the flame is to detect the chemiluminescence signal of the combustion intermediate species OH*. Thus, the OH*-chemiluminescence signal was recorded at high-pressure combustion tests to get more insight in the complex interactions between combustor design, operation conditions and combustion performance.
To minimize the impact of the measurement system on the combustion behavior, the optical access to the test rig was realized by using a water-cooled probe with an UV-transparent endoscope. The probe was located in the test rig side-wall, downstream of the burner outlet, viewing towards the burner with a 90° angle relative to the endoscope orientation. The experimental setup was completed by a combination of bandpass filters and an ICCD camera.
During the experiments acoustic pressure oscillations inside the combustion chamber were recorded simultaneously to the chemiluminescence images to allow for phase-sorting of the recorded images during the image post-processing. The post-processed images then were correlated with the pressure oscillations to investigate the relationship of the heat release to the pressure oscillations.
The measurements were carried out during single burner gas turbine combustion tests at realistic gas turbine operation conditions at a scaled pressure of 9 bar.
This paper presents selected test results and discusses how they give new insight in the complex combustion processes at full-scale high-pressure gas turbine combustion tests.
Durability of Oxide/Oxide CMCs in Gas Turbine Combustors
Authors: van Roode Mark , Bhattacharya Arun K. 
Published: July 2013

An integrated creep rupture strength degradation and water vapor degradation model for gas turbine oxide-based CMC (Ceramic Matrix Composite) combustor liners was expanded with heat transfer computations to establish maximum TRIT (Turbine Rotor Inlet Temperature) for gas turbines with 10:1 pressure ratio. Recession rates and average CMC operating temperatures were calculated for an existing baseline N720/A (N720/Al2O3) CMC combustor liner system, with and without protective Al2O3 FGI (Friable Graded Insulation) for 30,000-h liner service life.
The potential for increasing TRIT by YAG (Y3Al5O12) substitution for the fiber, matrix and FGI constituents of the CMC system was explored, because of the known superior creep and water vapor degradation resistance of YAG compared to Al2O3.
It was predicted that uncoated N720/A can be used as a combustor liner material up to a TRIT of ∼1200°C, offering no TRIT advantage over a conventional metal + TBC (Thermal Barrier Coating) combustor liner. A similar conclusion was previously reached for a SiC/SiC CMC liner with BSAS-type EBC (Barium Strontium Aluminum Silicate Environmental Barrier Coating). The existing N720/A + Al2O3 FGI combustor liner system can be used at a maximum TRIT of ∼1350°C, a TRIT increase over metal + TBC and uncoated N720/A of ∼150°C. Replacing the Al2O3 with YAG is predicted to increase the maximum allowable TRIT. Substitution of the fiber or matrix in N720/A increases TRIT by ∼100°C.
A YAG FGI improves the TRIT of the 720/A + Al2O3 FGI by ∼50°C, enabling a TRIT of ∼1400°C, similar to that predicted for SiC/SiC CMCs with protective rare earth monosilicate EBCs.
Performance and Techno-Economic Investigation of On-Wing Compressor Wash for a Short-Range Aero Engine
Authors: Giesecke Daniel , Igie Uyioghosa , Pilidis Pericles , Ramsden Kenneth , Lambart Paul 
Published: July 2013

The lower atmosphere is known to be relatively more concentrated with airborne pollutant. Short-range aircrafts are particularly more affected due to the altitude they operate, making them more susceptible to compressor fouling degradation. This usually leads to the demand for more fuel that increases the emissions, to make up for the reduction in the thrust. Compressor fouling is therefore a concern for aircraft operators due to increasing fuel cost and emission-based landing fees which impact the direct operating costs of an aircraft.
Highlighting the performance and economic benefits of compressor washing are the key aims of this study. An economic model is developed and the benefit is calculated for different wash intervals, which are based on the usual aircraft checks. The clean and fouled engine performance is simulated, indicating the impact of compressor fouling degradation on the twin-spool engine model. To emphasise the benefit of compressor washing, the degraded engine is compared to its new condition after washing. It was observed that it is impossible to fully recover the performance of a fouled compressor by an on-wing compressor wash.
This study concludes that compressor washing has a significant improvement on engine performance, as well as cost benefit in monetary terms. The results also suggest that compressor washing can reduce the unwanted thrust loss and exhaust gas temperature increase due to fouling, by half. A gross recovery of almost £75,000 per year for a short-range engine is shown to be achievable with a marginal increase of the total washing cost.
Total Pressure Correction of a Sub-Miniature Five-Hole Probe in Areas of Pressure Gradients
Authors: Town J. , Akturk A. , Camcı C. 
Published: July 2013

Five-hole probes, being a dependable and accurate aerodynamic tools, are excellent choices for measuring complex flow fields. However, total pressure gradients can induce measurement errors. The combined effect of the different flow conditions on the ports causes the measured total pressure to be prone to a greater error. This paper proposes a way to correct the total pressure measurement.
The correction is based on the difference between the measured total pressure data of a Kiel probe and a sub-miniature prism-type five-hole probe. By comparing them in a ducted fan related flow field, a line of best fit was constructed. The line of best fit is dependent on the slope of the line in a total pressure versus span and difference in total pressure between the probes at the same location.
A computer program, performs the comparison and creates the correction equation. The equation is subsequently applied to the five-hole probe total pressure measurement, and the other dependent values are adjusted. The validity of the correction is then tested by placing the Kiel probe and the five-hole probe in ducted fans with a variety of different tip clearances.
Effects of Advanced Engine Technology on Open Rotor Cycle Selection and Performance
Authors: Perullo Christopher A. , Tai Jimmy C. M. , Mavris Dimitri N. 
Published: July 2013

Recent increases in fuel prices and increased focus on aviation’s environmental impacts have reignited focus on the open rotor engine concept. This type of architecture was extensively investigated in previous decades but was not pursued through to commercialization due to relatively high noise levels and a sudden, sharp decrease in fuel prices. More recent increases in fuel prices and increased government pressure from taxing carbon-dioxide production mean the open rotor is once again being investigated as a viable concept. Advances in aero-acoustic design tools have allowed industry and academia to re-investigate the open rotor with an increased emphasis on noise reduction while retaining the fuel burn benefits due to the increased propulsive efficiency. Recent research with conceptual level multidisciplinary considerations of the open rotor has been performed [1], but there remains a need for a holistic approach that includes the coupled effects of the engine and airframe on fuel burn, emissions, and noise.
Years of research at Georgia Institute of Technology have led to the development of the Environmental Design Space (EDS) [2]. EDS serves to capture interdependencies at the conceptual design level of fuel burn, emissions, and noise for conventional and advanced engine and airframe architectures. Recently, leveraging NASA Environmentally Responsible Aviation (ERA) modeling efforts, EDS has been updated to include an open rotor model to capture, in an integrated fashion, the effects of an open rotor on conventional airframe designs. Due to the object oriented nature of EDS, the focus has been on designing modular elements that can be updated as research progresses. A power management scheme has also been developed with the future capability to trade between fuel efficiency and noise using the variable pitch propeller system. Since the original GE open rotor test was performed using a military core, there is interest in seeing the effect of modern core-engine technology on the integrated open rotor performance. This research applies the modular EDS open rotor model in an engine cycle study to investigate the sensitivity of thermal efficiency improvements on open rotor performance, including the effects on weight and vehicle performance. The results are that advances in the core cycle are necessary to enable future bypass ratio growth and the trades between core operating temperatures and size become more significant as bypass ratio continues to increase. A general benefit of a 30% reduction in block fuel is seen on a 737–800 sized aircraft.
A More Realistic Scheme of Deviation Error Representation for Gas Turbine Diagnostics
Authors: Loboda Igor , Yepifanov Sergiy , Feldshteyn Yakov 
Published: July 2013

Gas turbine diagnostic algorithms widely use fault simulation schemes, in which measurement errors are usually given by theoretical random number distributions, like the Gaussian probability density function. The scatter of simulated noise is determined on the basis of known information on maximum errors for every sensor type. Such simulation differs from real diagnosis because instead of measurements themselves the diagnostic algorithms work with their deviations from an engine baseline. In addition to simulated measurement inaccuracy, the deviations computed for real data have other error components. In this way, simulated and real deviation errors differ by amplitude and distribution. As a result, simulation-based investigations might result in too optimistic conclusions on gas turbine diagnosis reliability.
To understand error features, deviations of real measurements are analyzed in the present paper. To make error presentation more realistic, it is proposed to extract an error component from real deviations and to integrate it in fault description. Finally, the effect of the new noise representation mode on diagnostic reliability is estimated. It is shown that the reliability change due to inexact error simulation can be significant.
Experimental Measurements and Numerical Simulations in Isothermal Turbulent Flows
Authors: Florean Florin G. , Popescu Jeni A. , Porumbel Ionut , Carlanescu Cristian , Dumitrascu Gheorghe 
Published: July 2013

The paper presents Particle Image Velocimetry experimental measurements and an ANSYS CFX numerical simulation of the mean and fluctuating velocity field in a turbulent, isothermal flow downstream of a V-shaped bluff body flame stabilizer equipping a post-combustion system installed downstream of a Garrett 30-67 gas turbine engine.
The post-combustion system used as experimental model is described in detail, and the main characteristics of the Garrett 30-67 gas generator are included in the paper. Also, the instrumentation used on the experimental rig, including the Particle Image Velocimetry equipment, is briefly described.
The presence of a bluff body inside the high speed flow creates a re-circulating wake structure, clearly seen in the experimental data. In the near field, the re-circulating region’s length and width of are captured reasonably well by the numerical simulation, but the momentum rate transfer further downstream is over-predicted, as the grid resolution worsens.
An overall over prediction of the axial velocity by the numerical simulation is noted by comparing the numerical simulation results to the experimental data, explained by an over estimated inlet velocity in the numerical simulation, provided by Particle Image Velocimetry experimental measurements in the free exhaust jet behind the gas generator, without the installation of the post-combustion system, thus neglecting the effect of the latter on the operating regime of the gas turbine engine.
Component Map Tuning Procedure Using Adaptive Modeling
Authors: Verbist Michel L. , Visser Wilfried P. J. , Pecnik Rene , van Buijtenen Jos P. 
Published: July 2013

Performance models are effective tools for analysis of engine condition throughout the life cycle of a gas turbine engine. Component maps necessary for accurate performance modeling are typically not provided by the original equipment manufacturers. To compensate for the missing information, available maps of similar components are scaled to match component performance at one or more reference points. Although scaled maps can provide sufficiently accurate results close to the reference points, modeling errors tend to increase further away from these reference points. For applications such as gas path analysis, the resulting modeling errors can be of the same order of magnitude as the deterioration to be detected. This severely limits the application of such techniques.
This article presents a component map tuning procedure that tunes maps with more detail than just scaling. The tuned maps are a closer match to real component performance. The tuning procedure combines the adaptive modeling capability of the Gas turbine Simulation Program (GSP) and on-wing measured engine performance data. On-wing measured engine performance data allows map tuning over a wider range of power settings compared to engine performance data measured in a test cell. Effects of measurement uncertainty and scatter, and effects of compressor bleed flows on the map tuning procedure are analyzed and discussed. The tuned component maps enabled more accurate component condition estimations, mainly characterized by less scatter. By improving the accuracy of gas path analysis with on-wing measured performance data, this work has enabled more effective use of performance diagnostic techniques in the aero-engine maintenance industry.