A Formulation of Multidimensional Growth Models for the Assessment and Forecast of Technology Attributes Travis W. Danner 229 Pages Directed by Dr. Dimitri Mavris This research proposes the formulation of multidimensional growth models as an approach to simulating the advancement of multi-objective technologies towards their upper limits. These multidimensional growth models are formulated by noticing and exploiting the correlation between technology growth models and technology frontiers. Both are frontiers in actuality. The technology growth curve is a frontier between capability levels of a single attribute and time, while a technology frontier is a frontier between the capability levels of two or more attributes. Multidimensional growth models are formulated by exploiting the mathematical significance of this correlation. The result is a model that can capture both the interaction between multiple system attributes and their expected rates of improvement over time. The fundamental nature of technology development is maintained and interdependent growth curves are generated for each system metric with minimal data requirements. Being founded on the basic nature of technology advancement, relative to physical limits, the availability for further improvement can be determined for a single metric relative to other system measures of merit. A byproduct of this modeling approach is a single n-dimensional technology frontier linking all n system attributes with time. This provides an environment capable of forecasting future system capability in the form of advancing technology frontiers. In addition to formulating the multidimensional growth model, this research provides a systematic procedure for applying it to specific technology architectures. Researchers and decision-makers are able to investigate the potential for additional improvement within that technology architecture and estimate the expected cost of each incremental improvement relative to the cost of past improvements. In this manner, multidimensional growth models provide the necessary information to set reasonable program goals for the further development of a particular technological approach or to establish the need for new technological approaches in light of the constraining limits of conventional approaches.

This thesis addresses the need for an optimization method which can simultaneously optimize and balance an aerothermodynamic cycle. The method developed is be able to control cycle design variables at all operating conditions to meet the performance requirements while controlling any additional variables which may be used to optimize the cycle and maintaining all operating limits and engine constraints. The additional variables represent degrees of freedom above what is needed for conservation of mass and energy in the engine system. The motivation for such a method is derived from variable cycle engines, however it is general enough to use with most engine architectures. The method is similar to many optimization algorithms but differs in its implementation to an aircraft engine by combining the cycle balance and optimization using a Newton-Raphson cycle solver to efficiently find cycle designs for a wide range of engine architectures with extra degrees of freedom not needed to balance the cycle. Combination of the optimization with the cycle solver greatly speeds up the design and optimization process. A detailed process description for implementation of the method is provided as well as a proof of concept using several analytical test functions. Finally, the method is demonstrated on a separate flow turbofan model. Limitations and applications of the method are further explored including application to a multi-design point methodology.

Le besoin permanent de réduire le bruit des moteurs d’avion constitue un véritable engouement pour le développement de nouveaux traitements acoustiques. Les traitements traditionnels de type résonateur continuent d’être utilisé et permettent d’atténuer le son sur une bande de fréquence restreinte malgré l’augmentation du nombre de degré de liberté. Une alternative possible est l’utilisation de matériaux poreux, dit à réaction non localisée, qui permettent d’élargir le spectre d’atténuation. Ce rapport est consacré à la modélisation d’un traitement acoustique basé sur des matériaux poreux dans les conditions d’une manche d’entrée d’air de turboréacteur. Un modèle semi-analytique a donc été développé pour le calcul de la perte par transmission d’un conduit cylindrique traité en paroi et soumis à un écoulement uniforme. Une étude paramétrique a ensuite été réalisée afin de cibler les caractéristiques du traitement optimal pour une configuration aéronautique donnée. Des résultats expérimentaux sur une veine à échelle réduite sont également montrés et témoignent d’un certain accord avec le calcul. Enfin, dans le but d’approfondir les connaissances théoriques sur le problème, une étude préliminaire sur les effets d’une couche limite est réalisée et montre que sa prise en compte parait indispensable pour bien choisir les traitements acoustiques, surtout à haute fréquence. / The constant need to reduce noise emissions from aircraft engine leads to a real demand for developing new acoustic treatments. Conventional liners based on resonatorlike structure continue to be used and provide narrow-band attenuation in spite of an increasing degree of freedom. A possible alternative is the use of porous materials (nonlocally reacting), which offer the possibility of broadening the attenuation spectrum. This report deals with the modelling of an acoustic treatment based on porous materials for aeroengine nacelle inlet. A semi-analytical model is developed for predicting the transmission loss of a treated cylindrical duct containing uniform mean flow. Then, a parametrical study is carried out in order to target the optimal liner characteristics for a given turbofan duct application. Also, experiments have been performed on a small-scale duct and have shown agreement with the simulation. Finally, for a better theoretical unv derstanding of the problem, a preliminary study on the effect of a boundary layer is conducted and shows that its consideration seems to be essential for optimal choice of acoustic lining, espacially at high frequencies.

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Gas turbine performance models are usually created in an object oriented manner, where different standard components are connected to form the complete model. The characteristics of these components are often represented by component maps and empirical correlations. However, engine specific component characteristics are seldom available to anyone outside of the manufacturers. It is therefore very common for researchers to use publicly accessible or generic component maps instead. But in order to reduce prediction errors the maps have to be modified to fit any specific engine. This thesis work investigates the process of adapting a parametric turbofan engine model to a limited amount of test-data using the propulsion program EVA. Steady state test-data was generated using an initial reference model with SLS operating conditions. Another engine model with different fan, compressor and turbine maps was then used in the adaptation. An initial on-design model was adapted to the highest power test-data point. This model is based on aerothermodynamic equations and is used as a reference to scale the generic component maps to. A sensitivity analysis was done at this point in order to find dependencies between unknown component parameters and test data. These were then included in the cycle solver which employs a version of the Newton-Raphson method. After the fan and compressor maps had been scaled to the design point they were adapted to test-data by adjusting the mass flow parameters in a direct search optimizer. Finally, speed lines in the fan and compressor maps were relabeled to reduce rotor speed errors. The adapted performance model was then validated against the reference model at a few flying conditions. The performance model results demonstrate that it is possible to greatly reduce prediction errors by only adjusting the corrected mass flow in fan and compressor maps. Additionally, rotor speed errors could successfully be corrected as a final step in the adaptation by relabeling speed lines in the component maps. When validated, the adapted model had a maximum parameter error of 1.5%.

No description available.

As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation. For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility. A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Computational full-field total pressure recovery and swirl angle profiles matched within approximately 1% and 2.5° (RMSD), respectively. Experimental turbofan engine ground test results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Experimental full-field total pressure recovery and swirl angle profiles matched within approximately 1.25% and 3.0° (RMSD), respectively. Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased by approximately 5% in the presence of distortion. While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion. / Doctor of Philosophy / As a contribution to advancing turbofan engine ground test technology in support of propulsion system integration in modern conceptual aircraft, a novel inlet distortion generator (ScreenVaneTM) was invented. The device simultaneously reproduces combined inlet total pressure and swirl distortion elements in a tailored profile intended to match a defined turbofan engine inlet distortion profile. The device design methodology was intended to be sufficiently generic to be utilized in support of any arbitrary inlet distortion profile yet adequately specific to generate high-fidelity inlet distortion profile simulation. For the current investigation, a specific inlet distortion profile was defined using computational analysis of a conceptual boundary layer ingesting S-duct turbofan engine inlet. The resulting inlet distortion profile, consisting of both total pressure and swirl distortion elements, was used as the objective profile to be matched by the ScreenVane in a turbofan engine ground test facility. A ScreenVane combined inlet total pressure and swirl distortion generator was designed, computationally analyzed, and experimentally validated. The design process involved specifying a total pressure loss screen pattern and organizing a unique arrangement of swirl inducing turning vanes. Computational and experimental results indicated that the ScreenVane manufactured distortion profile matched the predicted S-duct turbofan engine inlet manufactured distortion profile with excellent agreement in pattern shape, extent, and intensity. Following the successful reproduction of the S-duct turbofan engine inlet manufactured distortion profile, a turbofan engine response evaluation was conducted using the validated ScreenVane inlet distortion generator. Flow measurements collected at discrete planes immediately upstream and downstream of the fan rotor isolated the component for performance analysis. Based on the results of this particular engine and distortion investigation, the adiabatic fan efficiency was negligibly altered while operating with distorted inflow conditions when compared to nominal inflow conditions. Fuel flow measurements indicated that turbofan engine inlet air mass flow specific fuel consumption increased in the presence of distortion. While a single, specific turbofan engine inlet distortion profile was studied in this investigation, the ScreenVane methodology, design practices, analysis approaches, manufacturing techniques, and experimental procedures are applicable to any arbitrary, realistic combined inlet total pressure and swirl distortion.

This master thesis covers the realization of a CFD calibration study regarding the numerical simulation of the ground vortex phenomenon. As part of InVIGO project (EU CleanSky2), wind tunnel test campaigns were performed using a nacelle model combined with a movable raised floor. Those tests allowed the gathering of an unprecedented amount of experimental data regarding this phenomenon, via the implementation of pressure measurements and stereo-PIV acquisition. Measured velocity fields were also post-processed to compute vortex characterization indices. In the context of this study, several Python scripts are developed to process this data through the computation of total pressure and velocity maps. A calculation matrix of 28 representative points is chosen among test cases and the corresponding nacelle geometry is prepared to generate a series of high resolution structured mesh (18.22 million cells) using an existing ICEM macro. For every case, corresponding CFD calculations are implemented on elsA using the RANS resolution method associated with the Spalart-Allmaras turbulence model. In parallel, a series of CFD post-processing scripts are developed using Cassiopée modules, in order to extract pressure/velocity maps and compute both distortion and convergence indices. Also, a Safran AE in-house script dedicated to vortex characterization is not only implemented to compute vortex indices, but also upgraded with an additional vortex profile extraction functionality. The cross-analysis of numerical and experimental data thus collected is performed, allowing the assessment of the great calculation representativeness regarding the overall inlet flow topology as well as the vortex apparition. The comparison of vortex characterization indices displays common tendencies, however disparities regarding the implemented processing methodologies along with convergence difficulties encountered for some vortex cases highlights the need to put the relevance of those results into perspective.

Attenuation of low frequency turbofan engine noise has been a challenging task in an industry that requires low weight and tightly-packed solutions. Without innovative advancements, the technology currently used will not be able to keep up with the increasingly stringent requirements on aircraft noise reduction. A need exists for novel technologies that will pave the way for the future of quiet aircraft. This thesis investigates acoustic metamaterials and their ability to achieve superior transmission loss characteristics not found in traditional honeycomb liners. The acoustic metamaterials investigated are an array of Helmholtz resonators with and without coupled cavities periodically-spaced along a duct wall. Analytical, numerical, and experimental developments of these acoustic metamaterial systems are used herein to study the effects of this technology on the transmission loss. Particularly focusing on analytical modeling will aid in understanding the underlying physics that governs their interesting transmission loss behavior. A deeper understanding of the physics will be used to aid in future acoustic metamaterial liner design. A parameter study is performed to understand the effects of the geometry, spacing, and number of resonators, as well as resonator cavity coupling on performance. Increased broadband transmission loss, particularly in low frequencies, is achieved through intelligent manipulation of these parameters. Acoustic metamaterials are shown to have appealing noise cancellation characteristics that prove to be effective for aircraft engine liner applications. / Master of Science / Aircraft noise reduction is an ongoing challenge for the aerospace industry. Without innovative advancements, the next generation of aircraft will not be able to keep up with increasingly stringent noise regulations; novel acoustic technology is needed to pave the way for a future of quieter aircraft. This thesis investigates acoustic metamaterials and their ability to achieve superior noise reduction over traditional methods. Modeling techniques were developed, and experimental tests were conducted to quantitatively evaluate the effectiveness of a new acoustic metamaterial system. The acoustic metamaterial design explored herein was proven to reduce noise effectively and shows promise for a world of quieter aircraft.

An innovative application of existing technology is proposed for attenuating the effects of transient phenomena, such as rotor-stator and rotor-strut interactions, linked to noise and fatigue failure in turbomachinery environments. A computational study was designed to assess the potential of Passive Porosity Technology as a mechanism for alleviating interaction effects and radiated noise by reducing the fluctuating forces acting on the vane surfaces. The study involved a typical high bypass fan stator airfoil immersed in a subsonic free field and exposed to the effects of a transversely moving wake. Time histories of the primitive aerodynamic variables obtained from Computational Fluid Dynamics (CFD) calculations were input into an acoustic prediction code to estimate noise levels at a radial distance of ten chords from the stator airfoil. This procedure was performed on the solid airfoil to obtain a baseline, and on approximately fifty porous configurations in order to isolate those that would yield maximum noise reductions without compromising the aerodynamic performance of the stator. It was found that, for a single stator immersed in a subsonic flow field, communication between regions of high pressure differential - made possible by the use of passive porosity - tends to induce a time-dependent oscillatory pattern of small inflow-outflow regions near the stator leading edge (LE), which is well established before wake effects come into play. The oscillatory pattern starts at the LE, and travels downstream on both suction and pressure sides of the airfoil. The amplitude of the oscillations seemed to be proportional to the extension of the porous patch on the pressure side. Regardless of this effect, which may not have occurred if the airfoil were placed within a stator cascade, communication between regions of high pressure differential is necessary to significantly alter the noise radiation pattern of the stator airfoil. Whether those changes result in noise abatement or enhancement depends primarily on the placement and extension of the porous patches. For most viable configurations, porosity reduced loading noise but increased thickness noise. Variations in nominal porosity were of secondary importance. In general, the best aerodynamic performers (i.e., those configurations that were able to reduce unsteady lift without severely altering the lift and/or drag characteristics of the solid airfoil) were also the best acoustic performers. As a result of using passive surface porosity, overall peak radiated noise was reduced by approximately 1.0 dB. This reduction increased to about 2.5 dB when the effects of loading noise alone were considered. / Ph. D.