Aircraft noise mitigation has been an ongoing challenge for the aeronautics research community. In response to this challenge, aircraft concepts have been developed in which the propulsion system is integrated with the airframe to shield the noise from the observer. These concepts exhibit situations where the jet exhaust interacts with an airframe surface. Jet flows interacting with nearby surfaces exhibit a complex behavior in which acoustic and aerodynamic characteristics are altered. The physical understanding and accurate modeling of these characteristics are essential to designing future low-noise aircraft. In this thesis, an alternative approach is created for predicting jet mixing noise that utilizes an acoustic analogy and the solution of the steady Reynolds-Averaged Navier-Stokes (RANS) equations using a two equation turbulence model. A tailored Green's function is used in conjunction with the acoustic analogy to account for the propagation effects of mixing noise due to a nearby airframe surface. The tailored Green's function is found numerically using a newly developed ray tracing method. The variation of the aerodynamics, acoustic source, and far- field acoustic intensity are examined as a large flat plate is moved relative to the nozzle exit. Steady RANS solutions are used to study the aerodynamic changes in the field-variables and turbulence statistics. To quantify the propulsion airframe aeroacoustic (PAA) installation effects on the aerodynamic source, a non-dimensional number is formed that can be used as a basic guide to determine if the aerodynamic source is affected by the airframe and if additional noise produced by the airframe surface is present. The aerodynamic and noise prediction models are validated by comparing results with Particle Image Velocimetry (PIV) and far-field acoustic data respectively. The developed jet noise scattering methodology is then used to demonstrate the shielding effects of the Hybrid Wing Body (HWB) aircraft. The validation assessment shows that the acoustic analogy and tailored Green's function provided by the ray tracing method are capable of capturing jet shielding characteristics for multiple configurations and jet exit conditions. / Master of Science

京都大学 / 0048 / 新制・課程博士 / 博士(理学) / 甲第18770号 / 理博第4028号 / 新制||理||1580(附属図書館) / 31721 / 京都大学大学院理学研究科数学・数理解析専攻 / (主査)教授 山田 道夫, 教授 玉川 安騎男, 准教授 竹広 真一 / 学位規則第4条第1項該当 / Doctor of Science / Kyoto University / DFAM

To mitigate climate change, all sectors must contribute their part. Electric propulsion system is a promising approach to reduce emissions from the aviation industry. A major challenge is however the low energy density in today’s battery technology. Even with today’s leading li-ion batteries, the specific energy density in jet fuel is 48 times larger. Because of the low energy density in any battery technology today, hybrid electric propulsion system could be a bridge between conventional and fully electric propulsion systems. The purpose of this degree project is to explore different designs of a parallel hybrid electric mixed turbofan to minimize the impact aircraft have on the environment. The engine has been designed in Modelon Impact and MATLAB has been used to evaluate the thrust requirement and, the performance and weight calculation of the electrical power system. Furthermore, the conventional engine was validated against GasTurb. Three different designs were evaluated, i.e., design #1 (reference engine with hybridization), #2 (decreased OPR) and #3 (increased BPR and decreased FPR) with two different weights of the electrical power system. None of the designs showed a reduction in terms of the total fuel consumption during the whole mission profile. However, design #3 showed the most beneficial results in terms of reducing the specific fuel consumption and could reduce the fuel consumption in the climb segment the most among the three different designs.

The purpose of this research was to operate a bluff body flame holder with the objective of stabilizing a flame at elevated pressures over a range of equivalence ratios. The ability to have a ground-based test rig capable of maintaining stable flames at high pressures and temperatures is critical in understanding flames present in modern jet engines and gas turbine technologies. The facility was reconfigured multiple times and the resultant flame was imaged within the optical test section. A converging nozzle was utilized to choke the flow and vary the operating pressures up to 5 atm. By regulating mass flow rates of both the fuel and air, the target range of equivalence ratios was achieved. Jet fuel was successfully ignited on the bluff body and a flame was maintained in the recirculation zone. Visualization of flames during the flights of any aircraft is limited due to material and weight requirements, therefore, performing these studies in ground- based facilities is required. Further analysis was performed to characterize C2* and CH* radicals in fuel lean and rich flames.

<p>Hypersonic flow over two generic missile configurations was investigated using CFD meth-</p> <p>ods. CFD results were compared with experimental results obtained by the hypersonic flight</p> <p>lab at Texas A&M University. Baseline RANS computations involving the missile configurations at a zero deg angle-of-attack were performed, along with computations at higher angles-of-attack. As the angle-of-attack was increased, complex vortex interactions were observed in the region between the fins. Increasing the angle-of-attack generally increased heating on the windward side of the missile geometries, especially on wall surface regions</p> <p>adjacent to the fin-root vortices. The results presented highlight observed fin region vortices and regions of intense heating on the body surface. DES simulations methods were also used to explore unsteady aspects of flow around the two generic missile configurations through time-accurate CFD simulations. Power spectral plots were generated to quantify the dominant frequencies of large-scale unsteadiness.</p>

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This paper has investigated if utilising Ion-Propulsion on a solar sunshade would lead to a reduction in total mass. The mechanism leading to this effect is that by using IP, the sunshade could be positioned closer to Earth. Positioning the sunshade closer to Earth would in turn allow for a smaller absolute size in the sunshade (whilst remaining a fixed angular size), and thus a decrease in total mass. This would lead to a decrease in launch costs and manufacturing costs.  The results indicate that fitting satellites with Ion-Propulsion would lead to a reduction in the total mass for the sunshade. More specifically they showcase the extent of which using IP has an effect on total mass, for example certain minimum ISP values that are needed to be reached in order for significant reductions of mass to arise. In a similar manner, the duration of which the satellites can remain in orbit also heavily impacts the total mass of the sunshade. If Ion-Propulsion where to be utilised it is quintessential to understand how the technology impacts the total lifetime of the satellites. Additional investigation may also be made in investigating if Ion-Propulsion could be used in such a way where the satellites do not remain stationary in respect to the distance between Earth and the Sun, perhaps this could lead to additional reductions of mass.

Hybrid rocket propulsion technology shows promise for the next generation of sounding rockets and small launch vehicles. This paper seeks to provide details on the process of developing hybrid propulsion systems to the academic and amateur rocket communities to assist in future research and development. Scaling hybrid rocket motors for use in sounding rockets has been a challenge due to the inadequacies in traditional boundary layer analysis. Similarity scaling is an amendment to traditional boundary layer analysis which is helpful in removing some of the past scaling challenges. Maintaining geometric similarity, oxidizer and fuel similarity and mass flow rate to port diameter similarity are the most important scaling parameters. Advances in composite technologies have also increased the performance through weight reduction of sounding rockets through and launch vehicles. Technologies such as Composite Overwrapped Pressure Vessels (COPV) for use as fuel and oxidizer tanks on rockets promise great advantages in flight performance and manufacturing cost. A small scale COPV, carbon fiber ablative nozzle and a N class hybrid rocket motor were developed, manufactured and tested to support the use of these techniques in future sounding rocket development. The COPV exhibited failure within 5% of the predicted pressure and the scale motor testing was useful in identifying a number of improvements needed for future scaling work. The author learned that small scale testing is an essential step in the process of developing hybrid propulsion systems and that ablative nozzle manufacturing techniques are difficult to develop. This project has primarily provided a framework for others to build upon in the quest for a method to easily develop hybrid propulsion systems sounding rockets and launch vehicles.

Analysis of the aerodynamic effects of a propeller flowfield on bodies downstream of the propeller is a complex task. These interaction effects can have serious repercussions for many aspects of the vehicle, including drag changes resulting in larger power requirements, stability changes resulting in adjustments to stabilizer sizing, and lift changes requiring wing planform adjustments. Historically it has been difficult to accurately account for these effects at any stage during the design process. More recently methods using Euler solvers have been developed that capture interference effects well, although they don't provide an ideal tool for early stages of aircraft design, due to computational cost and the time and expense of setting up complex volume grids. This research proposes a method to fill the void of an interference model useful to the aircraft conceptual and preliminary designer. The proposed method combines a flexible and adaptable tool already familiar to the conceptual designer in the aerodynamic panel code, with a pseudo-steady slipstream model wherein rotational effects are discretized onto vortex particle point elements. The method maintains a freedom from volume grids that are so often necessary in the existing interference models. In addition to the lack of a volume grid, the relative computational simplicity allows the aircraft designer the freedom to rapidly test radically different configurations, including more unconventional designs like the channel wing, thereby providing a much broader design space than otherwise possible. Throughout the course of the research, verification and validation studies were conducted to ensure the most accurate model possible was being applied. Once the vortex particle scheme had been verified, and the ability to model an actuator disk with vortex particles had been validated, the overall product was compared against propeller-wing wind tunnel results conducted specifically as benchmarks for numerical methods. The method discussed in this work provides a glimpse into the possibility of pseudo-steady interference modeling using vortex particles. A great groundwork has been laid that already provides reasonable results, and many areas of interest have been discovered where future work could improve the method further. The current state of the method is demonstrated through simulations of several configurations including a wing and nacelle and a channel wing.

A method for sizing remotely powered unmanned aerial vehicles is presented to augment the conventional design process. This method allows for unconventionally powered aircraft to become options in trade studies during the initial design phase. A design matrix is created that shows where, and if, a remotely powered vehicle fits within the design space. For given range and power requirements, the design matrix uses historical data to determine whether an internal combustion or electrical system would be most appropriate. Trends in the historical data show that the break in the design space between the two systems is around 30 miles and 1 kW. Electrical systems are broken into subcategories of onboard energy sources and remote power sources. For this work, only batteries were considered as an onboard energy source, but both lasers and microwaves were considered for remote power transmission methods. The conventional sizing method is adjusted to so that it is based on energy consumption, instead of fuel consumption. Using the manner in which microwaves and laser propagate through the atmosphere, the weight fraction of a receiving apparatus is estimated. This is then used with the sizing method to determine the gross takeoff weight of the vehicle. This new sizing method is used to compare battery systems, microwave systems, and laser systems.