記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： 10.2322/tjsass.66.164

リポジトリ公開URL： https://hdl.handle.net/2324/7172135

記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： https://doi.org/10.1063/5.0148974

リポジトリ公開URL： https://hdl.handle.net/2324/7172134

記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： https://doi.org/10.1016/j.ast.2022.107931

リポジトリ公開URL： https://hdl.handle.net/2324/7172131

記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： https://doi.org/10.1016/j.ast.2022.107879

リポジトリ公開URL： https://hdl.handle.net/2324/7172130

記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： https://doi.org/10.2514/1.B38854

リポジトリ公開URL： https://hdl.handle.net/2324/7172133

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Cavities are commonly employed in scramjet combustors for flameholding and mixing enhancement, but the mechanism used to enhance mixing is absent at high supersonic Mach numbers, limiting their operational scope. The present study investigates the ability of crescent-shaped cavities placed upstream of a fuel injector to enhance mixing through vorticity generation, a mixing enhancement mechanism that is also effective at high supersonic Mach numbers. The mixing performance of five crescent cavity designs, two of which incorporate hybrid fuelling, is investigated using unsteady Reynolds-Averaged Navier–Stokes (URANS) computations of a chemically frozen flow with hydrogen as the fuel. It is found that the crescent cavities enhance mixing by up to 22.6% without the hybrid fuelling arrangement and by up to 90.1% with the hybrid fuelling arrangement. While vertical jet penetration is lower for all cavity cases, lateral penetration is higher and the cavity cases incur no or negligible total pressure loss compared to the baseline at the domain outflow, within the margin of error. Wall drag is also lower than in the baseline for some cavity cases. The primary mechanism driving mixing is found to be enhanced streamwise vorticity in the vicinity of the cavity, caused by the cavity vortex leaving the cavity and wrapping around the injector. The cavity flowfields are also found to be oscillatory in nature, although the oscillations are lateral and the harmonic frequencies are much lower than those of the longitudinal oscillations characteristic of conventional cavity flow. The mechanisms driving these oscillations are discussed, as are the flowfields for the best performing cavity cases. Several flowfield features of the crescent cavities are also highlighted and discussed, demonstrating how the hybrid injection cavity cases enhance mixing.

DOI： 10.1016/j.actaastro.2020.08.022

記述言語：英語   掲載種別：研究論文（学術雑誌）  

A multiobjective design optimization (MDO) study has been conducted to characterize and maximize the performance of the downscaled cusped field thruster with respect to thrust, total efficiency, and specific impulse characterized by common design parameters: anode voltage, anode current, mass flow rate, and geometric configuration. Particle-in-cell (PIC) simulations have been performed for the selected design points identified in MDO study for verification by accurately accounting for phenomena and performance losses that originate from uncertainties and complexities associated with the thruster design and physics. The fidelity of the models employed in MDO has been enhanced owing to the new physical insights gained from the detailed analysis of the PIC results. The prediction errors associated with uncertainties such as the beam current and divergence angle have been reduced to within 5%.

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Cavities are commonly used to provide flame-holding in scramjets. While the injector is generally placed inside or upstream of the cavity, placement of the cavity behind the injector limits the influence of the cavity on the jet interaction and limits cavity-induced mixing enhancement. The current study investigates a geometry in which the cavity is placed directly upstream of the injector and examines its effect on scramjet combustor mixing performance. Specifically, enhancement in jet mixing and penetration is considered using chemically frozen hydrogen fuel. The influence of three different thermal boundary conditions (isothermal 300 K, isothermal 1800 K and adiabatic) on the flowfield and mixing was also examined. The upstream cavities are found to improve mixing efficiency and jet penetration relative to a baseline flat plate configuration for most configurations, while they do incur a total pressure loss up to 2% higher than in the baseline. The magnitude of these effects is found to depend on the cavity geometry and wall thermal model. The primary mechanism behind the performance improvement is the shielding of the barrel shock by the cavity recirculation, which introduces extra vorticity into the flowfield and reduces the strength of the bow shock. Increased shielding provided by the cavity is found to enhance mixing by up to 9%. An optimum cavity aspect ratio is observed to exist at a cavity length-to-depth ratio of L/D=15, for which performance is maximum compared to the baseline for all wall treatments. Wall heat flux increases in configurations with cavities, particularly on the aft wall of the cavity, while fuel drawn into the cavity is seen to contribute to wall cooling in case of high wall temperatures. This can reduce wall cooling requirements and simplify combustor design. In general the enhanced mixing and jet penetration induced by the cavity could allow for shorter combustor designs, which in turn allows for more compact flight vehicle design.

DOI： 10.1016/j.actaastro.2019.12.033

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Electric propulsion attracts increasing attention in contemporary space missions as an interesting alternative to chemical propulsion because of the high efficiency it offers. The High-Efficiency Multistage Plasma thruster, a class of cusped field thruster, is able to operate at different anode voltages and operation points and thereby generate different levels of thrust in a stable and efficient way. Since experiments of such thrusters are inherently expensive, multi-objective design optimization (MDO) is of great interest. Several optimized thruster designs have resulted from a MDO model based on a zero-dimensional (0D) power balance model. However, the MDO solutions do not warrant self-consistency due to their dependency on estimation from empirical modelling based on former experimental studies. In this study, one of the optimized thruster designs is investigated by means of particle-in-cell (PIC) analysis to examine the predicted performance characteristics with self-consistent simulations. The 0D power balance model is used to develop additional diagnostics for the PIC simulations to improve the physics analysis. Using input parameters for the 0D power balance model from the PIC simulations allows further improvement for the design optimization.

DOI： 10.1002/ctpp.201900028

記述言語：英語   掲載種別：研究論文（学術雑誌）  

High-performance hydrocarbon-fuelled scramjet engines require efficient fuel-air mixing due to the relatively short flow residence time through the combustor. At high temperatures, hydrocarbon fuels react endothermically and absorb thermal energy from the surroundings. The process known as cracking becomes essential at high Mach numbers to increase the total heat-sink capacity of the fuel. This study presents the results of chemically frozen numerical simulations that investigate the mixing characteristics of cracked gaseous heavy hydrocarbon fuels injected through a circular, flush-wall porthole injector. The mixing characteristics of fuel compositions representing cracking efficiencies ranging from 0 to 100% are investigated. The mixing rates and flow structures are found to change with fuel compositions. As the cracking increases, the mixing and streamwise circulation increase for an injectant. However, the jet penetration and stagnation pressure losses decrease. The streamwise circulation is found to have a strong influence on the mixing, the injection pressure on the jet penetration and the strength of the bow shock on stagnation pressure losses. Overall, it is shown that there are mixing benefits to be gained by injecting cracked hydrocarbon fuels compared to heavy uncracked fuels in scramjets.

DOI： 10.1016/j.actaastro.2019.06.010

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The shock structure around a fuel jet drives most of the initial mixing in a scramjet combustor, making it of interest for mixing enhancement studies. The effect of a cavity placed upstream of a fuel injector on the jet interaction of a transverse jet in a supersonic crossflow was examined numerically in this work. The cavity was found to significantly alter the structure of the typical supersonic cross-flow jet interaction. The typical horseshoe vortices were found to be absent and the barrel shock was found to be larger and more upright than in the no-cavity case. This was caused by the cavity recirculation shielding the fuel jet. The shock structure around the cavity was found to decrease the strength of the bow shock, reducing total pressure loss in the flowfield close to the injector. A small region of fluid above the cavity circulation was found to be the origin of vortical structures in the jet interaction, as opposed to the wall boundary layer in the conventional jet interaction. The presence of the fuel jet was found to alter the flow behaviour inside the cavity from closed cavity flow to open cavity flow, with the recirculation rising out of the cavity. This transition from closed to open cavity flow was found to be turbulence model dependent, however the main flow features and behaviour were shown to be maintained across turbulence models. Fuel was entrained in the cavity for the configuration under investigation, however this was dependent on fuel injection pressure, with no fuel entering the cavity at lower injection pressures.

DOI： 10.1016/j.actaastro.2018.12.032

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Efficient and reliable space transportation systems are fundamental to the future success of routine scientific, commercial and strategic space missions. RBCC (rocket-based combined cycle) comprising rocket, ramjet and scramjet engines is a propulsion technology that offers promise for reusable space launch systems owing to advantages over traditional propulsion in various aspects including efficiency and flexibility. However, the viability of RBCC-powered access-to-space requires careful consideration and its assessment represents a challenge to conventional design approaches due to the highly complex and coupled characteristics of the system associated with multi-mode propulsion and multi-stage launch. This study has been undertaken to examine the performance of a conceptual RBCC-based TSTO (two-stage-to-orbit) system and identify the key requirements and design factors to achieve space launch via this system. Multi-objective design optimization has been conducted with respect to important design criteria by means of evolutionary algorithms assisted by surrogate modeling and trajectory optimization. The influence of RBCC engine characteristics particularly in terms of performance and operation as well as the scaling of the vehicle and propulsion on the overall performance of the TSTO system and its feasibility has been quantified and examined. Furthermore, a comparative study with a rocket-only-based TSTO system has also been conducted, verifying the advantages of the RBCC-based TSTO system. The results highlight complex aerodynamic characteristics for both concepts as well as highly nonlinear propulsion characteristics for the RBCC-based concept.

DOI： 10.1016/j.actaastro.2018.12.034

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The identification of nonlinear systems in aeroelasticity poses a significant challenge for practitioners, often hampered by the complex nature of aeroelastic response data which may contain multiple forms of nonlinearity. Characterizing and quantifying nonlinearities is further hampered when the response is obtained at a location which is away from the nonlinear source and/or the response is contaminated by noise. In the present paper, a three-degree-of-freedom airfoil with a freeplay nonlinearity located in the control surface and exposed to transonic flow is investigated. In this Part I paper the main form of analysis is via higher-order spectra techniques to unveil features of the nonlinear mechanism which result from i) structural nonlinearities (freeplay) in isolation and ii) freeplay with Euler derived nonlinear inviscid aerodynamic phenomena (transition between Tijdeman Type-A and Type-B shock motion). It is shown that the control surface structural freeplay nonlinearity is characterized by strong cubic phase-coupling between linear and nonlinear modes. On the other hand, nonlinear inviscid flow phenomena are shown to be characterized by quadratic phase-coupling between linear and nonlinear modular modes, the strength of which is related to the strength of the aerodynamic nonlinearity (amplitude of the shock motion). The nonlinear inviscid flow phenomena do not appear to affect the identification of the freeplay nonlinearity. Conjectures are made which address the transition between aperiodic, quasi-periodic and periodic behavior (pre-flutter), further physical support towards these conjectures is provided in Part II [1]. The limitations of the higher-order spectra approach are assessed, in particular, the analysis demonstrates the difficulty in extracting natural frequencies with this approach.

DOI： 10.1016/j.ymssp.2018.05.053

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Viscous effects on centreline shock reflection in an axisymmetric flow are studied numerically using Navier-Stokes and direct simulation Monte Carlo solvers. Computations at low Reynolds numbers have resulted in a configuration consisting of two shock waves, in contrast to the inviscid theory. On the other hand, computations at high Reynolds numbers have yielded a three-shock configuration in qualitative agreement with the inviscid theory prediction. This behaviour is explained by the presence of the so-called non-Rankine-Hugoniot zone, which accounts for the deviation of the shock structure from the inviscid paradigm. At Reynolds numbers on the verge of the transition from a two-shock to three-shock configuration, extremely high pressure that would be unattainable with the classical Rankine-Hugoniot relation for any shock configuration may occur. An analogy to the Guderley singularity in cylindrical shock implosion has been deduced for the shock behaviour from a mathematical viewpoint.

DOI： 10.1063/1.5085267

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Progress in spaceflight research has led to the introduction of various manned and unmanned reusable space vehicle concepts, opening up uncharted opportunities for the newborn space transport industry. For future space transport operations to be technically and commercially viable, it is critical that an acceptable level of safety is provided, requiring the development of novel mission planning and decision support tools that utilize advanced Communication, Navigation and Surveillance (CNS) technologies, and allowing a seamless integration of space operations in the current Air Traffic Management (ATM) network. A review of emerging platform operational concepts is conducted, highlighting both the challenges and the opportunities brought in by the integration with conventional atmospheric air transport. Common launch and re-entry planning methodologies are then discussed, where the physical and computational limitations of these approaches are identified and applicability to future commercial space transport operations is assessed. Attention is then turned to the on-orbit phase, where the unique hazards of the space environment are examined, followed by an overview to the necessary elements required for space object de-confliction and collision avoidance modelling. The regulatory framework evolutions required for spacecraft operations are then discussed, with a focus on space debris mitigation strategies and operational risk assessment. Within the atmospheric domain, possible extensions and alternatives to the conventional airspace segregation approaches are identified including promising Air Traffic Flow Management (ATFM) techniques to facilitate the integration of new-entrant platforms. Lastly, recent modelling approaches to meet on-orbit risk criteria are discussed and evolutionary requirements to improve current operational procedures are identified. These insights will inform future research on CNS/ATM and Avionics (CNS + A) systems and associated cyber-physical architectures for Space Traffic Management (STM).

DOI： 10.1016/j.paerosci.2018.10.006

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The Hilbert–Huang Transform is used to analyze the nonlinear aeroelastic response of a 2D 3DOF aeroelastic airfoil system with control surface freeplay under transonic flow conditions. Both static and dynamic aerodynamic conditions, i.e., for accelerating freestream speed, are considered using a linearized aerodynamic model. The main aim of this paper is to provide an in-depth physical understanding of the observed transition between periodic and aperiodic behavior, and the presence of a stable periodic region well below the domain characterized by stable limit cycles. Physical insights towards the forward and backward abrupt transition between aperiodic/chaotic and periodic behavior types appear to be the result of an internal resonance (IR) phenomenon between linear modes followed by a lock-in between linear and nonlinear modes. More specifically, initially a 2:1 IR between linear modes leads to a shift in the frequency composition and dynamic behavior of the system. A secondary effect of the IR can be observed immediately after the exact point of 2:1 IR such that a nonlinear mode locks into a subharmonic of the linear mode which in-turn drives a finite stable periodic region.

DOI： 10.1016/j.ymssp.2018.04.039

記述言語：英語   掲載種別：研究論文（学術雑誌）  

In the present paper, a numerical study on the flow control over a backward facing step using dielectric barrier discharge (DBD) plasma actuators has been conducted by means of high-fidelity computational fluid dynamic and multi-objective design optimization (MDO) based on surrogate-assisted evolutionary algorithms. The main objectives of this study are minimizing total pressure loss and reattachment length while maximizing the flow uniformity index. It employs four decision variables including input voltage, frequency, width of generated plasma and distance from the flow inlet to the start of the generated plasma. Sensitivity analysis has been performed with the aid of surrogate modeling; the results have been evaluated through MDO with evolutionary algorithms. It has revealed major impact of the DBD plasma actuator on the behavior of the flow and major improvements on the objective functions in relation to the decision variables. In particular, flow separation has been suppressed considerably while maintaining reasonable levels of flow uniformity and total pressure loss.

DOI： 10.1007/s42405-018-0045-z

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjets are a class of hypersonic airbreathing engine that offer promise for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen. An after-burning scheme can be used to effectively utilize the remaining hydrogen by supplying additional oxygen into the nozzle, aiming to augment the thrust. This paper presents the results of a single-objective design optimization for a strut fuel injection scheme considering four design variables with the objective of maximizing thrust augmentation. Thrust is found to be augmented significantly owing to a combination of contributions from aerodynamic and combustion effects. Further understanding and physical insights have been gained by performing variance-based global sensitivity analysis, scrutinizing the nozzle flowfields, analyzing the distributions and contributions of the forces acting on the nozzle wall, and examining the combustion efficiency.

DOI： 10.1016/j.actaastro.2018.04.012

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The authors regret that an error was identified in the expression of the mixing vorticity efficiency defined in Eq. (5) of the article. The correct expression is [Formula presented], where the effective injector diameter D must be in the numerator rather than in the denominator according to dimensional analysis. The author regrets this error, while the correct definition was used for the calculation of mixing vorticity efficiency in the study, and so this correction does not affect the results, discussions, or conclusions. The authors would like to apologise for any inconvenience caused.

DOI： 10.1016/j.actaastro.2018.01.024

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The Cusped Field Thruster (CFT) concept has demonstrated significantly improved performance over the Hall Effect Thruster and the Gridded Ion Thruster; however, little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective design optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by five design parameters (i.e., anode voltage, anode current, mass flow rate, and magnet radii), simultaneously aiming to maximize three objectives; that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. The research indicates that the anode current and the Outer Magnet Radius have the greatest effect on the performance parameters. An optimal value for the anode current is determined, and a trend towards maximizing anode potential and mass flow rate is observed.

DOI： 10.3390/aerospace4040055

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Higher-Order Spectra (HOS) are used to characterise the nonlinear aeroelastic behaviour of a plunging and pitching 2-degree-of-freedom aerofoil system by diagnosing structural and/or aerodynamic nonlinearities via the nonlinear spectral content of the computed displacement signals. The nonlinear aeroelastic predictions are obtained from high-fidelity viscous fluid-structure interaction simulations. The power spectral, bi-spectral and tri-spectral densities are used to provide insight into the functional form of both freeplay and inviscid/viscous aerodynamic nonlinearities with the system displaying both low- and high-amplitude Limit Cycle Oscillation (LCO). It is shown that in the absence of aerodynamic nonlinearity (low-amplitude LCO) the system is characterised by cubic phase coupling only. Furthermore, when the amplitude of the oscillations becomes large, aerodynamic nonlinearities become prevalent and are characterised by quadratic phase coupling. Physical insights into the nonlinearities are provided in the form of phase-plane diagrams, pressure coefficient distributions and Mach number flowfield contours.

DOI： 10.1017/aer.2017.88

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The development and accelerated use of optimization frameworks in aircraft design is a testament to their ability to identify optimal and often non-intuitive shapes as a result of multi-disciplinary design objectives. Airfoil design is a continuously revised multi-disciplinary problem, and is pivotal to illustrate the performance of optimization frameworks involving numerical simulation, flexible shape parametrization, and intelligent evolutionary algorithms. An often overlooked component of this classic problem is to consider the dynamic aeroelastic behavior under trim conditions, which can generate explicit boundaries to the flight envelope. Trim introduces a significantly strong coupling with objectives governing static performance, e.g. aerodynamic and/or structural, thereby resulting in a highly nonlinear and discontinuous design space. In this paper, a multi-objective particle swarm optimization framework for multi-disciplinary performance improvement is presented, pertaining to aerodynamic, structural and aeroelastic design criteria at trim conditions. The framework is assisted by the construction of adaptive Kriging surrogates, which is cooperatively used with the numerical solver to identify optimal solutions within a computational constraint. Designer preferences are introduced to reflect the optimal compromise between the objectives. Results of the optimization process indicate a large spread in design variable influence and interaction, and a subtle yet clear distinction between all objectives is illustrated through the catalog of final airfoil candidates obtained.

DOI： 10.1177/1045389X16679019

記述言語：英語   掲載種別：研究論文（学術雑誌）  

A precooling heat exchanger has been the key mechanism for realizing turbine-based combined-cycle engines at high flight Mach numbers. A multi-objective design optimization coupling surrogate-assisted evolutionary algorithms with numerical simulation has been carried out with respect to three design objectives, that is, maximization of heat transfer effectiveness, total pressure recovery, and compactness. Physical insights into the underlying compressible aerodynamic and aerothermal phenomena in the bare tube bank geometry have been gained through scrutinizing the flowfields for the representative cases. With the nature of operating conditions having relatively high inflow velocity, tube bank configurations that are potentially prone to have flow choking are removed in the optimization process. The results from the optimization have been investigated by analyzing the selected individuals on the Pareto-optimal front and performing sensitivity analysis with the aid of surrogate models. The effects of uncertainties in the design parameters on the precooler performance have been examined. The unconventional tube profiles comprising elliptic and obround sections are found to effectively reduce unfavorable flow separation and permit smaller tube spacing ratios, yielding higher heat transfer rate per unit volume than the conventional circular tube bank.

DOI： 10.2514/1.T4921

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Effective fuel injection and mixing is of crucial importance for reliable operation of scramjet engines, where fuel must be injected into high-speed crossflow and mixed with air at an extremely short timescale. This paper presents the results of a numerical study that investigates the effects of the injection angle and pressure for various orifice shapes on fuel mixing characteristics into hypersonic airflow at Mach 5, aiming at the application to scramjet operation with upstream fuel injection at Mach 10. The mixing performance has been evaluated with respect to the mixing efficiency, total pressure recovery, fuel penetration, and streamwise circulation. Significant influence of the injection angle and intensity on the mixing has been observed in conjunction with the geometric features of the injector orifice. An additional performance parameter, namely the mixing vorticity effectiveness, has been found to be an effective measure to quantify the contribution of the streamwise vorticity in mixing enhancement.

DOI： 10.1016/j.actaastro.2016.08.008

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen which under ideal conditions, can be utilized to introduce an after-burning scheme. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a single-objective design optimization considering three design variables with the objective of producing maximum thrust augmentation. It is found that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of variance-based global sensitivity analysis, force contribution breakdowns, analysis of the nozzle flowfields, analysis of the surface pressure and shear stress distributions acting on the nozzle wall and analysis of the combustion efficiency.

DOI： 10.1016/j.actaastro.2015.07.007

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable access-to-space and high-speed atmospheric transport. Afterburning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a numerical study of three parameters and the effect that they have on thrust augmentation. These parameters include the injection pressure, injection angle and streamwise injection position. It is shown that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of a force contribution breakdown, analysis of the nozzle flowfields and finally the analysis of the surface pressure and shear stress distributions acting upon the nozzle wall.

DOI： 10.12989/aas.2015.2.2.183

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Efficient fuel/air mixing plays a crucial role in successful operation of hypersonic airbreathing engines, particularly scramjets, where fuel must be injected into high-speed crossflow and mixed with air at an extremely short timescale. This paper presents the results of a numerical study that investigates the effects of various orifice shapes on fuel mixing characteristics into hypersonic airflow at Mach 5, aiming at the application to scramjet operation with upstream fuel injection at Mach 10. The performance of the injectors at an inclination angle of 45 deg are assessed with respect to various criteria such as the mixing efficiency, streamwise circulation, total pressure recovery, fuel penetration, and spread. Streamwise slot injectors have been found to yield higher mixing efficiency than the other injectors tested (namely, square, circular, diamond, and triangular injectors), owing to the buffering effects. Apparent higher total pressure recovery has been obtained with these rectangular injectors, but their advantages have diminished significantly with the alignment of the trailing-edge position. The highest vertical penetration has been achieved by the square injector, whereas the performances of other injectors with a sharp leading edge have been found to be affected considerably by the axis switch phenomenon due to inclined injection at higher injection pressure. A major influence of the fore and aft shapes of the orifice has been observed on the three-dimensional bow shock formation and wake recirculation, respectively.

DOI： 10.2514/1.B35581

記述言語：英語   掲載種別：研究論文（学術雑誌）  

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Efficient mixing of fuel and air is a major objective for the successful operation of scramjet engines, in which the supersonic combustion process must occur within an extremely short time frame in the hypersonic flight. This paper presents the insights gained into the key design factors and underlying mechanism for fuel/air mixing enhancement with streamwise vortices introduced by alternating wedges called hypermixers. The results of a parametric numerical study and sensitivity analysis suggest that narrow spanwise intervals between the alternating wedges and large fuel/air equivalence ratios are beneficial for effective mixing, while higher total pressure recovery is associated with shallower ramp angles and lower fuel/air equivalence ratios. Higher fuel penetration is achieved with steeper ramp angles and higher fuel/air equivalence ratios. Streamwise vortex circulation increases with wider spanwise spacing and steeper ramp angles, but effective mixing enhancement is observed only in the latter case.

DOI： 10.2514/1.B35638

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Fuel injection and mixing into air play a crucial role in the operation of hypersonic airbreathing propulsion systems, particularly scramjet engines featuring upstream fuel injection. This study applies an advanced design methodology combining computational fluid dynamics and evolutionary algorithms assisted by surrogate modeling to a multi-objective optimization for fuel injection in a Mach 5.7 crossflow after the initial compression in a scramjet intake operating at Mach 7.6. Optimization is performed for elliptical injector configurations defined by four design parameters (i.e., the injection angle, spanwise spacing, aspect ratio, and radius of the injector), simultaneously aiming to maximize three objectives, that is, fuel/air mixing, total pressure saving, and fuel penetration into the crossflow. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. Major effects of the injection angle and aspect ratio have been observed on all considered design criteria. The spanwise spacing has been found to have considerable influence on the total pressure recovery, fuel penetration, and lateral spread when the injection pressure is adjusted to maintain a constant fuel/air equivalence ratio. Low-angle fuel injection through a highly elliptic orifice with wide spanwise spacing demonstrated the most comprehensive advantages in overall aspects.

DOI： 10.2514/1.B35661

記述言語：日本語   掲載種別：研究論文（学術雑誌）  

Hypersonic air-breathing propulsion, in particular, scramjet (supersonic combustion ramjet) engines, is a promising technology for efficient and economical access-to-space and atmospheric transport. Axisymmetric air intakes based on the Busemann geometry offer appreciable efficiency with maximum total pressure recovery and minimum shock loss, but the inherently long geometry incurs large skin friction drag and structural weight, requiring shortening by some means. Two distinctly different configurations of Mach reflection are found to exist at the centerline for identical inflow conditions and intake lengths in the course of shortening by axial contraction (stunting). Parametric studies with steady and transient numerical simulations are performed to examine the inviscid transient flowfields with variations in the shortening length and freestream Mach number. This paper presents the results and flowfields with focus on the variations of the exit Mach number and temperature as well as intake drag and discusses the hysteresis observed in the stunting and reverse (stretching) process of the Busemann intakes.

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Designing high-performance air intakes is of crucial importance for the successful operation of hypersonic scramjet propulsion. This paper investigates the performance of axisymmetric intakes obtained by applying two shortening methods, namely, leading-edge truncation and stunting (longitudinal contraction), to the full Busemann intake in inviscid and viscous flowfields. The primary aim is to identify the key design factors and underlying flow physics in order to achieve the optimum performance with minimum weight by striking the balance between viscous and shock losses. The effects of intake shortening on performance are similar for the two methods for moderate length reduction (25%), conducing to considerable reduction in intake weight. Even reduced to half length, truncated intakes can produce reasonable compression (60% of the full Busemann intake) with the original level of total pressure recovery maintained in viscous flowfields. Stunted intakes, on the other hand, can enhance compression with considerable total pressure penalty, eventually leading to intake unstart at a certain point due to the emergence of Mach reflection, which makes this method potentially useful in situations where locally high pressure and temperature are desired near the centreline at the cost of total pressure recovery.

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjet propulsion is a promising hypersonic airbreathing technology that offers the potential for efficient and flexible access to space and high-speed atmospheric transport. Robust nozzle design over a range of operating conditions is of critical importance for successful scramjet operation. In this paper, shape optimization has been performed with surrogate-assisted evolutionary algorithms to maximize the thrust generated by an axisymmetric scramjet nozzle configuration, including the base flow and external surface for cruise conditions at Mach 8 at two altitudes with and without fuel. The optimization results have been examined in a coupled numerical/analytical approach in order to identify the key design factors and investigate the effects of design parameters. It has been found that the optimum nozzle geometries are characterized by bell-type shapes for the fuel-on conditions, whereas the optima for the fuel-off case feature nearly conical shapes. Their robustness in thrust production has been demonstrated by cross- referencing the optimum geometries at off-design altitudes. The nozzle length and radius have been found to be the most influential parameters in all considered conditions, with their optimum values determined based on the balance between inviscid and viscous force components, whereas the other parameters have minor impact on the total axial force.

DOI： 10.2514/1.B34482

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjet propulsion is a promising technology for reliable and economical access to space and high-speed atmospheric transport. The inlet plays a key role in determining the performance of scramjets, in particular for the axisymmetric class of scramjet engines that are currently explored due to their advantages in numerous aspects. In the present study, a multi-objective design optimization based on evolutionary algorithms has been conducted with respect to four major inlet design criteria, that is, compression efficiency, drag, adverse pressure gradient, and exit temperature, where the former three criteria are used as the objective functions and the last one is the constraint function. The flowfields have been examined for representative geometries, and sensitivity analysis has been performed with the aid of surrogate modeling, revealing the major impact of the inlet exit radius, advantages of Busemann-type geometries in various aspects, and direct correlation of the inlet drag, exit temperature, and surfaceheat transfer. The insight gained here can be usefully applied to the design of high-performance scramjet inlets.

DOI： 10.2514/1.J051644

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Reliable in-flight starting of the inlet is of critical importance for the successful operation of scramjet engines, particularly axisymmetric configurations with high-contraction inlets. The present research is undertaken to examine the capability of various inlet starting methods based on two principles: unsteady flow effects and variable geometries. Time-accurate viscous computations have been performed to investigate the transitional flowfields introduced by a variety of methods that are applicable to axisymmetric geometries. Parametric studies have been conducted for instantaneous rupture of conical diaphragms and addition of bleed slots, which induce highly unsteady flow phenomena. Several methods employing variable inlet geometries have been tested for the latter principle, including opening doors and sliding doors (or diaphragm erosion). Successful inlet starting has been achieved as a result of unsteady transition induced by diaphragm rupture and quasi-steady transition, due to the sliding-door opening process. In particular, a bleed addition to the diaphragm rupture method has been found to be highly effective and pronounced flow stability has been observed in the sliding-door process.

DOI： 10.2514/1.48284

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Scramjet airbreathing propulsion is a promising technology for efficient and economical access-to-space. Flow compression in the inlet and fuel combustion in the combustor play a major role in scramjet mechanism, their efficiencies crucially influencing the overall scramjet performance. A double-objective shape optimisation for an axisymmetric inlet and combustor configuration using hydrogen as fuel premixed into air has been performed for minimum total pressure loss and maximum combustion efficiency in the present study. A state-of-the-art MDO (multi-objective design optimisation) capability with surrogate-assisted evolutionary algorithms has been employed, coupled with a CFD solver for inviscid flowfields involving chemical reactions represented by Evans & Schexnayder's model. The obtained Pareto optimal front suggests the possibility of substantial improvement in efficiency and the counteracting nature of the two objective functions. Geometries with higher combustion efficiency are characterised by a higher compression inlet with larger leading-edge radius and a longer combustor, whereas opposite trends are observed for configurations with smaller total pressure loss.

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Flow control can be applied to shock wave/boundary layer interactions to achieve two different goals; the delay of shock-induced separation and/or the reduction of stagnation pressure losses, which cause wave drag or inlet inefficiencies. This paper introduces the principles and main techniques for both approaches and assesses their relative suitability for practical applications. While boundary layer suction is already in wide use for separation control, the most promising novel device is the micro-vortex generator, which can deliver similar benefits to traditional vortex generators at much reduced device drag. Shock control is not yet used on practical applications for a number of reasons, but recent research has focused on three-dimensional devices which promise to deliver flow control with improved offdesign behaviour. Furthermore, there are some indications that a new generation of control devices may be able to combine the benefits of shock and boundary layer control and reduce shock-induced stagnation pressure losses as well as delay shock-induced separation.

DOI： 10.1007/s00193-008-0149-7

記述言語：英語   掲載種別：研究論文（学術雑誌）  

Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established λ-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30 %, especially when significant wave drag is present Counter-rotating stream wise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.

DOI： 10.2514/1.32049

記述言語：英語   掲載種別：研究論文（学術雑誌）  

The use of wind-tunnel setup for study of normal shock wave/boundary layer interaction control, was investigated. The rectangular working section that was 114 mm wide, and 178 mm high at the straight downstream of the nozzle was used. The incoming airflow was partitioned by a plate of 6 mm thickness to overcome the problem of shock wave instability. The height of the upper and lower passage was maintained at 91 and 122 mm respectively. The incoming boundary layer thickness was 5.7 mm and the Reynolds number based on boundary-layer displacement thickness was approximately 25,000. It was observed that shock can be located above 3-D bump and large λ-shock structure whose front shock leg starts at the onset of control cab be analyzed. Result shows that wind-tunnel setup can be used to test various types of shock control at positions where conventional setups are unable to hold shock system due to shock instability.

DOI： 10.2514/1.24370

記述言語：英語   掲載種別：研究論文（学術雑誌）  

An analytical expression is proposed to estimate the wave drag of an aerofoil equipped with shock control. The analysis extends the conventional approach for a single normal shock wave in the absence of control, based on the knowledge that all types of successful shock control on transonic aerofoils cause bifurcated λ-shock structures. The influence of surface curvature on the λ-shock structure has been taken into account. The extended method has been found to produce fairly good agreement with the results obtained by CFD methods while requiring negligible computational effort. This new formulation is expected to be beneficial in the industrial design process of transonic aerofoils and wings where a large number of computational simulations have to be performed.

DOI： 10.1016/j.ast.2005.08.001

記述言語：英語   掲載種別：研究論文（学術雑誌）  

A two-dimensional blunt body was submerged in a compressible turbulent boundary layer and the surface pressure was measured. It was found the fore-body drag was reduced due to the velocity defect in the boundary layer and the associated re-circulation behind an oblique shock wave. A small two-dimensional shock generator ahead of the blunt body further reduced the drag of the blunt body.

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Daegu   国名：大韓民国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Daegu   国名：大韓民国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Hobart   国名：オーストラリア連邦  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Baku   国名：アゼルバイジャン共和国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Hainan   国名：中華人民共和国  

開催年月日： 2024年4月

記述言語：日本語   会議種別：口頭発表（招待・特別）  

開催地：名古屋   国名：日本国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2024年4月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：秋田   国名：日本国  

開催年月日： 2024年4月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：神奈川   国名：日本国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Orlando, FL   国名：アメリカ合衆国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Orlando, FL   国名：アメリカ合衆国  

開催年月日： 2024年4月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Busan   国名：大韓民国  

開催年月日： 2023年7月

記述言語：日本語   会議種別：口頭発表（一般）  

国名：日本国  

開催年月日： 2023年6月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：東京   国名：日本国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sendai   国名：日本国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：National Harbor, MD   国名：アメリカ合衆国  

開催年月日： 2023年6月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：滋賀   国名：日本国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Kanazawa   国名：日本国  

開催年月日： 2023年6月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：角田   国名：日本国  

開催年月日： 2023年6月

記述言語：日本語  

開催地：オンライン   国名：日本国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Bruges   国名：ベルギー王国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Paris   国名：フランス共和国  

開催年月日： 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Sydney   国名：オーストラリア連邦  

開催年月日： 2023年6月

記述言語：日本語  

開催地：オンライン   国名：日本国  

開催年月日： 2023年5月 - 2023年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Bangalore   国名：インド  

開催年月日： 2022年6月 - 2023年7月

記述言語：日本語   会議種別：口頭発表（一般）  

国名：日本国  

開催年月日： 2022年5月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Toulouse   国名：フランス共和国  

開催年月日： 2022年3月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：米子   国名：日本国  

開催年月日： 2022年3月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：米子   国名：日本国  

開催年月日： 2022年3月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：米子   国名：日本国  

開催年月日： 2022年2月 - 2022年3月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Beppu, Oita (online)   国名：日本国  

開催年月日： 2022年2月 - 2022年3月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Beppu, Oita (online)   国名：日本国  

開催年月日： 2022年1月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Jeju   国名：大韓民国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年10月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年8月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年8月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年6月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2021年2月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年2月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年2月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年2月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年1月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2021年1月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Virtual   国名：アメリカ合衆国  

開催年月日： 2020年12月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Assam (online)   国名：インド  

開催年月日： 2020年11月

記述言語：日本語   会議種別：口頭発表（一般）  

開催地：オンライン   国名：日本国  

開催年月日： 2020年10月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Lisbon (virtual)   国名：ポルトガル共和国  

開催年月日： 2020年10月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：仙台（オンライン）   国名：日本国  

開催年月日： 2020年9月

記述言語：日本語  

開催地：オンライン   国名：日本国  

開催年月日： 2020年8月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：South Lake Tahoe, CA (online)   国名：アメリカ合衆国  

開催年月日： 2020年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Tokushima   国名：日本国  

開催年月日： 2020年6月

記述言語：英語   会議種別：シンポジウム・ワークショップ パネル（公募）  

開催地：Fukuoka   国名：日本国  

開催年月日： 2020年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Montreal   国名：カナダ  

開催年月日： 2020年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Montreal   国名：カナダ  

開催年月日： 2020年6月

記述言語：英語   会議種別：口頭発表（一般）  

開催地：Tokushima   国名：日本国  

開催年月日： 2019年1月

記述言語：英語  

開催地：San Diego   国名：アメリカ合衆国  

Electric Propulsion (EP) is a suitable propulsion technology for satellite and space missions, offering advantages over chemical propulsion in various aspects including fuel consumption hence launch cost. The cusped field thruster (CFT) offers advantages over other types of EP such as the gridded ion thruster and Hall effect thruster, with enhanced electron confinement owing to the magnetic mirror and reduced particle loss effects at the dielectric wall. The increasing demand for performance improvement of the propulsion system while downscaling for micro-satellite class platforms has led to considerable efforts dedicated to physical modeling and performance characterization of downsized CFT. In the present study a multi-objective design optimization (MDO) study has been conducted to characterize the performance to maximize three performance objectives of downscaled CFT, namely, thrust, total efficiency, and specific impulse defined by common design parameters, namely, anode voltage, anode current, mass flow rate and geometric configuration. Particle-in-cell simulations have been performed for the selected design points identified in MDO studies for verification by accurately accounting for phenomena and performance losses that originate from uncertainties and complexities associated with the thruster design and physics.

開催年月日： 2018年10月

記述言語：英語  

開催地：Chengdu   国名：中華人民共和国  

The cusped field thruster (CFT) is a class of advanced electric propulsion (EP) technology for satellite and space missions, offering advantages over other types of EP including enhanced electron confinement owing to the magnetic mirror and reduced particle loss effects at the dielectric wall. The increasing demand for downscaling for micro-satellite class platforms while keeping performance at similar level has led to considerable efforts dedicated to physical modeling and performance characterization of downsized CFT. Multi-objective design optimization is conducted in this study by employing performance parameters of downscaled CFT, namely, thrust, total efficiency, and specific impulse as the objective functions to maximize and design parameters including anode voltage and current, mass flow rate, and inner and outer magnet radii as the decision variables. Two geometric configurations are considered, i.e., those comprising three magnets with fixed thickness and four magnets with variable thickness to gain insights into the influence of magnet thickness on the performance. Considerable effects of magnet thickness on the performance have been found, including thrust increase of up to approximately 20% and increase in specific impulse by up to approximately 10%, as compared to the configuration with fixed thickness magnets.

開催年月日： 2018年10月

記述言語：英語  

開催地：Chengdu   国名：中華人民共和国  

The aerodynamic behaviour of transonic flow around a supercritical aerofoil is strongly influenced by shock-wave/boundary-layer interaction (SBLI) due to compressible and viscous effects. SBLI causes undesirable effects in various manners including flow instability, drag rise, and buffet, which crucially limit the flight envelop hence operation. In this paper, a numerical investigation is conducted for an OAT15A supercritical aerofoil under a typical buffet onset condition. Unsteady Reynolds-Averaged Navier-Stokes (URANS) equation is used to simulate the compressible, viscous flowfield. A two-dimensional (2D) surface bump based on preceding research on SBLI control is employed as a flow control device. It is placed on the suction side of the aerofoil relative to the shock position, with a fixed location of 27% of the chord length. A freestream condition of Mach 0.73 and a 3.5° angle of attack have been considered for the unsteady flowfield. It has been found that the trailing edge vortices within the separation bubble have considerable influence on self-sustained shock oscillation by scrutinising the flowfields in the presence/absence of bump control. The establishment of a λ-shock structure effectively restricts the motion of the front shock leg without incurring significant re-expansion generated by the moving rear shock leg. This subsequently suppresses flow separation at the trailing edge within an acceptable range, and attenuates the periodic lift fluctuation associated with the oscillating shock movement.

開催年月日： 2018年9月

記述言語：英語  

開催地：Orlando   国名：アメリカ合衆国  

High-performance hydrocarbon-fuelled scramjet engines require efficient fuel-air mixing due to the relatively short flow residence time through the combustor. At high temperatures, hydrocarbon fuels react endothermically and absorb thermal energy from the surroundings. The process known as cracking becomes essential at high Mach numbers to increase the total heat-sink capacity of the fuel. This study presents the results of numerical simulations that investigate the mixing characteristics of cracked gaseous heavy hydrocarbon fuels injected through a circular, flush-wall porthole injector inclined at 45-deg to the freestream. The mixing characteristics of six fuel compositions representing various cracking efficiencies ranging from 0-100% are investigated. The mixing rates and flow structures are found to change with fuel compositions. As the cracking increases, the mixing and streamwise circulation increase for an injectant. However, the jet penetration and stagnation pressure losses decrease. The density gradients determine the strength of vorticity in the vicinity of the injector. The streamwise circulation is found to have a strong influence on the mixing and the strength of bow shock on the jet penetration. Overall, it is shown that there are mixing benefits to be gained by injecting cracked hydrocarbon fuels compared to heavy uncracked fuels in scramjets.

開催年月日： 2018年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

The Hilbert-Huang Transform is used to analyze the nonlinear aeroelastic response of a 2D 3DOF aeroelastic airfoil system with control surface freeplay under transonic flow conditions. Both static aerodynamic condition and dynamic aerodynamic condition, that is for accelerating freestream speed, are considered using a linearized aerodynamic model. The main aim of this paper is to provide physical insights as to the observed transition between periodic and aperiodic behavior, and the presence of a stable periodic region well below the region characterized by stable limit cycles. Physical insights towards this transition are presented by showing that multiple internal resonances exist. It is shown that the abrupt transition from aperiodic / chaotic to periodic behavior is a result of multiple internal resonances between linear and nonlinear modes. Initially a 2:1 IR between linear modes leads to a shift in the frequency composition and dynamic behavior of the system, then immediately a secondary 2:1 IR occurs between of linear and nonlinear modes which drives a stable periodic region.

開催年月日： 2018年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

Rocket-based Combined-Cycle (RBCC) engines offer a promise for efficient and flexible propulsion over a large Mach number range by combining rocket and ramjet/scramjet technology. To achieve this, an RBCC engine uses four different modes of operation: rocket, ramjet, scramjet and dual-mode. During operation, the engine must make the transition from subsonic to supersonic combustion, i.e., ramjet to scramjet mode. The objective of this study is to gain physical insight into the ramjet-scramjet-ramjet mode transition by elucidating the underlying mechanics. Numerical simulations with chemical reactions have been performed for the transient flowfields of a two-dimensional RBCC combustor by using an unsteady Reynolds-Averaged Navier-Stokes solver. Mode transition is effected by changing the flow rate of the secondary hydrogen fuel injectors installed on the top and bottom walls of the combustor. A parametric study was conducted to investigate the characteristic and behavior of RBCC combustion in mode transition. The results indicated that transition is affected considerably by the presence and development of flow separation and pseudo-shock structures near fuel injectors. The complex effects of aerodynamic and aerothermal interactions on the transient flowfields and performance, along with a hysteresis observed between the scramjet-to-ramjet and ramjet-to-scramjet transition processes.

開催年月日： 2018年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

In the present research, the effects of dielectric barrier discharge plasma actuators on the flow control behind a backward facing step with various step angles are analyzed. A multi-objective optimization study has been conducted by means of evolutionary algorithms assisted by surrogate modelling coupled with computational fluid dynamics. A set of decision variables including input voltage, frequency, width of generated plasma and distance from the flow inlet to the start of the generated plasma are employed for an optimization problem aiming to simultaneously minimize the total pressure loss and reattachment length, while maximizing the uniformity of the flow. The flowfields have been analyzed for selected individuals from the Pareto optimal front in comparison with the baseline and reference results in the absence of a plasma actuator. Global sensitivity analysis has been performed to identify key design parameters for flow control. It has revealed major impact of the design parameters of the DBD plasma actuator on the behavior of the flow and major improvements in performance for all three angle configurations. Flow separation has been suppressed considerably while achieving moderate improvements in the flow uniformity and total pressure loss.

開催年月日： 2017年9月

記述言語：英語  

開催地：Adelaide   国名：オーストラリア連邦  

A synopsis of missions to icy bodies, in particular ones involving the resampling of the plumes of Enceladus, is offered. Following the speculation of others, it is assumed that any microbes possibly existing in the putative water mantle of Enceladus could be swept-up by the plume. After consideration of instrument capabilities, it is concluded that no planned or proposed hypervelocity plume fly-through missions is likely to result in a definitive confirmation of such possible life. Instead, it is suggested that plume particle collection with relative encounters at ∼200 ms^-1 with more relevant detection instruments, is far more likely to result in definitive confirmation. An illustrative model of the Enceladus plume is also presented to justify the need for collection of larger ice particles at altitudes of ∼3 km. To provide the recommended encounters at low velocity and altitude the potential use of tethered collection systems is briefly introduced. It is concluded that an Enceladus orbiter should be prioritized.

開催年月日： 2017年9月

記述言語：英語  

開催地：Adelaide   国名：オーストラリア連邦  

Efficient and flexible space transportation systems are fundamental to the future success of routine scientific, commercial and strategic space missions. RBCC (rocket-based combined cycle) systems comprising rockets and airbreathing engines such as ramjet and scramjets are a reusable launch propulsion concept, offering considerable advantages for space transport over traditional propulsion in various aspects, including efficiency and flexibility. Multiobjective design optimisation studies have been performed for a conceptual RBCC-based TSTO (two-stage-to-orbit) system by means of evolutionary algorithms and trajectory optimisation with respect to important design criteria. The results highlight complex and nonlinear aerodynamic and propulsion characteristics, associated with numerous design characteristics, whilst evaluating the feasibility of the transportation system and providing physical insight into key design parameters with predominant influences identified by sensitivity analysis.

開催年月日： 2017年9月

記述言語：英語  

開催地：Adelaide   国名：オーストラリア連邦  

There is a clear demand for advanced electric propulsion systems for current and future satellite applications for a variety of commercial and research missions to reduce launch costs. The HEMP-T/CFT propulsion systems are very complex and require a detailed analysis of how design criteria influence performance. By downscaling the CFT, further propellant and system reductions would result in lower costs while maintaining high performance. The CFT concept has demonstrated significantly improved performance over the HET and GIT, however little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by 5 design parameters (i.e., Φ_a, I_a, m_a, and magnet radii), simultaneously aiming to maximize 3 objectives, that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the 3 design objectives and additional performance measures. Significant effects of the anode power and magnet radii have been observed on the considered design criteria with the anode current exhibiting the most significant degree of influence on all 3 objectives. Several optimum design points were analyzed and one has demonstrated the most comprehensive advantages in important design criteria.

開催年月日： 2017年1月

記述言語：英語  

開催地：Grapevine   国名：アメリカ合衆国  

The identification of nonlinear systems in aeroelasticity poses a significant challenge for practitioners, often hampered by the complex nature of aeroelastic response data which may contain multiple forms of nonlinearity. Characterizing and quantifying nonlinearity is further complicated when the dynamic oscillations are of a high–amplitude limit cycle form, masking the underlying nonlinear contributions. In the present paper, a three– degree–of–freedom airfoil with freeplay in the control surface and transonic aerodynamics is investigated. The main form of analysis is via higher–order spectra to unveil the form of nonlinearity that i) freeplay with linearized aerodynamics and ii) freeplay with nonlinear inviscid aerodynamic phenomena will produce. It is shown that the freeplay nonlinearity with linearized aerodynamics is characterized solely by cubic interactions, i.e., the quadratic interactions are negligible. However, when considering the Euler-based CFD simulations, as the amplitude of the of the oscillations increases and Type-B shock motion becomes more apparent, the strength of the quadratic interactions becomes prominent. The findings demonstrate how the interaction between the different types of nonlinearity and the different solution methods affect the nonlinear spectral content of the system and how different forms of nonlinearity can be characterized by their higher–order spectra.

開催年月日： 2016年12月

記述言語：英語  

開催地：Perth   国名：オーストラリア連邦  

The flowfield characteristics of a relatively small quasi-axisymmetric scramjet vehicle at a Mach 6 flight condition have been investigated numerically. To this end, three-dimensional, compressible, turbulent, reacting flow calculations with a finite rate chemistry model consisting of 33 reactions and 13 species and two-equation SST k-ω RANS model have been performed. Hydrogen is used as the fuel and the injection pressure of fuel is varied from 1 to 7 MPa in order to study the effect of the injection pressure on the flowfields in the scramjet model. The combustion length has been found to decrease as the injection pressure increases. However, rather little thrust gain has been achieved by combustion heat release in the present configurations.

開催年月日： 2016年9月

記述言語：英語  

開催地：Daejeon   国名：大韓民国  

Higher-order spectra (HOS) are utilized to investigate the nonlinear flutter behavior of a two-dimensional pitch/plunge airfoil system in transonic flow with freeplay and aerodynamic nonlinearity. It is shown that on the route to diverging flutter the system begins with the pitching and plunging modes being uncoupled and undergoes various evolutions exhibiting quasi-periodic behavior as it moves towards an ordered state of high-amplitude periodic limit cycle behavior. New physical insights come from the bispectral densities being computed at critical stages as the system evolves from through different stages of periodicity with the coalescence of the pitching and plunging frequencies into a single coupled limit cycle. Based on these estimates the nonlinear interaction mechanisms which occur within the nonlinear aeroelastic system prior to diverging flutter can be characterized. Furthermore, aerodynamic nonlinearity in the form of Tijdeman Type-B shock motion is quantified and its effect in combination with freeplay assessed.

開催年月日： 2016年1月

記述言語：英語  

開催地：San Diego   国名：アメリカ合衆国  

Higher-order spectra (HOS) analysis is utilized to analyze the nonlinear utter aspects of the AGARD 445.6 wing with pitching and plunging freeplay nonlinearities installed at the root. High-fidelity fluid-structure interaction simulations are conducted and the transient responses are monitored at the root and tip of the wing. The transient responses are truncated such that only the static portion of the signal remains and and the HOS are evaluated. The pitching/plunging freeplay ranges of motion and initial setting angle are varied, and the effect on nonlinear interactions and transient response are monitored. It is demonstrated that the presence of pitching freeplay is a major driver in the nonlinear phenomena, that the amplitude of the LCO is highly sensitive to the plunging freeplay range of motion whilst the presence of nonlinear interactions are dependent on interactions between the pitching and plunging freeplay effects. The application of this methodology from a practical perspective includes structural health monitoring, and the ability to identify and mitigate undesirable nonlinear aeroelastic phenomena in the design phase.

開催年月日： 2015年11月

記述言語：英語  

開催地：Cairns   国名：オーストラリア連邦  

In this work the effect of the tube geometry on the performance of a staggered tube cross-flow heat exchanger, similar to the inlet air pre-coolers in hypersonic aircrafts and air-breathing launch vehicles, is studied using computational fluid dynamics (CFD) at sea level conditions with air velocity of up to 50 m/s. Single tube simulations showed a decrease in drag coefficient and increased utilization of tube surface area for heat exchange when elliptical or obround tubes are used instead of a circular tube. In a tube bank, the heat transfer coefficient is enhanced and the friction factor is reduced by using the elliptical and obround tubes as compared to circular tubes. Overall, for the same heat duty, elliptical and obround tube pre-coolers are lighter and compact but the pressure drop is higher than a circular tube pre-cooler.

開催年月日： 2015年10月

記述言語：英語  

開催地：Kobe   国名：日本国  

Higher-order spectra analysis is used to analyze the nonlinear flutter characteristics of the AGARD 445.6 wing benchmark model, modified to include freeplay structural nonlinearities in pitch and plunge at the wing root. High-fidelity coupled CFD-CSD fluid-structure interaction simulations are conducted and bispectral densities are computed from selected aeroelastic wing dynamic response signals to identify the presence of quadratic nonlinearities within the available time histories. Initially, various freeplay conditions are investigated, before focusing on the intensity of nonlinear behavior at various wing spanwise stations. The considerable effect of freeplay is clearly demonstrated as the wing enters a bounded limit cycle oscillation (LCO) with amplitudes several orders of magnitude larger than that of the case without freeplay. Nonlinear coupling is demonstrated via self-interaction of the LCO frequency and coupling between the LCO frequency and the first torsional mode, this is shown to be strongest at the root where the pitch / plunge mechanism is installed and torsion is most prominent, dissipating spanwise from the pitch / plunge mechanism as the torsional motion becomes less significant. Modal analysis, alongside analysis of the leading-trailing edge phase change and wing deformation shapes supports the findings of the HOS, generating understanding of the underlying physical attributes leading to the observed nonlinear aeroelastic phenomena.

開催年月日： 2015年10月

記述言語：英語  

開催地：Kobe   国名：日本国  

The ongoing use and development of optimization frameworks for aircraft design is due to their ability to identify optimal and often non-intuitive shapes pertaining to the multi-disciplinary design criteria. Airfoil design is a continuously revised multi-disciplinary problem in the literature, and is pivotal to illustrate the performance of optimization frameworks involving numerical simulation, flexible shape parameterization, and intelligent evolutionary algorithms. An often overlooked component of this classic problem is to consider the dynamic aeroelastic behavior under trim conditions, which can impose explicit boundaries to the flight envelope. In this paper, a multiobjective particle swarm optimization framework is presented, pertaining to aerodynamic and aeroelastic design criteria at the trim condition. Designer preferences are used to reflect the optimal compromise between the objectives. Results of the optimization process indicate a large spread in design variable influence and interaction, and a subtle yet clear distinction between all objectives is illustrated through the final airfoil shapes obtained.

開催年月日： 2015年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

A multi-objective design optimization study has been conducted for upstream fuel injection through porous media applied to the first ramp of a two-dimensional scramjet intake. The optimization has been performed by coupling evolutionary algorithms assisted by surrogate modeling and computational fluid dynamics with respect to three design criteria, that is, the maximization of the mixing efficiency, total pressure saving, and fuel penetration. A distinct Pareto optimal front has been obtained, highlighting the counteracting behavior of the total pressure against fuel penetration, while the mixing performance crucially depends on the fuel/air equivalence ratio, as suggested by an additional optimization using the absolute mixing quantity. The Darcian and Forchheimer coefficients in the porous flow direction have been identified as the key design parameters in conjunction with the geometric parameters as a result of a sensitivity analysis. Flowfield visualization has revealed the presence of local hot pockets with intensely high pressure and temperature offered by a long injector positioned upstream due to augmented shock interactions.

開催年月日： 2015年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

Scramjets are a class of hypersonic airbreathing engine that are associated with realizing the technology required for economical, reliable and high-speed access-to-space and atmospheric transport. The expanding flow in the scramjet nozzle comprises of unburned hydrogen which under ideal conditions, can be utilized to introduce an after-burning scheme. After-burning augments the thrust produced by the scramjet nozzle and creates a more robust nozzle design. This paper presents a single-objective design optimization considering three design variables with the objective of producing maximum thrust augmentation. It is found that significant levels of thrust augmentation are produced based upon contributions from increased pressure, mass flow and energy in the nozzle. Further understanding of the phenomenon by which thrust augmentation is being produced is provided in the form of variance-based global sensitivity analysis, force contribution breakdowns, analysis of the nozzle flowfields, analysis of the surface pressure and shear stress distributions acting on the nozzle wall and analysis of the combustion efficiency.

開催年月日： 2015年1月

記述言語：英語  

開催地：Kissimmee   国名：アメリカ合衆国  

Plasma actuators are all electrical devices capable of altering flow paths and preventing separated boundary layers. The application to low pressure turbine stages in axial turbine engines offers further fuel eficiency improvements at low Reynolds numbers. Validated against wind-tunnel experiments, a numerical plasma model via steady-state RANS tur- bulence modeling simulations was developed. Reasonable agreement has been observed in the calibration of the numerical plasma model to experimental data. Further parametric studies are considered, aiming to optimize control, reduce pressure loss for turbine stages, and offset actuator power consumption. Results indicate that optimum flow control and positioning can prevent flow separation, reducing aerodynamic drag by 70% and offset power consumption inclusive of the actuator by up to 20 ± 4 %.

開催年月日： 2014年12月

記述言語：英語  

開催地：Melbourne   国名：オーストラリア連邦  

Plasma actuators are all electrical devices capable of altering flow paths and reattaching separated boundary layers. The application to low pressure turbine stages in axial turbine engines indicates further fuel efficiency improvements at low Reynolds numbers. Using wind tunnel experiments, a numerical plasma model through steady-state simulations with RANS turbulence modeling was developed. Reasonable agreement has been observed in the calibration of the numerical plasma model to experimental data. Further parametric studies are considered, aiming to optimize control, reduce pressure loss for turbine stages, and reduce actuator power consumption.

開催年月日： 2014年12月

記述言語：英語  

開催地：Melbourne   国名：オーストラリア連邦  

A numerical study of scramjet after-burning through the injection of liquid oxygen into the nozzle is conducted. A maximum thrust augmentation of 300% is found. An understanding of the thrust augmentation phenomenon is provided in the form of a force contribution breakdown, analysis of the nozzle flowfields and finally the analysis of the surface pressure and shear stress distributions acting upon the nozzle wall.

開催年月日： 2014年12月

記述言語：英語  

開催地：Melbourne   国名：オーストラリア連邦  

Preliminary thrust tests carried-out on a 0.5 metre diameter, helical rotor based on a sketch by Leonardo da Vinci are presented, along with the results of computational flow simulation. Some of the key geometric variables that define the helical rotor and their influence on the maximum achievable thrust level are also discussed. It is concluded that with sufficient input power, da Vinci's helical rotorcraft could achieve hovering flight; however, the technical challenges involved in a developing a free-flight hovering demonstrator would be formidable.

開催年月日： 2014年9月

記述言語：英語  

開催地：Shanghai   国名：中華人民共和国  

Rocket-based combined cycle (RBCC) engines are an airbreathing propulsion technology that offers considerable potential for efficient access-to-space. Successful design of RBCC-powered space transport systems requires reliable databases for both vehicle and engine performance, calling for an effective sampling method to accurately resolve non-linear characteristics in vast design space. This paper presents an optimal sampling strategy based on the function gradients to realize efficient database construction based on evolutionary algorithms and assesses its effectiveness by applying the methodology to various test functions with multiple objectives as well as surrogate models representing scramjet intake characteristics for validation.

開催年月日： 2014年9月

記述言語：英語  

開催地：St. Petersburg   国名：ロシア連邦  

Fuel injection into crossflow behind a backwardfacing step is studied by means of multi-objective design optimization, aiming at fuel/air mixing for supersonic combustion of scramjet propulsion. A variety of injector configurations have been examined in the optimization process using evolutionary algorithms in conjunction with local search methods and surrogate modeling. Data mining has been performed by applying statistical techniques including variance-based sensitivity analysis to the surrogate models constructed with solutions from computational fluid dynamics. The injection angle and backward step height have been found to be the most influential design parameters on the mixing performance for the configurations considered in this study.

開催年月日： 2014年6月

記述言語：英語  

開催地：Atlanta, GA   国名：アメリカ合衆国  

A multi-objective design optimization study coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) comprising airbreathing components besides rocket engines, aiming to examine its feasibility to achieve efficient access to space, particularly to the international space station. The optimization has been performed with respect to three important design criteria, that is, the maximization of the final velocity, altitude, and mass at the terminus of the orbiter trajectory under certain constraints of acceleration and dynamic pressure. The results have revealed complex interactions of numerous design parameters and a counteractive trend between the final velocity and mass. Most influential parameters have been identified from trajectory investigation and sensitivity analysis, providing insights into the design requirements needed to fulfill the desired mission with the vehicle and propulsion configurations considered here.

開催年月日： 2014年1月

記述言語：英語  

開催地：National Harbor, MD   国名：アメリカ合衆国  

In this study, a multi-objective design optimization coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for the first stage of two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) engine which combines rockets and ramjets by blending two kinds of vehicle configurations with different aerodynamic characteristics. The design criteria include the minimization of fuel consumption and the maximization of the final Mach number up to a separation of the TSTO system at the maximum altitude under certain ranges of acceleration and dynamic pressure. The optimization results reveal a counteractive trend between the final Mach number and fuel mass ratio and the major impact of effective specific impulse on those two objectives, which is mainly controlled by thrust throttling parameter within the trajectory optimization. In addition, the RBCC-powered vehicle tends to fly at lower altitude to attain the minimum fuel mass ratio, in contrast to the case for maximum final Mach number, which is attributed to the hybrid aerodynamic performance of the two configurations. The insight gained here can be usefully applied to the design of high-performance RBCCpowered vehicles.

開催年月日： 2014年1月

記述言語：英語  

開催地：National Harbor   国名：アメリカ合衆国  

In this study, a multi-objective design optimization coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for the first stage of two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) engine which combines rockets and ramjets by blending two kinds of vehicle configurations with different aerodynamic characteristics. The design criteria include the minimization of fuel consumption and the maximization of the final Mach number up to a separation of the TSTO system at the maximum altitude under certain ranges of acceleration and dynamic pressure. The optimization results reveal a counteractive trend between the final Mach number and fuel mass ratio and the major impact of effective specific impulse on those two objectives, which is mainly controlled by thrust throttling parameter within the trajectory optimization. In addition, the RBCC-powered vehicle tends to fly at lower altitude to attain the minimum fuel mass ratio, in contrast to the case for maximum final Mach number, which is attributed to the hybrid aerodynamic performance of the two configurations. The insight gained here can be usefully applied to the design of high-performance RBCCpowered vehicles.

開催年月日： 2013年1月

記述言語：英語  

開催地：Grapevine, TX   国名：アメリカ合衆国  

Efficient fuel/air mixing plays a crucial role in successful operation of hypersonic airbreathing engines, particularly scramjets, where fuel must be injected into high-speed crossflow and mixed with air at an extremely short timescale. This paper presents the results of a numerical study that investigates the effects of various orifice shapes on fuel mixing into hypersonic airflow at Mach 5, aiming at the application to scramjet operation at Mach 10. The performance of the injectors are assessed with respect to various criteria such as the mixing efficiency, streamwise circulation, total pressure recovery, and fuel penetration. The injection flowfields are scrutinized in order to identify the key geometric features and underlying flow mechanism that are responsible for mixing enhancement, with particular focus on the interactions of the jet plume with streamwise vortices.

開催年月日： 2012年12月

記述言語：英語  

開催地：Launceston   国名：オーストラリア連邦  

Scramjets are a promising hypersonic airbreathing technology for economical access-to-space and atmospheric transport. Reliable scramjet inlet starting is of crucial importance for successful scramjet operation. It is shown that unsteady flow, resulting from the rupture of a diaphragm, can result in a steady started inlet flow. Time-accurate computations illustrate flow and wave motion during the unsteady starting process. The effects of viscosity and diaphragm mass on the inlet flowfields are investigated by time-accurate computational simulations.

開催年月日： 2012年9月

記述言語：英語  

開催地：Tours   国名：フランス共和国  

Scramjet propulsion is a promising hypersonic airbreathing technology that can enable flexible and economical systems for access-to-space and atmospheric cruise in both civilian and strategic applications. Fuel injection and efficient mixing play a crucial role in scramjet operation, which depends critically on the sequential flow process, particularly for scramjet configurations featuring upstream fuel injection. However, designing high-performance injection systems represents a substantial challenge to conventional design approaches due to complex flowfields introduced by highly coupled aerodynamic phenomena. The present study is conducted by applying an advanced methodology combining computational fluid dynamics and evolutionary algorithms assisted by surrogate modeling to a multi-objective optimization problem for high-performance injector design, assuming scramjet operation at Mach 7.6. Optimization is performed for elliptical injector configurations defined by four design parameters, namely, injection angle, spacing, aspect ratio, and hole radius, simultaneously aiming to maximize three objectives, that is, fuel/air mixing, total pressure recovery and fuel penetration into air. Flowfields are scrutinized for selected injector configurations and global sensitivity analysis is applied to the surrogate models trained during the optimization, in order to gain physical insight into underlying flow mechanism and to identify key design factors for mixing enhancement. It has been found that the injection angle and aspect ratio are primarily responsible for fuel/air mixing efficiency, while fuel penetration largely depends on the injector spacing in conjunction with the injection angle.

開催年月日： 2012年9月

記述言語：英語  

開催地：Tours   国名：フランス共和国  

A 2D numerical investigation was undertaken utilising transient, time-accurate, RANS computational fluid dynamics, with the aim of investigating the influence of viscous and unsteady effects when attempting to start a hypersonic scramjet inlet. A variable geometry inlet starting mechanism incorporating sliding doors was utilised, with the two doors initially extending upstream from the scramjet leading edge. After establishing a steady-state solution, the doors were retracted up and over the inlet and the resulting flow field captured. A started flow regime was achieved only when retracting the doors at such a speed (0.1ms total retraction time) that at no point could quasi-steady Kantrowitz assumptions be applied to the internal flow field. Viscous effects were seen to play a significant role in the inlet starting process, with unstart being obtained earlier for lower Reynolds number conditions. Shockwave boundary layer interactions (SWBLI) were found to be the key driver behind inlet unstart.

開催年月日： 2012年9月

記述言語：英語  

開催地：Brisbane   国名：オーストラリア連邦  

Hypersonic airbreathing engines offer great potential for reliable and economical access-tospace and high-speed atmospheric cruise for both civilian and strategic applications. Scramjet (supersonic combustion ramjet) propulsion, in particular, is a promising technology to materialize efficient and flexible transport systems by removing the need to carry oxidizers and other limitations of rocket engines. In the actual development procedure of aerospace applications, the design for fabrication is carefully determined in consideration of various requirements and criteria, based on the optimal results obtained in numerical analysis and experiments. This stage typically involves a significant amount of engineering activities and human-related factors, which may well give rise to unexpected errors, uncertainty, and accuracy loss in the design values. Discrepancies in the design between the numerical/analytical solutions and the actual product can also arise during operation, originating from various factors such as structural deformation due to aerodynamic loads and surface ablation due to aerothermal heating in hypersonic flight. It is desirable that such discrepancies do not drastically affect the performance of the design. In this work, we present our findings of Multi-objective robust design optimization of the nozzle and external contour of an axisymmetric scramjet. We analyze the effect that uncertainties in the design variables can have on the final solutions and try to understand the behavior from the physical point of view.

開催年月日： 2011年4月

記述言語：英語  

開催地：San Francisco, CA   国名：アメリカ合衆国  

Scramjet propulsion is a hypersonic airbreathing technology that offers the great potential for economical and flexible access to space and high-speed atmospheric transport. This paper presents the results and physical insight obtained in a design optimization study conducted for axisymmetric scramjet engines. A single-objective optimization has been performed to maximize the performance of a full flow-path scramjet configuration in the presence of nonuniform upstream fuel injection represented by a set of geometric parameters and injection scaling factors. Use is made of a state-of-the-art design methodology coupling a high-fidelity CFD code with an advanced optimization capability based on evolutionary algorithms assisted by surrogate modeling. The flowfields have been investigated to identify key factors and gain insight into underlying physics particularly in comparison with premixed fuel/air flowfields. The optimum flowfield with fuel injection is characterized by a mixed supersonic/subsonic flow regime similar to a transitional mode. The optimum design has been achieved by maximizing the nozzle thrust while reducing skin friction drag by boundary-layer combustion, although it represents a sensitive flowfield due to the delicate mechanism upon which it relies.

開催年月日： 2011年4月

記述言語：英語  

開催地：San Francisco, CA   国名：アメリカ合衆国  

Scramjet-based launch systems offer considerable promise for safe, reliable and economical access to space. Through both flight and ground tests, leveraging Australia's world leadership in scramjet R&D, the SCRAMSPACE project is designed to answer key scientific and technological questions and build an industry-ready talent pool for a future Australian space industry.

開催年月日： 2010年9月

記述言語：英語  

開催地：Nice   国名：フランス共和国  

Scramjet propulsion is a promising technology for reliable and economical access to space and high-speed atmospheric transport. The inlet plays a key role in determining the performance of scramjets, in particular for the axisymmetric class of scramjet engines that are currently explored due to their advantages in numerous aspects. In the present study a multi-objective design optimisation (MDO) has been conducted with respect to four major inlet design criteria: compression efficiency, drag, adverse pressure gradient, and exit temperature. The former three criteria are used as the objective functions and the last as the constraint function to evaluate the inlet flowfields in the state-of-the-art coupled CFD / MDO approach. The influential parameters and key physics have been identified by scrutinising the flowfields that have been obtained as an outcome of the optimisation.

開催年月日： 2009年10月

記述言語：英語  

開催地：Bremen   国名：ドイツ連邦共和国  

Reliable in-flight starting of the inlet is of critical importance for the successful operation of scramjet engines, particularly axisymmetric configurations with high contraction inlets. The present research is undertaken to examine the capability of various inlet starting methods based on two principles: unsteady flow effects and variable geometries. Timeaccurate viscous computations have been performed to investigate the transitional flowfields introduced by a variety of methods that are applicable to axisymmetric geometries. Parametric studies have been conducted for instantaneous rupture of conical diaphragms and addition of bleed slots, which induce highly unsteady flow phenomena. Several methods employing variable inlet geometries have been tested for the latter principle, including opening doors, rocket plugs and sliding doors (or diaphragm erosion). Successful inlet starting has been achieved as a result of unsteady transition induced by diaphragm rupture and quasi-steady transition due to the sliding door opening process. In particular, a bleed addition to the diaphragm rupture method has been found to be highly effective and pronounced flow stability has been observed in the sliding door process.

開催年月日： 2008年1月

記述言語：英語  

開催地：Reno, NV   国名：アメリカ合衆国  

The effects of three dimensional bump control on separated shock-wave / boundarylayer interactions have been investigated experimentally at M _∞ = 1.5, aiming to reduce total pressure losses incurred by strong normal shock / boundary-layer interactions. Such a situation occurs, for example, inside mixed-compression supersonic engine inlets or on external-compression inlets. Various bump configurations have been tested and their effects have been evaluated in various aspects such as total pressure recovery and boundary-layer thickness. Without control, the normal shock caused a region of relatively two-dimensional separation. Three-dimensional bumps commonly broke up this separation into individual regions of attached and separated flow, giving an overall reduction of the magnitude of separation. Fairly two-dimensional λ-shock structures and streamwise vortex pairs have been observed in all tested configurations. A configuration comprising multiple long, high, narrow rounded bumps has demonstrated the best performance, producing an appreciable total pressure saving as much as 30% owing to a λ-shock benefit with minimum viscous penalty from localised separation.

開催年月日： 2007年1月

記述言語：英語  

開催地：Reno, NV   国名：アメリカ合衆国  

Three-dimensional bumps have been developed and investigated, aiming at the two major objectives of shock-wave / boundary-layer interaction control, i.e. drag reduction and suppression of separation, simultaneously. An experimental investigation has been conducted for a default rounded bump in channel now at University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at University of Stuttgart. Observed in both cases are wave drag reduction owing to A-shock structures produced by three-dimensional surface bumps and mild control effects on the boundary layer. The effects of rough surface and tall extension have been investigated as well as several geometric variations and multiple bump configurations. A double configuration of narrow rounded bumps has been found to best perform amongst the tested, considerably reducing wave drag through a well-established A-shock structure with little viscous penalty and thus achieving substantial overall drag reduction. Counter-rotating streamwise vortex pairs have been produced by some configurations as a result of local flow separation, but they have been observed to be confined in relatively narrow wake regions, expected to be beneficial in suppressing large-scale separation under off-design condition despite increase of viscous drag. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically, and experimental investigation of bumps fitted on a transonic aerofoil or wing is suggested toward practical application.

開催年月日： 2006年6月

記述言語：英語  

開催地：San Francisco, CA   国名：アメリカ合衆国  

The application of shock control to transonic airfoils and wings has been demonstrated widely to have the potential to reduce wave drag. Most of the suggested control devices are two-dimensional, that is they are of uniform geometry in spanwise direction. Examples of such techniques include contour bumps and passive control. Recently it has been observed that a spanwise array of discrete three-dimensional controls can have similar benefits but also offer advantages in terms of installation complexity and drag. This paper describes research carried out in Cambridge into various three-dimensional devices, such as slots, grooves and bumps. In all cases the control device is applied to the interaction of a normal shock wave (M=1.3) with a turbulent boundary layer. Theoretical considerations are proposed to determine how such fundamental experiments can provide estimates of control performance on a transonic wing. The potential of each class of three-dimensional device for wave drag reduction on airfoils is discussed and surface bumps in particular are identified as offering potential drag savings for typical transonic wing applications under cruise conditions.

開催年月日： 2006年1月

記述言語：英語  

開催地：Reno, NV   国名：アメリカ合衆国  

A novel supersonic wind tunnel setup is proposed to enable the investigation of control on a normal shock wave. Previous experimental arrangements were found to suffer from shock instability. Wind tunnel tests with and without control have confirmed the capability of the new setup to stabilise a shock structure at a target position without changing the nature of the shock wave / boundary layer interaction flow at M_∞ = 1.3 and M _∞ = 1.5. Flow visualisation and pressure measurements with the new setup have revealed detailed characteristics of shock wave / boundary layer interactions and a λ-shock structure as well as benefits of control in total drag reduction in the presence of 3D bump control.

開催年月日： 2005年1月

記述言語：英語  

開催地：Reno, NV   国名：アメリカ合衆国  

An analytical expression is proposed to estimate the wave drag of an aerofoil equipped with shock control. The analysis extends the conventional approach for a single normal shock wave, based on the knowledge that all types of successful shock control on transonic aerofoils cause bifurcated λ-shock structures. The influence of surface curvature on the λ-shock structure has been taken into account. The extended method has been found to produce fairly good agreement with the results obtained by CFD methods while requiring negligible computational effort. This new formulation is expected to be beneficial in the industrial design process of transonic aerofoils and wings where a large number of computational simulations have to be performed.

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