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Aerospace, Volume 3, Issue 4 (December 2016)

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Editorial

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Open AccessEditorial Recent Advances in Aeroacoustics
Aerospace 2016, 3(4), 40; doi:10.3390/aerospace3040040
Received: 8 November 2016 / Revised: 8 November 2016 / Accepted: 11 November 2016 / Published: 23 November 2016
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Abstract
Acoustics is one of the oldest examples of applied research, long before the term was even coined: [...]
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(This article belongs to the Special Issue Recent Advances in Aeroacoustics)

Research

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Open AccessArticle Numerical Study of Transition of an Annular Lift Fan Aircraft
Aerospace 2016, 3(4), 30; doi:10.3390/aerospace3040030
Received: 14 August 2016 / Revised: 12 September 2016 / Accepted: 17 September 2016 / Published: 23 September 2016
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Abstract
The present study aimed at studying the transition of annular lift fan aircraft through computational fluid dynamics (CFD) simulations. The oscillations of lift and drag, the optimization for the figure of merit, and the characteristics of drag, yawing, rolling and pitching moments in
[...] Read more.
The present study aimed at studying the transition of annular lift fan aircraft through computational fluid dynamics (CFD) simulations. The oscillations of lift and drag, the optimization for the figure of merit, and the characteristics of drag, yawing, rolling and pitching moments in transition are studied. The results show that a two-stage upper and lower fan lift system can generate oscillations of lift and drag in transition, while a single-stage inner and outer fan lift system can eliminate the oscillations. The characteristics of momentum drag of the single-stage fans in transition are similar to that of the two-stage fans, but with the peak of drag lowered from 0.63 to 0.4 of the aircraft weight. The strategy to start transition from a negative angle of attack −21° further reduces the peak of drag to 0.29 of the weight. The strategy also reduces the peak of pitching torque, which needs upward extra thrusts of 0.39 of the weight to eliminate. The peak of rolling moment in transition needs differential upward thrusts of 0.04 of the weight to eliminate. The requirements for extra thrusts in transition lead to a total thrust–weight ratio of 0.7, which makes the aircraft more efficient for high speed cruise flight (higher than 0.7 Ma). Full article
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Open AccessArticle Comparison of Power Requirements: Flapping vs. Fixed Wing Vehicles
Aerospace 2016, 3(4), 31; doi:10.3390/aerospace3040031
Received: 23 May 2016 / Revised: 31 August 2016 / Accepted: 1 September 2016 / Published: 28 September 2016
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Abstract
The power required by flapping and fixed wing vehicles in level flight is determined and compared. Based on a new modelling approach, the effects of flapping on the induced drag in flapping wing vehicles are mathematically described. It is shown that flapping causes
[...] Read more.
The power required by flapping and fixed wing vehicles in level flight is determined and compared. Based on a new modelling approach, the effects of flapping on the induced drag in flapping wing vehicles are mathematically described. It is shown that flapping causes a significant increase in the induced drag when compared with a non-flapping, fixed wing vehicle. There are two effects for that induced drag increase; one is due to tilting of the lift vector caused by flapping the wings and the other results from changes in the amount of the lift vector during flapping. The induced drag increase yields a significant contribution to the power required by flapping wing vehicles. Furthermore, the power characteristics of fixed wing vehicles are dealt with. It is shown that, for this vehicle type, the propeller efficiency plays a major role. This is because there are considerable differences in the propeller efficiency when taking the size of vehicles into account. Comparing flapping and fixed wing vehicles, the conditions are shown where flapping wing vehicles have a lower power demand and where fixed wing vehicles are superior regarding the required power. There is a tendency such that fixed wing vehicles have an advantage in the case of larger size vehicles and flapping wing vehicles have an advantage in the case of smaller size ones. Full article
(This article belongs to the Special Issue Flapping Wings)
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Open AccessArticle An Empirical Study of Overlapping Rotor Interference for a Small Unmanned Aircraft Propulsion System
Aerospace 2016, 3(4), 32; doi:10.3390/aerospace3040032
Received: 25 July 2016 / Revised: 23 September 2016 / Accepted: 28 September 2016 / Published: 10 October 2016
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Abstract
The majority of research into full-sized helicopter overlapping propulsion systems involves co-axial setups (fully overlapped). Partially overlapping rotor setups (tandem, multirotor) have received less attention, and empirical data produced over the years is limited. The increase in demand for compact small unmanned aircraft
[...] Read more.
The majority of research into full-sized helicopter overlapping propulsion systems involves co-axial setups (fully overlapped). Partially overlapping rotor setups (tandem, multirotor) have received less attention, and empirical data produced over the years is limited. The increase in demand for compact small unmanned aircraft has exposed the need for empirical investigations of overlapping propulsion systems at a small scale (Reynolds Number < 250,000). Rotor-to-rotor interference at the static state in various overlapping propulsion system configurations was empirically measured using off the shelf T-Motor 16 inch × 5.4 inch rotors. A purpose-built test rig was manufactured allowing various overlapping rotor configurations to be tested. First, single rotor data was gathered, then performance measurements were taken at different thrust and tip speeds on a range of overlap configurations. The studies were conducted in a system torque balance mode. Overlapping rotor performance was compared to an isolated dual rotor propulsion system revealing interference factors which were compared to the momentum theory. Tests revealed that in the co-axial torque-balanced propulsion system the upper rotor outperforms the lower rotor at axial separation ratios between 0.05 and 0.85. Additionally, in the same region, thrust sharing between the two rotors changed by 21%; the upper rotor produced more thrust than the lower rotor at all times. Peak performance was recorded as a 22% efficiency loss when the axial separation ratio was greater than 0.25. The performance of a co-axial torque-balanced system reached a 27% efficiency loss when the axial separation ratio was equal to 0.05. The co-axial system swirl recovery effect was recorded to have a 4% efficiency gain in the axial separation ratio region between 0.05 and 0.85. The smallest efficiency loss (3%) was recorded when the rotor separation ratio was between 0.95 and 1 (axial separation ratio was kept at 0.05). Tests conducted at a rotor separation ratio of 0.85 showed that the efficiency loss decreased when the axial separation ratio was greater than 0.25. The lower rotor outperformed the upper rotor in the rotor separation ratio region from 0.95 to 1 (axial separation ratio was kept at 0.05) at an overall system thrust of 8 N, and matched the upper rotor performance at the tested overall thrust of 15 N. Full article
(This article belongs to the collection Unmanned Aerial Systems)
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Open AccessArticle Linearized Euler Equations for the Determination of Scattering Matrices for Orifice and Perforated Plate Configurations in the High Mach Number Regime
Aerospace 2016, 3(4), 33; doi:10.3390/aerospace3040033
Received: 19 July 2016 / Revised: 18 September 2016 / Accepted: 11 October 2016 / Published: 17 October 2016
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Abstract
The interaction of a plane acoustic wave and a sheared flow is numerically investigated for simple orifice and perforated plate configurations in an isolated, non-resonant environment for Mach numbers up to choked conditions in the holes. Analytical derivations found in the literature are
[...] Read more.
The interaction of a plane acoustic wave and a sheared flow is numerically investigated for simple orifice and perforated plate configurations in an isolated, non-resonant environment for Mach numbers up to choked conditions in the holes. Analytical derivations found in the literature are not valid in this regime due to restrictions to low Mach numbers and incompressible conditions. To allow for a systematic and detailed parameter study, a low-cost hybrid Computational Fluid Dynamic/Computational Aeroacoustic (CFD/CAA) methodology is used. For the CFD simulations, a standard kϵ Reynolds-Averaged Navier–Stokes (RANS) model is employed, while the CAA simulations are based on frequency space transformed linearized Euler equations (LEE), which are discretized in a stabilized Finite Element method. Simulation times in the order of seconds per frequency allow for a detailed parameter study. From the application of the Multi Microphone Method together with the two-source location procedure, acoustic scattering matrices are calculated and compared to experimental findings showing very good agreement. The scattering properties are presented in the form of scattering matrices for a frequency range of 500–1500 Hz. Full article
(This article belongs to the Special Issue Recent Advances in Aeroacoustics)
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Open AccessArticle A Two-Temperature Open-Source CFD Model for Hypersonic Reacting Flows, Part One: Zero-Dimensional Analysis
Aerospace 2016, 3(4), 34; doi:10.3390/aerospace3040034
Received: 2 August 2016 / Revised: 27 September 2016 / Accepted: 29 September 2016 / Published: 18 October 2016
Cited by 2 | PDF Full-text (1282 KB) | HTML Full-text | XML Full-text
Abstract
A two-temperature CFD (computational fluid dynamics) solver is a prerequisite to any spacecraft re-entry numerical study that aims at producing results with a satisfactory level of accuracy within realistic timescales. In this respect, a new two-temperature CFD solver, hy2Foam, has been developed
[...] Read more.
A two-temperature CFD (computational fluid dynamics) solver is a prerequisite to any spacecraft re-entry numerical study that aims at producing results with a satisfactory level of accuracy within realistic timescales. In this respect, a new two-temperature CFD solver, hy2Foam, has been developed within the framework of the open-source CFD platform OpenFOAM for the prediction of hypersonic reacting flows. This solver makes the distinct juncture between the trans-rotational and multiple vibrational-electronic temperatures. hy2Foam has the capability to model vibrational-translational and vibrational-vibrational energy exchanges in an eleven-species air mixture. It makes use of either the Park TTv model or the coupled vibration-dissociation-vibration (CVDV) model to handle chemistry-vibration coupling and it can simulate flows with or without electronic energy. Verification of the code for various zero-dimensional adiabatic heat baths of progressive complexity has been carried out. hy2Foam has been shown to produce results in good agreement with those given by the CFD code LeMANS (The Michigan Aerothermodynamic Navier-Stokes solver) and previously published data. A comparison is also performed with the open-source DSMC (direct simulation Monte Carlo) code dsmcFoam. It has been demonstrated that the use of the CVDV model and rates derived from Quantum-Kinetic theory promote a satisfactory consistency between the CFD and DSMC chemistry modules. Full article
(This article belongs to the Special Issue State-of-the-Art Aerospace Sciences and Technologies in Europe)
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Open AccessArticle Numerical Investigation of Effect of Parameters on Hovering Efficiency of an Annular Lift Fan Aircraft
Aerospace 2016, 3(4), 35; doi:10.3390/aerospace3040035
Received: 14 August 2016 / Revised: 8 October 2016 / Accepted: 10 October 2016 / Published: 19 October 2016
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Abstract
The effects of various parameters on the hovering performance of an annular lift fan aircraft are investigated by using numerical scheme. The pitch angle, thickness, aspect ratio (chord length), number of blades, and radius of duct inlet lip are explored to optimize the
[...] Read more.
The effects of various parameters on the hovering performance of an annular lift fan aircraft are investigated by using numerical scheme. The pitch angle, thickness, aspect ratio (chord length), number of blades, and radius of duct inlet lip are explored to optimize the figure of merit. The annular lift fan is also compared with a conventional circular lift fan of the same features with the same disc loading and similar geometry. The simulation results show that the pitch angle of 27°, the thickness of 4% chord length, the aspect ratio of 3.5~4.0, 32 blades, and the radius of inlet lip of 4.7% generate the maximum figure of merit of 0.733. The optimized configuration can be used for further studies of the annular lift fan aircraft. Full article
(This article belongs to the Special Issue Aircraft Design)
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Open AccessArticle Wing Tip Drag Reduction at Nominal Take-Off Mach Number: An Approach to Local Active Flow Control with a Highly Robust Actuator System
Aerospace 2016, 3(4), 36; doi:10.3390/aerospace3040036
Received: 19 July 2016 / Revised: 8 October 2016 / Accepted: 9 October 2016 / Published: 19 October 2016
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Abstract
This paper discusses wind tunnel test results aimed at advancing active flow control technology to increase the aerodynamic efficiency of an aircraft during take-off. A model of the outer section of a representative civil airliner wing was equipped with two-stage fluidic actuators between
[...] Read more.
This paper discusses wind tunnel test results aimed at advancing active flow control technology to increase the aerodynamic efficiency of an aircraft during take-off. A model of the outer section of a representative civil airliner wing was equipped with two-stage fluidic actuators between the slat edge and wing tip, where mechanical high-lift devices fail to integrate. The experiments were conducted at a nominal take-off Mach number of M = 0.2. At this incidence velocity, separation on the wing section, accompanied by increased drag, is triggered by the strong slat edge vortex at high angles of attack. On the basis of global force measurements and local static pressure data, the effect of pulsed blowing on the complex flow is evaluated, considering various momentum coefficients and spanwise distributions of the actuation effort. It is shown that through local intensification of forcing, a momentum coefficient of less than c μ = 0.6 % suffices to offset the stall by 2.4°, increase the maximum lift by more than 10% and reduce the drag by 37% compared to the uncontrolled flow. Full article
(This article belongs to the Special Issue State-of-the-Art Aerospace Sciences and Technologies in Europe)
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Open AccessArticle Results of Long-Duration Simulation of Distant Retrograde Orbits
Aerospace 2016, 3(4), 37; doi:10.3390/aerospace3040037
Received: 31 July 2016 / Revised: 23 October 2016 / Accepted: 26 October 2016 / Published: 8 November 2016
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Abstract
Distant Retrograde Orbits in the Earth–Moon system are gaining in popularity as stable “parking” orbits for various conceptual missions. To investigate the stability of potential Distant Retrograde Orbits, simulations were executed, with propagation running over a thirty-year period. Initial conditions for the vehicle
[...] Read more.
Distant Retrograde Orbits in the Earth–Moon system are gaining in popularity as stable “parking” orbits for various conceptual missions. To investigate the stability of potential Distant Retrograde Orbits, simulations were executed, with propagation running over a thirty-year period. Initial conditions for the vehicle state were limited such that the position and velocity vectors were in the Earth–Moon orbital plane, with the velocity oriented such that it would produce retrograde motion about Moon. The resulting trajectories were investigated for stability in an environment that included the eccentric motion of Moon, non-spherical gravity of Earth and Moon, gravitational perturbations from Sun, Jupiter, and Venus, and the effects of radiation pressure. The results indicate that stability may be enhanced at certain resonant states within the Earth–Moon system. Full article
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Open AccessFeature PaperArticle Climate-Compatible Air Transport System—Climate Impact Mitigation Potential for Actual and Future Aircraft
Aerospace 2016, 3(4), 38; doi:10.3390/aerospace3040038
Received: 30 August 2016 / Revised: 14 October 2016 / Accepted: 25 October 2016 / Published: 17 November 2016
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Abstract
Aviation guarantees mobility, but its emissions also contribute considerably to climate change. Therefore, climate impact mitigation strategies have to be developed based on comprehensive assessments of the different impacting factors. We quantify the climate impact mitigation potential and related costs resulting from changes
[...] Read more.
Aviation guarantees mobility, but its emissions also contribute considerably to climate change. Therefore, climate impact mitigation strategies have to be developed based on comprehensive assessments of the different impacting factors. We quantify the climate impact mitigation potential and related costs resulting from changes in aircraft operations and design using a multi-disciplinary model workflow. We first analyze the climate impact mitigation potential and cash operating cost changes of altered cruise altitudes and speeds for all flights globally operated by the Airbus A330-200 fleet in the year 2006. We find that this globally can lead to a 42% reduction in temperature response at a 10% cash operating cost increase. Based on this analysis, new design criteria are derived for future aircraft that are optimized for cruise conditions with reduced climate impact. The newly-optimized aircraft is re-assessed with the developed model workflow. We obtain additional climate mitigation potential with small to moderate cash operating cost changes due to the aircraft design changes of, e.g., a 32% and 54% temperature response reduction for a 0% and 10% cash operating cost increase. Hence, replacing the entire A330-200 fleet by this redesigned aircraft ( M a c r = 0.72 and initial cruise altitude (ICA) = 8000 m) could reduce the climate impact by 32% without an increase of cash operating cost. Full article
(This article belongs to the Special Issue Aircraft Design)
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Open AccessCommunication Effect of Leading-Edge Slats at Low Reynolds Numbers
Aerospace 2016, 3(4), 39; doi:10.3390/aerospace3040039
Received: 5 August 2016 / Revised: 10 October 2016 / Accepted: 13 November 2016 / Published: 17 November 2016
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Abstract
One of the most commonly implemented devices for stall control on wings and airfoils is a leading-edge slat. While functioning of slats at high Reynolds number is well documented, this is not the case at the low Reynolds numbers common for small unmanned
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One of the most commonly implemented devices for stall control on wings and airfoils is a leading-edge slat. While functioning of slats at high Reynolds number is well documented, this is not the case at the low Reynolds numbers common for small unmanned aerial vehicles. Consequently, a low-speed wind tunnel investigation was undertaken to elucidate the performance of a slat at Re = 250,000. Force balance measurements accompanied by surface flow visualization images are presented. The slat extension and rotation was varied and documented. The results indicate that for small slat extensions, slat rotation is deleterious to performance, but is required for larger slat extensions for effective lift augmentation. Deployment of the slat was accompanied by a significant drag penalty due to premature localized flow separation. Full article
(This article belongs to the collection Unmanned Aerial Systems)
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Open AccessFeature PaperArticle On the Importance of Morphing Deformation Scheduling for Actuation Force and Energy
Aerospace 2016, 3(4), 41; doi:10.3390/aerospace3040041
Received: 1 October 2016 / Revised: 15 November 2016 / Accepted: 19 November 2016 / Published: 25 November 2016
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Abstract
Morphing aircraft offer superior properties as compared to non-morphing aircraft. They can achieve this by adapting their shape depending on the requirements of various conflicting flight conditions. These shape changes are often associated with large deformations and strains, and hence dedicated morphing concepts
[...] Read more.
Morphing aircraft offer superior properties as compared to non-morphing aircraft. They can achieve this by adapting their shape depending on the requirements of various conflicting flight conditions. These shape changes are often associated with large deformations and strains, and hence dedicated morphing concepts are developed to carry out the required changes in shape. Such intricate mechanisms are often heavy, which reduces, or even completely cancels, the performance increase of the morphing aircraft. Part of this weight penalty is determined by the required actuators and associated batteries, which are mainly driven by the required actuation force and energy. Two underexposed influences on the actuation force and energy are the flight condition at which morphing should take place and the order of the morphing manoeuvres, also called morphing scheduling. This paper aims at highlighting the importance of both influences by using a small Unmanned Aerial Vehicle (UAV) with different morphing mechanisms as an example. The results in this paper are generated using a morphing aircraft analysis and design code that was developed at the Delft University of Technology. The importance of the flight condition and a proper morphing schedule is demonstrated by investigating the required actuation forces for various flight conditions and morphing sequences. More importantly, the results show that there is not necessarily one optimal flight condition or morphing schedule and a tradeoff needs to be made. Full article
(This article belongs to the Special Issue Adaptive/Smart Structures and Multifunctional Materials 2016)
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Open AccessArticle Analysis of Pilot-Induced-Oscillation and Pilot Vehicle System Stability Using UAS Flight Experiments
Aerospace 2016, 3(4), 42; doi:10.3390/aerospace3040042
Received: 15 August 2016 / Revised: 6 November 2016 / Accepted: 14 November 2016 / Published: 29 November 2016
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Abstract
This paper reports the results of a Pilot-Induced Oscillation (PIO) and human pilot control characterization study performed using flight data collected with a Remotely Controlled (R/C) unmanned research aircraft. The study was carried out on the longitudinal axis of the aircraft. Several existing
[...] Read more.
This paper reports the results of a Pilot-Induced Oscillation (PIO) and human pilot control characterization study performed using flight data collected with a Remotely Controlled (R/C) unmanned research aircraft. The study was carried out on the longitudinal axis of the aircraft. Several existing Category 1 and Category 2 PIO criteria developed for manned aircraft are first surveyed and their effectiveness for predicting the PIO susceptibility for the R/C unmanned aircraft is evaluated using several flight experiments. It was found that the Bandwidth/Pitch rate overshoot and open loop onset point (OLOP) criteria prediction results matched flight test observations. However, other criteria failed to provide accurate prediction results. To further characterize the human pilot control behavior during these experiments, a quasi-linear pilot model is used. The parameters of the pilot model estimated using data obtained from flight tests are then used to obtain information about the stability of the Pilot Vehicle System (PVS) for Category 1 PIOs occurred during straight and level flights. The batch estimation technique used to estimate the parameters of the quasi-linear pilot model failed to completely capture the compatibility nature of the human pilot. The estimation results however provided valuable insights into the frequency characteristics of the human pilot commands. Additionally, stability analysis of the Category 2 PIOs for elevator actuator rate limiting is carried out using simulations and the results are compared with actual flight results. Full article
(This article belongs to the collection Unmanned Aerial Systems)
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Open AccessArticle Gaskinetic Modeling on Dilute Gaseous Plume Impingement Flows
Aerospace 2016, 3(4), 43; doi:10.3390/aerospace3040043
Received: 9 November 2016 / Revised: 4 December 2016 / Accepted: 6 December 2016 / Published: 9 December 2016
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Abstract
This paper briefly reviews recent work on gaseous plume impingement flows. As the major part of this paper, also included are new comprehensive studies on high-speed, collisionless, gaseous, circular jet impinging on a three-dimensional, inclined, diffuse or specular flat plate. Gaskinetic theories are
[...] Read more.
This paper briefly reviews recent work on gaseous plume impingement flows. As the major part of this paper, also included are new comprehensive studies on high-speed, collisionless, gaseous, circular jet impinging on a three-dimensional, inclined, diffuse or specular flat plate. Gaskinetic theories are adopted to study the problems, and several crucial geometry-location and velocity-direction relations are used. The final complete results include impingement surface properties such as pressure, shear stress, and heat flux. From these surface properties, averaged coefficients of pressure, friction, heat flux, moment over the entire flat plate, and the distance from the moment center to the flat plate center are obtained. The final results include accurate integrations involving the geometry and specific speed ratios, inclination angle, and the temperature ratio. Several numerical simulations with the direct simulation Monte Carlo method validate these analytical results, and the results are essentially identical. The gaskinetic method and processes are heuristic and can be used to investigate other external high Knudsen (Kn) number impingement flow problems, including the flow field and surface properties for a high Knudsen number jet from an exit and flat plate of arbitrary shapes. The results are expected to find many engineering applications, especially in aerospace and space engineering. Full article
(This article belongs to the Special Issue Fluid-Structure Interactions)
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Open AccessArticle Continuation Methods for Nonlinear Flutter
Aerospace 2016, 3(4), 44; doi:10.3390/aerospace3040044
Received: 27 October 2016 / Revised: 3 December 2016 / Accepted: 5 December 2016 / Published: 9 December 2016
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Abstract
Continuation methods are presented that are capable of treating frequency domain flutter equations, including multiple nonlinearities represented by describing functions. A small problem demonstrates how a series of continuation processes can find all limit-cycle oscillations within a specified region with a reasonable degree
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Continuation methods are presented that are capable of treating frequency domain flutter equations, including multiple nonlinearities represented by describing functions. A small problem demonstrates how a series of continuation processes can find all limit-cycle oscillations within a specified region with a reasonable degree of confidence. Curves of the limit-cycle amplitude variation with velocity, indicating regions of stability and instability with colors, give a compact view of the nonlinear behavior throughout the flight regime. A continuation technique for reducing limit-cycle amplitudes by adjusting various system parameters is presented. These processes are economical enough to be a routine part of aircraft design and certification. Full article
(This article belongs to the collection Feature Papers in Aerospace)
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Open AccessArticle A Two-Temperature Open-Source CFD Model for Hypersonic Reacting Flows, Part Two: Multi-Dimensional Analysis †
Aerospace 2016, 3(4), 45; doi:10.3390/aerospace3040045
Received: 30 September 2016 / Revised: 25 November 2016 / Accepted: 8 December 2016 / Published: 14 December 2016
Cited by 2 | PDF Full-text (2676 KB) | HTML Full-text | XML Full-text | Supplementary Files
Abstract
hy2Foam is a newly-coded open-source two-temperature computational fluid dynamics (CFD) solver that has previously been validated for zero-dimensional test cases. It aims at (1) giving open-source access to a state-of-the-art hypersonic CFD solver to students and researchers; and (2) providing a foundation for
[...] Read more.
hy2Foam is a newly-coded open-source two-temperature computational fluid dynamics (CFD) solver that has previously been validated for zero-dimensional test cases. It aims at (1) giving open-source access to a state-of-the-art hypersonic CFD solver to students and researchers; and (2) providing a foundation for a future hybrid CFD-DSMC (direct simulation Monte Carlo) code within the OpenFOAM framework. This paper focuses on the multi-dimensional verification of hy2Foam and firstly describes the different models implemented. In conjunction with employing the coupled vibration-dissociation-vibration (CVDV) chemistry–vibration model, novel use is made of the quantum-kinetic (QK) rates in a CFD solver. hy2Foam has been shown to produce results in good agreement with previously published data for a Mach 11 nitrogen flow over a blunted cone and with the dsmcFoam code for a Mach 20 cylinder flow for a binary reacting mixture. This latter case scenario provides a useful basis for other codes to compare against. Full article
(This article belongs to the Special Issue State-of-the-Art Aerospace Sciences and Technologies in Europe)
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Other

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Open AccessCorrection Correction: Iemma, U. Theoretical and Numerical Modeling of Acoustic Metamaterials for Aeroacoustic Applications. Aerospace 2016, 3, 15
Aerospace 2016, 3(4), 46; doi:10.3390/aerospace3040046
Received: 14 December 2016 / Revised: 14 December 2016 / Accepted: 20 December 2016 / Published: 22 December 2016
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Abstract
The author regrets that this paper [1] contains a typographical error in Equation (1) [...]
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