With the actual tendency of space exploration, hypersonic flight have gain a significant relevance, taking the attention of many researchers over the world. This work aims to present a numerical tool to solve hypersonic gas dynamic flows for space propulsion geometries. This will be done by validating the code using two well-known hypersonic test cases, the double cone and the hollow cylinder flare. These test cases are part of NATO Research and Technology Organization Working Group 10 validation of hypersonic flight for laminar viscous-inviscid interactions. During the validation process several important flow features of hypersonic flow are captured and compared with available CFD and numerical data. Special attention is taken to the phenomenon of vibrational excitation of the molecules. Different vibrational non-equilibrium models are used and compared with the available data. The pressure and the heat flux along the surfaces are also analyzed.
Flow separation is among the major causes of aerodynamic drag experience by wings. Vortex generators are regularly used as a means of flow separation control in wings, their use leading to delayed flow separation and drag reduction. A disadvantage of external vortex generators has been observed to be high momentum loss and inefficiency in vortex generation. Internal vortex generators minimize the penalty of momentum loss and generate vortices closer to the surface. In this paper, the impact of the length of internal vortex generators on the aerodynamic characteristics of a wing have been investigated. Internal vortex generators have been placed at 30% chord distance on the suction side of a NACA 0012 airfoil. Analysis is carried out using the Computational Fluid Dynamics software ANSYS Fluent. The length of the vortex has been varied between H and 5H, H being the thickness of the boundary layer, at air flow Reynolds Number between 1,000,000 and 5,000,000.
Vortex generators are aerodynamic devices generally used to delay local air separation and stalling. Conventional vortex generators are external and located normal to the surface with a yaw angle against the flow. However, external vortex generators lead to high momentum loss in the boundary layer, producing inefficient vortices which separate from the surface. They hence do not reenergise the boundary layer to a large extent, in order to allow for delayed flow separation. In order to reduce this loss, internal vortex generators may be used. The effect of internal vortex generators has been investigated on a NACA 0012 airfoil using the Computational Fluid Dynamics software ANSYS Fluent. As the effect of a vortex on the boundary layer is inherently three-dimensional, the numerical analysis of an internal vortex generator is limited to a three-dimensional simulation of the flow.
In development of more electric aircraft applications, it is important to discuss aircraft energy management on various level of aircraft operation. This paper presents a computationally efficient optimization model for evaluating flight efficiency on global and interval flight ranges. The model is described as an optimal control problem with an objective functional subjected to state condition and control input constraints along a flight path range. A flight model consists of aircraft point-mass equations of motion including engine and aerodynamic models. The engine model generates the engine thrust and fuel consumption rate for operation condition and the aerodynamic model generates the drag force and lift force of an aircraft for flight conditions. These models is identified by data taken from a published literature as an example. First, approximate optimization process is performed for climb, cruise, decent and approach as each interval range path.
This paper presents the design and construction of a high force density tubular permanent-magnet (PM) linear motor. A strut structure of a tubular PM linear motor developed to improve resistance to impurities and structural rigidity is described. In the design, computationally efficient two-dimensional finite-element analysis is used to estimate the motor force density. The motor’s design is optimized for the major pole number/slot number combinations of 8/24, 16/24, 20/24, 28/24, 32/24, and 40/24. The optimized motor design of a three-phase 16/24 combination with one-layer winding achieved the highest force-to-mass density. The force-to-mass density of the designed motor is higher than that of the first prototype motor by a factor of 5. The validity of the proposed design method and the expected drive characteristics are experimentally verified using the prototype.
In partially premixed combustion engines high octane number fuels are injected into the cylinder during the late part of the compression cycle, giving the fuel and oxidizer enough time to mix into a desirable stratified mixture. If ignited by auto-ignition such a gas composition can react in an ignition wave-front dominated combustion mode. 3D-CFD modeling of such a combustion mode is challenging as the reaction speed is dependent on both mixing history and turbulence acting on the reaction wave. This paper presents a large eddy simulation (LES) study of the effects of energetic turbulence scale on the fuel/air mixing and on the propagation of reaction wave. The results are compared with optical experiments to validate both pressure trace and ignition location. The studied case is a closed cycle simulation of a single cylinder of a Scania D13 engine running PRF81 (81% iso-octane and 19% n-heptane).
Porous medium approach is widely used in modelling high resistance devices such as heat exchangers, automotive catalysts or filters, where details of flow distribution inside the channels are not important. This reduces the computational time considerably, as the whole length of the monolith does not need to be modelled, and the thin boundary layers in each channel do not need to be resolved. The drawback of the approach is compromised accuracy of the flow predictions downstream of the monolith, because the mixing of the individual jets coming out of the monolith channels is not accounted for. Very few studies exist where this issue has been addressed. The methods include artificial turbulence generation, inferring turbulence information from upstream, or using hybrid modelling approach to separate the flow into channels.
Pre-chamber spark ignition can stabilize and improve thermal efficiency of lean burn natural gas engines. During the compression stroke, homogeneous lean mixture is introduced into the pre-chamber that separates electrodes of spark plug from turbulent flow field. After the mixture in the pre-chamber is ignited, the burnt gas jet is discharged through multi-hole nozzles which promotes combustion of the lean mixture in the main chamber due to turbulence caused by high speed jet and multi-points ignition. However, details mechanism in the process has not been elucidated. To design pre-chamber geometries and to achieve stable combustion under lean condition for such engines, it is important to understand the fundamental aspects of the combustion process. In experiments, a rapid compression and expansion machine is used to visualize OH* self-luminosity by using a high-speed video camera with a 306 nm band pass filter and an image intensifier.
The LES hybridization of standard two-equation turbulence closures is often achieved leaving formally unchanged the turbulent viscosity expression in the URANS and LES modes of operation. Although generally convenient in terms of ease of implementation, this choice leaves some theoretical consistency questions unanswered, the most obvious being the actual meaning of the two transported turbulent scalars and their exact role in the modeled viscosity build-up. A possible remedy to this is represented by the simultaneous modification of one or both the turbulent transport equations and of the turbulent viscosity formula, for which a standard LES behavior is enforced whenever needed. The present work compares a conventional DES-based hybrid model with a consistency-enforcing modified variant for turbulent fuel spray simulation. In our case, LES-mode consistency is accomplished by excluding the second turbulent scalar quantity from the viscosity calculation.
Flexible, reliable and consistent combustion models are necessary for the improvement of the next generation spark-ignition engines. Different approaches have been proposed and widely applied in the past. However, the complexity of the process involving ignition, laminar flame propagation and transition to turbulent combustion need further investigations. Purpose of this paper is to compare two different approaches describing turbulent flame propagation. The first approach is the one-equation flame wrinkling model by Weller, while the second is the Coherent Flamelet Model (CFM). Ignition is described by a simplified deposition model while the correlation from Herweg and Maly is used for the transition from the laminar to turbulent flame propagation. Validation of the proposed models was performed with experimental data of a natural-gas, heavy duty engine running at different operating conditions.
Lockheed Martin successfully flight tested the AGM-183A Air-Launched Rapid Response Weapon (ARRW) on a U.S. Air Force Boeing B-52 Stratofortress. The captive carry flight – announced during the 2019 International Paris Air Show – marks Lockheed Martin’s most recent demonstration of hypersonic technology development.
The performance of an electrically heated aircraft ice protection system for a composite leading edge was evaluated. The composite leading edge of the model is equipped with a Ni alloy resistance heater. A state-of-the-art icing code, FENSAP-ICE, was used for the analysis of the electrothermal de-icing system. Computational results, including detailed information of conjugate heat transfer, were validated with experimental data. The computational model was then applied to the composite leading edge wing section at various metrological conditions selected from FAR Part 25 Appendix C.
A 3D CFD methodology is presented to simulate ice build-up on propeller blades exposed to known icing conditions in flight, with automatic blade pitch variation at constant RPM to maintain the desired thrust. One blade of a six-blade propeller and a 70-passenger twin-engine turboprop are analyzed as stand-alone components in a multi-shot quasi-steady icing simulation. The thrust that must be generated by the propellers is obtained from the drag computed on the aircraft. The flight conditions are typical for a 70-passenger twin-engine turboprop in a holding pattern in Appendix C icing conditions: 190 kts at an altitude of 6,000 ft. The rotation rate remains constant at 850 rpm, a typical operating condition for this flight envelope.
Many engineering systems operating in a cold environment are challenged by ice accretion, which unfavorably affects their aerodynamics and degrades both their performance and safety. Precise characterization of ice adhesion is crucial for an effective design of ice protection system. In this paper, a fracture mechanics-based approach incorporating single cantilever beam test is used to characterize the near mode-I interfacial adhesion of a typical ice/aluminum interface with different surface roughness. In this asymmetric beam test, a thin layer of ice is formed between a fixed and elastically deformable beam subjected to the applied loading. The measurements showed a range of the interfacial adhesion energy (GIC) between 0.11 and 1.34 J/m 2, depending on the substrate surface roughness. The detailed inspection of the interfacial ice fracture surface, using fracture surface replication technique, revealed a fracture mode transition with the measured macroscopic fracture toughness.