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.
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.
Delayed Detached Eddy Simulations (DDES) of separating-reattaching flows on the suction side of an ice-contaminated airfoil were conducted. A single-section straight-wing NACA23012 airfoil with leading-edge ice was studied. The geometry represents a realistic glaze horn-ice contamination obtained during the icing test campaigns described in , which has aerodynamic data for comparison. The three-dimensional transient flow behavior was simulated using the open-source flow solver OVERFLOW, version 2.2l  developed by NASA Langley Research Center. Configurations at three angles of attack that exhibit unsteady flow behavior starting with the bursting angle were examined at Mach number of 0.18 and Reynolds number of 1.8x106. As the stall angle was approached the aerodynamic performance parameters displayed large-scale unsteadiness where periods of attached and separated flows were observed. The time-averaged results show good agreement with the aerodynamic test data.
This paper presents an aerodynamic degradation study of an iced airfoil, using the Lattice Boltzmann approach with the commercial software PowerFLOW. Three-dimensional numerical simulations were performed with an extruded constant section of the GLC-305 airfoil with a leading-edge double-horn ice shape using periodic boundary conditions. The freestream Reynolds number, based on the chord, is 3.5 million and the Mach number is 0.12. An extensive comparison of the main flow features with experimental data is performed, including aerodynamic coefficients, pressure coefficient distributions, velocity and turbulence contours along with its profiles at several positions, and stagnation streamlines. The drag coefficient agrees well with experiments, in spite of a small shift. Two different wind tunnel measurements, using different measurement techniques, were compared to the CFD results, which mostly stayed in between the experimental data.
The objective of this paper is to present a numerical method to rank thick ice shapes for aircraft by comparing the ice accretion effects for different icing scenarios in order to determine the more critical ice shape. This ranking allows limiting the demonstration of the aerodynamic characteristics of the aircraft in iced condition during certification to a reduced number of ice shapes. The usage of this numerical method gives more flexibility to the determination of the critical ice shapes, as it is not dependent of the availability of physical test vehicles and/or facilities. The simulation strategy is built on the Lattice Boltzmann Method (LBM) and is validated based on a representative test case, both in terms of aircraft geometry and ice shapes. Validation against existing experimental results shows the method exhibits an adequate level of reliability for the ranking of thick ice shapes.
Many uncertainties in an in-flight ice shape prediction are related to convection heat transfer coefficient, which in turn depends on the flow, turbulence and laminar/turbulent transition models. The height of ice roughness element used to calculate the Equivalent Sand Grain Roughness height (ESGR) is a very important input of the turbulence model as it strongly influences the shape of the accreted ice. Unfortunately, for in-flight icing, the ESGR is unknown and generally calculated using semi-empirical models or empirical correlations based on a particular ice shape prediction code. Each ice shape prediction code is unique due to the models and correlations used and the numerical implementation. Ice roughness correlations do not have the same effect in each ice shape prediction code. A new approach to calculate the ESGR correlation taking into consideration the particularities of the ice shape prediction code is developed, calibrated and validated.
Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University’s Walter H. Beech Memorial Wind Tunnel utilizing an 8.9% scale model, and high Reynolds number tests were conducted at ONERA’s F1 wind tunnel utilizing a 13.3% scale model. Several identical geometrically-scaled ice shapes were tested at both facilities, and the results were compared at overlapping Reynolds and Mach numbers. This was to ensure that the results and trends observed at low Reynolds number could be applied and continued to high, near-flight Reynolds number. The data from Wichita State University and ONERA F1 agreed well at matched Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations.
Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing was carried out for 8.9% and 13.3% scale semispan wing models based upon the Common Research Model airplane configuration. Various levels of geometric fidelity of an artificial ice shape representing a realistic glaze-ice accretion on a swept wing were investigated. The highest fidelity artificial ice shape reproduced all of the three-dimensional features associated with the glaze ice accretion. The lowest fidelity artificial ice shapes were simple, spanwise-varying horn ice geometries intended to represent the maximum ice thickness on the wing upper surface.