18 research outputs found
Aerodynamics Flapping-Flight Robotic Bird using Unsteady Lifting-Line Method
The Robird is a bird-like drone, or ornithopter, that generates lift and thrust by flapping and pitching its wings, which performance resembles that of a Peregrine falcon. This paper describes an extension, from steady flow to unsteady flow, of Prandtl’s Lifting-Line method to predict the unsteady lift, thrust, pitching-moment, root-bending moment, required-power and propulsive-efficiency of the robotic bird. The extension comprises the derivation of the Kutta-Joukowski Theorem for unsteady flow, an unsteady trailing-edge Kutta condition and the representation of the wake as a stationary transpiration-type of planar surface carrying a time-dependent dipole distribution. Its instantaneous strength is obtained from the spanwise distribution of the circulation of the lifting line at earlier times. For the cases considered, the numerical method predicts that the section-lift, section-thrust, section-pitching-moment and section-required-power of the wing vary in time. During flapping flight, the cycle-averaged section-lift and section-thrust, as well as the cycle-averaged overall lift and thrust, are mostly positive. The spanwise distributions of cycle-averaged sectional aerodynamic quantities like circulation, lift, etc., as well as the corresponding span-integrated overall quantities and the propulsive efficiency, depend on flight parameters Strouhal number, but not all on pitch amplitude, cycle-averaged effective angle-of-attack nor phase difference between pitching and flapping. The topology of the wake in terms of the unsteady wake dipole distribution, as well as its corresponding vortex distribution, predicted by the unsteady-lifting-line method depend on all flight parameters. The paper provides the relation between wake topology and the generation of lift and thrust
Pulse actuation and its effects on separated lagrangian coherent structures for flow over a cambered airfoil
Numerical Simulation of Flow Control by Synthetic Jet Actuation
Numerical simulations of active flow control have been carried out for the flow around the NACA0018 profile for Mach = 0.15, Re = 2× 106, a = 15o using the Unsteady Reynolds Averaged Navier-Stokes (URANS) equations. Two types of flow control, zero-net-mass jets (synthetic jets) and continuously blowing jets, have been considered to delay the onset of separation. The synthetic jets have been applied to the 2D situation, i.e. infinitely long slits in the spanwise, for which the angle between the jet and surface normal has been varied to study the effect on the separation. For the continuously blowing jets the effect of 3D mixing is taken into account and an optimization of several jet parameters has been carried out to obtain the best result possible
Dual Injection in Supersonic Crossflow: Analysis Jet Shear Layer from Schlieren Images
In supersonic-combustion ramjets (scramjets), fuel is injected, which should mix rapidly with the supersonic crossflow to minimize the length of the scramjet. Tandem dual-jet injection has shown improved mixing performance over single-jet injection. However, experiments on tandem dual-jet injection have not addressed the jet shear layer, in which the mixing occurs, yet. The present study investigates the jet shear layer, as well as the bow shocks in front of the jets, in a continuous air-indraft supersonic wind tunnel at Mach number 1.55. A schlieren setup has been used for visualizing the flow features. A largely automated algorithm for processing schlieren images has been developed to determine the location of the upper boundary of the jet shear layer. The penetration of the jet is studied as a function of 1) J, the ratio of the momentum of the jet and that of the crossflow, and 2) the dimensionless distance S between the dual-jets. An empirical similarity relation has been established for the time-averaged location of the jet upper shear layer as function of J and S, covering the investigated conditions (J ∈ 2.8;3.8;4.8, S ∈ 0∶9.87). This empirical similarity relation provides Sopt, the spacing for maximal penetration of the jets as function of J
Aircraft icing in flight: Effects of impact of super cooled large droplets
In this study a computational method is presented which simulates the presence of a liquid layer on an airfoil and its effect on splashing of Supercooled Large Droplets (SLD). The thin liquid film is expected to have a significant influence on the impact behaviour of SLD. It will arise when the impacting droplets freeze only partially and leave behind a layer of runback water on top of the ice layer. The liquid film is modelled using the wall shear stress and by assuming a linear velocity profile within the water layer. The shear stress is calculated by coupling an integral boundary-layer method to a potential flow method. The SLD splashing model is extended with a deposition model that accounts for impact on a liquid film and includes the solidification time of the droplets. This solidification time is obtained using multiple approaches which are based on either planar solidification or dendritic solidification. Planar solidification is controlled by diffusion and based on the Stefan problem for heat conduction. Dendritic solidification is more rapid and mostly governed by kinetics. The comparison of the catching efficiency with experimental results for a NACA-23012 airfoil shows a significant improvement employing the new deposition model. Also, good agreement is found with the experimental results for the ice accretion on a NACA-0012 airfoil
Splashing model for impact of supercooled large droplets on a thin liquid film
Compared to conventional icing additional droplet phenomena have to be accounted for in icing caused by supercooled large droplets (SLD) such as splashing, rebound, breakup and deformation. In this study the effect of the presence of a thin liquid film of water on the surface has been investigated. This liquid layer can arise when SLD droplets freeze only partially following impact on the airfoil. The effect of the liquid film is simulated by using the wall shear stress and by assuming a linear velocity profile in the liquid layer. The shear stress is calculated by coupling an integral boundary-layer method to a potential ow method. An improved splashing model has been implemented in the existing com-putational method. This splashing model consists of a deposition model that accounts for splashing during impact of droplets on a liquid layer. In an extension to this model different solidification models have been investigated to estimate the time of solidification of a liquid splat produced on the surface after impact. One is a planar solidification model which is described by the Stefan problem for heat conduction and which is mostly controlled by diffusion. The second model is based on dendritic solidification, which is rapid and gov-erned by kinetics. The results of the deposition model on SLD ice accretion are compared with data from experiments on a NACA-23012 airfoil and on a NACA-0012 airfoil. Good agreement is found
Power VCSEL driven Schlieren visualization for cascaded injection in supersonic flow
Extended abstract Results are presented of a study on utilising Vertical-Cavity-Surface-Emitting-Laser (VCSEL) driven Schlieren visualization of cascaded injection in a supersonic flow. The background of the study is fuel injection within a supersonic combustion ramjet (scramjet). The scramjet is a ramjet airbreathing jet engine in which combustion takes place in a supersonic air flow. Scramjets promise significant economic advantages over rocket-based flight travel. However, at hypersonic flight speeds the compressibility effects delay shear layer mixing. In order to maintain for scramjets, the fuel-air mixture required for high combustion efficiency, the combustor becomes relatively long 1. In the present case of cascaded injection, the downstream injector benefits from the shielding effect induced by the smaller upstream injector. This provides a reduction of the momentum in the flow, allowing better penetration of fuel in the air stream over a shorter length. Validation of theoretical and computational results for the flow in scramjets requires a high spatial and temporal resolution of the flow field. In the present study Schlieren visualisation is employed to investigate the flow field. In our previous studies pulsed LED-driven Schlieren visualization was employed. However, for a Mach 1.6 free stream, the limited pulse width of 130 ns of LEDs creates a motion blur of roughly a pixel per second. Therefore, LED-based Schlieren visualisation is not adequate for Schlieren imaging of flows at higher Mach numbers. Furthermore, using an appropriate knife-edge filter, the power of 6 W/mm 2 of the LED employed, is only just sufficient to obtain acceptable Schlieren images. Fig. 1 Left: Schematic of Schlieren setup. Right: Example of Schlieren image obtained for dual-jet injection in a Mach = 1.6 cross-flow. Diameter orifices 1 mm (upstream) and 2 mm (downstream). Visualised are the tandem jets, 20 mm apart, momentum ratio J = 1.37, each featuring a Mach barrel at their exit; the two bow shocks induced by the jets; the boundary layer along the walls and their interaction with the shocks. Also visible are Mach waves originating from small slope discontinuities of the walls of the wind tunnel Power VCSELs provide the high-pulse modulation speeds necessary for high temporal resolution Schlieren imaging of high-speed flow fields. VCSELs consist of very small, densely packed, laser diodes, ordered on a chip in a 2D array, emitting light perpendicular to the chip's surface. Each VCSEL is around 25 micrometres in size and although its Continuous Wave (CW) power is limited, in our case to 10 mW, pulsing the laser increases the power up to a factor of 10. Furthermore, by integrating 600 lasers per square millimetre, th