Blowing is often used to alleviate the intense heating rates on blunt noses of hypersonic vehicles. This flow efflux at the leading edge transforms the flow field in the blunt-nose regions with implications on the dynamic stability of the vehicles. As a demonstrative exercise, the flow fields past blunt-nosed and truncated-nosed conical bodies under blowing and no-blowing conditions were perturbed to obtain the unsteady effects using the shock expansion method to recover the unsteady pressure coefficient. Static and pitching moment derivatives were then duly obtained by integrating the differential of the unsteady pressure coefficient with respect to the pitch angle (α) or the pitch rate ( $\dot \theta $ ) together with the moment arm with reference to the centre of gravity. The results obtained for blunt-nose and truncated conical bodies show a noticeable drop in dynamic stability. Even when the flow is transformed from a tangential blowing at the shoulder of the blunt-nosed vehicle shows some degradation in dynamic stability.

A computational-based investigation has been carried out to examine the heat transfer effects of normal and tangential blowing on blunt surfaces exposed to high Mach number flows. Experimental results from such studies where a controlled outflow is imposed upon a surface under hypersonic conditions are few and far between. Simple 3D axisymmetric configurations supporting flat front surfaces suited to accommodate uniform frontal flow or round-nosed conical bodies with a stepped shoulder to provide uniform parallel flow to the conical surface were selected from a NASA experiment for validation purposes. Uniform outflow of 0.55 to 1.0kg/s was applied in the presence of the prevailing free stream of M=6.8 to assess the extent of heat transfer rate alleviation. The numerical simulation confirmed the experiment that the application of outflow in both cases leads to the surface heat transfer relief.

Abrupt changes in wind velocities over small distances in a lateral or vertical direction can produce wind shear which is known to have serious effects upon the performance of an aircraft. Brought about by large-scale changes in the atmospheric conditions, it is a three-dimensional flow phenomenon imposing severe velocity gradients on an aircraft from all possible directions. While it would be difficult to model an instantaneous velocity gradient in a lateral plane, a vortical flow impinging from the sides which represents a wind shear in a vertical direction is imposed on a forward-moving aircraft to investigate the effect on the aerodynamic performance. The maximum shear wind speed from the side was fixed at 0.3 times the forward velocity. After due validations under no-wind shear conditions on simpler half-reflection plane models, a BGK airfoil-based full 3D wing and the ONERA M6 3D wing model were selected for preliminary studies. The investigation was concluded using the ARA M100 wing-fuselage model.

The Reynolds-averaged Navier–Stokes equations with one-equation turbulence models are used to simulate the flow field past a cone–flare geometry in the Mach number range from 5 to 8 with emphasis on the interaction region of the flare shock with the upstream boundary layer. A model based on the physics of shock unsteadiness is used to correct the standard Spalart–Allmaras turbulence model to improve the prediction of the extent of the separation bubble arising from the shock/turbulent boundary-layer interaction and its accompanying peak pressure and aerothermal loads on the surface. The computed results are validated against the experimental data. The limitations of the shock-unsteadiness model and the extent of the improvement in predicting the heat flux are discussed.

ABSTRACTA numerical simulation has been carried out to investigate the effects of leading edge blowing upon heat alleviation on the surface of hypersonic vehicles. The initial phase of this work deals with the ability of the present CFD-based techniques to solve hypersonic flow field past blunt-nosed vehicles at hypersonic speeds. Towards this end, the authors selected three re-entry vehicles with published flow field data against which the present computed results could be measured. With increasing confidence on the numerical simulation techniques to accurately resolve the hypersonic flow, the boundary condition at the solid blunt surface was then equipped with the ability to blow the flow out of the solid boundary at a rate of at least 0.01–0.1 times the free stream (ρ∞u∞) mass flow rate. The numerical iterative procedure was then progressed until the flow at the surface matched this new ‘inviscid like’ boundary condition. The actual matching of the flow field at the ejection control surface was achieved by iterating the flow on the adjacent cells until the flow conformed to the conditions prescribed at the control surface. The conditions at the surface could be submitted as a ρ∞u∞at the surface or could be equipped as a simple static pressure condition providing the desired flow rate. The comparison between the present engineering approach and the experimental data presented in this study demonstrate its ability to solve complex problems in hypersonic.

Purpose The purpose of this paper is to use a computational technique to simulate the flow in a two-dimensional (2D) wind tunnel where the effect of the solid walls facing the model has been addressed using a porous geometry so that interference arriving at the solid walls are duly damped and a flow suction procedure has been adopted at the side wall to minimize the span-wise effect of the growing side wall boundary layer. Design/methodology/approach A CFD procedure based on discretization of the Navier–Stokes equations has been used to model the flow in a rectangular volume with appropriate treatment for solid walls of the confined volume in which the model is placed. The rectangular volume was configured by stacking O-Grid sections in a span-wise direction using geometric growth from the wall. A porous wall condition has been adapted to counter the wall interference signatures and a separate suction procedure has been implemented for reducing the side wall boundary layer effects. Findings It has been shown that through such corrective measures, the flow in a wind tunnel can be adequately simulated using computational modeling. Computed results were compared against experimental measurements obtained from IAR (Institute for Aerospace, Canada) and NAL (National Aeronautical Laboratory, Japan) to show that indeed appropriate corrective means may be adapted to reduce the interference effects. Research limitations/implications The solutions seemed to converge a lot better using relatively coarser grids which placed the shock locations closer to the experimental values. The finer grids were more stiff to converge and resulted in reversed flow with the two equation k-w model in the region where the intention was to draw out the fluid to thin down the boundary layer. The one equation Spalart–Allmaras model gave better result when porosity and wall suction routines were implemented. Practical implications This method could be used by industry to point check the results against certain demanding flow conditions and then used for more routine parametric studies at other conditions. The method would prove to be efficient and economical during early design stages of a configuration. Originality/value The method makes use of an O-grid to represent the confined test section and its dual treatment of wall interference and blockage effects through simultaneous application of porosity and boundary layer suction is believed to be quite original.

Purpose The purpose of this paper is to obtain close form expressions for the dynamic stability of conical wave riders with flat surfaces which could be equipped with lifting surfaces on its plain flat surface. Numerical simulation would require very large meshes to resolve flows at subscale level and the experimental evaluations would be equally difficult, requiring expensive measurement facilities with challenging procedures to secure such vehicles in confined test sections to obtain satisfactory wind on and wind off oscillations. Design/methodology/approach The design method uses appropriate pressure fields using small disturbance theory, which, in turn, is perturbed using the unsteady shock expansion theory to recover suitable expressions for the dynamic stability behaviour. Findings It was observed that the dynamic stability of the standard half-cone-type wave riders with flat upper surfaces deteriorates with the axis position measured from the pointed apex reaching a minimum at around x/co = 0.666. The half-cone wave rider with flat upper surfaces is dynamically less stable than a pure cone. Research limitations/implications The method is typically less accurate when the similarity parameter M∞θ ≤ 1 = 1 or if the angle of attack is not small. Practical implications With renewed interest in hypersonic, future hypersonically would be designed as fast lifting bodies whose shapes would be close to the configurations of hypersonic wave riders, especially if they are designed to operate at upper atmosphere altitudes. Originality/value The analytic approach outlined in this paper for the evaluation of dynamic and static stability derivatives is original, drawing from the strengths of the small disturbance theory and shock expansion techniques. The method is particularly important, as there are no reported theoretical, numerical or experimental results in the literature.

Purpose - The purpose of this paper is to identifying ways to reduce the effects of wing-vortex interaction by applying surface porosity on selected areas of the exposed surface. A number of papers recently have investigated the aerodynamic implication of free-stream vortices impinging upon airfoils. Design/methodology/approach - The free-stream disturbance in these studies were represented by planting a vortex ahead of the wing or using some other disturbance invoking mechanism like von-Karman vortices in the wake of a cylinder or using a flipping plate to invoke a discrete vortex. In the present work, a well-defined method was used to germinate a system of controlled vortices of known strength, size and frequency ahead of the wing, and the impact of the subsequent interaction was studied with and without the presence of the surface porosity. The simulations tackled a number of cases when porosities of up to 20 and 22 per cent were applied to selected regions near the leading edge, with vortices of controlled strengths directed at the wing surface. Findings - The results showed that the effects of large vortices spanning the entire lengths of the wing can indeed be damped when porosity is selectively applied at strategic regions. Practical implications - Surface porosity application at strategic regions of a wing may dampen the effects of the unsteadiness of the incoming flow. This has profound implications on flight safety and structural damage prevention. Further implications could possibly be extended to UAV and wind turbines that operate at heavy gusting environment. Originality/value - Implementation of this particular method resolves some of the issues arisen when an airplane encounters atmospheric turbulence.

Abstract This paper presents the optimum design of a PID controller for the Adaptive Torsion Wing (ATW) using the genetic algorithm (GA) optimiser. The ATW is a thin-wall, two-spar wingbox whose torsional stiffness can be adjusted by translating the spar webs in the chordwise direction inward and towards each. The reduction in torsional stiffness allows external aerodynamic loads to deform the wing and maintain its shape. The ATW is integrated within the wing of a representative UAV to replace conventional ailerons and provide roll control. The ATW is modelled as a two-dimensional equivalent aerofoil using bending and torsion shape functions to express the equations of motion in terms of the twist angle and plunge displacement at the wingtip. The full equations of motion for the ATW equivalent aerofoil were derived using Lagrangian mechanics. The aerodynamic lift and moment acting on the aerofoil were modelled using Theodorsen’s unsteady aerodynamic theory. The equations of motion are then linearised around an equilibrium position and the GA is employed to design a PID controller for the linearised system to minimise the actuation power require. Finally, the sizing and selection of a suitable actuator is performed.

Dynamic stability data on axially symmetric pointed and blunt cones, parabolic profiles and other ogive and blunt cylindrical shapes is readily available in literature; the dynamic stability on elliptic blunt paraboloids has not been studied at any great lengths in the past. Both numerical and experimental results are scarce. The present paper uses the shock expansion method to obtain the unsteady pressure distribution on blunt elliptic conical bodies at small angles-of-attack. The resulting unsteady pressure distribution is suitably integrated Over the surface of the elliptic body to obtain appropriate analytic expressions for static and dynamic stability. Owing to scarcity of meaningful numerical or measured data for elliptic bodies, the results are compared in qualitative terms against published dynamic stability data on pointed elliptical cones or other axisymmetric blunt cones.

King Abdul Aziz University of Saudi Arabia (KAU) and Tokai University in Japan have collaborated to design and manufacture a solar powered unmanned air vehicle (UAV), Sun Falcon, which has capability of continuous one day flight with intended design extension towards night flight. The project is a student-driven endeavour involving some 30 students. Both universities are equally involved in the actual design studies of the performance characteristics, aerodynamic design, propulsion and structural analysis. Tokai University is in charge of the actual on-site supervision and examination of on-going manufacturing processes and ultimate fabrication of the prototype model.The conceptual design of the Sun Falcon was meticulously worked out in consideration of the operational mission, which included such flight characteristics as the cruising velocity, flight altitude, payload, flight time, rate of climb, power requirements and so on. The weather condition patterns in Saudi Arabia, which remained fairly supportive of the solar cell performance, were also deemed crucial in the design process. However, the design of a solar plane had other challenges in terms of power unit accommodation and payload consideration in comparison to other conventional UAVs. In this paper, an outline of the design features of the Sun Falcon is presented and other notable design features particular to solar UAVs are discussed. It was learnt in this exercise that the selection of the base aerofoil is perhaps one of the most important design items, as the Reynolds number for such UAVs understandably drops notably lower than conventional aeroplanes and such features as the camber curvature and wing area must cater for the installation of solar panels whose size, strength and quantity must respect local weather conditions. For the Sun Falcon, the actual design process examined two candidate aerofoils FX74-CL5-140 (FX74) and SD7037-092-88(SD7037) both of which were abundantly suitable for furnishing the required aerodynamic characteristics. SD7037 was ultimately chosen as it provided the best geometry and camber line in terms of accommodation and placement of the solar panels. Further scrutiny demonstrated that this latter aerofoil provided better take-off performance and superior L/D behaviour under cruise conditions. In order to check out the aerodynamic performance in general and overall stability and control characteristics, a preflight test under battery power (a 2,500 mAh li-po 4-cell 14.8v) was achieved on 4 June 2013. Other solar powered based tests are currently under way at present.

This paper discusses the concept of simple blowing directly from the trailing edge of the wing of an aircraft to provide increased lift, especially during the take-off and landing maneuvers. Although a noticeable increment in coefficient of lift has been obtained to replace or even augment the effect of classic flaps, its application as a means to exert immediate and instantaneous control on stability and control would be an added benefit. Indeed, if such a concept could truly be optimized to replace the physical presence of traditional flaps in an aircraft, the benefits in weight savings and comparative simplicity in design and operational adoptability would be enormous. The jet blowing has been applied at the trailing edge of an airfoil and a 3-D wing at a variety of jet-ejection vector angles with respect to the chord line; a noticeable increase in lift has been recorded using both configurations.

Structural response models addressing one-degree-of-freedom (1DOF) and twodegree-of-freedom (2DOF) aeroelastic oscillations were coupled with an in-house incompressible CFD code to perform large-eddy-based simulations (LES) for flows past rigid airfoils in free-to-rotate and free-to-rotate-and-heave conditions at low Reynolds numbers. As observed in experiments, the numerical simulations confirmed the presence of the self-sustained low-amplitude limit-cycle oscillations (LCOs) of the airfoils. It is understood that this behavior in the transitional Reynolds number regime results from the unsteadiness of the laminar boundary layer separation and its delayed recovery when compared to the corresponding static conditions.

Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Karman shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.

This paper presents recent progress in an investigation of the aeromechanical aspects of unsteady flapping wings for nano-air vehicles (NAV). Numerical results obtained using the in-house code INSflow for insect-sized airfoils and wings are discussed. The calculations revealed that, at hover conditions, the vortices formed during the airfoil plunging motion may remain in the near region and affect the new vortex formation and thus the integral aerodynamic performance. In addition, the flow around the insect-like wing is fully threedimensional. The tip flow affects the flow separations and is shown to reduce the separation intensity. Nomenclature b = length of the wing span c = chord length f = frequency of the flapping motion H = plunging amplitude (peak value) of the point about which the airfoil pitches H = normalized plunging amplitude, / H Hc = c k = reduced frequency, / c kf c U π ∞ = Re = Reynolds number based on airfoil chord, Re /( / ) ref Uc μρ = t = time ref U

Presently at the Von Karman Institute for Fluid Dynamics (VKI), a comprehensive research investigation is being performed on the lateral-directional stability characteristics of a BizJet-class Blended-Wing-Body (BWB). This paper presents the designing attempts using lower order methods and vortex lattice based methodology to get a comprehensive understanding of the aerodynamic performance, followed by and compared with more accurate CFD methods to fine tune the performance. Specific interest is placed in the take-o and landing flight regimes where the BWB is prone stall and departure due to assymetric wind gusts. Dierential control eectors on the bottom of the wing, called belly-flaps, are investigated for lateral-directional control.

This paper presents recent progress in a continuing investigation of the aeromechanical aspects of unsteady flapping wings for micro air vehicles (MAV). Numerical simulations were performed for two-dimensional (2D) pitching-plunging airfoils and three-dimensional (3D) flapping wings, mainly at hover conditions, using an in-house code called INSflow. The results were compared with available experimental data obtained in the water tunnel at the NRC-IAR. The investigation revealed that, at hover conditions, the vortices formed during the airfoil plunging motion may remain near the airfoil and affect new vortex formations, and thus the integral aerodynamic performance. In addition, the flow around the 3D insect-like wing is fully three-dimensional. The tip flow affects the flow separation, reducing the separation intensity. Two-dimensional calculations may over-predict the separation and the shedding vortices, thus affecting the generation of aerodynamic forces.

This paper presents recent progress of the research performed as part of the NRC’s participation in the NATO AVT-149 Unsteady Low-Re Aerodynamics Task Group. Results obtained using the NRC-IAR in-house code INSflow on an SD7003 airfoil compared with experimental results from the US Air Force Research Laboratory (AFRL) indicate that fully three-dimensional (3D) flow separation and stall occur during the flapping motions.

This paper furnishes further computed results for the unmanned combat air vehicle (UCAV) 1303 baseline model to be compared against the measurements provided by the Defence Science and Technology Laboratory (DSTL), UK. The subsonic/transonic computation fluid dynamics (CFD) computations were performed using both NPARC and FLOWer codes with Mach numbers M = 0.35 and 0.85 and a selection of angles of attack ranging from α = 2.0 to 18 degrees with Reynolds numbers based on a mean aerodynamic chord (MAC) of ReMAC = 4.1×10 and 7.2×10. Comparisons between computed results were best at lower angles of attack.