Search results

This research study aims to minimize autonomous flight cost and maximize autonomous flight performance of a slung load carrying rotary wing mini unmanned aerial vehicle (i.e. UAV) by stochastically optimizing autonomous flight control system (AFCS) parameters. For minimizing autonomous flight cost and maximizing autonomous flight performance, a stochastic design approach is benefitted over certain parameters (i.e. gains of longitudinal PID controller of a hierarchical autopilot system) meanwhile lower and upper constraints exist on these design parameters.

A rotary wing mini UAV is produced in drone Laboratory of Iskenderun Technical University. This rotary wing UAV has three blades main rotor, fuselage, landing gear and tail rotor. It is also able to carry slung loads. AFCS variables (i.e. gains of longitudinal PID controller of hierarchical autopilot system) are stochastically optimized to minimize autonomous flight cost capturing rise time, settling time and overshoot during longitudinal flight and to maximize autonomous flight performance. Found outcomes are applied during composing rotary wing mini UAV autonomous flight simulations.

By using stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads over previously mentioned gains longitudinal PID controller when there are lower and upper constraints on these variables, a high autonomous performance having rotary wing mini UAV is obtained.

Approval of Directorate General of Civil Aviation in Republic of Türkiye is essential for real-time rotary wing mini UAV autonomous flights.

Stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads is properly valuable for recovering autonomous flight performance cost of any rotary wing mini UAV.

Establishing a novel procedure for improving autonomous flight performance cost of a rotary wing mini UAV carrying slung loads and introducing a new process performing stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads meanwhile there exists upper and lower bounds on design variables.

To ensure the stability of the flying wing layout unmanned aerial vehicle (UAV) during flight, this paper uses the radial basis function neural network model to analyse the stability of the aforementioned aircraft.

This paper uses a linear sliding mode control algorithm to analyse the stability of the UAV's attitude in a level flight state. In addition, a wind-resistant control algorithm based on the estimation of wind disturbance with a radial basis function neural network is proposed. Through the modelling of the flying wing layout UAV, the stability characteristics of a sample UAV are analysed based on the simulation data. The stability characteristics of the sample UAV are analysed based on the simulation data.

The simulation results indicate that the UAV with a flying wing layout has a short fuselage, no tail with a horizontal stabilising surface and the aerodynamic focus of the fuselage and the centre of gravity is nearby, which is indicative of longitudinal static instability. In addition, the absence of a drogue tail and the reliance on ailerons and a swept-back angle for stability result in a lack of stability in the transverse direction, whereas the presence of stability in the transverse direction is observed.

The analysis of the stability characteristics of the sample aircraft provides the foundation for the subsequent establishment of the control model for the flying wing layout UAV.

For a delta wing it is not sufficient to consider spanwise bending and chordwise rotation only. Chordwise bending must be added. It is therefore necessary to calculate the aerodynamic coefficients accordingly. The theory of Lawrence and Gerber dealing with aerodynamic coefficients of oscillating delta wings in incompressible flow is extended to give the coefficients at any station of the wing. It is shown that for practical reasons the assumption is made that induction may be neglected. This means that the coefficients are theoretically only correct for a rigid wing in pitching and plunging. However they will be used for a flexible wing with spanwise bending, chordwise bending and rotation. For the oscillating delta wing with subsonic leading edges in supersonic flow the theory of Watkins and Berman will be discussed. Here again the original report is extended to give the coefficients at any station of the wing. The calculation of the natural modes of vibration of the wing, based on the methods of Scanlan and Rosenbaum, is presented for completeness. Finally it is shown how the coefficients and the modes may be combined to give the aerodynamic forces. As an appendix the differences between the British and the American techniques for calculating the aerodynamic coefficients are discussed.

The aim of this paper is to investigate the effect of wing kinematics change on force generation produced by flapping wings.

Forces produced by flapping wings are measured using a load cell and compared for the investigation. The measured forces are validated by estimation using an unsteady blade element theory.

From the measurement and estimation, the authors found that flapping wings produced positive and negative lifts when the wings are attached with the +30° and −30°, respectively.

The authors quantified the characteristics of change in the force generation by flapping wings for three wing kinematics. The wing kinematics was modified by changing the initial wing attachment angle.

The result may be applicable to design of control mechanism for an insect‐mimicking flapping‐wing micro air vehicle, which has only wings without control surfaces at its tail.

The preliminary work may provide an insight for design strategy of flapping‐wing micro air vehicles with compact and handy configurations, because they may perform controlled flight even without control surfaces at their tails.

The work included here is the first attempt to quantify the force generation characteristics for different wing kinematics. The suggested way of wing kinematics change can provide a concept for control mechanism of a flapping‐wing micro air vehicle.

A semi‐empirical relationship is derived for the structural weight of wings, applicable to a wide range of subsonic aircraft. The method is based on a generalized expression for the material required to resist the root bending moment due to wing lift in a specified flight condition. Appropriate factors make the result applicable to cantilever and braced wings, for passenger and general aviation aircraft and for freighters. An assessment of the accuracy, based on actual wing weights of 46 aircraft, indicates that a standard deviation of 9·64 per cent is achieved. The weight formula presented allows for the effects of variations in the main wing dimensions and operational limits of the airplane and is therefore suited to parametric design studies.

IN considering the size of wings, which aero‐plane designers require to lift a given weight, the fact is very apparent that lifting surfaces have become smaller as the art of aeroplane design has advanced. Fig. 1 shows the trend of this development from pre‐war days up to now, expressed by a steady increase in wing loading (lb. per sq. ft.). How is this development likely to go on, and where will it end ?

Thus, consider an unstaggered biplane at, say, 40 dug. incidence. At values of rotational speed likely to be expected in a spin it will have a iairly large positive value of rolling moment, which means, as we have seen, an inward sideslip. This means in turn less sideslip at the tail leading to difficulty in preserving the spin at a low incidence. ? monoplane spinning at the same incidence would do so with outward sideslip, thereby augmenting the sideslip due to rotation. Inertia moments enter into the question as well, and, in fact, the aerodynamic and inertia moments and the eitects of sideslip are so interwoven as to make any simple separation of cause and effect extremely difficult. An opinion has been expressed that the ordinaiy autorotation experiment has no bearing on the fully developed spin at all but may be important as regards the incipient spin. In the wind tunnel spins have been observed on monoplanes at incidences quite outside the autorotation range.

The purpose of this paper is to investigate the effect a variety of different boundary layers have on a wing in ground‐effect.

Experiments were carried out in the University of Southampton's 3′×2′ wind tunnel. A variable length splitter plate was designed and manufactured in order to generate four boundary‐layer thicknesses at a selected measurement position. A single element inverted GA(W)‐1 aerofoil was then introduced to the flow at varying heights above the plate. Laser Doppler anemometry (LDA) and surface static pressure measurements (both on the aerofoil surface and on the splitter plate) were recorded.

The flow beneath the wing is found to be affected considerably by the presence of the boundary layer. As the boundary‐layer thickness is increased, the under‐wing pressure is observed to increase, hence resulting in decreased suction. Further, the LDA results indicate a modification to the wake profile. In particular, at low wing heights, the wake is observed to become entrained in the boundary layer, to differing degrees dependant on the boundary layer present and the wing height.

The acquisition of force values from the tests will have allowed further understanding of the “real world” implications of the presence of the boundary‐layer thicknesses on a wing in ground‐effect but this is not possible in the test facility used.

The aerodynamics of a wing in ground‐effect are of great interest for both lifting surfaces for aircraft and downforce generation in motorsport applications. The implications of this paper enhance the importance of understanding the boundary conditions present when wind tunnel testing for these applications.

Although the influence of the boundary layer on low ground clearance objects has been well documented, the methods used here, in particular the use of the pressure tapped splitter plate and LDA, allow a further insight into the explanations behind this influence.

An aircraft wing comprises one or more sections each foldablc about a fore‐and‐aft substantially horizontal axis and a power‐operated linearly moving member which works in a guideway and is connected through a link to a foldablc wing section such that with the foldablc section folded the link is substantially at right angles to the movable part so that, aided by the friction of that part with respect to its guide‐way, it prevents self‐return movement of the foldablc wing section. The folding wing section 1 is pivotally attached to the fixed wing stub 6 by brackets 2, 3, the bracket 2 being extended to form a member 7 having pivoted thereto one end of a link 8 the other end of which is pivoted to a rectangular slide 9 located in a straight open‐top guideway 10 carried by the wing section 6. The piston rod 12 of a hydraulic jack 11 is connected to the slide 9. An extension 13 of the wing section 1 is apertured at 13 so positioned with respect to a pair of apertured lugs 14 formed on a bracket 4 secured to the wing section 6 that a preferably power‐operated pin may be inserted therebetween to lock the folded wing in its spread position. When the wing section is folded the link 8 is positioned substantially at right angles to the guideway 10, the axis about which folding takes place being triangulated with respect to the axes of the pins passing through the ends of the link 8 so that return movement of the wing caused by wind gusts is prevented. An additional lock may be provided to hold one or both wings in a partial or fully folded position by providing a plurality of holes in the guideway 10 into any one ofwhich holes a pin may be inserted to prevent return movement of the slide 9. Specifications 635,260 and 635,261 are referred to.

Several empirical and semi‐empirical formulae exist for predicting wing weights, and it is the purpose of this note to summarize the results of a study of accuracy associated with some of these methods, with a view to establishing their comparative reliability. The modus operandi for effecting this has been to collect together the weights of wing structures which have actually been weighed, thus giving values of true wing weight, and to use each formula or method to ‘predict’ the weights of the same wings, and then to compare the estimated results with the true wing weights. Deviations between the true and estimated values have been tabulated and subjected to statistical analysis and the comparative measures of accuracy derived from that analysis.