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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.

Ground effect is one of the important factors in the enhancement of wing aerodynamic performance. This study aims to investigate the aerodynamic forces and performance of a flapping wing with the bending deflection angel under the ground effect.

In this study, the wing and flapping mechanism were designed and manufactured based on the seagull flight and then assembled. It is worth noting that this mechanism is capable of wing bending in the upstroke flight as big birds. Finally, the model was examined at bending deflection angles of 0° and 107° and different distances from the surface, flapping frequencies and velocities in forward flight in a wind tunnel.

The results revealed that the aerodynamic performance of flapping wings in forward flight improved due to the ground effect. The effect of the bending deflection mechanism on lift generation was escalated when the flapping wing was close to the surface, where the maximum power loading occurred.

Flapping wings have many different applications, such as maintenance, traffic control, pollution monitoring, meteorology and high-risk operations. Unlike fixed-wing micro aerial vehicles, flapping wings are capable of operating in very-low Reynolds-number flow regimes. On the other hand, ground effect poses positive impacts on the provision of aerodynamic forces in the take-off process.

Bending deflection in the flapping motion and ground effect are two influential factors in the enhancement of the aerodynamic performance of flapping wings. The combined effects of these two factors have not been studied yet, which is addressed in this study.

The purpose of the work is to design a flapping wing that generates net positive propulsive force and vertical force over a flapping cycle operating at a given freestream velocity. In addition, an optimal wing is designed based on the comparison of the force estimated from the quasi‐steady theory, with the wind‐tunnel experiments. Based on the designed wing configuration, a flapping wing ornithopter is fabricated.

This paper presents a theoretical aerodynamic model of the design of an ornithopter with specific twist distribution that results generation of substantial net positive vertical force and thrust over a cycle at non‐zero advance ratio. The wing has a specific but different twist distribution during the downstroke and the upstroke that maintains the designed angle of attack during the strokes. The wing is divided into spanwise strips and Prandtl's lifting line theory is applied to estimate aerodynamic forces with the assumptions of quasi‐steady flow and the wings are without any dihedral or anhedral. Spanwise circulation distribution is obtained and hence lift is calculated. The lift is resolved along the freestream velocity and perpendicular to the freestream velocity to obtain vertical force and propulsive thrust force. Experiments are performed in a wind tunnel to find the forces generated in a flapping cycle which compares well with the theoretical estimation at low flying speeds.

The estimated aerodynamic force indicates whether the wing geometry and operating conditions are sufficient to carry the weight of the vehicle for a sustainable flight. The variation of the aerodynamic forces with varying flapping frequencies and freestream velocities has been illustrated and compared with experimental data that shows a reasonable match with the theoretical estimations. Based on the calculations a prototype has been fabricated and successfully flown.

The theory does not take into account the unsteady effects and estimates the aerodynamic forces at wing level condition. It doesn’t predict stall and ignores structural deformations due to aerodynamic loads. The airfoil section is only specified by the chord, zero lift angle of attack, lift slope, profile drag coefficient and angle of attack as given inputs. To fabricate a light weight wing that maintains a very accurate geometric twist and camber distribution as per the theoretical requirement is challenging.

Useful for designing ornithopter wing (preferably bigger) involving an unswept rigid spar with flapping and twisting.

The novelty of the present wing design is the appropriate spanwise geometric twisting about the leading edge spar.

Flapping-wing vehicles show various advantages as compared to fixed wing vehicles, making flapping-wing vehicles' study necessary in the current scenario. The present study aims to provide guidelines for fixing geometric parameters for an initial engineering design by a simple aerodynamic and flight dynamic parametric study.

A mathematical analysis was performed to understand the aerodynamics and flight dynamics of the micro-air vehicle (MAV). Only the forces due to the flapping wing were considered. The flapping motion was considered to be a combination of the pitching and plunging motion. The geometric parameters of the flapping wing were varied and the aerodynamic forces and power were observed. Attempts were then made to understand the flight stability envelope of the MAV in a forward horizontal motion in the vertical plane with similar parametric studies as those conducted in the case of aerodynamics.

From the aerodynamic study, insights were obtained regarding the interaction of design parameters with the aerodynamics and feasible ranges of values for the parameters were identified. The flapping wing was found to have neutral static stability. The flight dynamic analysis revealed the presence of an unstable oscillatory mode, a stable fast subsidence mode and a neutral mode, in the forward flight of the MAV. The presence of unstable modes highlighted the need for active control to restore the MAV to equilibrium from its unstable state.

The study does not take into account the effects of control surfaces and tail on the aerodynamics and flight dynamics of the MAV. There is also a need to validate the results obtained in the study through experimental means which shall be taken up in the future.

The parametric study helps us to understand the extent of the impact of the design parameters on the aerodynamics and stability of the MAV. The analysis of both aerodynamics and dynamic stability provides a holistic picture for the initial design. The study incorporates complex mathematical equations and simplifies such to understand the aerodynamics and flight stability of the MAV from an engineering perspective.

The study adds to already existing knowledge on the design procedures of a flapping wing.

The theory of rotor dynamics given in Ref. 1 is extended to include the effects of coupling between feathering and flapping (δ3 angle) and flapping hinge offset. Both introduce considerable modification to the classic equations, and instead of simple explicit equations for flapping amplitudes, coning angle, collective pitch and inflow angles, five simultaneous equations have now to be solved. Data sheets have been constructed which enable this to be done quickly and accurately for any design of linearly tapered and twisted blade. It is suggested that the intelligent use of such data sheets is of great assistance in a design office, not only because of the very considerable time savings achieved, but also because they eliminate the most fruitful sources of error in numerical calculation. It is shown that a high offset rotor enables much higher speeds to be achieved with a conventional helicopter—an effect which has already been fairly well publicized. A penalty is paid for this in the form of hub pitching moments which have to be balanced out externally; either by the use of two rotors, offset C.G., aerodynamic surfaces, or inclination of the mechanical axis. These effects will be considered in detail in a further article. Finally, equations are developed for a convenient method of calculating blade elemental angle of attack which is claimed to be superior to classic methods for design office purposes.

The purpose of this paper is to demonstrate the uncontrolled vertical takeoff of an insect-mimicking flapping-wing micro air vehicle (FW-MAV) of 12.5 cm wing span with a body weight of 7.36 g after installing batteries and power control.

The forces were measured using a load cell and estimated by the unsteady blade element theory (UBET), which is based on full three-dimensional wing kinematics. In addition, the mean aerodynamic force center (AC) was determined based on the UBET calculations using the measured wing kinematics.

The wing flapping frequency can reach to 43 Hz at the flapping angle of 150°. By flapping wings at a frequency of 34 Hz, the FW-MAV can produce enough thrust to over its own weight. For this condition, the difference between the estimated and average measured vertical forces was about 7.3 percent with respect to the estimated force. All parts for the FW-MAV were integrated such that the distance between the mean AC and the center of gravity is close to zero. In this manner, pitching moment generation was prevented to facilitate stable vertical takeoff. An uncontrolled takeoff test successfully demonstrated that the FW-MAV possesses initial pitching stability for takeoff.

This work has successfully demonstrated an insect-mimicking flapping-wing MAV that can stably takeoff with initial stability.

THE continual development of helicopter rotor systems has so far resulted in the use of about six main types, and it will be of value briefly to recapitulate their advantages and disadvantages in order to obtain a balanced picture against which the stiff‐hinged rotor can be judged.

The purpose of this paper is to present the procedure and results from instrumented flight test performed on the flapping MAVs being developed by the authors. A test is performed using a test measurement system to obtain the real characteristics of the flapping vehicles during their flight.

The test is performed in an indoor flight test facility, equipped with a motion capture system and tracking cameras. Spatial position data are obtained from the vehicles with retro‐reflective tracking markers attached. A quantitative analysis is carried out through the investigation and interpretation of the test data for the flight performance assessment of the vehicles.

The finding of the analysis addresses that the test enabled the numerical measurement of vehicles' flying performance and shows the present vehicles have combined characteristics of both birds and insects.

The test metrics attempted in the present study are applicable to the test and evaluation of general flapping micro air vehicles. Thus, this testing method will be useful for the development of future micro air vehicle system.

Full‐scale instrumented flight test and measurement of performance parameters of flapping micro air vehicles other than visual observation are unprecedented and expected to present the guideline of systematic test and evaluation of flapping micro air vehicles.

The purpose of this study is to experimentally analyse the effect of flexible and stiffened membrane wings in the lift generation of flapping micro air vehicle (MAV).

This is analysed by the rectangle wing made up of polyethylene terephthalate sheets of 100 microns. MAV is tested for the free stream velocity of 2 m/s, 4 m/s, 6 m/s and k^* of 0, 0.25, 1, 3, 8. This test is repeated for flapping MAV of the free flapping frequency of 2 Hz, 4 Hz, 6 Hz, 10 Hz and 12 Hz.

This study shows that the membrane wing with proper stiffeners can give better lift generation capacity than a flexible wing.

Only a normal force component is measured, which is perpendicular to the longitudinal axis of the model.

In MAVs, the wing structures are thin and light, so the effect of fluid-structure interactions is important at low Reynold’s numbers. This data are useful for the MAV developments.

The effect of chord-wise flexibility in lift generation is the study of the effect of a flexible wing and rigid wing in MAV. It is analysed by the rectangle wing. The coefficient of normal force at different free stream conditions was analysed.

The component of velocity, μ0ΩR, along the x‐axis is specified, and also X, the angle of inclination of the flight path. The concept of the ‘internal’ and ‘external’ parameters of the motion is introduced, the ‘internal’ parameters consisting of certain combinations of the control and flapping angles θ0, θ1, θ2, β0, β1, β2 and the ‘external’ parameters consisting of μ0, Λ0, Θ, X, ∈ and z*, where Θ is the inclination of the x‐axis to the horizontal, and e the inclination to the vertical of the rotor resultant force‐vector. For a given centre of gravity position of the helicopter the ‘external’ parameters are shown to be determinable independently of the ‘internal’ parameters, subject to certain assumptions regarding the fuselage drag coefficient. The fundamental ‘internal’ parameter then emerges as (β1—Θ), which physically is the fore or aft inclination to the vertical of the rotor cone axis. The value of this parameter is found as the root of a quadratic equation, and not a linear equation as hitherto. It is shown that at high values of μ0 the rotor force‐vector docs not coincide with the rotor cone axis (as it is commonly supposed to do), the difference between (β1—Θ) and ∈ amounting possibly to 5 deg., of which no more than 1½ deg. can be accounted for by the rotor mean drag force component. Particular attention is paid to the case of horizontal flight, in which the true speed is μ0ΩR see Θ, the factor sec Θ being important. It is shown that at a given true speed the following quantities are independent of the centre of gravity position of the helicopter: (i) ∈, (ii) the quantity (λ1—Λ0+λ0β1), i.e. the component of the air velocity along the rotor cone axis, (iii) the horse‐power of the engine. In addition the following ‘internal’ parameters, (iv) (β1—Θ), (v) (θ2—β1), (vi) (θ1—β2), are very nearly independent of C.G. position, their variations with C.G. position only just becoming noticeable at large values of μ0. This feature enables a much simplified solution to the trim problem to be obtained by working in terms of a fictitious C.G. position chosen to make Θ zero. The genuinely C.G.‐sensitive parameters θ1, θ2, β1, β2 and Θ are subsequently converted to their true values corresponding to the actual C.G. position. The effect of variation in blade moment of inertia is examined through variations in the associated parameter γ, and it is shown that both the ‘external’ and ‘internal’ conditions are substantially unaltered, save for (θ1+ β2) and the mean coning angle β0, which do vary considerably with γ. The analysis is accurate at least so far as μ05 and the results are shown to be consistent with an energy equation. The only sources of error in the dynamics of the problem lie in the rejection of higher powers of μ0 and of flapping harmonics of β beyond the first. In order to assess the effects of this rejection, an analysis is made of a rigid rotor system free only to feather without flapping, and an exact dynamical solution for horizontal flight is obtained, subject to the same aerodynamical assumptions as before. At a given value of μ0 see Θ the results differ slightly from the case of the flapping rotor because now the purely feathering rotor applies a pitching couple mechanically to the fuselage via the shaft, which the flapping rotor is not able to do. However, by calculating the value of this couple, and supposing that the helicopter with the flapping rotor experiences either an aerodynamic fuselage pitching moment of the same amount or alternatively an appropriate C.G. shift, the two cases are reduced to essentially equivalent physical systems, particularly if the mean coning angle, β0, is deliberately arranged to be zero by proper choice of γ. A recalculation of the flapping case is then made; if its theory were exact, its solution would coincide with the known exact solution of the purely feathering rotor case. The discrepancies are shown to be very small, and thus the validity of rejecting μ06 is established for the earlier flapping rotor analysis. Part III is concerned with the theoretical solution of the feathering‐cumflapping rotor case, and also with a geometrical account of the trim configuration. Part IV contains the solution for the purely feathering rotor together with a discussion of the large number of numerical results and related tables, obtained on a Pegasus Computer.