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Due to size limitations and the working environment near the ground, small low-altitude unmanned aerial vehicles (UAVs) cannot carry large-area solar cells and energy storage devices. Solar energy alone has only a limited effect on the improvement of their endurance. To solve this problem, this study proposes the combination of solar energy technology and dynamic soaring technology to improve the endurance of small UAVs. Then, the optimal energy acquisition strategy based on the combination of solar energy and dynamic soaring technology is analyzed in terms of energy. The motion equations and energy acquisition and consumption models of a small solar UAV based on horizontal wind shear are established. Moreover, an optimal trajectory planning problem with the maximum charging power of rechargeable batteries as the optimization objective is proposed. The optimal trajectory is solved based on the hp adaptive pseudo-spectral method, and the optimal trajectory combining the two technologies is simulated and compared with the constant-altitude constant-velocity (CACV) trajectory of the UAV using only solar technology. The simulation results show that under the flight background of tmission = 6:00 (summer sunrise) and β = 0.125 s−1 (wind gradient of the wind shear), the optimized O-type trajectory decreased the energy consumption power by 39.86% and increased the solar energy acquisition power by 86.34% compared to the CACV trajectory. Consequently, the small UAV stores more energy in the rechargeable batteries to improve the endurance performance. Compared to the environment where the sun is located on the leeward side of the wind shear, the environment where the sun is located on the windward side of the wind shear can enable the UAV to acquire more energy through the optimized O-type trajectory.

This study investigates the characteristics of a shedding vortex around the Coanda surface and its impact on circulation control (CC) performance. Delay detach-eddy simulation based on the two-equation [Formula: see text] shear stress transport (SST) turbulence model is performed to solve the flow field. The simulation results show that near the Coanda trailing edge, large vortices periodically shed from the lip above the jet slot, and the shedding vortex is surrounded by a series of strip vortices during its downstream movement. The shedding vortex transfers the disturbance to the jet boundary layer and affects the near-wall velocity. Dynamic mode decomposition shows that the shedding vortex and its high-order harmonic flow structures attenuate very slowly and dominate the flow field. The mode with sub-harmonic frequency of the shedding vortex displays the strip vortices dragged out by the shedding vortex. A larger shedding vortex can intensify the mixing of the jet shear and boundary layers, making the jet velocity decay faster. The size of the shedding vortex can be reduced by decreasing the lip height, thus slowing down the attenuation of the jet velocity and improving the CC performance.

Previous attempts at active flutter suppression have been based on driving the deflection of multiple pairs of discontinuous mechanical control surfaces. Here, we explore the effects of trailing-edge Circulation Control (CC) for flutter control on flexible wings. To avoid the problem that the nonlinear aeroelastic model is difficult to establish accurately, we trained a closed-loop control strategy based on the model-free deep reinforcement learning algorithm through aeroelastic wind tunnel testing. The results show that the strategy can intelligently select the appropriate jet intensity according to the real-time state of the flexible wing. The oscillation amplitude of flutter can be reduced by 92%. The air consumption required for unsteady CC to suppress flutter is reduced by 37% compared to steady CC. This study aims to provide an innovative control method and strategy for active flutter suppression of large aspect ratio flexible wings.

A special vortex evolution process was found near the Coanda surface for circulation control airfoil with pulsed actuation. For the jet-on process, two successive shedding vortices from the lower lip edge and the separation vortex on the Coanda surface collide and interact with each other, and periodic vortex shedding forms after several vortex shedding and collision cycles. For the jet-off process, the periodic shedding vortices were replaced by a row of vortices with reverse rotation direction from the upper lip edge, and then reverse vortices broke into small-scale vortices.

The linear and nonlinear evolutions and breakdown of the second modes in hypersonic boundary layers (HBLs) on a flared cone are investigated using Rayleigh-scattering flow visualization and fast-response pressure sensors. Based on two spatially separated pressure signals, cross-bicoherence analysis that permits the distinction of sum- and difference-interactions is utilized to identify the nonlinear interactions. In addition, the visualization temporal and spatial resolution allows fine flow features to be captured to provide additional flow information. Amplitude correlation technique is used to estimate the nonlinear energy transfer between the modes. Our results show that nonlinear interactions between the second mode and the low-frequency wave contribute to the growth of the low-frequency wave, and the difference interactions between the second mode and its first harmonic play a dominant role in modulating the waves in the overall transition process. Amplitude correlation analysis reveals that the spectral energy is nonlinearly transferred from the second mode into its first harmonic and into low-frequency wave, in agreement with the cross-bicoherence analysis. The amplitude modulation of the second mode caused by the difference interaction between the second mode and its first harmonic will reduce the propagation speed of the second mode. However, at the final breakdown stage, this difference interaction vanishes, and the second-mode propagation velocity recovers quickly. Since the frequency of the second mode keeps almost unchanged over the entire transition process, a higher propagation velocity will result in a larger wavelength, indicating an elongation and deformation of the second mode. Eventually, the difference interaction between the second mode and the low-frequency wave accompanying the energy transfer from the second mode to low-frequency waves leads to the final breakdown of the HBLs into a turbulent state.

A monocopter, which is a biology-inspired aircraft based on the samara, has been proved to have passive flight stability. However, due to the asymmetry of its configurations and the constant rotation during flight, its flight dynamics equation is complex. In this paper, the longitudinal stability of a monocopter is systematically analyzed. The longitudinal motion of the monocopter is used as the main research object in this paper. By transforming the body axis coordinate frame to the semi-body axis coordinate frame, its longitudinal dynamics equation is greatly simplified. Then, a fourth-order state space matrix of the longitudinal motion of the monocopter is established, and its longitudinal stability is analyzed at different pitch angles of the wing. Furthermore, the fourth-order state space matrix is simplified into a third-order matrix, and the Rouse criterion is used to analyze the stability of the simplified state space. A set of sufficient conditions is obtained as the criterion for longitudinal stability. Based on this criterion, it is found that the product of inertia [Formula: see text] is a very important factor affecting the stability. Then, the inertial parameters of the aircraft are modified, which greatly expands the range of the pitch angle that maintains longitudinal stability. Finally, the six-degree-of-freedom nonlinear flight dynamic simulation is performed to verify the rationality of the longitudinal stability criterion.

The transient flow structures produced by a pulsed nanosecond plasma actuator and the mechanism by which they are generated are investigated experimentally and through simulations for the case of flow control on a non-slender delta wing. Phase-averaged particle image velocimetry reveals a phenomenon in which, after each discharge pulse, two sub-vortices are generated in sequence and move along the shear layer regardless of the angle of attack, and this is confirmed by hot-wire anemometry. However, at high actuation frequencies ( F+ = fc/ U∞ ≥ 6.435), this phenomenon of double sub-vortices is not observed, and only one sub-vortex is generated per period. The results of pressure measurements indicate that each sub-vortex gives rise to a distinct pressure fluctuation on the wing surface. Numerical simulations reveal a number of residual heats resulting from plasma thermal effects in the shear layer, each of which turns out to induce a corresponding sub-vortex. At low actuation frequencies ( F+ ≤ 4.29), there is a division of the initial residual heat into two independent residual heats and, hence, double sub-vortices per period, whereas at high actuation frequencies ( F+ ≥ 6.435), residual heats from two consecutive periods merge into one, resulting in just one sub-vortex per period, which provides an explanation for the experimentally observed flow behavior.

The plasma synthetic jet actuator (PSJA) has been proven to have potential for shock wave control in the supersonic/hypersonic flow. To evaluate the possibility for the PSJA applied in the hypersonic inlet under a Mach number greater than the design value and provide a deeper understanding of its control mechanism, the relevant two-dimensional numerical simulations are carried out. Based on the discussions about the numerical results, the effects are clear of the actuations generated by the PSJAs, including diffracted waves, synthetic jets, and discharged heat, on the oblique shock waves induced by the compression ramps and the aerodynamic performance of the hypersonic inlet. Concretely, the shapes of the oblique waves are changed by propagations of the diffracted waves, whose starting point moved due to the synthetic jets. A deceleration occurs near the surface of the hypersonic inlet caused by the movement of the discharged heat, leading to a thickener boundary layer. Furthermore, the incremental analysis about the total pressure recovery coefficient at the throat shows that the diffracted waves can increase the static pressure at the throat achieving a higher total pressure recovery coefficient, while the discharged heat makes the total pressure recovery coefficient drop due to a lower local Mach number at the throat. The effect of the discharged heat is more obvious than that of the diffracted waves.

Pulsed surface arc discharge (PSAD) is one of the essential techniques in flow control. In this paper, the effects of electrode configuration on heat transfer characteristics of PSAD and disturbance characteristics of PSAD on the flow field were investigated by thermal imaging technology and high-speed photography technology. By defining the curvature of curve (COC) of the electrode, we investigated the physical mechanism of the electrode configuration, affecting the PSAD disturbed flow field. The results show that the COC has the optimal solution COC opt. When COC ≤ COC opt, the smaller the COC, the more concentrated the PSAD disturbances to the flow field. When COC > COC opt, the electrode configuration will have an end point effect, resulting in a deviation between the theoretical COC and the real COC. The larger the COC, the stronger the end point effect, and the more concentrated the PSAD disturbances to the flow field. The COC affects the disturbance degree of PSAD to the flow field by distorting the electric field intensity distribution. The change in the electric field intensity causes the fluctuation frequency of the flow field to be inconsistent with the discharge frequency of the PSAD. The stronger the distortion degree of the electric field intensity, the stronger the high-frequency characteristics of the fluctuating frequency of the flow field, and the stronger the high-frequency characteristics of the flow mode of the flow field. In addition, we obtained the value range of COC opt within 0.5–0.7 through theoretical derivation and established a mathematical model of electrode structure's effect on the flow field's flow structure.

The complete flow structure and the composition of the flow structure generated by Pulsed Surface Arc Discharge in a Magnetic Field (M-PSAD) are the keys to explaining the plasma application mechanism. The experimental results show that the M-PSAD can produce plasma clusters and detonation waves. The thermal motion of charged particles in the plasma cluster is the main reason for the thermal effect of arc discharge. The detonation wave consists of the shock wave, Taylor expansion wave, and two different thickness zones of charged particles. The experimental results provide a profound physical mechanism for applying plasma.

A better understanding of the mechanism of flow separation suppression by a dielectric barrier discharge is essential for flow control. This paper investigates the mechanism of improving the aerodynamic performance of the airfoil by dielectric barrier discharge when the Reynolds number is in the range of 6 × 104–4 × 105. The results show that the disturbance of the gas discharge to the flow field will form a new flow structure. The fluctuating frequency of the new flow structure determines the ability of the plasma actuator to suppress flow separation. This investigation improves and develops the mechanism of plasma flow control.

Jet flow-control technology is a promising area of fluid research. In this work, the flow-control effect of a curved-surface jet in an incoming flow of Mach = 5 and its underlying control mechanism are experimentally studied using high-speed photography and dynamic force measurement. From the establishment of complete stability of the flow field, the evolutionary process can be roughly divided into five stages: two equilibrium stages (short and long term), jet acceleration stage, bow shock formation stage, interference removal stage, and stable state. By defining the pressure ratio (PR) as an independent variable, it is found that the flow control of the jet occurs through different wave-system structures. The interaction between the jet and the incoming flow produces an oblique shockwave and expansion waves. The shockwave generates thrust and forms a virtual rudder surface; the expansion waves interact with the backflow region and the separated shear layer to generate lift. Moreover, PR has an optimal solution of PR opt. When PR opt, the effect of flow control is related to α, Ve, and ρe, and the greater the PR, the stronger the flow-control effect. When PR > PR opt, the flow-control effect is related to α and ρe, and the larger the PR, the weaker the effect of the shockwave and the stronger the effect of the expansion waves but the slower the growth. In experiments, the thrust, pitching moment, and lift increased by 17.43%, 17.75%, and 9.45%, respectively, because of the appearance of wave-system structure when PR opt = 201.32.

Unmanned Aerial Vehicles (UAVs) can extract energy to improve their range and endurance due to thermal updrafts in the environment via static soaring. In this paper, a soaring control strategy is improved based on multi-hole probe technology. A vertical wind speed estimation method based on the multi-hole probe is proposed to design thermal updraft identification, mode switching logic, and the soaring controller. Turn logic is used to determine the correct turn direction for UAVs when a thermal updraft is detected. A simulation environment containing a six-degree-of-freedom glider model and a thermal updraft model is described. The induced roll effect caused by the asymmetric vertical velocity distribution of thermal updraft is included in the simulations. The results show an improved soaring capability from the proposed soaring strategy. Thus, the proposed method based on a multi-hole probe is more accurate and robust, and the turn logic allows a glider to enter a thermal updraft more quickly. The more energy a glider can obtain from thermal updraft enables a lower soaring speed during static soaring.

Hypersonic aircraft represent an important area of aerospace development. Raising the aircraft lift-to-drag ratio, improving the efficiency of attitude control, and optimizing aerodynamic performance are important goals. Active flow control technology can improve the aerodynamic performance of the aircraft while optimizing its control attitude. In the field of low speed, subsonic speed, and supersonic speed, many experiments and numerical simulations have been carried out to verify the excellent auxiliary or control effects of active flow control technology on aircraft. In this paper, a two-dimensional hypersonic airfoil is selected to analyze the feasibility and control effect of jet flow control technology under hypersonic inflow, and a flow control scheme is designed to make the jet eject from the wing surface. First, the influence of different jet excitation parameters on the scheme is explored through numerical simulation, mainly including the Mach number, the jet location, the jet angle, and the change with the angle of attack. After that, typical excitation states of numerical simulation are selected for hypersonic wind tunnel experiment verification, whose results are in good agreement with those of the numerical simulation. The results prove that, under the condition of high altitude and low pressure, if jet excitation is applied on the surface of the hypersonic airfoil, a local high pressure area will be formed in front of the excitation port, and the differential pressure between upper and lower wings will be changed. It can optimize the aerodynamic characteristics, increase or reduce lift, produce pitching moment simultaneously, enhance the efficiency of the rudder surface, and assist attitude control.

In this study, we investigated the effects of surface sliding discharge of a dielectric barrier discharge (SL-DBD) plasma actuator on the boundary layer flow using the time-resolved particle image velocimetry (PIV) technique. According to the PIV measurements, the peak-shaped mean jet flow is induced by the SL-DBD plasma actuator in quiescent air. Under the freestream of U∞ = 6 m/s on the flat plate, the results show that the streamwise velocity fluctuation induced by SL-DBD plasma is increased pronouncedly at the downstream of the actuator. Meanwhile, the mean flow velocity increases obviously in the downstream of the actuator. The instantaneous results show the formation of the vorticity at the bottom of the boundary layer after the plasma excitation. In addition, the vorticity has the behavior to lift up, possibly because of the induced peak-shaped jet.

Chemical vapor deposition (CVD) is an attractive technique for the metallization of organic thin films because it is selective and the thickness of the deposited film can easily be controlled. However, thermal CVD processes often require high temperatures which are generally incompatible with organic films. In this paper, we perform proof-of-concept studies of photochemical CVD to metallize organic thin films. In this method, a precursor undergoes photolytic decomposition to generate thermally labile intermediates prior to adsorption on the sample. Three readily available Ru precursors, CpRu(CO)2Me, (η3-allyl)Ru(CO)3Br, and (COT)Ru(CO)3, were employed to investigate the role of precursor quantum yield, ligand chemistry, and the Ru oxidation state on the deposition. To investigate the role of the substrate chemistry on deposition, carboxylic acid-, hydroxyl-, and methyl-terminated self-assembled monolayers were used. The data indicate that moderate quantum yields for ligand loss (φ ≥ 0.4) are required for ruthenium deposition, and the deposition is wavelength dependent. Second, anionic polyhapto ligands such as cyclopentadienyl and allyl are more difficult to remove than carbonyls, halides, and alkyls. Third, in contrast to the atomic layer deposition, acid-base reactions between the precursor and the substrate are more effective for deposition than nucleophilic reactions. Finally, the data suggest that selective deposition can be achieved on organic thin films by judicious choice of precursor and functional groups present on the substrate. These studies thus provide guidelines for the rational design of new precursors specifically for selective photochemical CVD on organic substrates.

Vortex control is a thriving research area, particularly in relation to flying wing or delta wing aircraft. This paper presents the topological structures of vortex flow on a flying wing aircraft controlled by a nanosecond plasma dielectric barrier discharge actuator. Experiments, including oil flow visualization and two-dimensional particle image velocimetry (PIV), were conducted in a wind tunnel with a Reynolds number of 0.5 × 106. Both oil and PIV results show that the vortex can be controlled. Oil topological structures on the aircraft surface coincide with spatial PIV flow structures. Both indicate vortex convergence and enhancement when the plasma discharge is switched on, leading to a reduced region of separated flow.