Low-speed Wind Tunnel Tests Research Articles

Fluid–acoustic–structure resonance mechanism of a plane cascade via a low-speed wind tunnel test

Acoustic resonance is an important factor that contributes to aeroengine compressor failure. In this study, a plane cascade of compressor blades was designed to reproduce acoustic resonance via a low-speed wind tunnel test. A high-frequency hot-wire, microphone and strain gauge were used to synchronously measure the fluid, acoustic and structural parameters. We analysed the variation in the amplitude and frequency of the multi-field parameters with increasing mean flow velocity and explored the multi-field interaction mechanism that induces the acoustic resonance of the plane cascade. The plane cascade effectively reproduced the acoustic resonance phenomenon. The first-order acoustic-mode frequency of the plane cascade flow duct, second-order torsional vibration mode frequency of the blade and shedding mode frequency of the tip clearance leakage vortex were equal under acoustic resonance. The fluid, acoustic and structural fields showed a strong interaction effect, achieving the maximum blade vibration amplitude and causing fatigue cracks of torsional vibration at the blade root. The frequency lock-in region of the compressor plane cascade was divided into an ‘acoustic–structure’ interaction region, a ‘fluid–acoustic–structure’ interaction region and a first-order acoustic-mode dominant region with increasing mean flow velocity, which demonstrates an interesting phenomenon in which the fluid–acoustic–structure modes compete: acoustic mode > blade vibration mode > vortex shedding mode. The results demonstrate a unique approach to the study of acoustic resonance that provides insight into the acoustic resonance mechanism in a cascade of compressor blades.

An optical-scan method for measuring the as installed surface port incidence angles for flush air data sensing (FADS) systems

The Flush Air Data Sensing (FADS) System, where air data are inferred from non-intrusive surface pressure measurements, uses natural contours of the vehicle forebody, wing leading edge, or probe. Although multiple methods have been developed to derive airdata from the sensed pressure matrix, all methods rely on accurate knowledge of local surface contours at the port locations. One of the most well-developed solution methods curve-fits the surface pressure distribution against the associated surface incidence angles using a quasi-Newtonian model. The well-known "Triples" algorithm extracts airdata from the curve-fit model. This solution method requires precise knowledge of as-installed incidence angles, i.e. the angles between the surface normal and the longitudinal axis of the vehicle. This study investigates the feasibility and accuracy of using an inexpensive optical-scanning system to measure the in-situ FADS pressure ports surface incidence angles. Here, two legacy 3-D printed probe shapes, as previously tested during a series of very low-speed wind tunnel tests, were used to develop and evaluate this method. The shapes 1) a hemispherical head cylindrical forebody, and 2) a Rankine-Body, were scanned along the longitudinal axis and the resulting point-cloud was edited using open-source software to generate three concentric "loops" surrounding each surface port. Each annular loop was assumed as co-planar with the surface port, and the singular-value decomposition (SVD) was used calculate the local surface gradient vector from the null-space solution. From the resulting gradient vector, geometric relationships calculate the port's polar coordinates including the surface incidence angle. For both body contours the resulting calculations are compared to the "known" design surface angles as prescribed for the 3-D prints. Error plots are presented for each individual ring-set, and for the collected set using all three ring together. For the collected data sets, the incidence angle calculations are accurate to within a quarter-degree.

Numerical and experimental study of the impact on aerodynamic characteristics of the NACA0012 airfoil

Abstract Using computational models and low-speed wind tunnel tests, the aerodynamic characteristics of the NACA 0012 airfoil with low Re numbers of (8 × 10 4 {10}^{4} , 2 × 10 5 {10}^{5} , 3 × 10 5 {10}^{5} , and 4 × 10 5 ) {10}^{5}) and angle of attack (AOA) ranging from 0° to 18° by two steps are examined. Using the same 3-D wind tunnel dimensions, numerical simulations were run. The software program ANSYS FLUENT was used to solve the mathematical model using the continuity equation, the Navier‒Stokes equations, and the k–ω shear‒stress transport turbulence model. Findings demonstrate that at all AOAs, there is a direct relationship between Reynolds numbers (Re), lift and drag coefficients, kinetic energy, and stall angle. The lift coefficient rises linearly as the AOA increases, peaking at 14°, the stall angle at higher Reynolds number. The lift coefficient was found to decline when the AOA was increased further, reaching its minimal value at an AOA of 18°. With a greater AOA, the airfoil’s drag coefficient rises, creating turbulent flow. The eddies produced by the turbulence cause the flow to start separating from the airfoil surface as turbulence increases. As a result, the airfoil lift coefficient drops, and its drag coefficient rises at the same time, leading to poor performance. The validation of the numerical results through wind tunnel experiments provided confidence in the findings of the study.

Research on low-speed aerodynamic characterictics of flying wing standard model with low-aspect ratio

In order to study the aerodynamic characteristics of low-aspect ratio flying wing configuration, low speed wind tunnel tests and numerical simulation were carried out base on CHN-F1 standard model. The research results show that the repeatability accuracy of the test data is good for being benchmark to CFD verification and validation. In the range of medium angle of attack (less than 14 °), the aerodynamic coefficient changes linearly, and the root mean square of the test data is small. With the increase of the angle of attack, the aerodynamic coefficient changes nonlinearly. It is found that the simulation results were in good agreement with test data in the lift linear section.

Powered Low-Speed Experimental Aerodynamic Investigation of an Over-Wing-Mounted Nacelle Configuration

Over-wing integration of ultrahigh bypass turbofan engines can be a solution for next-generation commercial transport aircraft since it eliminates the ground clearance issue and it has the potential to reduce ground noise due to acoustic shielding. Moreover, a unique characteristic of this installation type is the powered lift benefit at low-speed flight conditions. This paper aims to experimentally investigate the effect of the engine power setting on the low-speed aerodynamic performance of an over-wing-mounted nacelle configuration comprising a conventional tube-and-wing layout. Thus, low-speed wind tunnel tests were performed for a half-span powered scale model of the aforementioned configuration. The effect of the engine power setting on the wing lift and spanwise pressure distributions was investigated. The experiments were carried out for angles of attack varying from 0 to 6° and inlet mass flow ratios up to 2.4. The results were used to validate computational fluid dynamics simulations conducted for the same wind tunnel conditions. It has been shown that a significant powered lift benefit can be achieved for the studied configuration, without a penalty in the net propulsive force and that the lift increases linearly with the inlet mass flow ratio. Furthermore, it was observed that the engine power setting largely influences the pressure distributions along the wing, especially at the spanwise sections closer to the nacelle. The low-momentum zone created upstream of the engine at high power settings reduces the pressure at the wing’s upper surface, which is the main factor responsible for the increased lift. By taking advantage of such behavior, drag can potentially be reduced at takeoff and climb due to a lower flap setting required for the same lift.

The Aerodynamic Performance of a Novel Overlapping Octocopter Considering Horizontal Wind

This paper investigates the aerodynamic performance of an overlapping octocopter with the effect of horizontal wind ranging from 0 to 4 m/s using both low-speed wind tunnel tests and numerical simulations. The hovering efficiency and the potential control strategies of the octocopter under the effect of horizontal wind are also validated using blade element momentum theory. The velocity distribution, rotor pressure and vortex of the downwash flow with the horizontal wind are presented using the Computational Fluid Dynamics (CFD) method. Finally, wind tunnel tests were performed to obtain the thrust and power consumption with the rotor speed ranging from 1500 to 2200 rpm for horizontal winds at 0 m/s, 2.5 m/s and 4 m/s. The results showed that horizontal wind decreased the flight efficiency of the planar octocopter and had little effect on the coaxial octocopter. It is also interesting to note that horizontal wind is beneficial for thrust increments at a higher rotor speed and power decrements at a lower rotor speed for the overlapping octocopter. Specifically, the horizontal wind of 2.5 m/s for a lower rpm is presented with a power decrement with proper aerodynamic interference between the rotor blades. Additionally, the overlapping octocopter obtains a higher hover efficiency at 4 m/s compared to traditional octocopters, which is more suitable for flying in a cross wind with a more compact structure.

Investigation of Fin Buffet Characteristics of a Developmental Combat Aircraft Through Wind Tunnel Experiments

Airframe buffet is an important dynamic problem associated with combat aircraft flight, especially in the high Angle-of-Attack (AoA) regime. Occurrence of buffet due to flow separation at high angles of attack needs to be structurally cleared and a buffet boundary needs to be established. The methodology adopted to clear the aircraft from buffet during high AoA flight testing of a developmental light combat aircraft is focused in this article. Experimental unsteady pressure data from low speed wind tunnel tests conducted on a scaled down (1:10) model and airframe especially fin vibration data measured from flight tests formed the basis for this methodology. The reference spectrum has been formulated through unsteady pressures measured using pressure transducers mounted on the fin of the wind tunnel model. This spectrum (scaled up to 1:1) i.e. pressure Power Spectral Density (PSD) excitation was used as an input to the FE model of aircraft fin mounted on the rear fuselage. Buffeting response predicted in terms of acceleration PSD on the fin and rudder have been compared with the flight measured acceleration PSD data to validate the numerical model. The validated model is used to predict loads to clear the aircraft for further flight envelope expansion at high AoA.

Testing and Commissioning of a Low-Speed Wind Tunnel (LSWT) Test Section

The calibration of a low-speed wind tunnel (LSWT) test section had been made in the present work. The tunnel was designed and constructed at the Aerodynamics Lab. in the Mechanical Engineering Department/University of Baghdad. The test section design speed is 70 m/s. Frictional loses and uniformity of the flow inside the test section had been tested and calibrated based on the British standards for flow inside ducts and conduits. Pitot-static tube, boundary layer Pitot tube were the main instruments which were used in the present work to measure the flow characteristics with emphasize on the velocity uniformity and boundary layer growth along the walls of the test section. It is found that the maximum calibrated velocity for empty test section is 55 m/s. Three speeds are tested for uniformity and walls boundary layer at inlet and mid-section of test section. The results show that the flows are uniform at inlet and mid-section with turbulent flow from inlet to outlet.

Aerodynamic modelling of a tilt-wing transition corridor

With the recent rapid evolution of the UAM class concepts in past years, the re-emerged interest in VTOL concepts and their aerodynamics arises. The tilt-wing is one of the possible solutions for such a vehicle. Recent decades of tilt-wing aerodynamic research point to the importance of the understanding transition regime from cruise to hover. Especially the wing stall phenomena which affect descent capabilities and control during the landing phase. Understanding such a regime is crucial for optimal wing design. This paper presents several approaches to the modelling of aerodynamic forces during the critical phase of transition. Computations are aimed particularly to the edge of the flight envelope. The low to high fidelity methods are employed and compared with the results of low-speed wind tunnel testing. The results from high fidelity methods match well with experimental results but the proposed simple analytical model fails in prediction of the transition corridor in the most critical regimes.

Evaluation of the frequency characteristics of cntTSP measurement for unsteady low-speed flow

Carbon nanotube temperature-sensitive paint (cntTSP) is a fluid measurement technology that utilizes temperature-sensitive paint (TSP). It works as an optical temperature sensor, with a thin layer of carbon nanotubes to heat the TSP layer. This study investigated the frequency characteristics of cntTSP measurements in low-speed flow. The flow field on the flat plate was periodically changed by introducing intermittent local disturbances to the flat plate in a low-speed wind tunnel test. cntTSP measurement was conducted behind the local disturbance to evaluate changes in the temperature associated with periodic changes in the flow. The amplitude of the temperature change decreased approximately linearly with the frequency in a double-logarithmic graph. Moreover, the temperature amplitude at 25 Hz was 0.008 K, and it was necessary to detect a very small temperature change to evaluate high-frequency phenomena.

Plasma Gurney Flap Flight Control at Low Angle of Attack

A flight control concept without movable surfaces using electrofluidic actuators is assessed from low-speed scaled two-dimensional (2-D) rectangular wings up to three-dimensional (3-D) full-scale tapered swept wings at realistic aircraft operating conditions. The plasma Gurney flap concept consists of placing spanwise dielectric barrier discharge (DBD) plasma actuators on both sides of a wing or empennage near the trailing edge, and pointing in opposite directions, to change the flow curvature in this region, thus altering lift. Low-speed wind-tunnel tests and corresponding computational fluid dynamics (CFD) simulations are first carried out for this concept on scaled 2-D rectangular wings. Measured lift moment and flow streamlines are used to validate the simulation (CFD) tool. Simulations with this tool are then performed to assess the effect of wing taper and sweep on the effectiveness of the concept at low speed and to evaluate the concept for full-scale 3-D tapered swept-wing geometries for a single-aisle airliner. The results indicate that the plasma Gurney flap has better performance on tapered and swept wings relative to rectangular wings because of the reduction in the velocity component normal to the plasma actuator. Moreover, the presence of shocks at cruise flight conditions improves the effectiveness of the plasma Gurney flap concept.

Study on circulation control of flying wing based on Coanda effect

In recent years, it has been a hot spot to use flow control technology to generate control capability similar to mechanical rudder. Taking the flying wing configuration as the research object, based on CFD and low speed wind tunnel test, the circulation control of airfoil and wing based on Coanda effect is studied, and the parameter influence law and control effect are given. The results show that: when the rear edge of the wing is opened with the corresponding slot, the control capability of the upper-surface blowing is similar to that of the mechanical rudder; the control capacity of pitch and roll is increased with the decreasing of jet height and the increasing of air pressure ratio, the pitching and rolling capacity increases first and then decreases; the control effect of circular Coanda surface is better than elliptical Coanda surface; the pitching and rolling control effect of the optimized circulation-control scheme in Ma=0.2 is not lower than that of the mechanical rudder, the corresponding flow flux is about 0.05~0.07 kg/s.

Experiment on longitudinal aerodynamic characteristics of flying wing model with plasma flow control

Horizontal tail is eliminated from the flying wing layout for improving the low observable and aerodynamic efficiency, resulting in degrading longitudinal maneuverability and fight stability. The low speed wind tunnel test study of improving the longitudinal aerodynamic characteristics of large aspect ratio flying wing model is carried out by using plasma flow control technology. The flying wing model has a leading-edge sweep angle of 34.5° and an aspect ratio of 5.79. The reasons for deteriorating the static maneuverability and stability of the flying wing model and the mechanism of plasma control of the flow field and longitudinal aerodynamic characteristics are studied by particle image velocimetry (PIV) flow visualization and static force measurement test. The control law of plasma control of the flight maneuverability and stability of the flying wing model is studied through flight test. The fact that the flow separation of the outer wing of the flying wing model occurs earlier than the inner wing and the wing is swept back can result in the forward movement of the aerodynamic center and the deterioration of the longitudinal static stability. The shock disturbance induced by plasma can suppress the flow separation of the suction surface, thereby extending the linear section of the lift curve of the model, preventing the aerodynamic center from moving forward, and improving the longitudinal static stability. When the wind speed is 50 m/s, the plasma control improves the horizontal rudder efficiency at a high angle of attack of the flying wing model, increases the maximum lift coefficient of the model by about 0.1, and postpones the stall angle of attack by more than 4° at different rudder angles. The plasma control allows the flying model to follow the command movement better while flying, increases the flying pitch limit angle from 11.5° to 15.1°, reduces the amplitude of longitudinal disturbance motion by 2°, and reduces the oscillation attenuation time from 15 to 8 s, thereby improving the longitudinal flight maneuverability and stability of the flying wing model. It can be seen that plasma flow control technology has great potential applications in improving the flight quality of flying wing layout.

Aeropropulsive Coupling Effects on a General-Aviation Aircraft with Distributed Electric Propulsion

As distributed electric propulsion has continued to gain popularity as an enabling aircraft technology, further research is required to understand the complete implications of aeropropulsive coupling effects for distributed propulsion systems integrated into actual aircraft geometries. A set of low-speed wind-tunnel tests was performed at the University of Illinois at Urbana-Champaign to characterize the aerodynamic performance of a semispan Cirrus SR22T wing-body model that was equipped with an array of four ducted fans integrated into the upper-surface, trailing edge of the wing. Aerodynamic performance and aeropropulsive coupling effects were investigated through six-component load-cell measurements of the forces and moments experienced by the model, supplemented by chordwise pressure distributions at four stations along the span of the wing. Additionally, a set of unsteady force and moment measurements were taken to investigate time constants associated with the response of aerodynamic forces and moments to step changes in fan speed. It was found that the fan tip-speed ratio dictates the aerodynamic performance of the model, with primary effects including an increased lifting force, yawing moment, rolling moment, and nose-down pitching moment as the tip-speed ratio increased. Additionally, the lift-curve slope changed with tip-speed ratio as a function of thrust setting in relation to thrust required.

Research and Development of a Small-Scale Icing Wind Tunnel Test System for Blade Airfoil Icing Characteristics

In order to study the icing mechanism and anti-icing technology, a small low-speed reflux icing wind tunnel test system was designed and constructed. The refrigeration system and spray system were added to the small reflux low-speed wind tunnel to achieve icing meteorological conditions. In order to verify the feasibility of the test system, the flow field uniformity, temperature stability, and liquid water content distribution of the test section were tested and calibrated. On this basis, the icing tests of an aluminium cylinder, an NACA0018 airfoil, and an S809 airfoil were carried out, and the two-dimensional ice shape obtained by the test was compared with the two-dimensional ice shape obtained by the numerical simulation software. The results show that in the icing conditions and icing time studied, the parameters of the test system are stable, and the experimental ice shape is consistent with the simulated ice shape, which can meet the needs of icing research.

Structural concept of an adaptive shock control bump spoiler

Drag reduction technologies in aircraft design are the key enabler for reducing emissions and for sustainable growth of commercial aviation. Laminar wing technologies promise a significant benefit by drag reduction and are, therefore, under investigation in various European projects. However, of the established moveable concepts and high-lift systems thus far most do not cope with the requirements for natural laminar flow wings. To this aim, new leading edge high-lift systems have been the focus of research activities in the last 5 years. Such leading edge devices investigated in projects include a laminar flow-compatible Kruger flap (Schlipf (2011) Insect shielding Krüger—structural design for a laminar flow wing. In: DGLR Congress 2011, Bremen, pp 55–60) and the Droop Nose concept (Kintscher et al. Ground testof an enhanced adaptive droop nose device. In: European Congress on Computational Methods in Applied Sciences and Engineering, ECCOMAS 2016. ECCOMAS2016—VII European Congress on Computational Methods in Applied Sciences and Engineering, 5–10 June 2016, Crete Island, Greece; Kintscher et al. Low speed wind tunnel test of a morphing leading edge. In: AIDAA—Italian Association of Aeronautics and Astronautics XXII Conference, 09–12 Sept. 2013. Neapel, Italien) and these can be considered as alternatives to the conventional slat. Hybrid laminar flow concepts are also under investigation at several research institutes in Europe (Fischer. Stepless and sustainable research for the aircraft of tomorrow—from AFloNext to Clean Sky 2. In: 1st AFloNext Workshop Key Note Lecture No. 1, Delft, The Netherlands, 10 Sept 2015). Another challenge associated with laminar wings aside from the development of leading edge movables is the need to address the control of aerodynamic shocks and buffeting as laminar wings are sensitive to high flow speeds. Here, one possible method of decreasing the wave drag caused by the aerodynamic shock is through the use of shock control bumps (SCBs). The objective of SCBs is the conversion of a single strong shock into several smaller and weaker λ-shocks resulting in a drag benefit when deployed correctly. A particular desirable characteristic of SCBs is that they should be adaptable in position and height as the shock position changes with varying conditions such as speed, altitude, and angle of attack during the flight. However, as a fixed case, SCBs can also help to control laminar buffeting by fixing the shock into given positions at the SCBs location. In this paper, a structural concept for an adaptive shock control bump spoiler is presented. Based on a concept of a fixed bump SCB spoiler, a design for an adaptive spoiler with two conventional actuators is presented. Design drivers and interdependencies of important design parameters are discussed. The presented design is simple and aims for a high TRL without adding much complexity to the spoiler. It is robust and able to form a bump with a height of 0.6% chord length which position can be adapted in a range of 10% chord. This paper is a follow-up of a previous publication (Kintscher and Monner, SAE Tech Paper 10.4271/2017-01-2164, 2017) with extending the focus by a validation of computational results by experimental tests.

Numerical analysis of high-lift configurations with oscillating flaps

This study presents two-dimensional aerodynamic investigations of various high-lift configuration settings concerning the deflection angles of droop nose, spoiler and flap in the context of enhancing the high-lift performance by dynamic flap movement. The investigations highlight the impact of a periodically oscillating trailing edge flap on lift, drag and flow separation of the high-lift configuration by numerical simulations. The computations are conducted with regard to the variation of the parameters reduced frequency and the position of the rotational axis. The numerical flow simulations are conducted on a block-structured grid using Reynolds Averaged Navier Stokes simulations employing the shear stress transport k-omega turbulence model. The feature Dynamic Mesh Motion implements the motion of the oscillating flap. Regarding low-speed wind tunnel testing for a Reynolds number of 0.5 times 10^{6} the flap movement around a dropped hinge point, which is located outside the flap, offers benefits with regard to additional lift and delayed flow separation at the flap compared to a flap movement around a hinge point, which is located at 15 % of the flap chord length. Flow separation can be suppressed beyond the maximum static flap deflection angle. By means of an oscillating flap around the dropped hinge point, it is possible to reattach a separated flow at the flap and to keep it attached further on. For a Reynolds number of 20 times 10^6, reflecting full scale flight conditions, additional lift is generated for both rotational axis positions.

Effects of Wind Disturbance on the Aerodynamic Performance of a Quadrotor MAV during Hovering

Wind disturbance could render thrust and power variation or even causing roll which is difficult to maintain a steady flight in gust especially when the horizontal or vertical wind is involved. In this paper, the horizontal wind and vertical wind are presented to study the influence of wind disturbance on aerodynamic characteristics of the quadrotor aircraft in hovering by experiments and numerical simulations. First, the simplified aerodynamic model with the wind disturbance was analyzed in detail. Also, the low‐speed wind tunnel tests were performed to obtain the thrust and power variation of the quadrotor aircraft with rotor spacing ratio s = 1.1‐1.8 in both horizontal and vertical winds of 0‐5 m/s with the rotational speed ranging from 1500 to 2300 rpm. Finally, the simulations are performed by utilizing the Computational Fluid Dynamics (CFD) software ANSYS to study the flow field distribution of quadrotor with the influence of the wind disturbance. The comparison between experimental results and simulation results shows that the quadrotor achieves better aerodynamic performance with larger thrust and smaller power consumption at rotor spacing ratio s = 1.8. Additionally, the quadrotor can effectively resist the horizontal wind disturbance, which will bring larger power loading for the quadrotor, especially at 2.5 m/s. However, the vortices near blade‐tip move upwards and deform with the influence of vertical wind, resulting in the reduction of thrust and aerodynamic performance of the quadrotor.

Design of Low-Speed Slotted, Natural-Laminar-Flow Airfoil Reproducing Transonic Behaviors

Airfoils tested at low speeds generally produce surface pressure distributions that differ substantially from those measured at transonic conditions, which in turn results in very different boundary-layer characteristics even for the same Reynolds number. A methodology is presented for mapping transonic pressure gradients to an equivalent low-speed airfoil. It is applied to the design of a low-speed, slotted, natural-laminar-flow (SNLF) airfoil intended to emulate the pressure gradients of a SNLF airfoil designed for a transonic commercial transport with a Reynolds number of 13.2 million. The on-design cruise operating condition of the original airfoil was chosen as the reference condition; however, the design lift coefficient of the new airfoil was lowered to account for both compressibility effects and suitability for low-speed wind-tunnel testing. The overall design process mirrors the development of the transonic airfoil by beginning with a single-element airfoil and then modifying it to include an aft element. An inverse design method based on conformal mapping was developed and employed to determine the necessary geometry of the low-speed single-element airfoil. Analysis of the new SNLF airfoil confirms that the design objectives were met, and the airfoil is found to be well behaved down to a Reynolds number of 1 million.

Aerodynamic Characteristics of a Micro Multi-Rotor Aircraft with 12 Rotors Considering the Horizontal Wind Disturbance

Wind disturbance posed difficulties for the stability of the micro air vehicles (MAVs) with attitude variation. In this paper, the aerodynamic performance of a MAV with six coaxial rotor pairs considering the horizontal wind is investigated by both experiments and numerical simulations. First, the effect of the horizontal wind on the multi-rotor aircraft is analyzed in detail. Then, low-speed wind tunnel tests were performed to obtain the thrust and power consumption and the aerodynamic performance of the multi-rotor aircraft (l/D = 1.2 and h/D = 0.19) with the rotational speed of 1500–2300 r/min in the horizontal wind ranged from 0 to 5 m/s. Finally, the distribution of streamline, the pressure of the blade tip, and the velocity and the vortices in the flow field of a multi-rotor aircraft with horizontal wind disturbance, were simulated and studied using the computational fluid dynamics (CFD) method. Through the comparison of experimental results and simulation results, it can be seen that the horizontal wind disturbance will increase power consumption to weaken the aerodynamic performance at higher rotor speeds. However, larger thrust and better hover performance are obtained at lower rotational speeds with good wind resistance. Additionally, due to the mutual induction between rotor wakes, the interactions of downwash flows become more intense at higher rotational speeds or larger wind speeds where the vortexes at the blade tip deformed and moved along with the wind.