Experimental Investigation of effect Slotted leading Edge Circulation Control Aerofoil Aerodynamic and Acoustics

Experimental Investigation of effect Slotted leading Edge Circulation Control Aerofoil Aerodynamic and Acoustics | Research Square window.SnipcartSettings = { analytics: { enabled: false } }; (function() { var accessVector = localStorage.getItem('access_vector') || ''; window.dataLayer = window.dataLayer || []; if (accessVector) { window.dataLayer.push({ user: { profile: { profileInfo: { snid: accessVector } } } }); } })(); (function(w,d,s,l,i){w[l]=w[l]||[];w[l].push({'gtm.start':new Date().getTime(),event:'gtm.js'});var f=d.getElementsByTagName(s)[0],j=d.createElement(s),dl=l!='dataLayer'?'&l='+l:'';j.async=true;j.src='https://www.googletagmanager.com/gtm.js?id='+i+dl;f.parentNode.insertBefore(j,f);})(window,document,'script','dataLayer','GTM-K279D39R'); Browse Preprints In Review Journals COVID-19 Preprints AJE Video Bytes Research Tools Research Promotion AJE Professional Editing AJE Rubriq About Preprint Platform In Review Editorial Policies Our Team Advisory Board Help Center Sign In Submit a Preprint Cite Share Download PDF Short Report Experimental Investigation of effect Slotted leading Edge Circulation Control Aerofoil Aerodynamic and Acoustics Yasir Al-okbi, Mohammed A. Atiya, Qusay Rasheed Al-amir, Muhsin Jaber Jweeg, and 1 more This is a preprint; it has not been peer reviewed by a journal. https://doi.org/ 10.21203/rs.3.rs-5769420/v1 This work is licensed under a CC BY 4.0 License Status: Posted Version 1 posted You are reading this latest preprint version Abstract This study discusses the separation flow active flow control of the NACA 633618-il three-dimensional straight aerofoil, which is carried out by using the inclined exit synthetic jet exciter. The synthetic jet can effectively control the flow separation of the wings, increase the maximum lift coefficient by 12.4%, and delay the stall angle of attack by 6°. Using boundary layer testing techniques and The particle image velocimetry system is used to study and analyse the control mechanism of the synthetic jet separation flow. The results show that the boundary layer velocity pattern becomes full after the control. The shape factor decreases, the underlying energy increases, and the ability to resist adverse pressure gradients increases. The transient and real-time homogenised PIV test flow field map further proves the synthetic jet flow direction. After the main flow is injected and mixed with momentum, the main flow adheres to the aerofoil surface, the fluid turbulent kinetic energy and Reynolds shear stress near the aerofoil increase, and the separation point is pushed downstream. Later, flow separation is inhibited. Inclined outlet synthetic jet wing separation flow boundary layer particle image velocimetry Figures Figure 1 Figure 2 Figure 3 Figure 4 Figure 5 Figure 6 Figure 7 Figure 8 Figure 9 Figure 10 Figure 11 Figure 12 1 Introduction The flying wing layout integrates the fuselage/wing/tail and the fuselage/propulsion. The integration is highly integrated, and the horizontal tail and vertical tail are cancelled to minimise the adverse aerodynamic interference and radar scatter area of the small wing body/airframe/propulsion, significantly improving the cruise aerodynamic performance and enhancing the stealth combat capability of the new bombers, sensor aircraft, long-range, long-endurance drones, and stealth reconnaissance. Limited research has examined the temporal dynamics of a Coandă-type flow actuator included into a clean aerofoil, designed for GLA or MLA applications. Unsteady Reynolds-Averaged Navier-Stokes simulations An analysis of a symmetrical aerofoil indicated that sufficient activation of a Coandă jet can result in nearly constant lift when subjected to a medium-amplitude 1-cos gust [ 1 ]. Unsteady 2D-RANS simulations of a supercritical aerofoil under varying operating conditions shown that an impulsive activation of a Coandă jet can achieve lift reduction amplitudes within a response time enough for counteracting the briefest 1-cos gusts [ 2 ]. The rapid emergence of actuator-induced lift was additionally demonstrated by an experimental investigation of a cambered aerofoil at low Reynolds numbers. The low slot blowing demonstrated significant control authority and a notably rapid reaction time, effectively mitigating gust-induced loads [ 3 ]. The conventional aileron, a jet normal to the airfoil's surface (normal jet), and a tangential jet over a rounded trailing edge (Coandă jet) are the only actuation designs that combine high peak lift reductions with slight changes in drag and pitching moment over the entire flight envelope, according to two-dimensional Reynolds-Averaged Navier-Stokes (2D-RANS) simulations [ 4 ]. The design of advanced aircraft such as drones is of enormous significance [ 5 – 6 ]. But the design features of the wing-body integrating the lifting surface and the no-tail layout are also important features of the flying wing fabric. Adverse effect on the flight quality of the Bureau's aircraft—heading stability Insufficient, poor longitudinal stability characteristics, etc., all of the above will improve the flight control system. There are extremely high design requirements [ 7 – 8 ]. Due to the need for high lift during takeoff and landing, large aircraft often have to make the wing in a state of high angle of attack, and it is difficult to separate the flow of the aerofoil at this time. to avoid. Flow separation results in increased resistance, reduced rudder efficiency, and a series of problems such as body vibration, lift drop, and even stall, which seriously affect the safety and manoeuvrability of the aircraft during flight. Therefore, researchers have been looking for a variety of efficient and simple means to divide the wing-controlled flow. The passive control method has no external energy consumption, which is convenient for Reliable installation and performance, such as vortex generators, have been widely used in aircraft use. Active control means, by virtue of which it can fine-tune the moving fluid The advantages of accurate phase control and adjustment according to actual working conditions are obtained. People are paying more and more attention, such as when Boeing is researching and applying hybrid laminar flow control technology to the B787-9. As a new type of active flow control technology, synthetic jet has There are many advantages, such as no air source, compact structure, low cost, etc., so it quickly became a hot topic in the field of flow control. Its development can be traced back to 20. The discovery of the acoustic rectification effect in the middle and late nineteenth century [ 9 – 10 ]. A high-amplitude oscillating sound wave was used to obtain radiation with a maximum speed of about 10 m/s. Acoustic rectification, which was observed during sound waves controlling the separation flow. Subsequently, through software numerical simulation, experimental and theoretical analysis at home and abroad, A large number of studies have been carried out on the mechanism of synthetic jet technology [ 11 – 12 ]. Exist On the basis of mechanism research, people have made a lot of research on synthetic jet technology. A large number of experimental investigations have been carried out to develop its engineering application potential. synthetic shot Streaming has a wide range of applications: it can effectively delay separation, delay the stall [ 13 – 14 ], so as to greatly increase the lift and reduce the drag, improve flight performance, achieve thrust vector control [ 15 ], increase strong mixing, and improve the control force of the MAV [ 16 ]; body vortex and helicopter rotor dynamic stall control, etc. These experimental studies This study lays a robust foundation for further research. The PIV results of the flow field at the inclined outlet of the synthetic jet. After the analysis, it is pointed out that the oblique outlet exciter can make the surrounding gas directional. energy and mass transport control, convective flow fields appear as transverse flow along the wall movement transport characteristics [ 17 ]. The influence of the inclination angle of the outflow outlet on the flow control effect has been studied. The oblique outlet exciter applied to the S-shaped intake port [ 18 ]. In separation flow control, a good control effect is achieved. This paper is based on previous research. On the basis of research, the flow of the synthetic jet on the wing surface using the inclined outlet. The separation is controlled by focusing on the interaction between the jet and the main flow by means of PIV. The measurement of the aerofoil flow field after the action, while combining the jet to control the front and rear aerofoil, Variation of pressure distribution and boundary layer velocity profile, control of oblique outlet exciter. A preliminary study on the mechanism of flow separation was carried out. The current study aims to develop an actuator that has a high enough lift control authority to mitigate typical gust-induced loads while having negligible affects on the aerofoil's baseline performance. A low-speed wind tunnel was used to test a subsonic and modular wind tunnel model. This work focusses on both the actuation system's unsteady performance for impulsive switching between upper and lower blowing and the steady actuator performance attained for slot blowing. 2. Experimental and Physical Models 2.1 Aerofoil model The wing model adopts the NACA 633618-il aerofoil, an all-metal, straight machine. Wing, wing chord length C = 300 mm, span length S = 495 mm, aspect ratio Ar = 2.3. At a distance of 0.38c from the leading edge of the aerofoil, there are 6 uniformly distributed in the spanwise direction. Independent exciters with 60mm spacing. The exciter adopts acoustic excitation vibration. Moving diaphragm, power 5W; tangential angle between jet outlet and aerofoil 32°, slit length SL = 15mm, width W = 1.0mm. The edge at the half-span position of the wing A total of 60 pressure measuring holes are opened around the chord to measure the surface pressure of the aerofoil force distribution. The exit velocity of the synthetic jet exciter is mainly determined by the external excitation frequency Y and power amplifier drive voltage Y control. In this experiment, the total pressure probe was used first. The velocity characteristics of the jet were calibrated, and it was found that the fixed voltage R was Under the resonant frequency f = 300 Hz, the outlet velocity is the largest; under the rate n, the outlet velocity uo increases with the increase in the voltage v. The concept of jet-to-main flow ratio proposed by Smith [ 19 ] combines the exciter and the export velocity, which is dimensionless: The middle shape of the above formula is the number of exciters, u o is the jet outlet velocity, and b is the length and width of the exit slit, respectively. The non-dimensional jet excitation frequency is determined by: $$\:{F}^{+}=\frac{\frac{{w}_{jet}}{2\pi\:}}{\frac{{U}_{\infty\:}}{c}}=\frac{{w}_{jet}c}{2\pi\:{U}_{\infty\:}}$$ 1 The non-dimensional jet momentum coefficient is determined by: $$\:{C}_{u0}=2\frac{h}{c}{\left(\frac{{U}_{0}}{{U}_{\infty\:}}\right)}^{2},\:{C}_{um}=2\frac{h}{c}{\left(\frac{{U}_{m}}{\sqrt{2}{U}_{\infty\:}}\right)}^{2}$$ 2 Where c is the speed of sound and U m and U 0 are the amplitudes of the steady and unsteady components of velocity, respectively. Furthermore, a modification of the original pressure boundary condition near the jet orifice is required. The conservation of normal momentum can be used to determine the pressure boundary condition. The boundary condition can be changed as follows by ignoring the impact of viscosity and taking into account the influence of the instantly changing normal velocity component U n : $$\:\frac{\partial\:p}{\partial\:\eta\:}=-\rho\:\frac{\partial\:{U}_{n}}{\partial\:t}$$ 3 2.2 Force measuring system and pressure measuring system The experiment was conducted in the open type of the Department of Aerodynamics, Nanjing University of Aeronautics and Astronautics. A low-speed reflow wind tunnel. The wind tunnel has low turbulence, low noise, etc. The size of the experimental section is 0.5 m × 0.5 m, and the turbulence degree is 0.045%. The minimum stable wind speed of the cave is 2 m/s, and the maximum wind speed is 45 m/s. This experiment is under the wind velocity of v ∞ = 11 m/s; the variation range of the angle of attack α is -2°–30°. The Reynolds number (Re) based on the aerofoil chord length is Re = 2.5×10 5 . The force measuring system consists of a 6-component box-type pneumatic balance, a signal amplification device, a 16-bit data acquisition card, an acquisition control computer, and a special test software composition. After the transformation of the body/wind axis system, each component of the model is finally obtained, amounting to aerodynamic force and torque. The aerofoil surface pressure test system consists of a 60-channel differential pressure transmitter sensor, a 16-bit data acquisition card, an acquisition control computer, and special processing management software composition. The sensor range is 0.35 PSI, and the test system is integrated; the test accuracy is 0.06% FS. The boundary layer probe scanning system is self-developed by Brunel University London of Aeronautics and Aerodynamic Lap. It is mainly composed of a boundary layer probe, an optical coordinate frame, and a pressure sensor. It is composed of several parts, such as an acquisition board and acquisition and processing software, which can realise the probe's arbitrary, precise, and subtle changes in position in three-dimensional space. Probe ruler Small inch (thickness less than 0.3 mm), flat shape, can be approached at close range; the surface of the aerofoil has little interference to the flow field. 2.3 Particle image velocimetry of the flow field test system The PIV flow field test system used in the experiment is the American TSI company. The company's two-dimensional PIV mainly includes double-pulse Nd:YAG lasers, mutual related CCD cameras, synchronisers, tracer particles, operating systems, etc. That The maximum output power of the medium laser is 300 mJ/pulse, and the pulse duration is the pulse interval is 10 ns; the pulse interval is adjustable, and the repetition frequency of the double-pulse laser is 15 Hz. The laser scanning plane is 0.60 between the upper surface of the wing and the root of the wing. The cross section at the doubled length, the chordwise position range is from l = 0.30c to l = 0.70c, as shown in Fig. 3 . 2.4 Three-component balance Plint & Partners LTD employed a three-component balance to assess the aerodynamic forces generated by the aerofoil within the wind tunnel (Fig. 4 ). The balance is made up of a force plate and an aluminium mounting plate that is secured to the wind tunnel working portion. The force plate is able to revolve about the horizontal axis. The aerofoil model is attached to the force plate's model support using a Ø12 mm mounting stem. This support is free to rotate for adjustment of the angle of attack, while its position may be fixed by means of the incidence clamp. The forces acting on the force plate are conveyed by flexible cables to strain gauge load cells, which quantify the lift and the drag forces. The drag cable is horizontal and goes through the centre of the force plate stem, whilst the two lift cables are vertical and equidistant from the model support. Pitching moment is determined by multiplying the difference between fore and aft by 0.127 [ 20 ]. The readings are accurate to within ± 0.05 N. To ensure that the readings are accurate, the three-component balance is calibrated before each measurement. Calibration is performed by taking it from the wind tunnel and placing it on a mounting frame. Before applying a load, the load cells' zero readings are checked. The cells are adjusted by setting the adjustment screw until the display shows a value of zero. The calibration technique comprises the application of known lift and drag forces using dead weights. It is crucial that the force balance is completely balanced both vertically and horizontally during the calibration procedure. Any differences given from the forces are addressed by modifying the cable tensions until equivalent readings for both load components, aft and fore, are reached. Three calibrations are done to assess the linearity of the relationship between load and cell output. Microphone Acoustic Measurements Metric and Polar Array An array of eight G.R.A.S 0.5-inch microphones was positioned 1.2 meters from the leading edge of the aerofoil in a circular arc to assess far-field noise radiation, as illustrated in Fig. 4 . The microphone signals were recorded at 44 kHz for 20 seconds using a 16-bit National Instrument Analogue-Digital card PXI 1042, with a sampling frequency of 50 kHz. The noise spectra were computed with a window size of 1024 data points, yielding a frequency resolution of 48.83 Hz and a BT product of approximately 500, ensuring minimal variance in the spectral estimate. The acoustic pressure at the microphone locations was documented at mean flow velocities ( U ∞ ) ranging from 20 to 60 m/s, respectively. Acoustic pressure spectra and directivity patterns can be computed within an azimuthal range of 90 degrees (45° ≤ θ ≤ 135°, with θ = 0° aligned with the jet axis). The data acquisition is managed using a PC linked by an NI 8360 express card, and is then exported to Matlab. Assuming a cylindrical spreading of sound waves from the aerofoil leading edge, the microphone array in the current research can be used to compute the sound power level (PWL) and overall sound power level (OAPWL). This research examines noise radiation through the Sound Power Level spectrum (PWL), as delineated in Eq. 4, and measured between radiation angles of 50° and 110° (refer to Fig. 5 ), which are also utilised to assess the noise reduction attained by various edge treatments. \(\:{P}_{xx}\left(f\right)=2\pi\:\int\:{W}_{xx}\left(f,{\theta\:}_{i}\right).\varDelta\:\theta\:/\rho\:c0\) , where i = 1, ….,N (4) \(\:PWL\left(f\right)=10{\text{log}}_{10}\left(\frac{{P}_{xx}\left(f\right)}{{P}_{0}}\right)\) , 50°< θ < 110° (5) Where P 0 = 10 − 12 W/Hz, Wxx (f, θi) is the acoustic pressure PSD, measured at a polar angle Δθ, θ (rad) between adjacent radiation angles of the microphone, and Pxx (f) is the integrated sound power between the 50° and 110° radiation angles. The overall sound power level (OAPWL), which is a particular flow rate, can also be used to characterise aerofoil noise: \(\:OAPWL=10{\text{log}}_{10}\left({\int\:}_{f}^{}{P}_{xx}\left(f\right)df/{P}_{0}\right)\) (6) To find precise values for W_xx, a standard Brüel & Kjaer 4231 calibrator is utilised to calibrate and supply a steady 94 dB signal at 1 kHz for 20 seconds. 3. Experimental results 3.1 Discussion on the aerodynamic characteristics of the aerofoil Figure 6 shows the variation curve of the lift-drag coefficient with the angle of attack, which can be It can be seen that the NACA 633618-il straight wing has better slow stall characteristics. In the uncontrolled state, when α = 16°, the lift line is inclined, and the lift coefficient begins to decrease; when α = 18°, the lift coefficient reaches its maximum value. The force begins to decrease slowly, and until α = 25°, the lift decreases sharply, combined with the pressure distribution curve shown in Fig. 7 , which shows that the flow is completely separated. $$\:Cn=\frac{2\sum\:_{i=1}^{n}A{U}_{m}^{2}}{b{U}_{\infty\:}^{2}}$$ 7 3.2 Analysis of Boundary Layer of the Velocity On the basis of obtaining effective control, further explore the synthetic jet First, measure and analyze the near-wall flow field in the upper aerofoil. Under the condition of α = 16°, l = 0.50c, and 0.65c on the wing surface, respectively, the measurement of the velocity profile of the boundary layer with or without synthetic jet control was carried out, and the results are shown in Fig. 9. The boundary layer velocities in different states and shape parameters of type, where: From the velocity type test results of the boundary layer on the upper wing surface, it can be seen that the excitation After the exciter is turned on (Cn = 0.0025), within a certain distance downstream, the speed type of the shape becomes fuller and the form factor becomes smaller. The bottom energy of the boundary layer is obtained; the ability to overcome the adverse pressure gradient is enhanced. This is also the dynamometer knot in Section 2.1 . After the synthetic jet is controlled, the stall angle of attack of the aerofoil is delayed, and the lift coefficient increases. Based on achieving effective control, further explore the synthetic jet output. First, the near-wall flow field on the upper aerofoil must be measured and analysed to control the mechanism. In the state of α = 16°, x = 0.48 hu and 0.60 hu on the wing surface, respectively. The velocity pattern of the boundary layer controlled by a synthetic jet is measured at As shown in Fig. 9. Table 1 gives the boundary layer velocity under different conditions. It can be seen from the test results of the boundary layer velocity pattern on the upper wing of the wing that the exciting. After the actuator is turned on (Cn = 0.0025), the speed type shape becomes fuller and the form factor becomes smaller within a certain distance downstream. The bottom energy of the boundary layer is obtained to improve the ability to overcome adverse pressure gradients. This is also the force-measuring k not in Section 2.1 . As a result, after synthetic jet control, the wing stall angle of attack is delayed and the lift coefficient increases. 3.3 Particle image velocimetry (PIV) flow field test To further explore the interaction between the exciter synthetic jet and the mainstream. The process of using the PIV flow field test system at a typical angle of attack α = 25° The chordwise section of the upper wing surface at a distance of 0.59 from the aerofoil root. l = 0.30c ~ 0.70c to measure the flow state of the two-dimensional flow field. Wind velocity u ∞ = 25m/s. Figure 10 shows that when α = 25°, with or without synthesis the averaged flow field velocity contour when the jet controls the upper surface of the aerofoil. It can be seen that the air flow on the wall of the front wing is controlled gradually along the flow direction decreases, and the surface airflow velocity becomes zero at about 0.50K, and even appears reverse flow velocity, the boundary layer is separated from the wing surface, refer to Fig. 10 (a); After control, the airflow velocity increases at the same chord position on the wing surface. Exist within the test field of view, the main flow reattached the wall and the separation zone was delayed after l = 0.70c (due to the limitation of the experimental conditions of the PIV instrument, the Measure the flow field after l = 0.70c). Figure 10 is a partially enlarged PIV velocity vector diagram. Uncontrolled under the condition of α = 25°, 0.45c ≤ l ≤ 0.65c air velocity on the aerofoil surface The degree direction is disordered, the velocity value is small, the backflow area appears, and the typical flow appears detached state. The comparison of multiple transient speed vector diagrams can also be found, the separation. The position of the point is not fixed, it is about 0.45c ≤ l ≤ 0.50c back and forth oscillation, see Figs. 10 a-b. Due to the limited length of the article, other PIV test results at the moment are not listed here. Synthetic jet excitation after the device is turned on, within the field of view of 0.42c ≤ x ≤ 0.48c, the airflow is stable when the wall is fixed, the separation point is postponed to the outside of the field of view, as shown in Fig. 10 (b). PIV transient time-in-time homogenization test results reflect control of the wing surface Favourable changes in the global flow field, the surface flow separation point is delayed, the separation zone decrease, further confirming its effective control of the flow of the separation of the wings. After averaging multiple transient images, the aerofoil flow field control is calculated the front and rear turbulent kinetic energy k (turbulent kinetic energy), and the thunder nor shear stress τ, defined as follows: $$\:k=\frac{1}{2\left(\stackrel{-}{{u}^{{\prime\:}2}}+\stackrel{-}{{v}^{{\prime\:}2}}+\stackrel{-}{{w}^{{\prime\:}2}}\right)}$$ 8 $$\:I=\frac{{u}^{{\prime\:}}}{U}\approx\:\frac{1}{U}\sqrt{\frac{2k}{3}}$$ 9 $$\:\tau\:=-{u}^{{\prime\:}}{v}^{{\prime\:}}/{U}_{\infty\:}^{2}$$ 10 In the uncontrolled state (refer to Figs. 10 (a) and 10(a)), the upper aerofoil is approximately A strong turbulent kinetic energy band appears 0.4 g downstream from the leading edge, as shown by the dotted line in the figure region, and the (Re s )Reynolds shear stress in this region is also larger. (a) It can be seen that flow separation occurs in this area, indicating that the airflow in the separation area is turbulent. The speed pulsation is large and the flow is unstable, which adversely affects the aerodynamic performance of the airplane. After synthetic jet control (refer to Figs. 10 (a) and 10(b)), the strong separation shear layer disappears and is replaced by turbulent kinetic energy near the wall increases, the shear stress increases. Near the jet outlet due to the jet and the mainstream strong shear occurs during the mixing process, so the turbulent kinetic energy and Reynolds shear Increased shear stress; zone control downstream of jet outlet to separation point. There is basically no change before and after; x = 0.4k downstream, the turbulent boundary layer is formed, the energy of the bottom layer is increased, and the ability to resist the adverse pressure gradient is improved, which is consistent with Section 2.2 . The boundary layer velocity profile measurements are consistent. The original separation air flow formed reattach, the separation point is delayed backwards. Figure 10 shows that when the wing angle of attack α = 25°, no control and synthesis Pressure distribution curve of wing surface after jet control (Cn = 0.00125) Compared. It can be seen from the figure that if there is no control, the air on the upper surface of the wing when α = 25°. The flow will be completely separated and a stall will occur; while after the synthetic jet is controlled, the boundary layer. The energy is enhanced, the separation is effectively improved, and the leading edge of the wing has been eliminated. The lost suction peak reappears. Variation of pressure distribution on aerofoil surface it leads to the change of the macro aerodynamic force of the wing. Increased aerofoil maximum lift coefficient and delayed stall angle of attack. 3.4 Effect of Blowing rate ( Q՛ ) This section presents aeroacoustics studies concerning noise generated by turbulence–leading edge interactions, attributed to a freestream turbulence intensity of 4.5%. Figure 11 compares the sound power level PWL generated by the baseline aerofoil (i.e., Cn = 0-0.00375) with that of the blowing leading edge under the condition of Cn = 0.00375. ΔPWL denotes the disparity in noise spectra between the baseline and the blowing leading edges, expressed as a function of non-dimensional frequency (f.C/U∞). A positive result of ΔPWL indicates greater noise reduction compared to λ′10, and conversely. The jet velocity is U ∞ = 25 m/s, and the angle of attack is zero degrees. The background noise of the wind tunnel, when measuring far-field noise radiation related to the naked jet, is considerably lower than when the aerofoil is included. This verifies the elevated signal-to-noise ratio of the aeroacoustics findings in the present investigation. The broadband characteristic resulting from the turbulence interaction is easily observable at the low to mid-frequency ranges when analysing the PWL spectra generated by the baseline aerofoil. It is also hypothesised that the increased freestream turbulence strength triggered a bypass transition on the surface of the aerofoil, given there is no sign of the unstable tone noise in the spectra. The insertion of the blowing leading edge is clearly helpful, with a large reduction in the turbulent broadband noise (up to 9 dB in some circumstances). Interestingly, whereas the level of broadband noise reduction rises with Cn between 0.00125–0.00375 as illustrated in the ΔPWL spectra, further increase of Cn actually becomes deleterious. For example, no detectable noise reduction can be accomplished at Cn 0.0025. As a result, the results highlight that there is an ideal blow rate for cancelling out incoming freestream turbulence and lowering interaction broadband noise. However, based on our available data set, it can be tentatively observed that the condition of Cn for an effective reduction of the interaction broadband noise does not apply for greater freestream velocity (e.g. U ∞ = 45 m/s). Additionally, self-noise from the air jets from the leading edge orifices is a contributing factor to the increase in high frequency noise, usually when f > 10. This undesirable source of noise is outside the desired frequency range, but it is still undesirable. Moreover, it becomes less of a concern at higher freestream velocity, because of the increased degree of overall aerofoil noise radiation. Figure 12 illustrates the contour graphs of ∆ PWL as a function of f and U ∞ subjected to 0.00125 Cn 0.00375. Despite noise reduction typically occurring within a comparable frequency range, the example with the smaller value higher performance and a broader range of effective freestream velocity. The outcomes illustrated in Fig. 12 predominantly align with the ∆ PWL trend presented in Fig. 11 . The maximum broadband noise reduction across an extensive frequency range can be achieved with the optimal blowing leading edge. It is important to notice that noise levels rise in certain instances, particularly at non-dimensional frequencies beyond 8. The findings indicate that the Cn = 0.001 scenario cannot attain noise radiation across the full velocity spectrum, but exhibits only marginal enhancements at lower frequencies, as illustrated in Fig. 12 (d-e). It is noteworthy that an increase in noise at elevated frequencies is evident in all instances. Conclusions A kind of inclined outlet synthetic jet exciter is designed and developed, which can the NACA 633618-il straight wing model can effectively control the separation flow; The interaction characteristics of the synthetic jet and the mainstream were tested by PIV flow field and the Boundary layer scanning was used to analysis the flow control mechanism. Get the following conclusion Argument: (a) The inclined outlet synthetic jet exciter can effectively control the wing surface Surface flow separation. Within the experimental range of R, the higher the energy of the jet is than Cn is larger, the better the control effect is; when Cn = 0.0025, the maximum lift of the aerofoil is the coefficient is increased by 10.6%, and the stall angle of attack is delayed by 4°; (b) The synthetic jet at the oblique outlet is transported to the mainstream by the lateral momentum injecting energy to enhance the shear stress of the bottom layer of the boundary layer and lift the vicinity of the wall. Turbulent kinetic energy, increased underlying boundary layer energy, ability to resist adverse pressure gradients increase; the separation point is delayed backwards, the separation airflow is reattached, and the separation zone is turbulent flow energy drops. (c)The leading edge jet consistently counteracts and potentially reduces entering turbulent eddies by either changing the leading edge stagnation point of the aerofoil or establishing a "buffer zone" surrounding the airfoil's leading edge. It appears that one or both of these mechanisms may exhibit significant sensitivity to the blowing volume flow rate, which correlates with the exit jet velocity. A pressure difference of up to 7 dB can be attained by one of the leading edge blowing configurations when Cn = 0.0025. Declarations Author Declarations Conflict of Interest The authors have no conflicts to disclose. Data Availbility The data that supports the findings of this study are available within the article. Funding Information This research was supported by the by University London and Office of the Prime Minister of Iraq Higher Committee for Education Development in Iraq (HCED) Author Contribution Declaration of interests☒ The authors declare that they have no known competing financial interests or personal relationships that could have appeared to influence the work reported in this paper.☐The authors declare the following financial interests/personal relationships which may be considered as potential competing interests: Acknowledgments The authors highly appreciated the anonymous referees' insightful remarks and recommendations, which significantly enhanced the original manuscript. The Higher Committee for Education Development in Iraq (HCED) and the University of Baghdad. Any opinions, research, or suggestions are those of the authors and do not necessarily represent those of the supporting organizations. 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Flow-fieldcharacter- izationofanovelsyntheticjetflowcontrolators[J]. Acta Aerodynamica Sinica, 2007, 25(12): 107–113. (2007) Rumsey, C.L., Gatski, T.B., Sellers, W.L.: Summary of the 2004 CFD validation workshop synthetic jets and turbulent separa tion control[R]. AIAA 2004–2217. Hao, Lishu, Qiao Zhide:. Research on synthetic jet for airfoil separation flow control [J]. Journal of Northwestern Polytechnical University, 24(4): 528–531. Hao Lishu, Qiao Zhide. Maximizingtheeffectofsyntheticjet Onairfoilseparationflowcontrol[J]. JournalofNorthwestern Polytechnical University, 2006, 24(4): 528–531. (2006) Specifications, C.: Acceptable Means of Compliance for Large Aeroplanes CS-25. European Aviation Safety Agency, Amendment 24.10 (2020) Zhenbing, L., Bopeng, Z.: Xia Zhixun. Effect of Synthetic Jet Exciter on Jet Vector [J]. Propulsion Technology, 25(5): 405–410. Luo Zhenbing, Zhu Bopeng, Xia Zhixun, Jet vectoringu- sing synthetic jet actuators[J]. Journal of Propulsion Technology, 2004, 25(5): 405–410. (2004) Xiao, X., et al.: Photovoltaic data cleaning based on interpolation and Pearson correlation. Inform. Technol. 5 , 19–22 (2019) Gross, J., Lance, W.: Traub. Experimental and theoretical investigation of ground effect at low reynolds numbers. J. Aircr. 49 (2), 576–586 (2012) Gupta, S., et al.: CFD simulation of turbulent flow around multi-element airfoil. AIP Conference Proceedings. Vol. 2204. No. 1. AIP Publishing, (2020) Plint, Partners, L.T.D., Engineers: Operating Instructions TE.81/D, s.l. PLINT (1986) Additional Declarations No competing interests reported. Cite Share Download PDF Status: Posted Version 1 posted You are reading this latest preprint version Research Square lets you share your work early, gain feedback from the community, and start making changes to your manuscript prior to peer review in a journal. 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Also discoverable on Platform About Our Team In Review Editorial Policies Advisory Board Help Center Resources Author Services Accessibility API Access RSS feed Manage Cookie Preferences © Research Square 2026 | ISSN 2693-5015 (online) Privacy Policy Terms of Service Do Not Sell My Personal Information {"props":{"pageProps":{"initialData":{"identity":"rs-5769420","acceptedTermsAndConditions":true,"allowDirectSubmit":true,"archivedVersions":[],"articleType":"Short Report","associatedPublications":[],"authors":[{"id":398360783,"identity":"e91db2b0-6c8b-4846-b3e9-ee7db4e7f70b","order_by":0,"name":"Yasir Al-okbi","email":"data:image/png;base64,iVBORw0KGgoAAAANSUhEUgAAAZAAAAAyAQMAAABI0h/eAAAABlBMVEX///8AAABVwtN+AAAACXBIWXMAAA7EAAAOxAGVKw4bAAABC0lEQVRIiWNgGAWjYJCCA0Asx8DAw8DA2MAM5kqAEQEtxqRpAYHEBjQtuIHB8eMPD1cw2KVvON578OPXHdbyfAeYD97mYbCQbcCl5UyOwcEzDMm5G86cS5aWPZNuOPMAW7I1D4OEMU4tB3IYDjYwMOduuJFjIC3ZdphxwwEeM2mglkScWs4/fwDUUp9ucCPH+DdQi/2GA/zf8Gu5kWAA1HI4AajFTPJj2+FEoC1seLVI3ngD1GJw3HDmmXNp1oxt6ckzD7MZW84xwO0XvvPpjz82VFTL8x3vPXzzZ5u1bd/x5oc33lTU4QwxhQNg50E4zDxgEiLCiEuLPLIE4w9kDi4to2AUjIJRMOIAAM+0YJR8dUe/AAAAAElFTkSuQmCC","orcid":"","institution":"University of Florida | USA, Herbert Wertheim College of Engineering","correspondingAuthor":true,"prefix":"","firstName":"Yasir","middleName":"","lastName":"Al-okbi","suffix":""},{"id":398360784,"identity":"51f8bdf7-1466-4ee7-b3f0-e02297c9b355","order_by":1,"name":"Mohammed A. Atiya","email":"","orcid":"","institution":"University of Baghdad","correspondingAuthor":false,"prefix":"","firstName":"Mohammed","middleName":"A.","lastName":"Atiya","suffix":""},{"id":398360785,"identity":"346178cd-e949-4d57-ae4c-8c7190fc0641","order_by":2,"name":"Qusay Rasheed Al-amir","email":"","orcid":"","institution":"Al-Mustaqbal University","correspondingAuthor":false,"prefix":"","firstName":"Qusay","middleName":"Rasheed","lastName":"Al-amir","suffix":""},{"id":398360786,"identity":"870f2caf-d9b7-4777-b2ea-48f4a1d9750b","order_by":3,"name":"Muhsin Jaber Jweeg","email":"","orcid":"","institution":"Al-Farahidi University, Al Jadriyah Bridge","correspondingAuthor":false,"prefix":"","firstName":"Muhsin","middleName":"Jaber","lastName":"Jweeg","suffix":""},{"id":398360787,"identity":"0f0b870d-fb2a-4ed1-84a0-a36cf11cd63c","order_by":4,"name":"Rana Al-Dujele","email":"","orcid":"","institution":"Al-Nahrain University","correspondingAuthor":false,"prefix":"","firstName":"Rana","middleName":"","lastName":"Al-Dujele","suffix":""}],"badges":[],"createdAt":"2025-01-05 20:53:11","currentVersionCode":1,"declarations":"","doi":"10.21203/rs.3.rs-5769420/v1","doiUrl":"https://doi.org/10.21203/rs.3.rs-5769420/v1","draftVersion":[],"editorialEvents":[],"editorialNote":"","failedWorkflow":false,"files":[{"id":73270495,"identity":"536297dc-4257-4ee2-b951-86f443dd93bb","added_by":"auto","created_at":"2025-01-08 10:53:01","extension":"png","order_by":1,"title":"Figure 1","display":"","copyAsset":false,"role":"figure","size":70274,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eNACA 633618-il model.\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"1.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/1ef22969a6a5e5d6bdc20d4a.png"},{"id":73269607,"identity":"39a84fb1-699b-47d9-8362-7a8868571537","added_by":"auto","created_at":"2025-01-08 10:45:01","extension":"png","order_by":2,"title":"Figure 2","display":"","copyAsset":false,"role":"figure","size":168203,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eExperimental PIV setup.\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"2.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/8b8045a109052c92e663699b.png"},{"id":73269613,"identity":"72831b2a-bc0d-4d81-a500-179f761b0603","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":3,"title":"Figure 3","display":"","copyAsset":false,"role":"figure","size":98401,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eExperimental PIV measurements.\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"3.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/a56bf6178e1143e0f985dee7.png"},{"id":73269610,"identity":"cc3001fb-3f77-4266-9726-177b5e3d129b","added_by":"auto","created_at":"2025-01-08 10:45:01","extension":"png","order_by":4,"title":"Figure 4","display":"","copyAsset":false,"role":"figure","size":381411,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eShows a three-force balance system used to quantify lift and drag.\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"4.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/f5493af041a66367abab9d76.png"},{"id":73270496,"identity":"e3e02185-5790-4314-8e2d-e5d0ec529c00","added_by":"auto","created_at":"2025-01-08 10:53:02","extension":"png","order_by":5,"title":"Figure 5","display":"","copyAsset":false,"role":"figure","size":387090,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eNozzle, aerofoil model and polar array in the large anechoic chamber.\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"5.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/d90afbda98f7ceb0e75cbb92.png"},{"id":73269639,"identity":"1188e3a0-bfd3-4b28-aa97-67232c8acb14","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":6,"title":"Figure 6","display":"","copyAsset":false,"role":"figure","size":25482,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eLift coefficient vs. angle of attack\u003c/strong\u003e \u003cstrong\u003eat Re = 1.85 x 10\u003c/strong\u003e\u003csup\u003e\u003cstrong\u003e5\u003c/strong\u003e\u003c/sup\u003e\u003c/p\u003e","description":"","filename":"6.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/36372eaa025f928284cbf2e1.png"},{"id":73270503,"identity":"0fa22609-a98d-4606-92ca-dad878da1ac5","added_by":"auto","created_at":"2025-01-08 10:53:02","extension":"png","order_by":7,"title":"Figure 7","display":"","copyAsset":false,"role":"figure","size":28676,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eDrag coefficient vs. angle of attack\u003c/strong\u003e \u003cstrong\u003eat Re = 1.85 x 10\u003c/strong\u003e\u003csup\u003e\u003cstrong\u003e5\u003c/strong\u003e\u003c/sup\u003e\u003c/p\u003e","description":"","filename":"7.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/f80a7c476e28ffb3b368fa50.png"},{"id":73269651,"identity":"e502abe6-ba99-4254-bd85-20694e693139","added_by":"auto","created_at":"2025-01-08 10:45:03","extension":"png","order_by":8,"title":"Figure 8","display":"","copyAsset":false,"role":"figure","size":31211,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003ePressure distribution NACA 633618-il with angle of attack\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"8.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/f71c8dd8cd98d5570d48f61a.png"},{"id":73269611,"identity":"e07ab7d1-1916-4075-93aa-60cb53d0ff67","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":9,"title":"Figure 9","display":"","copyAsset":false,"role":"figure","size":53648,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eBoundary layer velocity profile with and without jet\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"9.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/c55a899e14d42577f687f8b3.png"},{"id":73269624,"identity":"08870f18-9550-4ff6-be27-bcf4fd7a3207","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":10,"title":"Figure 10","display":"","copyAsset":false,"role":"figure","size":310850,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eInstantaneous PIV measurement results\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"10.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/5bd68fee9d7b276dd4ce4562.png"},{"id":73269619,"identity":"570231e3-86c4-46d6-87c3-3bea20582bd8","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":11,"title":"Figure 11","display":"","copyAsset":false,"role":"figure","size":136718,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eComparison of PWL and ΔPWL as a function of normalised frequency\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"11.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/7e611a6c93b24ce8ae26a59d.png"},{"id":73269618,"identity":"060b88cb-6886-4869-b0b2-7cb5c783b9f4","added_by":"auto","created_at":"2025-01-08 10:45:02","extension":"png","order_by":12,"title":"Figure 12","display":"","copyAsset":false,"role":"figure","size":131889,"visible":true,"origin":"","legend":"\u003cp\u003e\u003cstrong\u003eComparison of ∆PWL as a function of normalised frequency\u003c/strong\u003e\u003cem\u003e\u003cstrong\u003e f\u003c/strong\u003e\u003c/em\u003e\u003cstrong\u003e \u0026nbsp;and \u003c/strong\u003e\u003cem\u003e\u003cstrong\u003eU∞\u003c/strong\u003e\u003c/em\u003e\u003cstrong\u003e when \u003c/strong\u003e\u003cem\u003e\u003cstrong\u003eTu\u003c/strong\u003e\u003c/em\u003e\u003cstrong\u003e = 4.5%, under Cn = 0.00375 (a), Cn=0.0025 (b), Cn = 0.0020 (c), Cn = 0.00175 (d) \u0026nbsp;and Cn = 0.0010025 (e).\u003c/strong\u003e\u003c/p\u003e","description":"","filename":"12.png","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/a8c7f0e106b05b73f264fed6.png"},{"id":81990439,"identity":"eec7c43d-8734-49fb-ae7d-16c7f1abe2d2","added_by":"auto","created_at":"2025-05-05 16:31:43","extension":"pdf","order_by":0,"title":"","display":"","copyAsset":false,"role":"manuscript-pdf","size":2889856,"visible":true,"origin":"","legend":"","description":"","filename":"manuscript.pdf","url":"https://assets-eu.researchsquare.com/files/rs-5769420/v1/14398849-3e29-4121-a2d1-1b8ba1f15706.pdf"}],"financialInterests":"No competing interests reported.","formattedTitle":"Experimental Investigation of effect Slotted leading Edge Circulation Control Aerofoil Aerodynamic and Acoustics","fulltext":[{"header":"1 Introduction","content":" \u003cp\u003eThe flying wing layout integrates the fuselage/wing/tail and the fuselage/propulsion. The integration is highly integrated, and the horizontal tail and vertical tail are cancelled to minimise the adverse aerodynamic interference and radar scatter area of the small wing body/airframe/propulsion, significantly improving the cruise aerodynamic performance and enhancing the stealth combat capability of the new bombers, sensor aircraft, long-range, long-endurance drones, and stealth reconnaissance. Limited research has examined the temporal dynamics of a Coandă-type flow actuator included into a clean aerofoil, designed for GLA or MLA applications. Unsteady Reynolds-Averaged Navier-Stokes simulations\u003c/p\u003e \u003cp\u003eAn analysis of a symmetrical aerofoil indicated that sufficient activation of a Coandă jet can result in nearly constant lift when subjected to a medium-amplitude 1-cos gust [\u003cspan citationid=\"CR1\" class=\"CitationRef\"\u003e1\u003c/span\u003e]. Unsteady 2D-RANS simulations of a supercritical aerofoil under varying operating conditions shown that an impulsive activation of a Coandă jet can achieve lift reduction amplitudes within a response time enough for counteracting the briefest 1-cos gusts [\u003cspan citationid=\"CR2\" class=\"CitationRef\"\u003e2\u003c/span\u003e]. The rapid emergence of actuator-induced lift was additionally demonstrated by an experimental investigation of a cambered aerofoil at low Reynolds numbers. The low slot blowing demonstrated significant control authority and a notably rapid reaction time, effectively mitigating gust-induced loads [\u003cspan citationid=\"CR3\" class=\"CitationRef\"\u003e3\u003c/span\u003e]. The conventional aileron, a jet normal to the airfoil's surface (normal jet), and a tangential jet over a rounded trailing edge (Coandă jet) are the only actuation designs that combine high peak lift reductions with slight changes in drag and pitching moment over the entire flight envelope, according to two-dimensional Reynolds-Averaged Navier-Stokes (2D-RANS) simulations [\u003cspan citationid=\"CR4\" class=\"CitationRef\"\u003e4\u003c/span\u003e].\u003c/p\u003e \u003cp\u003eThe design of advanced aircraft such as drones is of enormous significance [\u003cspan citationid=\"CR5\" class=\"CitationRef\"\u003e5\u003c/span\u003e\u0026ndash;\u003cspan citationid=\"CR6\" class=\"CitationRef\"\u003e6\u003c/span\u003e]. But the design features of the wing-body integrating the lifting surface and the no-tail layout are also important features of the flying wing fabric. Adverse effect on the flight quality of the Bureau's aircraft\u0026mdash;heading stability Insufficient, poor longitudinal stability characteristics, etc., all of the above will improve the flight control system. There are extremely high design requirements [\u003cspan citationid=\"CR7\" class=\"CitationRef\"\u003e7\u003c/span\u003e\u0026ndash;\u003cspan citationid=\"CR8\" class=\"CitationRef\"\u003e8\u003c/span\u003e]. Due to the need for high lift during takeoff and landing, large aircraft often have to make the wing in a state of high angle of attack, and it is difficult to separate the flow of the aerofoil at this time. to avoid. Flow separation results in increased resistance, reduced rudder efficiency, and a series of problems such as body vibration, lift drop, and even stall, which seriously affect the safety and manoeuvrability of the aircraft during flight. Therefore, researchers have been looking for a variety of efficient and simple means to divide the wing-controlled flow. The passive control method has no external energy consumption, which is convenient for Reliable installation and performance, such as vortex generators, have been widely used in aircraft use. Active control means, by virtue of which it can fine-tune the moving fluid The advantages of accurate phase control and adjustment according to actual working conditions are obtained. People are paying more and more attention, such as when Boeing is researching and applying hybrid laminar flow control technology to the B787-9.\u003c/p\u003e \u003cp\u003eAs a new type of active flow control technology, synthetic jet has There are many advantages, such as no air source, compact structure, low cost, etc., so it quickly became a hot topic in the field of flow control. Its development can be traced back to 20. The discovery of the acoustic rectification effect in the middle and late nineteenth century [\u003cspan citationid=\"CR9\" class=\"CitationRef\"\u003e9\u003c/span\u003e\u0026ndash;\u003cspan citationid=\"CR10\" class=\"CitationRef\"\u003e10\u003c/span\u003e]. A high-amplitude oscillating sound wave was used to obtain radiation with a maximum speed of about 10 m/s. Acoustic rectification, which was observed during sound waves controlling the separation flow. Subsequently, through software numerical simulation, experimental and theoretical analysis at home and abroad, A large number of studies have been carried out on the mechanism of synthetic jet technology [\u003cspan citationid=\"CR11\" class=\"CitationRef\"\u003e11\u003c/span\u003e\u0026ndash;\u003cspan citationid=\"CR12\" class=\"CitationRef\"\u003e12\u003c/span\u003e]. Exist On the basis of mechanism research, people have made a lot of research on synthetic jet technology. A large number of experimental investigations have been carried out to develop its engineering application potential. synthetic shot Streaming has a wide range of applications: it can effectively delay separation, delay the stall [\u003cspan citationid=\"CR13\" class=\"CitationRef\"\u003e13\u003c/span\u003e\u0026ndash;\u003cspan citationid=\"CR14\" class=\"CitationRef\"\u003e14\u003c/span\u003e], so as to greatly increase the lift and reduce the drag, improve flight performance, achieve thrust vector control [\u003cspan citationid=\"CR15\" class=\"CitationRef\"\u003e15\u003c/span\u003e], increase strong mixing, and improve the control force of the MAV [\u003cspan citationid=\"CR16\" class=\"CitationRef\"\u003e16\u003c/span\u003e]; body vortex and helicopter rotor dynamic stall control, etc. These experimental studies This study lays a robust foundation for further research. The PIV results of the flow field at the inclined outlet of the synthetic jet. After the analysis, it is pointed out that the oblique outlet exciter can make the surrounding gas directional. energy and mass transport control, convective flow fields appear as transverse flow along the wall movement transport characteristics [\u003cspan citationid=\"CR17\" class=\"CitationRef\"\u003e17\u003c/span\u003e]. The influence of the inclination angle of the outflow outlet on the flow control effect has been studied. The oblique outlet exciter applied to the S-shaped intake port [\u003cspan citationid=\"CR18\" class=\"CitationRef\"\u003e18\u003c/span\u003e]. In separation flow control, a good control effect is achieved. This paper is based on previous research. On the basis of research, the flow of the synthetic jet on the wing surface using the inclined outlet. The separation is controlled by focusing on the interaction between the jet and the main flow by means of PIV. The measurement of the aerofoil flow field after the action, while combining the jet to control the front and rear aerofoil, Variation of pressure distribution and boundary layer velocity profile, control of oblique outlet exciter. A preliminary study on the mechanism of flow separation was carried out. The current study aims to develop an actuator that has a high enough lift control authority to mitigate typical gust-induced loads while having negligible affects on the aerofoil's baseline performance. A low-speed wind tunnel was used to test a subsonic and modular wind tunnel model. This work focusses on both the actuation system's unsteady performance for impulsive switching between upper and lower blowing and the steady actuator performance attained for slot blowing.\u003c/p\u003e"},{"header":"2. Experimental and Physical Models","content":"\u003cdiv id=\"Sec3\" class=\"Section2\"\u003e \u003ch2\u003e2.1 Aerofoil model\u003c/h2\u003e \u003cp\u003eThe wing model adopts the NACA 633618-il aerofoil, an all-metal, straight machine. Wing, wing chord length C\u0026thinsp;=\u0026thinsp;300 mm, span length S\u0026thinsp;=\u0026thinsp;495 mm, aspect ratio Ar\u0026thinsp;=\u0026thinsp;2.3. At a distance of 0.38c from the leading edge of the aerofoil, there are 6 uniformly distributed in the spanwise direction. Independent exciters with 60mm spacing. The exciter adopts acoustic excitation vibration. Moving diaphragm, power 5W; tangential angle between jet outlet and aerofoil 32\u0026deg;, slit length SL\u0026thinsp;=\u0026thinsp;15mm, width W\u0026thinsp;=\u0026thinsp;1.0mm. The edge at the half-span position of the wing A total of 60 pressure measuring holes are opened around the chord to measure the surface pressure of the aerofoil force distribution. The exit velocity of the synthetic jet exciter is mainly determined by the external excitation frequency Y and power amplifier drive voltage Y control. In this experiment, the total pressure probe was used first. The velocity characteristics of the jet were calibrated, and it was found that the fixed voltage R was Under the resonant frequency f\u0026thinsp;=\u0026thinsp;300 Hz, the outlet velocity is the largest; under the rate n, the outlet velocity uo increases with the increase in the voltage v. The concept of jet-to-main flow ratio proposed by Smith [\u003cspan citationid=\"CR19\" class=\"CitationRef\"\u003e19\u003c/span\u003e] combines the exciter and the export velocity, which is dimensionless:\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eThe middle shape of the above formula is the number of exciters, u\u003csub\u003eo\u003c/sub\u003e is the jet outlet velocity, and b is the length and width of the exit slit, respectively.\u003c/p\u003e \u003cp\u003eThe non-dimensional jet excitation frequency is determined by:\u003cdiv id=\"Equ1\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ1\" name=\"EquationSource\"\u003e\n$$\\:{F}^{+}=\\frac{\\frac{{w}_{jet}}{2\\pi\\:}}{\\frac{{U}_{\\infty\\:}}{c}}=\\frac{{w}_{jet}c}{2\\pi\\:{U}_{\\infty\\:}}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e1\u003c/div\u003e\u003c/div\u003e\u003c/p\u003e \u003cp\u003eThe non-dimensional jet momentum coefficient is determined by:\u003cdiv id=\"Equ2\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ2\" name=\"EquationSource\"\u003e\n$$\\:{C}_{u0}=2\\frac{h}{c}{\\left(\\frac{{U}_{0}}{{U}_{\\infty\\:}}\\right)}^{2},\\:{C}_{um}=2\\frac{h}{c}{\\left(\\frac{{U}_{m}}{\\sqrt{2}{U}_{\\infty\\:}}\\right)}^{2}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e2\u003c/div\u003e\u003c/div\u003e\u003c/p\u003e \u003cp\u003eWhere c is the speed of sound and \u003cem\u003eU\u003c/em\u003e\u003csub\u003e\u003cem\u003em\u003c/em\u003e\u003c/sub\u003e and \u003cem\u003eU\u003c/em\u003e\u003csub\u003e\u003cem\u003e0\u003c/em\u003e\u003c/sub\u003e are the amplitudes of the steady and unsteady components of velocity, respectively.\u003c/p\u003e \u003cp\u003eFurthermore, a modification of the original pressure boundary condition near the jet orifice is required. The conservation of normal momentum can be used to determine the pressure boundary condition. The boundary condition can be changed as follows by ignoring the impact of viscosity and taking into account the influence of the instantly changing normal velocity component \u003cem\u003eU\u003c/em\u003e\u003csub\u003e\u003cem\u003en\u003c/em\u003e\u003c/sub\u003e:\u003cdiv id=\"Equ3\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ3\" name=\"EquationSource\"\u003e\n$$\\:\\frac{\\partial\\:p}{\\partial\\:\\eta\\:}=-\\rho\\:\\frac{\\partial\\:{U}_{n}}{\\partial\\:t}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e3\u003c/div\u003e\u003c/div\u003e\u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec4\" class=\"Section2\"\u003e \u003ch2\u003e2.2 Force measuring system and pressure measuring system\u003c/h2\u003e \u003cp\u003eThe experiment was conducted in the open type of the Department of Aerodynamics, Nanjing University of Aeronautics and Astronautics. A low-speed reflow wind tunnel. The wind tunnel has low turbulence, low noise, etc. The size of the experimental section is 0.5 m \u0026times; 0.5 m, and the turbulence degree is 0.045%. The minimum stable wind speed of the cave is 2 m/s, and the maximum wind speed is 45 m/s. This experiment is under the wind velocity of v\u003csub\u003e\u0026infin;\u003c/sub\u003e = 11 m/s; the variation range of the angle of attack α is -2\u0026deg;\u0026ndash;30\u0026deg;. The Reynolds number (Re) based on the aerofoil chord length is Re\u0026thinsp;=\u0026thinsp;2.5\u0026times;10\u003csup\u003e5\u003c/sup\u003e. The force measuring system consists of a 6-component box-type pneumatic balance, a signal amplification device, a 16-bit data acquisition card, an acquisition control computer, and a special test software composition. After the transformation of the body/wind axis system, each component of the model is finally obtained, amounting to aerodynamic force and torque. The aerofoil surface pressure test system consists of a 60-channel differential pressure transmitter sensor, a 16-bit data acquisition card, an acquisition control computer, and special processing management software composition. The sensor range is 0.35 PSI, and the test system is integrated; the test accuracy is 0.06% FS.\u003c/p\u003e \u003cp\u003eThe boundary layer probe scanning system is self-developed by Brunel University London of Aeronautics and Aerodynamic Lap. It is mainly composed of a boundary layer probe, an optical coordinate frame, and a pressure sensor. It is composed of several parts, such as an acquisition board and acquisition and processing software, which can realise the probe's arbitrary, precise, and subtle changes in position in three-dimensional space. Probe ruler Small inch (thickness less than 0.3 mm), flat shape, can be approached at close range; the surface of the aerofoil has little interference to the flow field.\u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec5\" class=\"Section2\"\u003e \u003ch2\u003e2.3 Particle image velocimetry of the flow field test system\u003c/h2\u003e \u003cp\u003eThe PIV flow field test system used in the experiment is the American TSI company. The company's two-dimensional PIV mainly includes double-pulse Nd:YAG lasers, mutual related CCD cameras, synchronisers, tracer particles, operating systems, etc. That The maximum output power of the medium laser is 300 mJ/pulse, and the pulse duration is the pulse interval is 10 ns; the pulse interval is adjustable, and the repetition frequency of the double-pulse laser is 15 Hz. The laser scanning plane is 0.60 between the upper surface of the wing and the root of the wing. The cross section at the doubled length, the chordwise position range is from l\u0026thinsp;=\u0026thinsp;0.30c to l\u0026thinsp;=\u0026thinsp;0.70c, as shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig3\" class=\"InternalRef\"\u003e3\u003c/span\u003e.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec6\" class=\"Section2\"\u003e \u003ch2\u003e2.4 Three-component balance\u003c/h2\u003e \u003cp\u003ePlint \u0026amp; Partners LTD employed a three-component balance to assess the aerodynamic forces generated by the aerofoil within the wind tunnel (Fig.\u0026nbsp;\u003cspan refid=\"Fig4\" class=\"InternalRef\"\u003e4\u003c/span\u003e). The balance is made up of a force plate and an aluminium mounting plate that is secured to the wind tunnel working portion. The force plate is able to revolve about the horizontal axis. The aerofoil model is attached to the force plate's model support using a \u0026Oslash;12 mm mounting stem. This support is free to rotate for adjustment of the angle of attack, while its position may be fixed by means of the incidence clamp. The forces acting on the force plate are conveyed by flexible cables to strain gauge load cells, which quantify the lift and the drag forces. The drag cable is horizontal and goes through the centre of the force plate stem, whilst the two lift cables are vertical and equidistant from the model support. Pitching moment is determined by multiplying the difference between fore and aft by 0.127 [\u003cspan citationid=\"CR20\" class=\"CitationRef\"\u003e20\u003c/span\u003e]. The readings are accurate to within \u0026plusmn;\u0026thinsp;0.05 N. To ensure that the readings are accurate, the three-component balance is calibrated before each measurement. Calibration is performed by taking it from the wind tunnel and placing it on a mounting frame. Before applying a load, the load cells' zero readings are checked. The cells are adjusted by setting the adjustment screw until the display shows a value of zero. The calibration technique comprises the application of known lift and drag forces using dead weights. It is crucial that the force balance is completely balanced both vertically and horizontally during the calibration procedure. Any differences given from the forces are addressed by modifying the cable tensions until equivalent readings for both load components, aft and fore, are reached. Three calibrations are done to assess the linearity of the relationship between load and cell output.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eMicrophone Acoustic Measurements Metric and Polar Array An array of eight G.R.A.S 0.5-inch microphones was positioned 1.2 meters from the leading edge of the aerofoil in a circular arc to assess far-field noise radiation, as illustrated in Fig.\u0026nbsp;\u003cspan refid=\"Fig4\" class=\"InternalRef\"\u003e4\u003c/span\u003e. The microphone signals were recorded at 44 kHz for 20 seconds using a 16-bit National Instrument Analogue-Digital card PXI 1042, with a sampling frequency of 50 kHz. The noise spectra were computed with a window size of 1024 data points, yielding a frequency resolution of 48.83 Hz and a BT product of approximately 500, ensuring minimal variance in the spectral estimate. The acoustic pressure at the microphone locations was documented at mean flow velocities (\u003cem\u003eU\u003c/em\u003e\u003csub\u003e\u003cem\u003e\u0026infin;\u003c/em\u003e\u003c/sub\u003e) ranging from 20 to 60 m/s, respectively. Acoustic pressure spectra and directivity patterns can be computed within an azimuthal range of 90 degrees (45\u0026deg; \u0026le; θ\u0026thinsp;\u0026le;\u0026thinsp;135\u0026deg;, with θ\u0026thinsp;=\u0026thinsp;0\u0026deg; aligned with the jet axis). The data acquisition is managed using a PC linked by an NI 8360 express card, and is then exported to Matlab. Assuming a cylindrical spreading of sound waves from the aerofoil leading edge, the microphone array in the current research can be used to compute the sound power level (PWL) and overall sound power level (OAPWL).\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eThis research examines noise radiation through the Sound Power Level spectrum (PWL), as delineated in Eq.\u0026nbsp;4, and measured between radiation angles of 50\u0026deg; and 110\u0026deg; (refer to Fig.\u0026nbsp;\u003cspan refid=\"Fig5\" class=\"InternalRef\"\u003e5\u003c/span\u003e), which are also utilised to assess the noise reduction attained by various edge treatments.\u003c/p\u003e \u003cp\u003e \u003cdiv class=\"gridtable\"\u003e\u003ctable float=\"No\" id=\"Taba\" border=\"1\"\u003e \u003ccolgroup cols=\"2\"\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c1\" colnum=\"1\"\u003e\u003c/div\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c2\" colnum=\"2\"\u003e\u003c/div\u003e \u003cthead\u003e \u003ctr\u003e \u003cth align=\"left\" colname=\"c1\"\u003e \u003cp\u003e\u003cspan class=\"InlineEquation\"\u003e\u003cspan class=\"mathinline\"\u003e\\(\\:{P}_{xx}\\left(f\\right)=2\\pi\\:\\int\\:{W}_{xx}\\left(f,{\\theta\\:}_{i}\\right).\\varDelta\\:\\theta\\:/\\rho\\:c0\\)\u003c/span\u003e\u003c/span\u003e\u003c/p\u003e \u003c/th\u003e \u003cth align=\"left\" colname=\"c2\"\u003e \u003cp\u003e, where i\u0026thinsp;=\u0026thinsp;1, \u0026hellip;.,N (4)\u003c/p\u003e \u003c/th\u003e \u003c/tr\u003e \u003c/thead\u003e \u003ctbody\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003e\u003cspan class=\"InlineEquation\"\u003e\u003cspan class=\"mathinline\"\u003e\\(\\:PWL\\left(f\\right)=10{\\text{log}}_{10}\\left(\\frac{{P}_{xx}\\left(f\\right)}{{P}_{0}}\\right)\\)\u003c/span\u003e\u003c/span\u003e\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e, 50\u0026deg;\u0026lt; θ\u0026thinsp;\u0026lt;\u0026thinsp;110\u0026deg; (5)\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003c/tbody\u003e \u003c/colgroup\u003e \u003c/table\u003e\u003c/div\u003e \u003c/p\u003e \u003cp\u003eWhere P\u003csub\u003e0\u003c/sub\u003e\u0026thinsp;=\u0026thinsp;10\u0026thinsp;\u0026minus;\u0026thinsp;12 W/Hz, Wxx (f, θi) is the acoustic pressure PSD, measured at a polar angle Δθ, θ (rad) between adjacent radiation angles of the microphone, and Pxx (f) is the integrated sound power between the 50\u0026deg; and 110\u0026deg; radiation angles.\u003c/p\u003e \u003cp\u003eThe overall sound power level (OAPWL), which is a particular flow rate, can also be used to characterise aerofoil noise:\u003c/p\u003e \u003cp\u003e \u003cdiv class=\"gridtable\"\u003e\u003ctable float=\"No\" id=\"Tabb\" border=\"1\"\u003e \u003ccolgroup cols=\"2\"\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c1\" colnum=\"1\"\u003e\u003c/div\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c2\" colnum=\"2\"\u003e\u003c/div\u003e \u003ctbody\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003e\u003cspan class=\"InlineEquation\"\u003e\u003cspan class=\"mathinline\"\u003e\\(\\:OAPWL=10{\\text{log}}_{10}\\left({\\int\\:}_{f}^{}{P}_{xx}\\left(f\\right)df/{P}_{0}\\right)\\)\u003c/span\u003e\u003c/span\u003e\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e(6)\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003c/tbody\u003e \u003c/colgroup\u003e \u003c/table\u003e\u003c/div\u003e \u003c/p\u003e \u003cp\u003eTo find precise values for W_xx, a standard Br\u0026uuml;el \u0026amp; Kjaer 4231 calibrator is utilised to calibrate and supply a steady 94 dB signal at 1 kHz for 20 seconds.\u003c/p\u003e \u003c/div\u003e"},{"header":"3. Experimental results","content":"\u003cdiv id=\"Sec8\" class=\"Section2\"\u003e \u003ch2\u003e3.1 Discussion on the aerodynamic characteristics of the aerofoil\u003c/h2\u003e \u003cp\u003eFigure \u003cspan refid=\"Fig6\" class=\"InternalRef\"\u003e6\u003c/span\u003e shows the variation curve of the lift-drag coefficient with the angle of attack, which can be It can be seen that the NACA 633618-il straight wing has better slow stall characteristics. In the uncontrolled state, when α = 16°, the lift line is inclined, and the lift coefficient begins to decrease; when α = 18°, the lift coefficient reaches its maximum value. The force begins to decrease slowly, and until α = 25°, the lift decreases sharply, combined with the pressure distribution curve shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig7\" class=\"InternalRef\"\u003e7\u003c/span\u003e, which shows that the flow is completely separated.\u003c/p\u003e\u003cdiv id=\"Equ4\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ4\" name=\"EquationSource\"\u003e\n$$\\:Cn=\\frac{2\\sum\\:_{i=1}^{n}A{U}_{m}^{2}}{b{U}_{\\infty\\:}^{2}}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e7\u003c/div\u003e\u003c/div\u003e\u003cp\u003e\u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec9\" class=\"Section2\"\u003e \u003ch2\u003e3.2 Analysis of Boundary Layer of the Velocity\u003c/h2\u003e \u003cp\u003eOn the basis of obtaining effective control, further explore the synthetic jet First, measure and analyze the near-wall flow field in the upper aerofoil. Under the condition of α = 16°, l = 0.50c, and 0.65c on the wing surface, respectively, the measurement of the velocity profile of the boundary layer with or without synthetic jet control was carried out, and the results are shown in Fig.\u0026nbsp;9. The boundary layer velocities in different states and shape parameters of type, where: From the velocity type test results of the boundary layer on the upper wing surface, it can be seen that the excitation After the exciter is turned on (Cn = 0.0025), within a certain distance downstream, the speed type of the shape becomes fuller and the form factor becomes smaller. The bottom energy of the boundary layer is obtained; the ability to overcome the adverse pressure gradient is enhanced. This is also the dynamometer knot in Section \u003cspan refid=\"Sec3\" class=\"InternalRef\"\u003e2.1\u003c/span\u003e. After the synthetic jet is controlled, the stall angle of attack of the aerofoil is delayed, and the lift coefficient increases.\u003c/p\u003e \u003cp\u003eBased on achieving effective control, further explore the synthetic jet output. First, the near-wall flow field on the upper aerofoil must be measured and analysed to control the mechanism. In the state of α = 16°, x = 0.48 hu and 0.60 hu on the wing surface, respectively. The velocity pattern of the boundary layer controlled by a synthetic jet is measured at As shown in Fig.\u0026nbsp;9. Table\u0026nbsp;1 gives the boundary layer velocity under different conditions. It can be seen from the test results of the boundary layer velocity pattern on the upper wing of the wing that the exciting. After the actuator is turned on (Cn = 0.0025), the speed type shape becomes fuller and the form factor becomes smaller within a certain distance downstream. The bottom energy of the boundary layer is obtained to improve the ability to overcome adverse pressure gradients. This is also the force-measuring k not in Section \u003cspan refid=\"Sec3\" class=\"InternalRef\"\u003e2.1\u003c/span\u003e. As a result, after synthetic jet control, the wing stall angle of attack is delayed and the lift coefficient increases.\u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec10\" class=\"Section2\"\u003e \u003ch2\u003e3.3 Particle image velocimetry (PIV) flow field test\u003c/h2\u003e \u003cp\u003eTo further explore the interaction between the exciter synthetic jet and the mainstream. The process of using the PIV flow field test system at a typical angle of attack α = 25° The chordwise section of the upper wing surface at a distance of 0.59 from the aerofoil root.\u003c/p\u003e \u003cp\u003e \u003cem\u003el\u003c/em\u003e = 0.30c ~ 0.70c to measure the flow state of the two-dimensional flow field. Wind velocity \u003cem\u003eu\u003c/em\u003e\u003csub\u003e\u003cem\u003e∞\u003c/em\u003e\u003c/sub\u003e = 25m/s. Figure\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e shows that when α = 25°, with or without synthesis the averaged flow field velocity contour when the jet controls the upper surface of the aerofoil. It can be seen that the air flow on the wall of the front wing is controlled gradually along the flow direction decreases, and the surface airflow velocity becomes zero at about 0.50K, and even appears reverse flow velocity, the boundary layer is separated from the wing surface, refer to Fig.\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e(a); After control, the airflow velocity increases at the same chord position on the wing surface. Exist within the test field of view, the main flow reattached the wall and the separation zone was delayed after \u003cem\u003el\u003c/em\u003e = 0.70c (due to the limitation of the experimental conditions of the PIV instrument, the Measure the flow field after \u003cem\u003el\u003c/em\u003e = 0.70c). Figure\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e is a partially enlarged PIV velocity vector diagram. Uncontrolled under the condition of α = 25°, 0.45c ≤ \u003cem\u003el\u003c/em\u003e ≤ 0.65c air velocity on the aerofoil surface The degree direction is disordered, the velocity value is small, the backflow area appears, and the typical flow appears detached state. The comparison of multiple transient speed vector diagrams can also be found, the separation. The position of the point is not fixed, it is about 0.45c ≤ \u003cem\u003el\u003c/em\u003e ≤ 0.50c back and forth oscillation, see Figs.\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003ea-b. Due to the limited length of the article, other PIV test results at the moment are not listed here. Synthetic jet excitation after the device is turned on, within the field of view of 0.42c ≤ x ≤ 0.48c, the airflow is stable when the wall is fixed, the separation point is postponed to the outside of the field of view, as shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e(b). PIV transient time-in-time homogenization test results reflect control of the wing surface Favourable changes in the global flow field, the surface flow separation point is delayed, the separation zone decrease, further confirming its effective control of the flow of the separation of the wings. After averaging multiple transient images, the aerofoil flow field control is calculated the front and rear turbulent kinetic energy k (turbulent kinetic energy), and the thunder nor shear stress τ, defined as follows:\u003c/p\u003e\u003cdiv id=\"Equ5\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ5\" name=\"EquationSource\"\u003e\n$$\\:k=\\frac{1}{2\\left(\\stackrel{-}{{u}^{{\\prime\\:}2}}+\\stackrel{-}{{v}^{{\\prime\\:}2}}+\\stackrel{-}{{w}^{{\\prime\\:}2}}\\right)}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e8\u003c/div\u003e\u003c/div\u003e\u003cdiv id=\"Equ6\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ6\" name=\"EquationSource\"\u003e\n$$\\:I=\\frac{{u}^{{\\prime\\:}}}{U}\\approx\\:\\frac{1}{U}\\sqrt{\\frac{2k}{3}}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e9\u003c/div\u003e\u003c/div\u003e\u003cdiv id=\"Equ7\" class=\"Equation\"\u003e\u003cdiv format=\"TEX\" class=\"mathdisplay\" id=\"FileID_Equ7\" name=\"EquationSource\"\u003e\n$$\\:\\tau\\:=-{u}^{{\\prime\\:}}{v}^{{\\prime\\:}}/{U}_{\\infty\\:}^{2}$$\u003c/div\u003e\u003cdiv class=\"EquationNumber\"\u003e10\u003c/div\u003e\u003c/div\u003e\u003cp\u003e\u003c/p\u003e \u003cp\u003eIn the uncontrolled state (refer to Figs.\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e(a) and 10(a)), the upper aerofoil is approximately A strong turbulent kinetic energy band appears 0.4 g downstream from the leading edge, as shown by the dotted line in the figure region, and the (Re\u003csub\u003es\u003c/sub\u003e)Reynolds shear stress in this region is also larger.\u003c/p\u003e \u003cp\u003e(a) It can be seen that flow separation occurs in this area, indicating that the airflow in the separation area is turbulent. The speed pulsation is large and the flow is unstable, which adversely affects the aerodynamic performance of the airplane. After synthetic jet control (refer to Figs.\u0026nbsp;\u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e(a) and 10(b)), the strong separation shear layer disappears and is replaced by turbulent kinetic energy near the wall increases, the shear stress increases. Near the jet outlet due to the jet and the mainstream strong shear occurs during the mixing process, so the turbulent kinetic energy and Reynolds shear Increased shear stress; zone control downstream of jet outlet to separation point. There is basically no change before and after; x = 0.4k downstream, the turbulent boundary layer is formed, the energy of the bottom layer is increased, and the ability to resist the adverse pressure gradient is improved, which is consistent with Section \u003cspan refid=\"Sec4\" class=\"InternalRef\"\u003e2.2\u003c/span\u003e. The boundary layer velocity profile measurements are consistent. The original separation air flow formed reattach, the separation point is delayed backwards.\u003c/p\u003e \u003cp\u003eFigure \u003cspan refid=\"Fig9\" class=\"InternalRef\"\u003e10\u003c/span\u003e shows that when the wing angle of attack α = 25°, no control and synthesis Pressure distribution curve of wing surface after jet control (Cn = 0.00125) Compared. It can be seen from the figure that if there is no control, the air on the upper surface of the wing when α = 25°. The flow will be completely separated and a stall will occur; while after the synthetic jet is controlled, the boundary layer. The energy is enhanced, the separation is effectively improved, and the leading edge of the wing has been eliminated.\u003c/p\u003e \u003cp\u003eThe lost suction peak reappears. Variation of pressure distribution on aerofoil surface it leads to the change of the macro aerodynamic force of the wing. Increased aerofoil maximum lift coefficient and delayed stall angle of attack.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003c/div\u003e \u003cdiv id=\"Sec11\" class=\"Section2\"\u003e \u003ch2\u003e3.4 Effect of Blowing rate (\u003cem\u003eQ՛\u003c/em\u003e)\u003c/h2\u003e \u003cp\u003eThis section presents aeroacoustics studies concerning noise generated by turbulence–leading edge interactions, attributed to a freestream turbulence intensity of 4.5%. Figure\u0026nbsp;\u003cspan refid=\"Fig10\" class=\"InternalRef\"\u003e11\u003c/span\u003e compares the sound power level PWL generated by the baseline aerofoil (i.e., Cn = 0-0.00375) with that of the blowing leading edge under the condition of Cn = 0.00375. ΔPWL denotes the disparity in noise spectra between the baseline and the blowing leading edges, expressed as a function of non-dimensional frequency (f.C/U∞). A positive result of ΔPWL indicates greater noise reduction compared to λ′10, and conversely. The jet velocity is U ∞ = 25 m/s, and the angle of attack is zero degrees. The background noise of the wind tunnel, when measuring far-field noise radiation related to the naked jet, is considerably lower than when the aerofoil is included. This verifies the elevated signal-to-noise ratio of the aeroacoustics findings in the present investigation.\u003c/p\u003e \u003cp\u003eThe broadband characteristic resulting from the turbulence interaction is easily observable at the low to mid-frequency ranges when analysing the PWL spectra generated by the baseline aerofoil. It is also hypothesised that the increased freestream turbulence strength triggered a bypass transition on the surface of the aerofoil, given there is no sign of the unstable tone noise in the spectra. The insertion of the blowing leading edge is clearly helpful, with a large reduction in the turbulent broadband noise (up to 9 dB in some circumstances). Interestingly, whereas the level of broadband noise reduction rises with Cn between 0.00125–0.00375 as illustrated in the ΔPWL spectra, further increase of Cn actually becomes deleterious. For example, no detectable noise reduction can be accomplished at Cn 0.0025. As a result, the results highlight that there is an ideal blow rate for cancelling out incoming freestream turbulence and lowering interaction broadband noise. However, based on our available data set, it can be tentatively observed that the condition of Cn for an effective reduction of the interaction broadband noise does not apply for greater freestream velocity (e.g. U ∞ = 45 m/s).\u003c/p\u003e \u003cp\u003eAdditionally, self-noise from the air jets from the leading edge orifices is a contributing factor to the increase in high frequency noise, usually when f \u0026gt; 10. This undesirable source of noise is outside the desired frequency range, but it is still undesirable. Moreover, it becomes less of a concern at higher freestream velocity, because of the increased degree of overall aerofoil noise radiation.\u003c/p\u003e \u003cp\u003eFigure \u003cspan refid=\"Fig11\" class=\"InternalRef\"\u003e12\u003c/span\u003e illustrates the contour graphs of \u003cb\u003e∆\u003c/b\u003ePWL as a function of \u003cem\u003ef\u003c/em\u003e and \u003cem\u003eU ∞\u003c/em\u003e subjected to 0.00125 Cn 0.00375. Despite noise reduction typically occurring within a comparable frequency range, the example with the smaller value higher performance and a broader range of effective freestream velocity. The outcomes illustrated in Fig.\u0026nbsp;\u003cspan refid=\"Fig11\" class=\"InternalRef\"\u003e12\u003c/span\u003e predominantly align with the \u003cb\u003e∆\u003c/b\u003ePWL trend presented in Fig.\u0026nbsp;\u003cspan refid=\"Fig10\" class=\"InternalRef\"\u003e11\u003c/span\u003e. The maximum broadband noise reduction across an extensive frequency range can be achieved with the optimal blowing leading edge. It is important to notice that noise levels rise in certain instances, particularly at non-dimensional frequencies beyond 8. The findings indicate that the Cn = 0.001 scenario cannot attain noise radiation across the full velocity spectrum, but exhibits only marginal enhancements at lower frequencies, as illustrated in Fig.\u0026nbsp;\u003cspan refid=\"Fig11\" class=\"InternalRef\"\u003e12\u003c/span\u003e(d-e). It is noteworthy that an increase in noise at elevated frequencies is evident in all instances.\u003c/p\u003e \u003c/div\u003e"},{"header":"Conclusions","content":"\u003cp\u003eA kind of inclined outlet synthetic jet exciter is designed and developed, which can the NACA 633618-il straight wing model can effectively control the separation flow;\u003c/p\u003e\u003cp\u003eThe interaction characteristics of the synthetic jet and the mainstream were tested by PIV flow field and the Boundary layer scanning was used to analysis the flow control mechanism. Get the following conclusion Argument:\u003c/p\u003e\u003cp\u003e(a) The inclined outlet synthetic jet exciter can effectively control the wing surface\u003c/p\u003e\u003cp\u003eSurface flow separation. Within the experimental range of R, the higher the energy of the jet is than Cn\u003c/p\u003e\u003cp\u003eis larger, the better the control effect is; when Cn = 0.0025, the maximum lift of the aerofoil is the coefficient is increased by 10.6%, and the stall angle of attack is delayed by 4°;\u003c/p\u003e\u003cp\u003e(b) The synthetic jet at the oblique outlet is transported to the mainstream by the lateral momentum injecting energy to enhance the shear stress of the bottom layer of the boundary layer and lift the vicinity of the wall. Turbulent kinetic energy, increased underlying boundary layer energy, ability to resist adverse pressure gradients increase; the separation point is delayed backwards, the separation airflow is reattached, and the separation zone is turbulent flow energy drops.\u003c/p\u003e\u003cp\u003e(c)The leading edge jet consistently counteracts and potentially reduces entering turbulent eddies by either changing the leading edge stagnation point of the aerofoil or establishing a \"buffer zone\" surrounding the airfoil's leading edge. It appears that one or both of these mechanisms may exhibit significant sensitivity to the blowing volume flow rate, which correlates with the exit jet velocity. A pressure difference of up to 7 dB can be attained by one of the leading edge blowing configurations when Cn = 0.0025.\u003c/p\u003e"},{"header":"Declarations","content":"\u003ch2\u003e \u003cb\u003eAuthor Declarations\u003c/b\u003e \u003c/h2\u003e \u003cp\u003e \u003cstrong\u003eConflict of Interest\u003c/strong\u003e \u003cp\u003eThe authors have no conflicts to disclose.\u003c/p\u003e \u003c/p\u003e\u003cp\u003e \u003ch2\u003eData Availbility\u003c/h2\u003e \u003cp\u003eThe data that supports the findings of this study are available within the article.\u003c/p\u003e \u003c/p\u003e\u003cp\u003e\u003cstrong\u003eFunding Information\u003c/strong\u003e\u003c/p\u003e\n\u003cp\u003eThis research was supported by the by\u003c/p\u003e\n\u003cp\u003eUniversity London \u0026nbsp; and Office of the\u003c/p\u003e\n\u003cp\u003e\u0026nbsp;Prime Minister of Iraq Higher Committee for\u003c/p\u003e\n\u003cp\u003e\u0026nbsp;Education Development in Iraq (HCED)\u003c/p\u003e\u003ch2\u003eAuthor Contribution\u003c/h2\u003e\u003cp\u003eDeclaration of interests☒ The authors declare that they have no known competing financial interests or personal relationships that could have appeared to influence the work reported in this paper.☐The authors declare the following financial interests/personal relationships which may be considered as potential competing interests:\u003c/p\u003e\u003ch2\u003eAcknowledgments\u003c/h2\u003e \u003cp\u003eThe authors highly appreciated the anonymous referees' insightful remarks and recommendations, which significantly enhanced the original manuscript. The Higher Committee for Education Development in Iraq (HCED) and the University of Baghdad. Any opinions, research, or suggestions are those of the authors and do not necessarily represent those of the supporting organizations.\u003c/p\u003e"},{"header":"References","content":"\u003col\u003e\u003cli\u003e\u003cspan\u003eLi, Y., Qin, N.: Airfoil gust load. alleviation circulation control Aerosp. Sci. Technol. \u003cb\u003e98\u003c/b\u003e, 105622 (2020)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eAsaro, S., Khalil, K., Andr\u0026eacute;, Bauknecht: Unsteady characterization of fluidic flow control devices for gust load alleviation. New Results in Numerical and Experimental Fluid Mechanics XIII: Contributions to the 22nd STAB/DGLR Symposium. Springer International Publishing, (2021)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eD\u0026uuml;ssler, S.: Florian Siebert, and Andr\u0026eacute; Bauknecht. Coandă-Type Flow Actuation for Load Alleviation. J. 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(1991)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eSmithBL, G.A.: Theformation and development of synthetic jets [J]. Physics. \u003cb\u003e10\u003c/b\u003e, 2281 (1998)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eShuster, J.M., Smith, D.R.: Experimentalstudyoftheformation and scalingofaroundsyntheticjet[J]. Phys. Fluids. \u003cb\u003e19\u003c/b\u003e, 045109045109\u0026ndash;045109045109 (2007)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eGordon, M., Cater, J.E., Soria, J.: Investmentofthemeanpassive scalarfieldinzero-net-mass-fluxjetsincross-flowingplanarlaser Inducedflorescence[J]. Phys. Fluids. \u003cb\u003e16\u003c/b\u003e(3), 794\u0026ndash;808 (2004)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eNishino, T.: Effect of jet nozzle lip momentum loss on circulation control airfoil performance. 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Technol. \u003cb\u003e5\u003c/b\u003e, 19\u0026ndash;22 (2019)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eGross, J., Lance, W.: Traub. Experimental and theoretical investigation of ground effect at low reynolds numbers. J. Aircr. \u003cb\u003e49\u003c/b\u003e(2), 576\u0026ndash;586 (2012)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eGupta, S., et al.: CFD simulation of turbulent flow around multi-element airfoil. AIP Conference Proceedings. Vol. 2204. No. 1. AIP Publishing, (2020)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003ePlint, Partners, L.T.D., Engineers: Operating Instructions TE.81/D, s.l. PLINT (1986)\u003c/span\u003e\u003c/li\u003e\u003c/ol\u003e"}],"fulltextSource":"","fullText":"","funders":[],"hasAdminPriorityOnWorkflow":false,"hasManuscriptDocX":true,"hasOptedInToPreprint":true,"hasPassedJournalQc":"","hasAnyPriority":false,"hideJournal":true,"highlight":"","institution":"","isAcceptedByJournal":false,"isAuthorSuppliedPdf":false,"isDeskRejected":"","isHiddenFromSearch":false,"isInQc":false,"isInWorkflow":false,"isPdf":false,"isPdfUpToDate":true,"isWithdrawnOrRetracted":false,"journal":{"display":true,"email":"[email protected]","identity":"researchsquare","isNatureJournal":false,"hasQc":true,"allowDirectSubmit":true,"externalIdentity":"","sideBox":"","snPcode":"","submissionUrl":"/submission","title":"Research Square","twitterHandle":"researchsquare","acdcEnabled":true,"dfaEnabled":false,"editorialSystem":"","reportingPortfolio":"","inReviewEnabled":false,"inReviewRevisionsEnabled":true},"keywords":"Inclined outlet, synthetic jet, wing separation flow, boundary layer, particle image velocimetry","lastPublishedDoi":"10.21203/rs.3.rs-5769420/v1","lastPublishedDoiUrl":"https://doi.org/10.21203/rs.3.rs-5769420/v1","license":{"name":"CC BY 4.0","url":"https://creativecommons.org/licenses/by/4.0/"},"manuscriptAbstract":"\u003cp\u003eThis study discusses the separation flow active flow control of the NACA 633618-il three-dimensional straight aerofoil, which is carried out by using the inclined exit synthetic jet exciter. The synthetic jet can effectively control the flow separation of the wings, increase the maximum lift coefficient by 12.4%, and delay the stall angle of attack by 6°. Using boundary layer testing techniques and\u003c/p\u003e\n\u003cp\u003eThe particle image velocimetry system is used to study and analyse the control mechanism of the synthetic jet separation flow. The results show that the boundary layer velocity pattern becomes full after the control.\u003c/p\u003e\n\u003cp\u003eThe shape factor decreases, the underlying energy increases, and the ability to resist adverse pressure gradients increases. The transient and real-time homogenised PIV test flow field map further proves the synthetic jet flow direction.\u003c/p\u003e\n\u003cp\u003eAfter the main flow is injected and mixed with momentum, the main flow adheres to the aerofoil surface, the fluid turbulent kinetic energy and Reynolds shear stress near the aerofoil increase, and the separation point is pushed downstream. Later, flow separation is inhibited.\u003c/p\u003e","manuscriptTitle":"Experimental Investigation of effect Slotted leading Edge Circulation Control Aerofoil Aerodynamic and Acoustics","msid":"","msnumber":"","nonDraftVersions":[{"code":1,"date":"2025-01-08 10:44:56","doi":"10.21203/rs.3.rs-5769420/v1","editorialEvents":[{"type":"communityComments","content":0}],"status":"published","journal":{"display":true,"email":"[email protected]","identity":"researchsquare","isNatureJournal":false,"hasQc":true,"allowDirectSubmit":true,"externalIdentity":"","sideBox":"","snPcode":"","submissionUrl":"/submission","title":"Research Square","twitterHandle":"researchsquare","acdcEnabled":true,"dfaEnabled":false,"editorialSystem":"","reportingPortfolio":"","inReviewEnabled":false,"inReviewRevisionsEnabled":true}}],"origin":"","ownerIdentity":"55f6346d-0c76-40f4-afa9-e7c230846f48","owner":[],"postedDate":"January 8th, 2025","published":true,"recentEditorialEvents":[],"rejectedJournal":[],"revision":"","amendment":"","status":"posted","subjectAreas":[],"tags":[],"updatedAt":"2025-05-05T16:23:32+00:00","versionOfRecord":[],"versionCreatedAt":"2025-01-08 10:44:56","video":"","vorDoi":"","vorDoiUrl":"","workflowStages":[]},"version":"v1","identity":"rs-5769420","journalConfig":"researchsquare"},"__N_SSP":true},"page":"/article/[identity]/[[...version]]","query":{"redirect":"/article/rs-5769420","identity":"rs-5769420","version":["v1"]},"buildId":"8U1c8b4HqxoKbykW_rLl7","isFallback":false,"isExperimentalCompile":false,"dynamicIds":[84888],"gssp":true,"scriptLoader":[]}