# Streamline Effect Improvement of Additive Manufactured Airfoil Utilizing Dynamic Stream Control Procedure

Authors: R. Srinath; R. Mukesh; Manish C. Poojari; Inamul Hasan; Wubetu Amare Alebachew
Journal: Advances in Materials Science and Engineering
 (2022)

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## Abstract

In the era of fast transport, to create inventive stream flow management solutions that are capable of diminishing the aerodynamic drag of the vehicles, there is a need to modify the flow characteristics over the vehicle by deferring or expelling the position of the flow partition. The objective of this study involves the parameterized design of an airfoil utilizing the Bezier curve technique with the assistance of the simulation program. For flow regulations, synthetic jet modules are ingrained at different percentages of the chord to manage the stall characteristics. The parametrization system, combined with the stream control method, can give a much better insight into flow re-energization and pave some way for the reduction of the wake. Digital fabrication technique (3d printing or Rapid Prototyping method) is used to fabricate the end product for aerodynamic testing. The comparative outcome showed a reduction in drag at certain angles of attack due to the surface finish obtained. By comparing the results, the aerodynamic efficiency showed a significant rise of 13.05% at lower angles of attack when compressed gas was used in the synthetic jet closer to the frontier edge of the airfoil. Near the stall angle of attack, the coefficient of lift (Cl) and coefficient of drag (Cd) values showed no progress.


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## Body

## 1. Introduction

Diminishment of drag over an airfoil is dealt with in two stages. The configuration file is described in the right detail first, and then the flow control device is added later. There are diverse holes available to model the rib profile curve. They are adjusted beneath a roof referred to as parameterization procedures. Parameterization techniques include Bezier curves, Class-Shape function Transformation, Hicks-Henne “Bump” function and polynomial method, Ferguson curve, and ?3 [1]. Each parameterization strategies have its own recompenses, Selection of the exact method guarantees the smoothness of the 2d bend, the more parameters the optimization process could seek out more aerofoil shapes. On the other hand, with more design parameters, it would be costly to seek design space unwanted curves. The parameterization technique used here is Bezier curve which comes with the disadvantage of no local control. With this change, the position of a control point affects the entire curve. Bezier curve parameterization is a spline curve with a parametric methodology (u, t) that can be used to more precisely model a 2D aerofoil. The advantage of the Bezier curve over others is the ease of computation, steadiness at the lower degrees of control points, and a Bezier curve can be turned and interpreted by performing operations on the control points, which comes with a disadvantage of no local control. With this change, the position of a control point affects the entire curve. Although the work has nothing to do with optimization, the use of parameterization approaches is described in detail with reference to optimization techniques. Optimization of the airfoil requires a parametrization technique to represent the shape of the airfoil and various such techniques are discussed in a review paper Salunke et al. [1]. A geometric comparison of seven different airfoil parameterization methods accomplished by Masters et al. [2] indicates that all techniques require 20–25 design variables to cover the proposed space within the geometric tolerance. Describing an airfoil with a lesser number of parameters proves to be vital in optimization problems. Optimization of NACA 2411 airfoil by Mukesh et al. [3] coupled with the PARASEC technique as a parametrization method indicated that the Coefficient of lift of the airfoil is escalated. Twelve parameters selected by the author served as design variables to characterize the geometry of the airfoil. A new parameterization technique Bezier-GAN proposed by Chen et al. [4] resulted in the well-ordered depiction of the airfoil and it proves to be an accurate fit for the optimization problem. Wei, et al. [5], obtained an optimized E-387 airfoil with enhancement and modification in L/D ratio as their objective function. The author used the Bezier curve as his parameterization technique, where eight Bezier control points were designated as design variables to represent the coordinates of the airfoil. A comparative study made on NACA airfoil using different parameterization methods by Selvan and concluded that by comparing Class-Shape function transformation, Hicks-Henee “Bump” function, Bezier and polynomial approaches to study the effect of shape parameterization, each method uses the different number of parameters to obtain the final geometry. They observed that the Bezier technique had an increase in the lift-to-drag ratio by 13.42857% Selvam [6]. The work by Parasaram and Charyulu [7] expresses that through the techniques of the Quantic Reverse Engineering of Bezier curve formula, it is easier to obtain camber control points from the existing camber cloud of points. In most of the optimization problems, the parameterization of geometry plays a crucial role in deciding the optimal design. Designs of wind turbine blades are complex in general compared to the baseline design of NACA profiles. Wind turbine blade parameterization and optimization do vary with the problems faced in aerospace industry due to the aerodynamic constraints on lift. Airfoil optimization technique for wind turbines done by Ribeiro claims that the wind turbine blade parameterized and optimized through the Bezier curve and genetic algorithm techniques perceived a 50% reduction in the computational time Ribeiro et al. [8]. The procedure for maximizing the form of transonic airfoils was introduced [9, 10] using the Bezier parametric curve with limited control points to precisely replicate the original design. The authors Yang and Zeng [11] suggested a strategy for improving the aerodynamic system to integrate the Bezier curve with the radial base function. The pressure fall has decreased since the refinement of the NACA 0012 airfoil, and the design result shows that the solution can significantly improve the efficiency of aerodynamics. The research suggests that the Bezier curve is a prevailing approach in numerous flight vehicles for aerodynamic development applications.Parameterized airfoil leads to better results when flow control devices are incorporated into the modeled airfoil. This will ensure a delay in the transition of the turbulent to the laminar one or a delay in flow separation. In fluid dynamics, the flow control procedure has gained a large indebtedness since its introduction to aero-related studies. They are categorized into active and passive flow control. Here, for our experimental and methodical studies, a type of active flow control (synthetic jet) method is chosen. Synthetic jets are typically formed by a flow roving back and forth across a small aperture. A jet is a fluid flux that blends with a surrounding medium. Jet flows vary depending on velocity, the diameter of the orifice, density, and viscosity of the fluid (Reynolds’s number). A synthetic jet is an adaptive control of airflow, with the use of an external source of fluid, such as channelized compressed air. Fuel consumption of a commercial aircraft could be hoarded to 8%, if the transition phase over the wing surface can be delayed to 50% and in general these flow patterns occur at low Reynolds numbers (Re) Serdar Genç, et al. [12]. Aerodynamic efficiency improvement in a wing was explored by utilizing “active slat” on a wind turbine aerofoil (DU96-W-180) Halawa et al. [13]. Periodically opening and latching the slat passage contributes to the blowing impact over the wing being created. As a result, they postponed the breakup of the boundary layer, decreased the drag power, and had 3% better aerodynamic efficiency relative to the clean airfoil design. Drag elimination is due to the inclusion of several cavities at the base of a blunt body. This multicavity has beneficial effects on both drag reduction and wake randomness in enhancing vehicle aerodynamics close to dimples on the golf ball to improve ball aerodynamics Martín-Alcántara et al. [14]. Flow control at low Reynolds numbers can be accomplished via actuator which converts the electrical signal to the required physical quantities. Cattafesta and Sheplak [15] classified the various types of actuators based on their specifications (Fluidic, plasma, etc.), design and aeronautical applications Prawin and Rose [16], announced that the maximum drag reduction obtained on the NACA6-series aerofoil configuration (NACA 63-XX8) with installed MJ actuators was about 20% at 0° AOA. For the past ninety years of research work on active control devices, no significant contributions have been attempted over aircraft applications due to technical challenges and the installation of additional systems. Jubran and Hamdan [17], represented the experimental investigation of the secondary flow injection in a cylinder. They have concluded that the velocity and location of injection reduces the vibration. Khalid et al. [18], in his work on flow control devices opines that jet flow actuators address the need for change in the lift at a required moment coefficient. The active piezoelectric transducer actuator engine mount provided better isolation performance, which could lessen around 80% vibration and performs well at both higher and lower frequencies Sui et al. [19]. Recent advanced research on high lift flow control methods via the combined effects of both the active method (porous upper surface of the wing) and passive method (VG) promises to exhibit superior outcomes in delaying transitions. Minimization of drag through active and passive flow regulation techniques reduces drag in a laminar flow airfoil rate 5243 using the optimization method and active and passive control techniques which yielded the reduction in drag of 3.95% and lift is comparatively increased by 5.03% Yagiz et al. [20]. In a real-world scenario, a subsonic wind tunnel test was conducted to substantiate the fluent mathematical analysis [21, 22]. For a wing, the effects of wake turbulence, pitching moment, lift, and drag forces for various angles of attack were measured by wind tunnel measurements Ghazi and Olwi [23]. The prototypes required for wind tunnel testing can be fabricated from any suitable material. Building a physical prototype is a critical step before producing a functional product. A prototype is developed to test different aspects of a design, or to illustrate its features. Early days traditional materials such as wood and other metals, Aluminum in general provided greater flexibility due to advancements in machining compared to former one is used for prototyping wind tunnel models. With the advancements in machining, rapid prototyping plays a crucial role as the materials used in it provide a better surface finish to the prototype compared to metals or wood. Rapid prototyping is a substitute for subtractive processes [24]. 3d printing, is a more generic name used for rapid prototyping, and is distinct in that they combine and bond materials in layers to create objects. where they can also be called additive fabrication, solid free form fabrication (SFF), and layered manufacturing, to name a few. In its most basic form, rapid prototyping (RP) is the manufacturing of three-dimensional (3D) parts using computer-aided design (CAD) data on a smaller time scale [25].The use of 3D printing technology will speed up manufacturing while decreasing prices. At the same time, the requirements will have a greater effect on production. 3D printing technologies can facilitate a more flexible and responsive manufacturing process, as well as better quality control. Various rapid manufacturing techniques are used to fabricate a wind tunnel testing model such as Stereolithography (SLA), Selective Laser Sintering (SLS), Fused Deposition Modelling (FDM) or Material Jetting, Selective Laser Melting (SLM) or Powder Bed Fusion, Laminated Object Manufacturing (LOM) or Sheet Lamination [26, 27].


## 2. Methodology

The symmetrical airfoil NACA 0024 is picked over an asymmetrical one in light of the fact that, the thicker airfoils tend to accelerate the flow on both surfaces. This reduced the pressure difference in turn resulting in a reduction of lift. NACA 0024 is thicker when compared to other symmetric airfoils. So, optimizing such an airfoil with control surfaces may improvise its application in many fronts [28]. The variation of the coefficient of drag at lower or zero AOA is inconsequential for this situation. NACA 0024 airfoil is mapped utilizing the generalized Bezier curve. Inkscape software was used to vectorize the projected airfoil. The ideal finishing and scaling of the vectorized graph are done with the help of the AutoCAD application. 2D airfoil is imported into CATIA for 3D design. Figure 1 indicates the steps followed to procure the expected outcome. The change is centered at about 16% of the front edge of the chord. The first synthetic jet modules were therefore placed just in front of the laminar to turbulent boundary layer transfer. Somewhere about 30% of the chord from the front edge was the second jet unit. As the current flows into the aft arm, the pressure transmission increases due to the boundary layer separation. As a result, the third jet module was installed at 75% from the front edge. Besides, the aerofoil model was evaluated for experimental flow trends such as separation, reattachment, and laminar-turbulent transition of the boundary layer over the aerofoil with and without synthetic jet modules for numerous angles of attack (0°, 5°, 10°, 15°) for differing approaches.Figure 1
Flowchart for aerodynamic efficiency of airfoil improvement using flow control device.


### 2.1. Bezier Curve

A Bezier curve is the technique chosen to obtain the airfoil shape since they are generally acknowledged in the representation of the airfoil. In recent years, the Bezier curve has been applied on a number of occasions for its benefits over both the straight and the curved description. A Bezier curve is a parametric curve composed of Bernstein basis polynomials. In a simple way, it is defined as the summation of the functions of the base weighted by the control points 𝑛:(1)pu=∑i=0nPiBi,nuu∈0,1,where n is the degree of polynomial (defined by the control points), n = no of control points—1, u is the parametric variable, and Bi,n = bases function:(2)Bi,nn!i!n−i!ui1−un−i.The parameter equation of a curve defined by four control points can be generated by equations (1) and (2) as a quadratic curve:(3)Pu=P01−u3+3P1u1−u2+3P2u21−u+P3u3.Airfoil curve acquired using the Bezier curve technique is shown in Figure2; which are characterized by its generalized properties as it always infiltrates the first and last control points; curves are restricted under the arched casing of the control point; they show up the Global Control point. the 2d airfoil curve is later transformed into a 3d model as in Figure 3. Figure 4(a) illustrates the airfoil with synthetic jet modules that are pricked as per the specification shown in Table 1, where Figure 4(b) illustrates the mesh applied over the airfoil.Figure 2
2D design of NACA 0024 using Bezier curve.Figure 3
3D design of NACA 0024 using Bezier curve.Figure 4
(a) Engineering drawing of an aerofoil with trickle channels. (b) Model after applying the mesh for NACA 0024 aerofoil.
(a)(b)Table 1
Specifications of wing and synthetic jet modules.
Wing chord300 mmWing span570 mmNumber of synthetic jet modules3Number of openings in each jet modules5Distance of synthetic jet module from the leading edge(a) 1st module48 mm(b) 2nd module90 mm(c) 3rd module225 mmDiameter of synthetic jet modules and openings8 mmReynolds number3.71229 ∗ 105Spanwise spacing of synthetic jet openings95 mm each


## 2.1. Bezier Curve

A Bezier curve is the technique chosen to obtain the airfoil shape since they are generally acknowledged in the representation of the airfoil. In recent years, the Bezier curve has been applied on a number of occasions for its benefits over both the straight and the curved description. A Bezier curve is a parametric curve composed of Bernstein basis polynomials. In a simple way, it is defined as the summation of the functions of the base weighted by the control points 𝑛:(1)pu=∑i=0nPiBi,nuu∈0,1,where n is the degree of polynomial (defined by the control points), n = no of control points—1, u is the parametric variable, and Bi,n = bases function:(2)Bi,nn!i!n−i!ui1−un−i.The parameter equation of a curve defined by four control points can be generated by equations (1) and (2) as a quadratic curve:(3)Pu=P01−u3+3P1u1−u2+3P2u21−u+P3u3.Airfoil curve acquired using the Bezier curve technique is shown in Figure2; which are characterized by its generalized properties as it always infiltrates the first and last control points; curves are restricted under the arched casing of the control point; they show up the Global Control point. the 2d airfoil curve is later transformed into a 3d model as in Figure 3. Figure 4(a) illustrates the airfoil with synthetic jet modules that are pricked as per the specification shown in Table 1, where Figure 4(b) illustrates the mesh applied over the airfoil.Figure 2
2D design of NACA 0024 using Bezier curve.Figure 3
3D design of NACA 0024 using Bezier curve.Figure 4
(a) Engineering drawing of an aerofoil with trickle channels. (b) Model after applying the mesh for NACA 0024 aerofoil.
(a)(b)Table 1
Specifications of wing and synthetic jet modules.
Wing chord300 mmWing span570 mmNumber of synthetic jet modules3Number of openings in each jet modules5Distance of synthetic jet module from the leading edge(a) 1st module48 mm(b) 2nd module90 mm(c) 3rd module225 mmDiameter of synthetic jet modules and openings8 mmReynolds number3.71229 ∗ 105Spanwise spacing of synthetic jet openings95 mm each


## 3. Results and Analysis

Parameters required for computational fluid dynamics simulations are itemized in Table2. [29, 30]. Simulations are used to study the downstream movement of fluids on different wing models. Here, the research is conducted for NACA 0024 wing design with and without synthetic jet modules. K-epsilon (k − ε) model is used to simulate the turbulent flow condition, which defines turbulence by two transport equations. The influence of the synthetic jet unit was tested in three separate cases by positioning it at three different locations on the upper surface of the aerofoil (i.e., at 16%, 30%, and 75% of the chord). Computational fluid analysis is performed on a wing to study the wing flow properties and to understand the flow pattern and flow separation theory.Table 2
Reference values used for analysis.
Reynolds number3.7123 ∗ 105Density1.225 kg/m3Dynamic viscosity1.7894 ∗ 10–5 kg/m ∗ sTemperature15°CPressure1.01325 ∗ 105 PaVelocity of free stream18.0756 m/sVelocity of air used in synthetic jets6 m/sA maximum of 4 cases were evaluated based on the computational results obtained from the experimental models, and these cases were numerically assessed using the analysis of the wind tunnel trials.


### 3.1. Computational Data

#### 3.1.1. Parameterized Wing without Synthetic Jet Modules

The stagnation point is located near the frontier portion of the airfoil, but its location alters with a different angle of attack. It is explicit from Figures5–7 that at the stagnation level, the stream encounters the maximum pressure and, as the air moves back from the stagnation stage, the pressure comparatively reduces on the top surface for an increased angle of attack. As the flow is nearing the trailing edge, there is a slight increase in the pressure distribution due to the adverse pressure gradient. The stream velocity rises as it moves away from the front edge and reduces again as it reaches the trailing edge owing to the divergence of the air. As the stream reaches the trailing edgeFigure 5
Total pressure at 0° AOA.Figure 6
Total pressure at 10° AOA.Figure 7
Total pressure at 15° AOA.For a higher angle of attack, the fluid divergence happens even faster than the lower angle of attack. Figure8 reveals that the stress of the wall shear differs along with the airfoil sheet, indicating the effect of the pressure gradient and the flow anomalies, such as isolation, reattachment, and laminar-turbulent movement of the boundary layer over the airfoil. It is observed that at 0° AOA, the current is added to the front edge and the wall shear stress is suddenly increased downstream, supporting the reattachment of the boundary layer (i.e. a thinner boundary layer is formed over the airfoil). As a consequence, the divergence level shifts downstream relative to 10° and 15° AOA. As the stream reaches the trailing end, the shear stress of the wall slowly reduces, signaling that the displacement of the boundary layer (i.e., the thicker boundary layer above the airfoil) has begun, culminating in the separation of the flow.Figure 8
Total pressure at 0° AOA for the 1st module.


#### 3.1.2. Parameterized Wing with 1st Synthetic Jet Module Open

From Figures8–10, the highest pressure is observed at the front edge. As the air enters near the trickle channel of the 1st synthetic jet module, a small amount of forwarding air is blocked, thus increasing the pressure around the synthetic jet module. Furthermore, it could be seen that the pressure over the airfoil decreases gradually, and again downstream, the pressure increases due to flow separation. As the free flow air passes via the 1st module due to the bleed air from the trickle channels, the velocity of the downstream air over the top surface of the wing is slightly reduced, culminating in a delay in the flow separation adjacent to the trailing edge. There is a sharp increase in the shear stress of the wall around the trickle channels of the airfoil due to the laminar-turbulent transition. As a result, a thinner boundary layer is formed near the trickle channel, and the separating point moves downstream compared to the parameterized wing without a synthetic jet module. With the increase in AOA, the acceleration of the flow closer to the foremost edge of the airfoil is reinforced, and a rather strong adverse gradient of pressure is formed along with the downstream airfoil. Therefore, the point of separation moves upstream in comparison with that of 0° AOA.Figure 9
Total pressure at 10° AOA for the 1st module.Figure 10
Total pressure at 15° AOA for the 1st module.


#### 3.1.3. Parameterized Wing with 2nd Synthetic Jet Module Open

Figures11–13 demonstrate that the influence of the jet stream from the 2nd module on fluid reattachment at a lower angle of attack is appreciable relative to that of a higher angle of attack. Pressure rises at 0° AOA as the fluid flows through the trickle channel, and the pressure is almost steady as it reaches the trailing side, thereby slowing the isolation of the flow. There is no overall change in pressure across the 2nd jet unit as the stream reaches it. It can also be inferred that there is no significant increase in pressure near the trailing edge at a higher angle of attack. The velocity is almost decreased across the trickle channel as it reaches the trailing edge at 0° AOA, but in the case of 10° and 15° AOA, the current detaches roughly in the center of the airfoil. For 10° and 15° AOA, although, the flow is almost disconnected along the leading edge, and a very insignificant amount of flow is added downstream of the 2nd moduleFigure 11
Total pressure at 0° AOA for the 2nd module.Figure 12
Total pressure at 10° AOA for the 2nd module.Figure 13
Total pressure at 15° AOA for the 2nd module.


#### 3.1.4. Parameterized Wing with 3rd Synthetic Jet Module Open

Figures14–16 indicate that the impact of the jet stream from the 3rd module increases the pressure at the trailing edge at the lower AOA (i.e., at 0° AOA). At the higher angle of attack, there are no significant pressure differences in the stream re-energization at the trailing edge. It is observed that the stream has already been isolated by a higher angle of attack along the front edge of the wing design. The module at the trailing edge is therefore not very critical for downstream flow control, while there is a slight variation in velocity near the trickle channels at the lower angle of attack. The wall shear stress value continues to decrease at the leading edge itself and approaches the minimum value (i.e., zero). It implies that the flow has dispersed, and there is no significant increase in shear stress downstream, indicating that there is no visible reattachment of the current. However, there is a trivial increase in shear stress near the run-off channel at a lower AOA.Figure 14
Total pressure at 0° AOA for the 3rd module.Figure 15
Total pressure at 15° AOA for the 3rd module.Figure 16
Total pressure at 10° AOA for the 3rd module.


### 3.2. Experimental Validation

#### 3.2.1. Wood Based Model

The model was built with a total of 7 ribs. The fabricated airfoil chord width is 300 mm, and the span is 570 mm. 10 triangular slots are created in each rib to position the spars along the spanwise. Thin copper tubings are mounted around the airfoil to calculate the distribution of pressure on the upper and lower surfaces of the airfoil. A maximum of 13 pressure ports have been used on the design. The airfoil was supported by a stainless steel rod measuring 0.75 inches in diameter. A typical copper tube synthetic jet unit has five trickle channels. A maximum of 3 modules comprising 15 trickle channels was used, which were mounted on 5 ribs at the bottom of the upper surface of the prototype. Eventually, the prototype was finished with a wooden lacquer as seen in Figure17.Figure 17
Airfoil wooden lacquer.The experimentation was performed using the subsonic open circuit, suction-type wind tunnel. Instruments used for experimentation were multi-tube manometers, Smoke generator, Pitot-static tubes, Air compressor, Digital anemometer, and Digital thermometer. Pressure variation over a model can be measured either by pressure probes through which a variety of measurements are being made, including force, airspeed, and pressure. Pitot tubes are used to measure pressure and wind speed on various surfaces of an aircraft or other structures, as well as to measure wind speed in a wind tunnel. It has 13 tubes mounted on the board, each with a different angle of inclination or by pressure sensors which has units that can accommodate up to 64 channels; transducers are mounted in 8-card sets that are self-configuring and hot-swappable. Kits of sensor cards are available for mixed-range operation, and sensing ranges as low as 160 Pa differential are available. Table3 lists the tunnel specifications utilized for testing. The stream pattern around the body was measured for 0°, 10°, and 15° AOA with and without bleeding air for the 1st, 2nd, and 3rd synthetic jet modules. The distribution of surface pressure around the model for 0°, 5°, 10°, 15°, and 20° AOA was tabulated by measuring the readings of the Multitude manometer for the constant Reynolds number (i.e., 371230). The smoke generator is used to visualize the flow across the models where paraffin is used for producing the dense white smoke. Essentially it consists of an electrical cartridge heater, overhead diesel bottle, sump bottle air supply pipe from blower, and smoke distributor. Table 4 displays the experimental values are calculated using the formulas given as follows:(4)ReynoldsNumberRe=ρVCμ,where, ρ—Density of air for given temperature, kg/m3, V—Velocity of air, m/s, C—Chord of an airfoil = 0.30 m, and μ—Dynamic viscosity of air for given temperature, Ns/m2.Table 3
Specifications of wind tunnel.
Type of tunnelSubsonic, open circuit, suction typeTest section1000 mm × 1000 mm (inside)Air speed (velocity)30 m/sContraction ratio9 : 1Contraction length2.25 mDriveAxial flow fan driven by AC motor with AC drive speed controllerOverall size4-meter × 4-meter × 13.5 meters (approximately)Power requirementA.C 15 HP (11 kW), 440 volts, electrical supply with neutral & earth connectionsTable 4
Cl and Cd values of at various AOA.
α (°)Basic aerofoilWith 1st module openWith 2nd module openWith 3rd module openClCdClCdClCdClCd00.110.070.150.070.110.050.180.0650.780.020.870.020.710.020.730.02101.020.021.190.020.960.031.010.02151.020.051.110.050.970.050.910.04200.660.060.540.080.630.070.660.05Velocity of air is as follows:(5)V=3.24∆h∵2P0−P∞ρ=V2,where ∆h—difference in Pitot-static tube manometer reading (mm)Coefficient of pressure is as follows:(6)Cp=hi−h∞h0−h∞=h∞−hih∞−h0,where hi—local static manometer reading measured around aerofoil, h∞—Free stream static manometer reading measured by pitot-static probe and h0—Free stream total manometer reading measured by pitot-static probe.Normal coefficientCn is as follows:(7)Σ∆Cp,u×dXC−Σ∆Cp,l×dXC,where X—distance of the respective pressure port from the leading edge, C—Chord of an airfoil,∆Cp,u—Average coefficient of pressure on upper surface and∆Cp,l—Average coefficient of pressure on the lower surface.From Figures18–20, it is apparent that for wings without bleed air and wings with the 1st module open, the coefficient of lift increases with an increase in the AOA, even though there is a slight decrease in the lift at 10° AOA. Furthermore, the lift coefficient increases to a certain maximum point known as the stall angle. The lift coefficient, however, reaches a peak at this angle and then decreases. Due to flow re-energization, there is a significant change in the pressure gradient in the case of the 1st module synthetic jet, which results in a transformation of the flow from laminar to turbulent, thereby slowing the separation of air.Figure 18
Plot betweenCl and angle of attack.Figure 19
Plot betweenCd and angle of attack.Figure 20
Plot betweenCl/Cd and angle of attack.


#### 3.2.2. 3d Printed Model

Different materials from plastic to metals can be used to prototype an object for rapid prototyping. In the case of aerodynamic testing in wind tunnels, materials such as Acrylonitrile Butadiene Styrene (ABS), Polylactic Acid (PLA), Polyethylene terephthalate glycol-modified (PETG), and Thermoplastic Polyurethane (TPU) are the most commonly used ones. It is also been evident from [31, 32] that the models fabricated using rapid manufacturing techniques reduce the roughness. On comparing the properties of the materials mentioned, it is been found that PLA is the most suited material for prototype since they exhibit a high tensile strength and more suitable for prototypes and toys [33, 34]. PLA is chosen for making a prototype based on its biodegradable quality which the other material fails to possess [35]. The printing temperature required to print a prototype using PLA material is around 180°C–200°C [36], where other materials require a higher temperature range. On the other hand, thermal resistance and flexibility are poor when compared to ABS and TEG, which makes it lease accountable for real-time aerospace models. Prototypes were made using both traditional (wood) and rapid prototyping (PLA) techniques to compare the force variation due to surface finishFused Deposition Modelling (FDM), is a rapid technique (extrusion based), is used to prototype the airfoil that operates on the layer-by-layer premise [33]. The FDM method is preferred over other methods because it is cost-effective, shortens lead times, and speeds up the prototyping process. Print larger objects and the design of FDM printers can be easily scaled for a lower cost-to-size ratio [37]. Fused Deposition Modelling takes standard STL data files as input and may employ various build materials in a build or support relationship. FDM machines are essentially CNC-controlled robots with miniature extruder heads. Solid things are constructed “string by string” by feeding the head with a plastic wire. Figure21, shows the 3d printed airfoil along with synthetic jets using fusion deposition modeling with PLA material. It was tested. The printed airfoil was tested in the wind tunnel shown in Figure 22 under the same conditions as the wood-based model and its results were compared.Figure 21
3d printed airfoil.Figure 22
Wind tunnel testing of 3d printed airfoil.Table5 portrays the CL and CD values of the prototype fabricated using wood and through rapid prototyping. The graphs, Figures 23 and 24 show that there is negligible change in the coefficient values when the material used for the prototype differs. This is due to the evenness of the surface of the prototype fabricated using RP. The wood based prototype was also coated with lacquer to improve its surface finish due to which errors in values are greatly diminishedTable 5
Table4Cl and Cd values of wood-based model and RP model.
α (°)Basic aerofoilWith 1st module openWoodRP modelWoodRP modelClCdClClCdCd00.110.070.150.120.070.0550.780.020.870.820.020.04101.020.021.191.260.020.02151.020.051.111.060.050.04200.660.060.540.610.080.07Figure 23
Plot betweenCl and angle of attack.Figure 24
Plot betweenCd and angle of attack.


## 3.1. Computational Data

### 3.1.1. Parameterized Wing without Synthetic Jet Modules

The stagnation point is located near the frontier portion of the airfoil, but its location alters with a different angle of attack. It is explicit from Figures5–7 that at the stagnation level, the stream encounters the maximum pressure and, as the air moves back from the stagnation stage, the pressure comparatively reduces on the top surface for an increased angle of attack. As the flow is nearing the trailing edge, there is a slight increase in the pressure distribution due to the adverse pressure gradient. The stream velocity rises as it moves away from the front edge and reduces again as it reaches the trailing edge owing to the divergence of the air. As the stream reaches the trailing edgeFigure 5
Total pressure at 0° AOA.Figure 6
Total pressure at 10° AOA.Figure 7
Total pressure at 15° AOA.For a higher angle of attack, the fluid divergence happens even faster than the lower angle of attack. Figure8 reveals that the stress of the wall shear differs along with the airfoil sheet, indicating the effect of the pressure gradient and the flow anomalies, such as isolation, reattachment, and laminar-turbulent movement of the boundary layer over the airfoil. It is observed that at 0° AOA, the current is added to the front edge and the wall shear stress is suddenly increased downstream, supporting the reattachment of the boundary layer (i.e. a thinner boundary layer is formed over the airfoil). As a consequence, the divergence level shifts downstream relative to 10° and 15° AOA. As the stream reaches the trailing end, the shear stress of the wall slowly reduces, signaling that the displacement of the boundary layer (i.e., the thicker boundary layer above the airfoil) has begun, culminating in the separation of the flow.Figure 8
Total pressure at 0° AOA for the 1st module.


### 3.1.2. Parameterized Wing with 1st Synthetic Jet Module Open

From Figures8–10, the highest pressure is observed at the front edge. As the air enters near the trickle channel of the 1st synthetic jet module, a small amount of forwarding air is blocked, thus increasing the pressure around the synthetic jet module. Furthermore, it could be seen that the pressure over the airfoil decreases gradually, and again downstream, the pressure increases due to flow separation. As the free flow air passes via the 1st module due to the bleed air from the trickle channels, the velocity of the downstream air over the top surface of the wing is slightly reduced, culminating in a delay in the flow separation adjacent to the trailing edge. There is a sharp increase in the shear stress of the wall around the trickle channels of the airfoil due to the laminar-turbulent transition. As a result, a thinner boundary layer is formed near the trickle channel, and the separating point moves downstream compared to the parameterized wing without a synthetic jet module. With the increase in AOA, the acceleration of the flow closer to the foremost edge of the airfoil is reinforced, and a rather strong adverse gradient of pressure is formed along with the downstream airfoil. Therefore, the point of separation moves upstream in comparison with that of 0° AOA.Figure 9
Total pressure at 10° AOA for the 1st module.Figure 10
Total pressure at 15° AOA for the 1st module.


### 3.1.3. Parameterized Wing with 2nd Synthetic Jet Module Open

Figures11–13 demonstrate that the influence of the jet stream from the 2nd module on fluid reattachment at a lower angle of attack is appreciable relative to that of a higher angle of attack. Pressure rises at 0° AOA as the fluid flows through the trickle channel, and the pressure is almost steady as it reaches the trailing side, thereby slowing the isolation of the flow. There is no overall change in pressure across the 2nd jet unit as the stream reaches it. It can also be inferred that there is no significant increase in pressure near the trailing edge at a higher angle of attack. The velocity is almost decreased across the trickle channel as it reaches the trailing edge at 0° AOA, but in the case of 10° and 15° AOA, the current detaches roughly in the center of the airfoil. For 10° and 15° AOA, although, the flow is almost disconnected along the leading edge, and a very insignificant amount of flow is added downstream of the 2nd moduleFigure 11
Total pressure at 0° AOA for the 2nd module.Figure 12
Total pressure at 10° AOA for the 2nd module.Figure 13
Total pressure at 15° AOA for the 2nd module.


### 3.1.4. Parameterized Wing with 3rd Synthetic Jet Module Open

Figures14–16 indicate that the impact of the jet stream from the 3rd module increases the pressure at the trailing edge at the lower AOA (i.e., at 0° AOA). At the higher angle of attack, there are no significant pressure differences in the stream re-energization at the trailing edge. It is observed that the stream has already been isolated by a higher angle of attack along the front edge of the wing design. The module at the trailing edge is therefore not very critical for downstream flow control, while there is a slight variation in velocity near the trickle channels at the lower angle of attack. The wall shear stress value continues to decrease at the leading edge itself and approaches the minimum value (i.e., zero). It implies that the flow has dispersed, and there is no significant increase in shear stress downstream, indicating that there is no visible reattachment of the current. However, there is a trivial increase in shear stress near the run-off channel at a lower AOA.Figure 14
Total pressure at 0° AOA for the 3rd module.Figure 15
Total pressure at 15° AOA for the 3rd module.Figure 16
Total pressure at 10° AOA for the 3rd module.


## 3.1.1. Parameterized Wing without Synthetic Jet Modules

The stagnation point is located near the frontier portion of the airfoil, but its location alters with a different angle of attack. It is explicit from Figures5–7 that at the stagnation level, the stream encounters the maximum pressure and, as the air moves back from the stagnation stage, the pressure comparatively reduces on the top surface for an increased angle of attack. As the flow is nearing the trailing edge, there is a slight increase in the pressure distribution due to the adverse pressure gradient. The stream velocity rises as it moves away from the front edge and reduces again as it reaches the trailing edge owing to the divergence of the air. As the stream reaches the trailing edgeFigure 5
Total pressure at 0° AOA.Figure 6
Total pressure at 10° AOA.Figure 7
Total pressure at 15° AOA.For a higher angle of attack, the fluid divergence happens even faster than the lower angle of attack. Figure8 reveals that the stress of the wall shear differs along with the airfoil sheet, indicating the effect of the pressure gradient and the flow anomalies, such as isolation, reattachment, and laminar-turbulent movement of the boundary layer over the airfoil. It is observed that at 0° AOA, the current is added to the front edge and the wall shear stress is suddenly increased downstream, supporting the reattachment of the boundary layer (i.e. a thinner boundary layer is formed over the airfoil). As a consequence, the divergence level shifts downstream relative to 10° and 15° AOA. As the stream reaches the trailing end, the shear stress of the wall slowly reduces, signaling that the displacement of the boundary layer (i.e., the thicker boundary layer above the airfoil) has begun, culminating in the separation of the flow.Figure 8
Total pressure at 0° AOA for the 1st module.


## 3.1.2. Parameterized Wing with 1st Synthetic Jet Module Open

From Figures8–10, the highest pressure is observed at the front edge. As the air enters near the trickle channel of the 1st synthetic jet module, a small amount of forwarding air is blocked, thus increasing the pressure around the synthetic jet module. Furthermore, it could be seen that the pressure over the airfoil decreases gradually, and again downstream, the pressure increases due to flow separation. As the free flow air passes via the 1st module due to the bleed air from the trickle channels, the velocity of the downstream air over the top surface of the wing is slightly reduced, culminating in a delay in the flow separation adjacent to the trailing edge. There is a sharp increase in the shear stress of the wall around the trickle channels of the airfoil due to the laminar-turbulent transition. As a result, a thinner boundary layer is formed near the trickle channel, and the separating point moves downstream compared to the parameterized wing without a synthetic jet module. With the increase in AOA, the acceleration of the flow closer to the foremost edge of the airfoil is reinforced, and a rather strong adverse gradient of pressure is formed along with the downstream airfoil. Therefore, the point of separation moves upstream in comparison with that of 0° AOA.Figure 9
Total pressure at 10° AOA for the 1st module.Figure 10
Total pressure at 15° AOA for the 1st module.


## 3.1.3. Parameterized Wing with 2nd Synthetic Jet Module Open

Figures11–13 demonstrate that the influence of the jet stream from the 2nd module on fluid reattachment at a lower angle of attack is appreciable relative to that of a higher angle of attack. Pressure rises at 0° AOA as the fluid flows through the trickle channel, and the pressure is almost steady as it reaches the trailing side, thereby slowing the isolation of the flow. There is no overall change in pressure across the 2nd jet unit as the stream reaches it. It can also be inferred that there is no significant increase in pressure near the trailing edge at a higher angle of attack. The velocity is almost decreased across the trickle channel as it reaches the trailing edge at 0° AOA, but in the case of 10° and 15° AOA, the current detaches roughly in the center of the airfoil. For 10° and 15° AOA, although, the flow is almost disconnected along the leading edge, and a very insignificant amount of flow is added downstream of the 2nd moduleFigure 11
Total pressure at 0° AOA for the 2nd module.Figure 12
Total pressure at 10° AOA for the 2nd module.Figure 13
Total pressure at 15° AOA for the 2nd module.


## 3.1.4. Parameterized Wing with 3rd Synthetic Jet Module Open

Figures14–16 indicate that the impact of the jet stream from the 3rd module increases the pressure at the trailing edge at the lower AOA (i.e., at 0° AOA). At the higher angle of attack, there are no significant pressure differences in the stream re-energization at the trailing edge. It is observed that the stream has already been isolated by a higher angle of attack along the front edge of the wing design. The module at the trailing edge is therefore not very critical for downstream flow control, while there is a slight variation in velocity near the trickle channels at the lower angle of attack. The wall shear stress value continues to decrease at the leading edge itself and approaches the minimum value (i.e., zero). It implies that the flow has dispersed, and there is no significant increase in shear stress downstream, indicating that there is no visible reattachment of the current. However, there is a trivial increase in shear stress near the run-off channel at a lower AOA.Figure 14
Total pressure at 0° AOA for the 3rd module.Figure 15
Total pressure at 15° AOA for the 3rd module.Figure 16
Total pressure at 10° AOA for the 3rd module.


## 3.2. Experimental Validation

### 3.2.1. Wood Based Model

The model was built with a total of 7 ribs. The fabricated airfoil chord width is 300 mm, and the span is 570 mm. 10 triangular slots are created in each rib to position the spars along the spanwise. Thin copper tubings are mounted around the airfoil to calculate the distribution of pressure on the upper and lower surfaces of the airfoil. A maximum of 13 pressure ports have been used on the design. The airfoil was supported by a stainless steel rod measuring 0.75 inches in diameter. A typical copper tube synthetic jet unit has five trickle channels. A maximum of 3 modules comprising 15 trickle channels was used, which were mounted on 5 ribs at the bottom of the upper surface of the prototype. Eventually, the prototype was finished with a wooden lacquer as seen in Figure17.Figure 17
Airfoil wooden lacquer.The experimentation was performed using the subsonic open circuit, suction-type wind tunnel. Instruments used for experimentation were multi-tube manometers, Smoke generator, Pitot-static tubes, Air compressor, Digital anemometer, and Digital thermometer. Pressure variation over a model can be measured either by pressure probes through which a variety of measurements are being made, including force, airspeed, and pressure. Pitot tubes are used to measure pressure and wind speed on various surfaces of an aircraft or other structures, as well as to measure wind speed in a wind tunnel. It has 13 tubes mounted on the board, each with a different angle of inclination or by pressure sensors which has units that can accommodate up to 64 channels; transducers are mounted in 8-card sets that are self-configuring and hot-swappable. Kits of sensor cards are available for mixed-range operation, and sensing ranges as low as 160 Pa differential are available. Table3 lists the tunnel specifications utilized for testing. The stream pattern around the body was measured for 0°, 10°, and 15° AOA with and without bleeding air for the 1st, 2nd, and 3rd synthetic jet modules. The distribution of surface pressure around the model for 0°, 5°, 10°, 15°, and 20° AOA was tabulated by measuring the readings of the Multitude manometer for the constant Reynolds number (i.e., 371230). The smoke generator is used to visualize the flow across the models where paraffin is used for producing the dense white smoke. Essentially it consists of an electrical cartridge heater, overhead diesel bottle, sump bottle air supply pipe from blower, and smoke distributor. Table 4 displays the experimental values are calculated using the formulas given as follows:(4)ReynoldsNumberRe=ρVCμ,where, ρ—Density of air for given temperature, kg/m3, V—Velocity of air, m/s, C—Chord of an airfoil = 0.30 m, and μ—Dynamic viscosity of air for given temperature, Ns/m2.Table 3
Specifications of wind tunnel.
Type of tunnelSubsonic, open circuit, suction typeTest section1000 mm × 1000 mm (inside)Air speed (velocity)30 m/sContraction ratio9 : 1Contraction length2.25 mDriveAxial flow fan driven by AC motor with AC drive speed controllerOverall size4-meter × 4-meter × 13.5 meters (approximately)Power requirementA.C 15 HP (11 kW), 440 volts, electrical supply with neutral & earth connectionsTable 4
Cl and Cd values of at various AOA.
α (°)Basic aerofoilWith 1st module openWith 2nd module openWith 3rd module openClCdClCdClCdClCd00.110.070.150.070.110.050.180.0650.780.020.870.020.710.020.730.02101.020.021.190.020.960.031.010.02151.020.051.110.050.970.050.910.04200.660.060.540.080.630.070.660.05Velocity of air is as follows:(5)V=3.24∆h∵2P0−P∞ρ=V2,where ∆h—difference in Pitot-static tube manometer reading (mm)Coefficient of pressure is as follows:(6)Cp=hi−h∞h0−h∞=h∞−hih∞−h0,where hi—local static manometer reading measured around aerofoil, h∞—Free stream static manometer reading measured by pitot-static probe and h0—Free stream total manometer reading measured by pitot-static probe.Normal coefficientCn is as follows:(7)Σ∆Cp,u×dXC−Σ∆Cp,l×dXC,where X—distance of the respective pressure port from the leading edge, C—Chord of an airfoil,∆Cp,u—Average coefficient of pressure on upper surface and∆Cp,l—Average coefficient of pressure on the lower surface.From Figures18–20, it is apparent that for wings without bleed air and wings with the 1st module open, the coefficient of lift increases with an increase in the AOA, even though there is a slight decrease in the lift at 10° AOA. Furthermore, the lift coefficient increases to a certain maximum point known as the stall angle. The lift coefficient, however, reaches a peak at this angle and then decreases. Due to flow re-energization, there is a significant change in the pressure gradient in the case of the 1st module synthetic jet, which results in a transformation of the flow from laminar to turbulent, thereby slowing the separation of air.Figure 18
Plot betweenCl and angle of attack.Figure 19
Plot betweenCd and angle of attack.Figure 20
Plot betweenCl/Cd and angle of attack.


### 3.2.2. 3d Printed Model

Different materials from plastic to metals can be used to prototype an object for rapid prototyping. In the case of aerodynamic testing in wind tunnels, materials such as Acrylonitrile Butadiene Styrene (ABS), Polylactic Acid (PLA), Polyethylene terephthalate glycol-modified (PETG), and Thermoplastic Polyurethane (TPU) are the most commonly used ones. It is also been evident from [31, 32] that the models fabricated using rapid manufacturing techniques reduce the roughness. On comparing the properties of the materials mentioned, it is been found that PLA is the most suited material for prototype since they exhibit a high tensile strength and more suitable for prototypes and toys [33, 34]. PLA is chosen for making a prototype based on its biodegradable quality which the other material fails to possess [35]. The printing temperature required to print a prototype using PLA material is around 180°C–200°C [36], where other materials require a higher temperature range. On the other hand, thermal resistance and flexibility are poor when compared to ABS and TEG, which makes it lease accountable for real-time aerospace models. Prototypes were made using both traditional (wood) and rapid prototyping (PLA) techniques to compare the force variation due to surface finishFused Deposition Modelling (FDM), is a rapid technique (extrusion based), is used to prototype the airfoil that operates on the layer-by-layer premise [33]. The FDM method is preferred over other methods because it is cost-effective, shortens lead times, and speeds up the prototyping process. Print larger objects and the design of FDM printers can be easily scaled for a lower cost-to-size ratio [37]. Fused Deposition Modelling takes standard STL data files as input and may employ various build materials in a build or support relationship. FDM machines are essentially CNC-controlled robots with miniature extruder heads. Solid things are constructed “string by string” by feeding the head with a plastic wire. Figure21, shows the 3d printed airfoil along with synthetic jets using fusion deposition modeling with PLA material. It was tested. The printed airfoil was tested in the wind tunnel shown in Figure 22 under the same conditions as the wood-based model and its results were compared.Figure 21
3d printed airfoil.Figure 22
Wind tunnel testing of 3d printed airfoil.Table5 portrays the CL and CD values of the prototype fabricated using wood and through rapid prototyping. The graphs, Figures 23 and 24 show that there is negligible change in the coefficient values when the material used for the prototype differs. This is due to the evenness of the surface of the prototype fabricated using RP. The wood based prototype was also coated with lacquer to improve its surface finish due to which errors in values are greatly diminishedTable 5
Table4Cl and Cd values of wood-based model and RP model.
α (°)Basic aerofoilWith 1st module openWoodRP modelWoodRP modelClCdClClCdCd00.110.070.150.120.070.0550.780.020.870.820.020.04101.020.021.191.260.020.02151.020.051.111.060.050.04200.660.060.540.610.080.07Figure 23
Plot betweenCl and angle of attack.Figure 24
Plot betweenCd and angle of attack.


## 3.2.1. Wood Based Model

The model was built with a total of 7 ribs. The fabricated airfoil chord width is 300 mm, and the span is 570 mm. 10 triangular slots are created in each rib to position the spars along the spanwise. Thin copper tubings are mounted around the airfoil to calculate the distribution of pressure on the upper and lower surfaces of the airfoil. A maximum of 13 pressure ports have been used on the design. The airfoil was supported by a stainless steel rod measuring 0.75 inches in diameter. A typical copper tube synthetic jet unit has five trickle channels. A maximum of 3 modules comprising 15 trickle channels was used, which were mounted on 5 ribs at the bottom of the upper surface of the prototype. Eventually, the prototype was finished with a wooden lacquer as seen in Figure17.Figure 17
Airfoil wooden lacquer.The experimentation was performed using the subsonic open circuit, suction-type wind tunnel. Instruments used for experimentation were multi-tube manometers, Smoke generator, Pitot-static tubes, Air compressor, Digital anemometer, and Digital thermometer. Pressure variation over a model can be measured either by pressure probes through which a variety of measurements are being made, including force, airspeed, and pressure. Pitot tubes are used to measure pressure and wind speed on various surfaces of an aircraft or other structures, as well as to measure wind speed in a wind tunnel. It has 13 tubes mounted on the board, each with a different angle of inclination or by pressure sensors which has units that can accommodate up to 64 channels; transducers are mounted in 8-card sets that are self-configuring and hot-swappable. Kits of sensor cards are available for mixed-range operation, and sensing ranges as low as 160 Pa differential are available. Table3 lists the tunnel specifications utilized for testing. The stream pattern around the body was measured for 0°, 10°, and 15° AOA with and without bleeding air for the 1st, 2nd, and 3rd synthetic jet modules. The distribution of surface pressure around the model for 0°, 5°, 10°, 15°, and 20° AOA was tabulated by measuring the readings of the Multitude manometer for the constant Reynolds number (i.e., 371230). The smoke generator is used to visualize the flow across the models where paraffin is used for producing the dense white smoke. Essentially it consists of an electrical cartridge heater, overhead diesel bottle, sump bottle air supply pipe from blower, and smoke distributor. Table 4 displays the experimental values are calculated using the formulas given as follows:(4)ReynoldsNumberRe=ρVCμ,where, ρ—Density of air for given temperature, kg/m3, V—Velocity of air, m/s, C—Chord of an airfoil = 0.30 m, and μ—Dynamic viscosity of air for given temperature, Ns/m2.Table 3
Specifications of wind tunnel.
Type of tunnelSubsonic, open circuit, suction typeTest section1000 mm × 1000 mm (inside)Air speed (velocity)30 m/sContraction ratio9 : 1Contraction length2.25 mDriveAxial flow fan driven by AC motor with AC drive speed controllerOverall size4-meter × 4-meter × 13.5 meters (approximately)Power requirementA.C 15 HP (11 kW), 440 volts, electrical supply with neutral & earth connectionsTable 4
Cl and Cd values of at various AOA.
α (°)Basic aerofoilWith 1st module openWith 2nd module openWith 3rd module openClCdClCdClCdClCd00.110.070.150.070.110.050.180.0650.780.020.870.020.710.020.730.02101.020.021.190.020.960.031.010.02151.020.051.110.050.970.050.910.04200.660.060.540.080.630.070.660.05Velocity of air is as follows:(5)V=3.24∆h∵2P0−P∞ρ=V2,where ∆h—difference in Pitot-static tube manometer reading (mm)Coefficient of pressure is as follows:(6)Cp=hi−h∞h0−h∞=h∞−hih∞−h0,where hi—local static manometer reading measured around aerofoil, h∞—Free stream static manometer reading measured by pitot-static probe and h0—Free stream total manometer reading measured by pitot-static probe.Normal coefficientCn is as follows:(7)Σ∆Cp,u×dXC−Σ∆Cp,l×dXC,where X—distance of the respective pressure port from the leading edge, C—Chord of an airfoil,∆Cp,u—Average coefficient of pressure on upper surface and∆Cp,l—Average coefficient of pressure on the lower surface.From Figures18–20, it is apparent that for wings without bleed air and wings with the 1st module open, the coefficient of lift increases with an increase in the AOA, even though there is a slight decrease in the lift at 10° AOA. Furthermore, the lift coefficient increases to a certain maximum point known as the stall angle. The lift coefficient, however, reaches a peak at this angle and then decreases. Due to flow re-energization, there is a significant change in the pressure gradient in the case of the 1st module synthetic jet, which results in a transformation of the flow from laminar to turbulent, thereby slowing the separation of air.Figure 18
Plot betweenCl and angle of attack.Figure 19
Plot betweenCd and angle of attack.Figure 20
Plot betweenCl/Cd and angle of attack.


## 3.2.2. 3d Printed Model

Different materials from plastic to metals can be used to prototype an object for rapid prototyping. In the case of aerodynamic testing in wind tunnels, materials such as Acrylonitrile Butadiene Styrene (ABS), Polylactic Acid (PLA), Polyethylene terephthalate glycol-modified (PETG), and Thermoplastic Polyurethane (TPU) are the most commonly used ones. It is also been evident from [31, 32] that the models fabricated using rapid manufacturing techniques reduce the roughness. On comparing the properties of the materials mentioned, it is been found that PLA is the most suited material for prototype since they exhibit a high tensile strength and more suitable for prototypes and toys [33, 34]. PLA is chosen for making a prototype based on its biodegradable quality which the other material fails to possess [35]. The printing temperature required to print a prototype using PLA material is around 180°C–200°C [36], where other materials require a higher temperature range. On the other hand, thermal resistance and flexibility are poor when compared to ABS and TEG, which makes it lease accountable for real-time aerospace models. Prototypes were made using both traditional (wood) and rapid prototyping (PLA) techniques to compare the force variation due to surface finishFused Deposition Modelling (FDM), is a rapid technique (extrusion based), is used to prototype the airfoil that operates on the layer-by-layer premise [33]. The FDM method is preferred over other methods because it is cost-effective, shortens lead times, and speeds up the prototyping process. Print larger objects and the design of FDM printers can be easily scaled for a lower cost-to-size ratio [37]. Fused Deposition Modelling takes standard STL data files as input and may employ various build materials in a build or support relationship. FDM machines are essentially CNC-controlled robots with miniature extruder heads. Solid things are constructed “string by string” by feeding the head with a plastic wire. Figure21, shows the 3d printed airfoil along with synthetic jets using fusion deposition modeling with PLA material. It was tested. The printed airfoil was tested in the wind tunnel shown in Figure 22 under the same conditions as the wood-based model and its results were compared.Figure 21
3d printed airfoil.Figure 22
Wind tunnel testing of 3d printed airfoil.Table5 portrays the CL and CD values of the prototype fabricated using wood and through rapid prototyping. The graphs, Figures 23 and 24 show that there is negligible change in the coefficient values when the material used for the prototype differs. This is due to the evenness of the surface of the prototype fabricated using RP. The wood based prototype was also coated with lacquer to improve its surface finish due to which errors in values are greatly diminishedTable 5
Table4Cl and Cd values of wood-based model and RP model.
α (°)Basic aerofoilWith 1st module openWoodRP modelWoodRP modelClCdClClCdCd00.110.070.150.120.070.0550.780.020.870.820.020.04101.020.021.191.260.020.02151.020.051.111.060.050.04200.660.060.540.610.080.07Figure 23
Plot betweenCl and angle of attack.Figure 24
Plot betweenCd and angle of attack.


## 4. Discussion

Aerodynamic efficiency is improved by 13.05% in the case of the 1st module synthetic jet. In the case of the 2nd module synthetic jet, it can be found that the lift coefficient rise is at 0° AOA, while the lift coefficient declines with an improvement in the AOA relative to the wing without bleeding air. In the case of the 3rd module synthetic jet, it is interesting that the lift coefficient is raised at 0° and 5° AOA.The research was aimed at investigating the effect of effective flow control on the airfoil’s aerodynamic properties growth. The Mat lab code for the Bezier curve technique contributed to the improvement of the smooth spline curve for the drawing of the airfoil. The study clearly shows that CFD results in a substantial interpretation of the flow pattern over the airfoil. Figures25–28 confirm that the smoke stream visualization contributes to a better understanding of the flow pattern, fluid mixing, and vortex produced by the synthetic jet due to the laminar-turbulent boundary layer transfer. The primary aim was to increase the aerodynamic efficiency by using a synthetic jet. Synthetic jets are useful aerodynamic devices that can be used by aircraft manufacturers to improve aircraft flying qualities. However, the complexities such as the pressure drop across the valve, high energy loss the system may experience and its structural issues may incur are not considered here [38]. Comparison between the different materials used to prototype the model also shows that the surface finish plays a substantial role in reducing the skin friction to certain extent. This can be further improved by using advanced polymers such as ABS or TEG, which provides additional flexibility and improved surface finish The prudent use of synthetic jets leads to optimal aerodynamic efficiency. The leading-edge synthetic jet unit has a significant contribution to slow stream separation and boost lift. Ultimately, and in view of the results obtained, the synthetic jet is a flexible flow control tool.Figure 25
Smoke flow visualization at 0° AOA without synthetic jet module.Figure 26
Smoke flow visualization at 0° AOA for the 1st module open.Figure 27
Smoke flow visualization at 10° AOA for the 1st module open.Figure 28
Smoke flow visualization at 15° AOA for the 1st module open.



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*Source: 1252681-2022-09-17.xml*