# Wind Tunnel Testing and Validation of Helicopter Rotor Blades Using Additive Manufacturing

Authors: Inamul Hasan; R. Mukesh; P. Radha Krishnan; R. Srinath; Dhanya Prakash Babu; Negash Lemma Gurmu
Journal: Advances in Materials Science and Engineering
 (2022)

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## Abstract

This research paper aims to validate the aerodynamic performance of rotor blades using additive manufacturing techniques. Wind tunnel testing is a technique used to find the flow characteristics of the body. Computational fluid dynamics (CFD) techniques are used for aerodynamic analysis, and validation should be done using wind tunnel testing. In the aerodynamic testing of models, additive manufacturing techniques help in validating the results by making models easily for wind tunnels. Recent developments in additive manufacturing help in the aerodynamic testing of models in wind tunnels. The CFD analysis of helicopter rotor blades was analyzed in this research, and validation was done using additive manufacturing techniques. Computational analysis was carried out for static analysis for the forward speeds of Mach numbers 0.3, 0.4, and 0.5. The results obtained were satisfactory to the previous results and were validated with wind tunnel testing. Results proved that the error percentage was lower, and the computational analysis was valid. In this research, models were designed using the FDM technique for wind tunnel testing as it is cost-effective and easy to manufacture.


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## Body

## 1. Introduction

A helicopter is a flying machine with many advantages over other flying objects, such as landing or take-off at almost any terrain, flying at a specific position (hovering), and flying backward and sidewards quickly. Helicopters can be termed any terrain vehicle as they have many advantages over aircraft [1, 2]. The helicopter is a type of rotorcraft currently used for many purposes such as rescue, firefighting, and agriculture. The helicopter’s design differs with various parameters for different applications [3]. Helicopters possess more capability than aircraft as it has the unique ability of vertical take-off and landing, and no particular runway. Helicopters have some disadvantages over aircraft in the case of speed and pilot operation. The helicopter speed is improved since the invention of the helicopter in recent days. Many factors are suppressing the speed of helicopters [4]. Helicopters can fly forward, backward, or sideways, stay at a position (hover), and vertically can go up or come down. In aircraft, lift is produced by the wings, and engines produce thrust. In helicopters, lift and thrust are produced by a rotor driven by one or two engines [5, 6].CFD software is used for analyzing the aerodynamic performance of moving vehicles to assess the aerodynamic performance. CFD is widely used in many fields, and improving CFD techniques helps improve results and reduce errors [7]. Even though CFD software provides good results, researchers prefer to compare the results obtained from CFD with wind tunnel testing. CFD results can be validated with the previous experimental results or wind tunnel tests [8, 9]. The model was made from wood or metals for wind tunnel tests in the previous days. Models created in wood require a high cost for skilled workers, and it takes a long time to make the models. Metals like aluminum were used to prepare the model using CNC machines are high cost and less affordable [10, 11].The evolution of additive manufacturing techniques helps in the easy and affordable manufacturing of models for various applications. Mainly for wind tunnel testing with the help of CAD design software, it is easy to design and manufacture with less cost [12, 13]. Advancements in additive manufacturing technologies help many fields in engineering, medicine, and research to test or demonstrate the models [14, 15]. Material qualities are improved by the advancements in the development of composite materials and help make low-cost models faster [16].


## 2. Experimental Details

### 2.1. CFD Analysis

Four-bladed helicopters are chosen as it is used in most high-speed helicopters. Analysis was carried out by ANSYS fluent software. It provides the numerical solution computationally with the help of the Navier–Stokes equation over the flow region. These computational techniques are beneficial for research on complex problems. Four-bladed rotors are designed by CATIA V5 software. Coordinates are downloaded from the UIUC airfoil data site, and the 2D profile can be generated from the airfoil data [17].


#### 2.1.1. Design and Preprocessing

The CAD model design is essential and a preliminary requirement for computational analysis. CAD models can be designed by any design software like Solid Works, Solid Edge, and Catia. [18]. Catia V5 is used in this research work to design the three-bladed and four-bladed helicopter blades. The same model is used for the experimental model and scaled down for wind tunnel size using the 3D printing method [19, 20]. The designed model will be imported to fluent for further analysis. After designing the model, the model is imported to the ANSYS CFD workbench for further processing and analysis. The first stage is preprocessing, where the meshing and boundary conditions are applied [21]. As it is a static analysis, C type domain was chosen for computational analysis. Unstructured meshing is chosen as it is a complex structure to provide accurate results in CFD analysis [22, 23]. The realizable k-ε model is chosen as the solver because it provides better results than other models, as mentioned in Figure 1.Figure 1
Turbulence model selection.The analysis was carried out for all the models and compared with experimental results to finalize the turbulence model. The analysis showed that realizablek − ε was near the experimental results [22, 24].An analysis was carried out for HH02 and NASA SC(2)-0714 rotor airfoils for Mach numbers of 0.3, 0.4, and 0.5. Blades are designed with the exact scaling of the Apache helicopter as the HH02 airfoil provides high performance. Parts of the rotors are designed separately as the leading edge, trailing edge, and center part, having a fine mesh for the analysis. Geometric parameters are followed in Table1. Blade designs are shown in Figures 2 and 3.Table 1
Design parameters for four-bladed analysis.
ParametersSymbolsHH02NASA SC(2)-0714The number of bladesNb44Rotor blade radiusR5.5745.574Chord lengthC0.1450.145Mach no.M0.3, 0.4, 0.50.3, 0.4, 0.5Blade aspect ratioAR16.916.9Rotor solidityσ0.0330.033Blade span (m)b5.0865.086Blade root chord (m)Cr0.31910.3191Blade root chord (m)Ct0.06630.0663Blade planform area (m2)S1.52991.5299Figure 2
HH02 rotor model.Figure 3
NASA SC(2)-0714 rotor model.Design parameters for the four-bladed configuration are shown in Table1. Required parameters to design the helicopter blades are mainly rotor solidity and aspect ratio of the wing. Analysis was carried out for the Mach numbers of 0.3, 0.4, and 0.5 as most high-speed helicopters operate at Mach number 0.4. HH02 and NASA SC(2)-0714 rotors were analyzed to obtain the flow characteristics.


#### 2.1.2. Grid Independence Study

The grid independence test was conducted to obtain the correct result for verification. Details of the grid independence test are represented in Table2. Y+ value is maintained at 30 to 50 for all analyses. The number of elements increased, and the lift and drag coefficients were obtained from the numerical results. The grid size is decreased, and the study was carried out until the result variation was less than 2% higher than the previous grid [25, 26]. For the elements of 10921474, the lift coefficient and drag coefficient values showed a 1.47% deviation. So, the number of elements is fixed, and the boundary conditions are applied. The plot showing the variation of lift coefficient with many elements is shown in Figure 4.Table 2
Grid independence test.
Domain namesDomain_01Domain_02Domain_03Domain 04The number of elements579200170552441005794710921474Lift coefficient0.898256380.991469611.04036251.0559107Drag coefficient0.0893241540.0822204940.0836408160.08252949Error %N/A9.40%4.69%1.47%Figure 4
Grid independence test for four-bladed analysis.


## 2.1. CFD Analysis

Four-bladed helicopters are chosen as it is used in most high-speed helicopters. Analysis was carried out by ANSYS fluent software. It provides the numerical solution computationally with the help of the Navier–Stokes equation over the flow region. These computational techniques are beneficial for research on complex problems. Four-bladed rotors are designed by CATIA V5 software. Coordinates are downloaded from the UIUC airfoil data site, and the 2D profile can be generated from the airfoil data [17].


### 2.1.1. Design and Preprocessing

The CAD model design is essential and a preliminary requirement for computational analysis. CAD models can be designed by any design software like Solid Works, Solid Edge, and Catia. [18]. Catia V5 is used in this research work to design the three-bladed and four-bladed helicopter blades. The same model is used for the experimental model and scaled down for wind tunnel size using the 3D printing method [19, 20]. The designed model will be imported to fluent for further analysis. After designing the model, the model is imported to the ANSYS CFD workbench for further processing and analysis. The first stage is preprocessing, where the meshing and boundary conditions are applied [21]. As it is a static analysis, C type domain was chosen for computational analysis. Unstructured meshing is chosen as it is a complex structure to provide accurate results in CFD analysis [22, 23]. The realizable k-ε model is chosen as the solver because it provides better results than other models, as mentioned in Figure 1.Figure 1
Turbulence model selection.The analysis was carried out for all the models and compared with experimental results to finalize the turbulence model. The analysis showed that realizablek − ε was near the experimental results [22, 24].An analysis was carried out for HH02 and NASA SC(2)-0714 rotor airfoils for Mach numbers of 0.3, 0.4, and 0.5. Blades are designed with the exact scaling of the Apache helicopter as the HH02 airfoil provides high performance. Parts of the rotors are designed separately as the leading edge, trailing edge, and center part, having a fine mesh for the analysis. Geometric parameters are followed in Table1. Blade designs are shown in Figures 2 and 3.Table 1
Design parameters for four-bladed analysis.
ParametersSymbolsHH02NASA SC(2)-0714The number of bladesNb44Rotor blade radiusR5.5745.574Chord lengthC0.1450.145Mach no.M0.3, 0.4, 0.50.3, 0.4, 0.5Blade aspect ratioAR16.916.9Rotor solidityσ0.0330.033Blade span (m)b5.0865.086Blade root chord (m)Cr0.31910.3191Blade root chord (m)Ct0.06630.0663Blade planform area (m2)S1.52991.5299Figure 2
HH02 rotor model.Figure 3
NASA SC(2)-0714 rotor model.Design parameters for the four-bladed configuration are shown in Table1. Required parameters to design the helicopter blades are mainly rotor solidity and aspect ratio of the wing. Analysis was carried out for the Mach numbers of 0.3, 0.4, and 0.5 as most high-speed helicopters operate at Mach number 0.4. HH02 and NASA SC(2)-0714 rotors were analyzed to obtain the flow characteristics.


### 2.1.2. Grid Independence Study

The grid independence test was conducted to obtain the correct result for verification. Details of the grid independence test are represented in Table2. Y+ value is maintained at 30 to 50 for all analyses. The number of elements increased, and the lift and drag coefficients were obtained from the numerical results. The grid size is decreased, and the study was carried out until the result variation was less than 2% higher than the previous grid [25, 26]. For the elements of 10921474, the lift coefficient and drag coefficient values showed a 1.47% deviation. So, the number of elements is fixed, and the boundary conditions are applied. The plot showing the variation of lift coefficient with many elements is shown in Figure 4.Table 2
Grid independence test.
Domain namesDomain_01Domain_02Domain_03Domain 04The number of elements579200170552441005794710921474Lift coefficient0.898256380.991469611.04036251.0559107Drag coefficient0.0893241540.0822204940.0836408160.08252949Error %N/A9.40%4.69%1.47%Figure 4
Grid independence test for four-bladed analysis.


## 2.1.1. Design and Preprocessing

The CAD model design is essential and a preliminary requirement for computational analysis. CAD models can be designed by any design software like Solid Works, Solid Edge, and Catia. [18]. Catia V5 is used in this research work to design the three-bladed and four-bladed helicopter blades. The same model is used for the experimental model and scaled down for wind tunnel size using the 3D printing method [19, 20]. The designed model will be imported to fluent for further analysis. After designing the model, the model is imported to the ANSYS CFD workbench for further processing and analysis. The first stage is preprocessing, where the meshing and boundary conditions are applied [21]. As it is a static analysis, C type domain was chosen for computational analysis. Unstructured meshing is chosen as it is a complex structure to provide accurate results in CFD analysis [22, 23]. The realizable k-ε model is chosen as the solver because it provides better results than other models, as mentioned in Figure 1.Figure 1
Turbulence model selection.The analysis was carried out for all the models and compared with experimental results to finalize the turbulence model. The analysis showed that realizablek − ε was near the experimental results [22, 24].An analysis was carried out for HH02 and NASA SC(2)-0714 rotor airfoils for Mach numbers of 0.3, 0.4, and 0.5. Blades are designed with the exact scaling of the Apache helicopter as the HH02 airfoil provides high performance. Parts of the rotors are designed separately as the leading edge, trailing edge, and center part, having a fine mesh for the analysis. Geometric parameters are followed in Table1. Blade designs are shown in Figures 2 and 3.Table 1
Design parameters for four-bladed analysis.
ParametersSymbolsHH02NASA SC(2)-0714The number of bladesNb44Rotor blade radiusR5.5745.574Chord lengthC0.1450.145Mach no.M0.3, 0.4, 0.50.3, 0.4, 0.5Blade aspect ratioAR16.916.9Rotor solidityσ0.0330.033Blade span (m)b5.0865.086Blade root chord (m)Cr0.31910.3191Blade root chord (m)Ct0.06630.0663Blade planform area (m2)S1.52991.5299Figure 2
HH02 rotor model.Figure 3
NASA SC(2)-0714 rotor model.Design parameters for the four-bladed configuration are shown in Table1. Required parameters to design the helicopter blades are mainly rotor solidity and aspect ratio of the wing. Analysis was carried out for the Mach numbers of 0.3, 0.4, and 0.5 as most high-speed helicopters operate at Mach number 0.4. HH02 and NASA SC(2)-0714 rotors were analyzed to obtain the flow characteristics.


## 2.1.2. Grid Independence Study

The grid independence test was conducted to obtain the correct result for verification. Details of the grid independence test are represented in Table2. Y+ value is maintained at 30 to 50 for all analyses. The number of elements increased, and the lift and drag coefficients were obtained from the numerical results. The grid size is decreased, and the study was carried out until the result variation was less than 2% higher than the previous grid [25, 26]. For the elements of 10921474, the lift coefficient and drag coefficient values showed a 1.47% deviation. So, the number of elements is fixed, and the boundary conditions are applied. The plot showing the variation of lift coefficient with many elements is shown in Figure 4.Table 2
Grid independence test.
Domain namesDomain_01Domain_02Domain_03Domain 04The number of elements579200170552441005794710921474Lift coefficient0.898256380.991469611.04036251.0559107Drag coefficient0.0893241540.0822204940.0836408160.08252949Error %N/A9.40%4.69%1.47%Figure 4
Grid independence test for four-bladed analysis.


## 3. Wind Tunnel Testing

The wind tunnel is a device used to study and analyze the flow characteristics of any object. The wind tunnel is an essential part of aerodynamics. The wind tunnel used a control stream of air to create an environment where the air flows over the model to study the effects of airflow. Wind tunnel testing is not only applicable to flying vehicles. It is also applicable for automobiles, buildings, or moving or flowing objects [9, 27].Wind tunnels are classified based on the speed of flow created in the test section. Also, based on the operation, it is classified as open section and closed section wind tunnels.Wind tunnels for rotorcraft differ from wind tunnels for aircraft or only wing sections. Static analysis can be carried out in any wind tunnel to analyze rotorcraft blade properties in static conditions [28, 29].This study carried out an open section low-speed subsonic wind tunnel analysis for static analysis of HH02 and NASA SC(2)-0714 rotors.Lift, drag, and moment can be measured by a three-component balance system attached to the wind tunnel. The three-component balance system directly measures the force and moment with the help of strain gauges inside the balance. In modern wind tunnels, a six-component balance is available to measure lateral force, longitudinal force, vertical force, pitching moment, rolling moment, and yawing moment [30].The wind tunnel used for testing and force balance measurement and control for wind tunnels are shown in Figures5(a) and 5(b).Figure 5
Wind tunnel and digital indicator for force measurement. (a) Wind tunnel used for testing; (b) force measurement.
(a)(b)In Figure6, 3D printed models are shown. Four-bladed models of HH02 and NASA SC(2)-0714 rotors are printed with the FDM method. The NASA SC(2)-0714 rotor is the black colored model, and the HH02 rotor is the white colored model. The models are used in the wind tunnel for testing in static conditions.Figure 6
3D printed models used for testing.Computational techniques have developed a lot in recent days. But experimental testing is essential for validating the computational results and rectifying the errors [31]. In previous days, making models are not accessible. Wood or steel models should be used for testing in wind tunnels. Nowadays, 3D printing models are developed, and models can be made easily for the requirement.Wind tunnels are available in various sizes and configurations. The full-scale wind tunnels where a complete aircraft or helicopter can be placed inside the test section. The scaled-down tunnel will have a lesser dimension test section. The actual model should be scaled down based on similarity parameters for scaled-down models. The model is a small one that is scaled-down, and the prototype is a real one with full dimensions. A geometric, kinematic, and dynamic similarity between the model and prototype should exist [32, 33].In this research, an analysis of an isolated blade was taken. The actual size taken for the analysis in CFD is 12m, but the test section size is 0.6 m × 0.6 m. The model was designed for 500 mm with 200 mm blades and a 100 mm hub for the attachment of blades. A scaled-down model of 1 : 24 size was taken for the analysis in a wind tunnel.Models were designed using 3D printing techniques. The development of 3D printing technologies helps easy printing and saves time. There are many techniques available for 3D printing. The fusion diffusion modeling (FDM) technique has been used for this research. FDM is a low-cost 3D printing technique in which surfaces will not be smooth. In FDM, the plastic elements are provided with thermal treatment and melted to the required shape. It has less resolution and accuracy. It can be used for basic research purposes. In stereolithography (SLA) techniques, the laser-cured photopolymer resin to obtain the required model. We can select from many materials, and the method is versatile. The model will be very accurate and high resolution. It helps work with prototypes, molds, and patterns. In the selective laser sintering (SLS) technique, polymer powder is fused with laser. It is less costly than SLA and has good mechanical properties [10, 34]. The SLS method is also used for working prototypes and is more expensive than the FDM model. In selective laser melting (SLM), procedures are similar to the SLS technique. In SLM melting, process and materials are different. In SLM, materials are used as a solidified material layer in an entirely molten state. The actual powder layer is completely or partially melted because of the solid laser technique [8].3D inkjet printing, 3D models are 2D cross section layers by slicing them on a computer. The droplet of the binder material to the powder is applied where the solidification is required as programmed in the computer. Other methods, such as digital light processing (DLP) and carbon DLS technology, are high-cost techniques mainly used for industrial purposes [29].Pressure and velocity are essential parameters in aerodynamics. In wind tunnels, pressure and velocity are measured with the help of a pitot-static tube and manometer. The pitot-static tube measures total or stagnation pressure and static pressure. Dynamic pressure relates the velocity and pressure, so from the relations, we can find the velocity with the help of pressure [27, 28].Pressure is related by(1)P0=P+12ρV2.P0 is the total pressure, and P is the static pressure.From the above relation, velocity can be found as the pitot-static tube measures total and static pressure.So, velocity can be found from(2)V=2P0−Pρ.The manometer measures the static pressure and stagnation pressure head using a pitot-static tube. The expression relates to pressure head (H), and pressure is(3)P=ρgH.So, pressure can be measured by the above relation for stagnation and static pressure. A digital velocity indicator measures velocity, and it is used to calibrate the wind tunnel.Figures7 and 8 show the models are fixed inside the wind tunnel test section. Static conditions are measured by varying the speed in the test section by varying the RPM of the motor. Models are tested at different speeds and validated.Figure 7
HH02 rotor model in the wind tunnel.Figure 8
NASA SC(2)-0714 rotor model in the wind tunnel.


## 4. Results and Discussion

### 4.1. Computational Analysis Results

The models were analyzed for Mach numbers 0.3, 0.4, and 0.5. The lift and drag values calculated for HH02 and NASA SC(2)-0714 airfoils gave comparable results to those reported in the literature. The results show that the supercritical airfoil has a better aerodynamic performance than the HH02 airfoil. The lift, drag, and lift-to-drag ratio values are given in Tables3 and 4 and are represented graphically in Figures 9 and 10. Postprocessing results obtained from fluent provided the aerodynamic characteristics of the rotor blade.Table 3
Result of HH02 rotor.
4Mach number0.30.40.5The number of elements121577471215774712157747Lift coefficient0.396601770.401461450.40501018Drag coefficient0.0589237480.0558159010.054278395L/D ratio6.7307627827.1926000087.461719898Max pressure (N/m2)4462791912350Max velocity (m/s)116.1154.7193.2Table 4
Result of NASA SC(2)-0714 rotor.
4Mach number0.30.40.5The number of elements109214741092147410921474Lift coefficient1.05591071.06597641.0717936Drag coefficient0.082529490.0804216250.078886333L/D ratio12.7943441813.2548478113.58655624Max pressure (N/m2)5023893313960Max velocity (m/s)124.5166207Figure 9
Lift coefficient vs. Mach number.Figure 10
Aerodynamic efficiency vs. Mach no.The NASA SC(2)-0714 rotor exhibits more lift than the HH02 rotor in comparing lift coefficient. But, drag values are also higher for NASA SC(2)-0714 rotor as we know that induced drag occurs in the flow and causes the higher drag values. In comparing aerodynamic efficiency, the NASA SC(2)-0714 rotor shows much higher efficiency than the HH02 rotor.The plot between maximum velocity and free-stream Mach number has been plotted. It shows that velocity is always higher for the supercritical airfoil rotor than the HH02 rotor for all three Mach numbers 0.3, 0.4, and 0.5, as shown in Figure11. From the Velocity values, we can understand that if the velocity is increased, the lift will be increased as the lift is directly proportional to the square of velocity from the basic formula for lift. Similarly, the maximum pressure will be on the lower part of the airfoil from the pressure values, which induces the lift. It proves that NASA SC(2)-0714 airfoil has better lifting capability than HH02 airfoil, as shown in Figure 12.Figure 11
Maximum velocity vs. Mach number.Figure 12
Maximum pressure vs. Mach number.The maximum pressure and velocity values clearly show that NASA SC(2)-0714 shows high pressure and pressure difference. A high-pressure difference indicates high lift-producing capability for NASA SC(2)-0714 rotors. Similarly, NASA SC(2)-0714 achieves more velocity over the surface. Boundary layers are formed around the blades, but we aim to find the aerodynamic performance of the blades. The effects of the boundary layers are not considered. It is also clear that it increases the critical and drag divergence Mach number and delays the shock creation. Owing to the change in the shape of the trailing edge in the supercritical airfoil, it has different characteristics from the HH02 airfoil. HH02 has a curve aft of the trailing edge at the upper surface, whereas NASA SC(2)-0714 has a curve at the bottom surface of the trailing edge.In comparing the aerodynamic performance, for Mach number 0.3, the NASA SC(2)-0714 rotor provides 47% better aerodynamic efficiency than the HH02 rotor. Similarly, for Mach number 0.4, the NASA SC(2)-0714 rotor offers 46% better aerodynamic efficiency than the HH02 rotor. Then, for Mach number 0.5,the NASA SC(2)-0714 rotor provides 45% better aerodynamic efficiency than the HH02 rotor. So, in static analysis, the NASA SC(2)-0714 rotor offers superior performance to the HH02 rotor for all speeds.


### 4.2. Wind Tunnel Results

Models are analyzed for static conditions, as shown in Figures7 and 8. Models are fitted to the three-component balance system attached to the wind tunnel, and the lift force, drag force, and the moment are taken from the force balance system. Readings are tabulated and represented in Tables 5 and 6. Based on the values obtained from wind tunnel testing, aerodynamic efficiency is calculated for NASA SC(2)-0714 and HH02 rotors.Table 5
Wind tunnel readings for HH02 model.
S. no.RPMVelocityLift forceDrag forcePitching momentAerodynamic efficiency (L/D)1500102.210.20.3611.052600143.280.310.5910.583700195.060.40.8912.654800257.020.561.2812.545900309.450.611.6915.49610003511.60.752.1615.46Table 6
Wind tunnel readings for the NASA SC(2)-0714 model.
S. no.RPMVelocityLift forceDrag forcePitching momentAerodynamic efficiency (L/D)1500102.620.150.4317.462600144.610.250.718.443700195.770.31.0119.234800257.890.411.4219.759003010.670.51.9520.92610003512.980.612.4821.18Tables5 and 6 provide lift, drag, and pitching moment values to show that the NASA SC(2)-0714 rotor produces higher lift, lesser drag, and higher aerodynamic efficiency.Figures13 and 14 represent the production of lift force in the wind tunnel for various RPMs for NASA SC(2)-0714 and HH02 rotor models. It clearly shows the velocity increase for an increase in RPM of the wind tunnel rotor fan. Similarly, the lift force is increased for increasing the velocity as the lift force is proportional to the square of the velocity. In comparing figures, the lift produced in the NASA SC rotor is higher than the HH02 rotor, like the computational analysis results obtained from CFD.Figure 13
Lift force measured for the HH02 model.Figure 14
Lift force measured for the NASA SC(2)-0714 model.Figures15 and 16 represent the generation of drag force in the wind tunnel for various RPMs for NASA SC(2)-0714 and HH02 rotor models. It clearly shows that drag force is increased for increasing the velocity as the drag force is proportional to the square of the velocity. In comparing figures and tables, the drag force produced in the NASA SC rotor is lesser than the HH02 rotor, like the computational analysis results obtained from CFD.Figure 15
Drag force measured for the HH02 model.Figure 16
Drag force measured for the NASA SC(2)-0714 model.Figures17 and 18 represent the pitching moment measured in the wind tunnel force measurement for the NASA SC rotor and HH02 rotor. It proves that the pitching moment is higher for the NASA SC rotor than the HH02 rotor as it has a higher lift and better aerodynamic efficiency.Figure 17
Moment measured for the HH02 model.Figure 18
Moment measured for NASA SC(2)-0714 model.A comparison of aerodynamic performance is shown in Table7.Table 7
Comparison of experimental aerodynamic efficiency.
S. no.VelocityL/D NASA SC(2)-0714L/D HH02% of improvement11017.4611.0536.7124921418.4410.5842.6247331919.2312.6534.2173742519.712.5436.3451853020.9215.4925.9560263521.1815.4627.00661The comparison table clearly shows that aerodynamic efficiency for the NASA SC(2)-0714 rotor was better than the HH02 rotor by around 30 to 40% for all speeds. The CFD results obtained for static analysis also showed an improved aerodynamic efficiency of approximately 40%, as discussed in Chapter 4. The experimental data also proved that the NASA SC(2)-0714 rotor provides better aerodynamic performance than the HH02 rotor.


### 4.3. Validation of Results

For the model’s dynamic similarity, force on the model should be equal to the force on the prototype.Force can be calculated from the relation for dynamic similarity as follows:(4)forceprototype=forcemodelF=m×a,where “m” is the mass of the body and “a” is the acceleration due to gravity.Mass flow rate is given as follows:(5)ṁ=ρAV.From the relation, it can be found that it is a scaled model from 120 cm to 5 cm. The scale ratio is 24.(6)Fprototype=641.45×Fmodel.The CFD and experimental values are compared from the calculation of force as in Tables8 and 9.Table 8
Validation of the NASA SC(2)-0714 rotor.
S. no.VelocityExp lift forceCFD lift force% of error1101680.599198015.122142957.085338412.613193701.167425112.934255061.041575412.045306844.27276009.946358326.02190518.01Table 9
Validation of HH02 rotor.
S. no.VelocityExp lift forceCFD lift force% of error1101417.605168415.812142103.956245014.123193245.737375113.474254502.979515212.595306061.703674210.096357440.8282019.27The validation of results was mentioned in Tables8 and 9. It compares the percentage of changes in the CFD and experimental results. It shows that the error percentage is around 15% for lower speed and much lesser at higher speed. It is around 9% for both NASA SC rotor and HH02 rotor. Figures 19 and 20 represent the validation of results and change of nature of the curve for both NASA SC(2)-0714 and HH02 rotors with the CFD and experimental results.Figure 19
Validation of the NASA SC(2)-0714 rotor.Figure 20
Validation of the HH02 rotor.


## 4.1. Computational Analysis Results

The models were analyzed for Mach numbers 0.3, 0.4, and 0.5. The lift and drag values calculated for HH02 and NASA SC(2)-0714 airfoils gave comparable results to those reported in the literature. The results show that the supercritical airfoil has a better aerodynamic performance than the HH02 airfoil. The lift, drag, and lift-to-drag ratio values are given in Tables3 and 4 and are represented graphically in Figures 9 and 10. Postprocessing results obtained from fluent provided the aerodynamic characteristics of the rotor blade.Table 3
Result of HH02 rotor.
4Mach number0.30.40.5The number of elements121577471215774712157747Lift coefficient0.396601770.401461450.40501018Drag coefficient0.0589237480.0558159010.054278395L/D ratio6.7307627827.1926000087.461719898Max pressure (N/m2)4462791912350Max velocity (m/s)116.1154.7193.2Table 4
Result of NASA SC(2)-0714 rotor.
4Mach number0.30.40.5The number of elements109214741092147410921474Lift coefficient1.05591071.06597641.0717936Drag coefficient0.082529490.0804216250.078886333L/D ratio12.7943441813.2548478113.58655624Max pressure (N/m2)5023893313960Max velocity (m/s)124.5166207Figure 9
Lift coefficient vs. Mach number.Figure 10
Aerodynamic efficiency vs. Mach no.The NASA SC(2)-0714 rotor exhibits more lift than the HH02 rotor in comparing lift coefficient. But, drag values are also higher for NASA SC(2)-0714 rotor as we know that induced drag occurs in the flow and causes the higher drag values. In comparing aerodynamic efficiency, the NASA SC(2)-0714 rotor shows much higher efficiency than the HH02 rotor.The plot between maximum velocity and free-stream Mach number has been plotted. It shows that velocity is always higher for the supercritical airfoil rotor than the HH02 rotor for all three Mach numbers 0.3, 0.4, and 0.5, as shown in Figure11. From the Velocity values, we can understand that if the velocity is increased, the lift will be increased as the lift is directly proportional to the square of velocity from the basic formula for lift. Similarly, the maximum pressure will be on the lower part of the airfoil from the pressure values, which induces the lift. It proves that NASA SC(2)-0714 airfoil has better lifting capability than HH02 airfoil, as shown in Figure 12.Figure 11
Maximum velocity vs. Mach number.Figure 12
Maximum pressure vs. Mach number.The maximum pressure and velocity values clearly show that NASA SC(2)-0714 shows high pressure and pressure difference. A high-pressure difference indicates high lift-producing capability for NASA SC(2)-0714 rotors. Similarly, NASA SC(2)-0714 achieves more velocity over the surface. Boundary layers are formed around the blades, but we aim to find the aerodynamic performance of the blades. The effects of the boundary layers are not considered. It is also clear that it increases the critical and drag divergence Mach number and delays the shock creation. Owing to the change in the shape of the trailing edge in the supercritical airfoil, it has different characteristics from the HH02 airfoil. HH02 has a curve aft of the trailing edge at the upper surface, whereas NASA SC(2)-0714 has a curve at the bottom surface of the trailing edge.In comparing the aerodynamic performance, for Mach number 0.3, the NASA SC(2)-0714 rotor provides 47% better aerodynamic efficiency than the HH02 rotor. Similarly, for Mach number 0.4, the NASA SC(2)-0714 rotor offers 46% better aerodynamic efficiency than the HH02 rotor. Then, for Mach number 0.5,the NASA SC(2)-0714 rotor provides 45% better aerodynamic efficiency than the HH02 rotor. So, in static analysis, the NASA SC(2)-0714 rotor offers superior performance to the HH02 rotor for all speeds.


## 4.2. Wind Tunnel Results

Models are analyzed for static conditions, as shown in Figures7 and 8. Models are fitted to the three-component balance system attached to the wind tunnel, and the lift force, drag force, and the moment are taken from the force balance system. Readings are tabulated and represented in Tables 5 and 6. Based on the values obtained from wind tunnel testing, aerodynamic efficiency is calculated for NASA SC(2)-0714 and HH02 rotors.Table 5
Wind tunnel readings for HH02 model.
S. no.RPMVelocityLift forceDrag forcePitching momentAerodynamic efficiency (L/D)1500102.210.20.3611.052600143.280.310.5910.583700195.060.40.8912.654800257.020.561.2812.545900309.450.611.6915.49610003511.60.752.1615.46Table 6
Wind tunnel readings for the NASA SC(2)-0714 model.
S. no.RPMVelocityLift forceDrag forcePitching momentAerodynamic efficiency (L/D)1500102.620.150.4317.462600144.610.250.718.443700195.770.31.0119.234800257.890.411.4219.759003010.670.51.9520.92610003512.980.612.4821.18Tables5 and 6 provide lift, drag, and pitching moment values to show that the NASA SC(2)-0714 rotor produces higher lift, lesser drag, and higher aerodynamic efficiency.Figures13 and 14 represent the production of lift force in the wind tunnel for various RPMs for NASA SC(2)-0714 and HH02 rotor models. It clearly shows the velocity increase for an increase in RPM of the wind tunnel rotor fan. Similarly, the lift force is increased for increasing the velocity as the lift force is proportional to the square of the velocity. In comparing figures, the lift produced in the NASA SC rotor is higher than the HH02 rotor, like the computational analysis results obtained from CFD.Figure 13
Lift force measured for the HH02 model.Figure 14
Lift force measured for the NASA SC(2)-0714 model.Figures15 and 16 represent the generation of drag force in the wind tunnel for various RPMs for NASA SC(2)-0714 and HH02 rotor models. It clearly shows that drag force is increased for increasing the velocity as the drag force is proportional to the square of the velocity. In comparing figures and tables, the drag force produced in the NASA SC rotor is lesser than the HH02 rotor, like the computational analysis results obtained from CFD.Figure 15
Drag force measured for the HH02 model.Figure 16
Drag force measured for the NASA SC(2)-0714 model.Figures17 and 18 represent the pitching moment measured in the wind tunnel force measurement for the NASA SC rotor and HH02 rotor. It proves that the pitching moment is higher for the NASA SC rotor than the HH02 rotor as it has a higher lift and better aerodynamic efficiency.Figure 17
Moment measured for the HH02 model.Figure 18
Moment measured for NASA SC(2)-0714 model.A comparison of aerodynamic performance is shown in Table7.Table 7
Comparison of experimental aerodynamic efficiency.
S. no.VelocityL/D NASA SC(2)-0714L/D HH02% of improvement11017.4611.0536.7124921418.4410.5842.6247331919.2312.6534.2173742519.712.5436.3451853020.9215.4925.9560263521.1815.4627.00661The comparison table clearly shows that aerodynamic efficiency for the NASA SC(2)-0714 rotor was better than the HH02 rotor by around 30 to 40% for all speeds. The CFD results obtained for static analysis also showed an improved aerodynamic efficiency of approximately 40%, as discussed in Chapter 4. The experimental data also proved that the NASA SC(2)-0714 rotor provides better aerodynamic performance than the HH02 rotor.


## 4.3. Validation of Results

For the model’s dynamic similarity, force on the model should be equal to the force on the prototype.Force can be calculated from the relation for dynamic similarity as follows:(4)forceprototype=forcemodelF=m×a,where “m” is the mass of the body and “a” is the acceleration due to gravity.Mass flow rate is given as follows:(5)ṁ=ρAV.From the relation, it can be found that it is a scaled model from 120 cm to 5 cm. The scale ratio is 24.(6)Fprototype=641.45×Fmodel.The CFD and experimental values are compared from the calculation of force as in Tables8 and 9.Table 8
Validation of the NASA SC(2)-0714 rotor.
S. no.VelocityExp lift forceCFD lift force% of error1101680.599198015.122142957.085338412.613193701.167425112.934255061.041575412.045306844.27276009.946358326.02190518.01Table 9
Validation of HH02 rotor.
S. no.VelocityExp lift forceCFD lift force% of error1101417.605168415.812142103.956245014.123193245.737375113.474254502.979515212.595306061.703674210.096357440.8282019.27The validation of results was mentioned in Tables8 and 9. It compares the percentage of changes in the CFD and experimental results. It shows that the error percentage is around 15% for lower speed and much lesser at higher speed. It is around 9% for both NASA SC rotor and HH02 rotor. Figures 19 and 20 represent the validation of results and change of nature of the curve for both NASA SC(2)-0714 and HH02 rotors with the CFD and experimental results.Figure 19
Validation of the NASA SC(2)-0714 rotor.Figure 20
Validation of the HH02 rotor.


## 5. Conclusion

The analysis was carried out using the CFD techniques, and the results were validated by wind tunnel testing. Results are validated, and the error percentage is 8% to 15%. Error percentages are within the acceptable limit, and the results are validated with the help of the experimental data. Additive manufacturing techniques are very useful in the aerodynamic testing and validation of models. The computational analysis was successfully validated using an additive manufacturing technique and proved to be the same with acceptable error limits. Error percentage can be minimized by using other methods like SLA or SLS techniques to reduce the error percentage in experimental results. FDM methods are economical, and it has some roughness over the model. SLS and SLA technique models will have a good finishing and smooth surface. The error in the experimental analysis may be due to the roughness of the FDM method in additive manufacturing. The analysis may be conducted with other additive manufacturing techniques and can be tested in the future for accurate results. Aerodynamic performance can be compared and studied for better additive manufacturing models.



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*Source: 4052208-2022-09-21.xml*