wind tunnel experiments on double delta wing model

Journal of Aeronautical -Science and Engineering-, Vol.4

November 30, 2015

1 Published by International Society of Ocean, Mechanical and Aerospace Scientists and Engineers

Wind Tunnel Experiments on Double Delta Wing Model

Mohamad Farhan Bin Mohamad Isaa, Kannan Perumalb, Mazuriah Saidc and Shabudin Bin Matd,*

a,b,c,d)

Department of Aeronatical, Automotive and Offfshore Engineering, University Teknologi Malaysia, Johor Bahru, Malaysia [email protected], [email protected], [email protected] *Corresponding author: [email protected] Paper History Received: 8-October-2015 Received in revised form: 10-November-2015 Accepted: 30-November-2015

ABSTRACT This paper presents the results of a static wind tunnel experiments carried out on a double delta wing of 650/250 configuration in Universiti Teknologi Malaysia Low Speed Wind Tunnel (UTM-LST). In this project, two measurement techniques were employed on the model. The fist technique was the static balance data followed by flow visualization technique using smoke and tufts. The experiments were carried out at 5 different angles of attack and 3 different Reynolds numbers of 0.5, 1.0 and 1.13 x 106 based on mean aerodynamic chord respectively. This paper provides a better insight into the complexity of flow behaviours above double delta-shaped wing. More experiments or numerical study using modern measurement techniques are necessary in the future. KEYWORDS : Double Delta, Wind Tunnel Test; Flow Visualization and Separation; Vortex, Forces and moments. NOMENCLATURE: Cl Coefficient of Lift Cd Coefficient of Drag CM Coefficient of Moment Re Reynolds Number α Angle of Attack (Pitch Angle) 1.0 INTRODUCTION

A delta wing is a wing shape when viewed from top like a Greek symbol (Δ) forms like a triangle. It sweeps sharply back from the

fuselage with the angle between the leading edge of the wing often high as 60 degrees and the angle between the fuselage and the trailing edge of the wing mostly around 90 degrees. The early work of this triangular shaped delta wing or was done in Germany during the 1930s. Nowadays, delta wing has been a common planform and is widely used in many different types of high speed civil and military aircraft. A double delta wing is essentially a delta wing with a “kink” in its leading edges that forms the shoulder where the leading edges of the strake (or Leading Edge Extension, LEX) and main wing intersect (Nettelbeck, 2008).The double delta wing consists of leading edge extensions or strakes in front of the main wing. The most important configuration for double delta wing is the sweep angle combinations in order to achieve sufficient supersonic cruise performance. The strakes usually are highly swept to provide a stable vortex and the main wing will be less leading edge sweep.

Figure 1 shows the configuration of the double delta wing 65o/25o developed in Universiti Teknologi Malaysia Low Speed Wind Tunnel Facility. This research is carried to estimate the aerodynamics characteristics of the model experimentally. Two different experiments were carried out in this research which were the balanced measurement test (forces and moments measurement) and flow visualization (smoke and tufts) technique. This research is important to estimate the aerodynamic characteristic of the double delta shaped wing model. Flow visualization techniques using smoke and tufts were also performed to predict the trend of primary vortex occurred above the main and strake wings. The experiments were also performed at several speeds in order to obtain the Reynolds number effects.


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Figure 1: 65º/25º Double Delta Wing model 2.0 DOUBLE DELTA WING FLOW FIELDS Aerodynamic investigation of flow over delta wing configurations have been performed for many years. Typical fact that known by previous research is that the flow starts to separate even at low angles of attack for the highly swept leading edges (Breitsamter, 2012). The flow over the upper surface of a delta wing is a vortex dominated flow field. The vortex formed attaches to the upper surface of the wing. The flow can be described as a movement of a part of the flow from the lower to the upper surface into spiral type of motion which then forms as a vortex flow over the upper surface as shown in Figure 2, (Hesamodin et al,1995). Verhagen, Jenkins, Kern, and Washburn (1992) in their study showed that when α < 10o, the two vortices remained separated and hardly interacts. Beyond this angle of attack, the interaction between the two vortices became more pronounced. This was believed to indicate that the breakdown of the strake vortex was causing the wing vortex to burst.

Figure 2: Vortex Flow around Double Delta Wing

(Hesamodin et al,1995).

Another important characteristic in delta wing study is the vortex breakdown. Lu Zhi-Yong[2004] in his study showed that there are two types of vortex breakdown – one is in bubble form and the other is called as spiral form. Bubble form of breakdown occurs because rapid expansion of the core forming a bubble-like structure that is nearly axisymmetric while for a spiral form, the

vortex centerline deforms into a spiral without inappreciable growth in core size. Figure 3 shows the spiral and bubble form of vortex breakdown. The main focus of this project were to estimate the flow pattern on the upper surface, aerodynamic loads and derivatives calculation such as coefficient of drag, lift and moments.

Figure 3: Form of vortex breakdown

Lu Zhiyong et al, 2004).

3.0 DESIGN PROCESS OF THE DOUBLE DELTA SHAPED WING MODEL

The design process of this double delta wing consists of three different phases which are conceptual design, preliminary design and detail design. The conceptual design process starts with a set of requirement and specification for the model. At this stage, several considerations such as specifications, parameters and wind tunnel dimensions were considered. Finally the 650/250 configuration was chosen for this research as the shape rarely used by other researches, thus the results can provide an initial data for further investigation in the future. Next was the preliminary design which involves the construction of the model. The detail design was done using Computer Aided Design software (both SolidWorks and AutoCAD). Table 1 showed the dimension of the model. It has the wing area of 0.57 m2 and wing span of 1.7716 m respectively. The model is mounted to UTM-LST 3 strut supports system shown in Figure 4.

Table 1 Dimensions of the double delta wing model

Dimensions Value

Leading edge sweep angle 65º/25º

Wing area (m2) 0.5711

Wing span (m) 1.7716

Root Chord, c (m) 1.0

Thickness, t (m) 0.036

Mean aerodynamic chord (m) 0.5


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Figure 4: The installation of UTM double delta wing in UTM – LST wind tunnel.

3.1 Model Fabrication

In this project, the model was fabricated from wood. Due to the size of the model, the model was machined from 3 main parts as shown in Figure 5. The final dimensions has the span of 1.77 meter and the maximum chord of 1.0 meter.

Figure 5: The different segments and dimensions (in mm) of the test model.

Each parts were assembled using adhesive wood, then several aluminum plates were attached to further strenghthen the model, this is shown shown in Figure 6 below. The final model was coated with a black paint for flow visualization studies. The final model is shown in Figure 7.

Figure 6 Complete assembled model and locations of aluminium plates.

Figure 7:The final model

4.0 WIND TUNNEL EXPERIMENTS

The studies were conducted in 2.0m (width) x 1.5m (height) x 5.8m (long), closed-return wind tunnel type, at Universiti Teknologi Malaysia Aeronautics Laboratory Low Speed Wind Tunnel (UTM-LST). The tunnel is able to generate airflow up to Mach 0.23 or 80 m/s wind speed inside the test section. It has the flow and temperature uniformity less than 0.15% to 0.20% respectively and also turbulence less than 0.06%.

During the experiments, the model was mounted to the test section through three struts support system shown in Figure 8. This three struts support system is linked with heavy capacity strain gauge measurement or called as external balance located underneath the test section. During the experiments, steady characteristics of the model were recorded by this strain gauge automatically. The complete installation of the model in the UTM Low Speed Wind Tunnel (UTM-LST) is shown below.


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Figure 8: The Installation of UTM double delta wing model in 2 x 1.5 meter Wind Tunnel

The experiments were conducted at three Reynolds numbers of 0.5 x 106, 1.0 x 106, and 1.13 x 106 based on the mean aerodynamic chord. The model was tested at 5 different angle of attack of α = 00, 50, 100,150,200.

During the experiments, two set of data were captured. There were;

i. Wind-off, the model’s pitching angles were varies from 0° to 20°. This test is required to obtain the tare values of the model for the forces and moments. The data were recorded simultaneously.

ii. Wind-on, at speed of 15.26 m/s, 30.51 m/s and 33m/s, the model’s pitching angles were varied from 0° to 20°, the data were also recorded simultaneously.

To obtain the aerodynamic data of the model, the wind-off configuration will be deducted from the wind on data. Wind tunnel blockage correction is then applied for all data based on the techniques discussed by Cooper, 2001.

The aerodynamic forces are much depend on the Reynolds number and Mach number which is related to the wind speed and the shape (dimension) of the model. It important to select the appropriate wind test speed to obtain the required Reynolds number. The Reynolds number is the ratio of inertial forces (resistant to motion) to viscous forces. Reynolds number can be expressed in equation form as below:

��

(1)

Where � is air density, V is velocity,Mac is mean aerodynamic chord of the model and is dynamic viscosity. The known values are the air density, mean aerodynamic chord and dynamic viscosity which are constant values. By manipulating the above equation, the required Reynolds number for certain experiment can be obtained. In this project, the model was tested at 3 different wind speeds of 15.26, 30.51 and 33 m/s that equivalent to 0.5 x 106 , 1.0 x 106 and 1.13 x 106 Reynolds numbers.

Table 2: Approximate wind speed for the wind tunnel experiments.

Reynolds Number, Re Wind speed approximation

0.5 x 106 15.26 m/s

1.0 x 106 30.51 m/s

1.13 x 106 33.00 m/s

4.1 Smoke Sheets Technique The smoke sheets technique requires a smoke generator to form a vaporised column of smoke that flows over the model’s surface. This will produce a visible smoke sheet over the upper surface of the model. Photos were taken to capture the surface flow patterns during and after the experiments. Videos also were taken for analysis to validate the experiments result. This method can be used to determine and estimate the location of the stagnation, attachment point and separation lines over a range of angle of attack under static test condition (Verhaagen, 1995). In this project, the smoke was injected in front of the model and the images were taken at every angle of attack. The results for this study are discussed in section 5.2

4.2 Tufts Flow Visualization Technique Tufts flow visualization is one the old visualization technique that used in flight test and wind tunnel experiments. Tufts are small length strings that are attached to the surface of the model by using some adhesive such as glue or tape. Common materials for tufts are monofilament nylon and polyester or cotton sewing thread. As the wind flow over the model, the tufts are blown and move in the direction of the wind flow. Photos and videos are also taken for comparison.

The advantages of this technique are easy to install and provides a view of the flow pattern over a large area while the disadvantage of this technique is that it does not provide a detailed flow pattern since they are constantly moving with the wind flow. In this project, the tufts were place at certain chordwise position of 0.2, 0.4, 0.6, and 0.8 percent of the trailing edge. The results for this study are discussed in section 5.3. 5.0 RESULT AND DISCUSSION 5.1 Balanced Measurement Data The uncorrected and corrected results in the co-efficient form are presented at three different Reynolds numbers and five different angles of attack (pitch angles) are tabulated in Tables 3, 4 and 5. The raw data were corrected accordingly using corrections methods discussed by Cooper, 2001.

Table 3: Wind Tunnel Experiment Data for Re=0.5 x 106

(α°) Uncorrected Corrected

CD CL CM CD CL CM 0 0.0697 0.1042 -0.0121 0.0854 0.1277 -0.0148 5 0.0723 0.6274 -0.0909 0.0885 0.7683 -0.1113 10 0.0982 0.8694 -0.1527 0.1201 1.0632 -0.1867 15 0.2320 1.0023 -0.2093 0.2817 1.2172 -0.2542 20 0.3397 1.0327 -0.2556 0.4102 1.2471 -0.3086


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Table 4: Wind Tunnel Experiment Data for Re=1.0 x 106 (α°) Uncorrected Corrected

CD CL CM CD CL CM 0 0.0656 -0.052 0.0307 0.0804 -0.064 0.0377 5 0.0720 0.2846 -0.0484 0.0881 0.3485 -0.0592 10 0.0901 0.5879 -0.1145 0.1102 0.7189 -0.1400 15 0.1842 0.7472 -0.1637 0.2238 0.9074 -0.1988 20 0.2990 0.8928 -0.2161 0.3611 1.0781 -0.2609

Table5: Wind Tunnel Experiment Data for Re=1.13 x 06 (α°) Uncorrected Corrected

CD CL CM CD CL CM 0 0.0674 0.0142 0.0199 0.0825 0.0174 0.0244 5 0.0705 0.2424 -0.0761 0.0863 0.2969 -0.0931 10 0.1027 0.5424 -0.1507 0.1256 0.6633 -0.1843 15 0.2048 0.8143 -0.2096 0.2488 0.9889 -0.2546 20 0.3296 0.9027 -0.2452 0.3980 1.0901 -0.2961

Figure 9, 10 and 11 show the steady characteristics of the

model at certain angle of attack. Figure 9 showed the uncorrected lift coefficients versus angle of attack for its corresponding Reynolds number. From the figure, the results showed the coefficient of lift increases with the decreasing in Reynolds number. Similar trend is also observed for the drag coefficient (In figure 10). This is consistent with S.Mat (2010), this situation happens because at low Reynolds number, the flow dominated by the laminar region rather than turbulent region. Lower ability of laminar separation to sustain the adverse pressure gradient has promotes early separation and enlarge the size of the primary vortex, thus increasing the Coefficient of Lift and drag. The characteristic of pitching moment is shown in Figure 11 (i & ii). The Figures show a consistent trend where the nose down pitching moment increases with angle of attack. The effects of Reynolds number cannot be explain at this condition, more experiments is required to verify this results. However, the negative slope of the graph indicates that the double delta wing in this experiment is statically stable.

Figure 9 i): Uncorrected CL VS α

Figure 9ii): Corrected CL VS α

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 5 10 15 20 25

Coe

ffici

ent o

f lift

Angle of attack

Graph of uncorrected CL vs angle of attack

Re=0.5MRe=1.0MRe1.13M

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

0 5 10 15 20 25

Coe

ffic

ient

of l

ift

Angle of attack

Graph of corrected CL vs angle of attack

Re=0.5M

Re=1.0M

Re=1.13M


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Figure 10 i): Uncorrected CD VS α

Figure 10 ii): Corrected CD VS α

Figure 11 i) : Uncorrected CM VS α

Figure 11 ii): Corrected CM VS α

5.2 Flow visualization Test Result Figure 12 shows the smoke sheets flow over the upper surface of the model at different angles of attack, the images presented in this paper is the experiment conducted at 1 x 106 Reynolds number. It can be seen that as the angle of attack is increased, the size of the primary vortex is also increased. The vortex breakdown can also be seen on the flow field structure at higher angles of attack starting from 15° and clearly observed at 20°.Within this region, the flow becomes stagnant and turns into an unsteady and unstructured flow. The vortex structure then starts to breakdown at certain point along the chord-line of the model. It is depends on the angle of attack and this phenomenon is called vortex breakdown. The vortex breakdown occurs at the trailing edge and moves upstream as the angle of attack increases.

0.000

0.050

0.100

0.150

0.200

0.250

0.300

0.350

0.400

0 5 10 15 20 25

Coe

ffic

ient

of D

rag

Angle of attack

Graph of uncorrected CD vs angle of attack

Re=0.5M

Re=1.0M

Re=1.13M

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0 5 10 15 20 25

Coe

ffic

ient

of d

rag

Angle of attack

Graph of corrected CD vs angle of attack

Re=0.5M

Re=1.0M

Re=1.13M

-0.30

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0 5 10 15 20 25

Coe

ffic

ient

of p

itchi

ng m

omen

t

Angle of attack

Graph of uncorrected CM vs angle of attack

Re=0.5M

Re=1.0M

Re=1.13M

-0.35

-0.30

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0 5 10 15 20 25

Coe

ffic

ient

of p

itchi

ng m

omen

t

Angle of attack

Graph of corrected CM vs angle of attack

Re=0.5M

Re=1.0M

Re=1.13M


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Figure 12: Smoke flow visualization at different angles of attack

at Reynolds number of 1 x 106 5.3 Tufts Flow visualization Results Tufts flow visualization was also conducted at different Reynolds number and angles of attack ranging from 0º to 20º, at interval of 5°. The sample of the images taken at the Reynolds number of 1 x 106 is shown in Figure 13. At low Reynolds number, there were not much tufts movement can be seen. However at higher wind speed, the tufts on the upper surface of the model were seen actively. The disadvantage of this tuft flow visualization technique is that it does not provide a detailed flow pattern. This technique only provides a general view of the flow and vortices.

Figure 13: Tuft flow visualization at different angle of attack at 1

x 106 Reynolds number 6.0 CONCLUSION A series of wind tunnel experiment methods were conducted to study the flow over a 65º/25º double delta wing model at the angles of attack ranging from 0º to 20º (at 5° interval) and Reynolds numbers varies from 0.5 x 106 to 1.13 x 106.

This experiment provides an initial data for the aerodynamics of double delta wing model in the future. The results are reasonable and provide a consistent trend with several papers published recently. However, more experiments using modern measurement techniques such as PIV and high speed camera are necessary to observe the complexity of the vortex.

REFERENCES

1. Breitsamter, A. F. C. (2012). Turbulent and Unsteady Flow Characteristics of Delta Wing Vortex Systems.

2. HesamodinEbnodinHamidi, M. R. (2011). Numerical Investigation of High Attach Angle Flow on 76o/45o Double Delta Wing in Incompressible Flow, World Academic of Science, Engineering and Technology.

3. K.R Cooper 2001, A summary of classical Blockage Correction for Aircraft Models in Closed-Wall Wind


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Tunnels, Aerodynamics laboratory, NRC, Otaawa Canada 00/06/01

4. Lu Zhiyong, Zhu Lirguo (February 2004), Study on Forms of Vortex Breakdown over Delta Wing

5. Mat, S. The analysis of flow on round-edged delta wing, PhD dissertation, University of Glasgow, UK 2010

6. Nettelbeck, C. (2008). Dynamic Analysis of a Double Delta Wing in Free Roll. School of Aerospace, Civil and Mechanical Engineering, (BE)

7. Verhaagen N. G., Jenkins L. N., Kern S. B., and Washburn A. E. A, Study of the vortex flow over a 76/40-deg double-delta wing. AIAA-1992-279 33rd Aerospace Sciences Meeting and Exhibit, Reno NV, 1995.

wind tunnel experiments on double delta wing model

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