Wind Tunnel Testing Laboratory Report Using Aerofoil Models
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Course
Instructor
Date Due
Table of contents
Introduction ………………………………………………………………………………….
Objective…………………………………………………………………………………….
Apparatus and Procedure……………………………………………………………………
Apparatus……………………………………………………………………………..
Procedure…………………………………………………………………………
Results……………………………………………………………………………………….
Discussion…………………………………………………………………………………..
Application question…………………………………………………………………………
Conclusion…………………………………………………………………………………….
List of tables
Table1: 20 m/s L/D and AoA………………………………………………………………….
Table 2: 30 m/s L/D and AoA…………………………………………………………………
Table 3: 20 m/s Coefficient of Lift……………………………………………………………
Table 4: 30 m/s Coefficient of Lift……………………………………………………………..
Table 5: Literature data on wind tunnel parameters………………………………………….
List of graphs
Graph 1: 20 m/s L/D against AoA……………………………………………………………
Graph 2: 30 m/s L/D against AoA……………………………………………………………
Graph 3: 20 m/s Coefficient of Lift…………………………………………………………
Graph 4: 30 m/s Coefficient of Lift…………………………………………………………..
Graph 5: A plot of CL against alpha, deg…………………………………………………….
Graph 6: A plot of CD against CL………………………………………………………
References…………………………………………………………………………………
Introduction
Wing tunnel is an aerodynamic structure that is responsible of the generation of lift upon coming into contact with molecules of air that are moving i.e. wind. The generation of the lift is greatly attributed to the unique shape of the wing. It is usually almost flat on the bottom surface and curved on the upper surface. This unique form ensures that the air go faster over the top of the wing than the bottom. This speed differences results in a difference in pressure between the bottom and top of the wing thereby exerting an upward net force on the wing creating an upward force referred to as the lift. However, the amount of lift achieved from the wing is greatly dependent on the aerofoil shape and its angle of incidence. Therefore, a relationship is usually created between the angle at which the wing is permanently inclined to the longitudinal axis of the airplane and the amount of lift that is generated. For instance, at small angles the lift is also small, and an increase in the angle of attack the lift increases; but at a certain points the wing drag dominates the lift causing the aircraft to stop. Each section of the wing consists of a certain aerofoil usually classified as either conventional or laminar.
This report is to present an introduction into the wing’s structure and theory of wings. It also includes background information concerning wind tunnels and wind tunnel testing as well as describing the testing procedure using aerofoil which presents a graphs and tables evaluating the data obtained through the conducted tests.
An airplane that is in flight is always in a tug of was that continuously maintained between four forces such as the lift, thrust, gravity force or weight, and drag. However, the drag and lift are regarded as the aerodynamic forces due to the fact that they exist as a result of the aircraft movement through the air. The plane is however pulled down by the weight opposing the lift that result from air flowing over the wing. Generation of thrust is caused by the propeller and opposes drag which is as a result of air resistance to the airplane’s frontal area. During the airplane’s take off, the drag must be overpowered by the thrust and lift must overcome the weight prior to the aircraft becoming airborne. In a constant speed level flight, thrust is exactly equivalent to the drag and lift is exactly equivalent to the gravity force or weight. For landings, a reduction of thrust must be ensured to be below the drag level and lift below the weight or gravity force level. This is mainly because the lift is produced by a lower pressure created on the airplane’s upper surface of the wing in comparison to the pressure on the airplane’s lower surface of the wing, causing the wing to be upwardly lifted. The special airplane wing’s shape (aerofoil) is designed so that air that flow over travels a greater distance faster, creating a lower pressure area thus leading to the lifting of the wing upward.
The aerofoils are in common use in the design of airplanes. These include the laminar flow and convectional aerofoils that were originally developed in order to make an airplane fly. Moreover, the angle of incidence is referred to as the angle at which the wing is inclined permanently to the longitudinal axis of the airplane. Choosing the angle of incidence appropriately can greatly improve flight visibility; enhance landing and take-off and landing as well as reducing drag in level flight. Therefore, the usually chosen angle of incidence is the angle of attack at which the lift-drag ratio is optimum. For instance, in most modern airplanes, the positive angle of incidence is small in order to make sure that the wing has a slight angle of attack when the airplane is in level cruising flight.
Moreover, wind tunnel testing is a crucial step in an aircraft design. This is due to the fact that it is likely to give accurate information on an aircraft’s performance or a section of an aircraft collecting data on a scale model. This is very crucial since it has the possibility of saving enormous amounts of money through testing models rather than using prototypes. In addition, it is also much safer to test aerofoils in a wind tunnel instead of testing them out in the open. Thus, it wind tunnel testing of the aerofoils has been of great importance in the design of aircrafts over a long period of time due to the fact that it offers the most appropriate alternative for making sure that the most effective ones are used.
Objective
The objective of the present low speed wind tunnel laboratory experiment is to reinforce the principles presented in the aerodynamics part of the lecture notes and illustrate the equations hold in the physical environment. The chosen set of wind tunnel models are simple 2D aerofoils with symmetric profile. Studying these aerodynamic bodies in the wind tunnel and collecting load and pressure data via the data acquisition system will provide basic exposure to standard facilities and experimental techniques used in aerospace engineering.
Apparatus and Procedure
Apparatus
Health & Safety
Risk Assessment, Ear Defenders, Emergency Stop
Wind Tunnel (Sheaf 4L20)
Contraction, Working Section, Diffuser, Turbine, Silencer, NACA0012 Wind Tunnel Models, Aerofoil
Diagnostics
3 Component Balance, Water Pressure Manometer, 36 Channels Pressure Transducers, Speed Controller, Pitot Static Tube & Vertical Positioning System, Differential Pressure, Transducers & Wind Speed Controller and Balance Calibration Unit & Coms
Representative Data
Aerodynamic Load Data (Lift & Drag), Pressure Distribution, Validation of Experimental Data with CFD and Data Acquisition System: Graphical User Interface
Procedure
- It was ensured that the electrical supply to the wind tunnel was disconnected
- One of the side panels in the working section had a large model holder that has three locking screws that was left in place, but the other panel was removed.
- It was also ensured that the pitot tubes that were in the working section were out of the way.
- The model support shaft was then slide into the model holder that was in the panel on the other side. The model was then rotated so that the leading edge faced into air flow (towards the wind tunnel’s inlet). The model was further rotated in order to make sure that the trailing edge was at same height as the model’s centre line which was nominally 153 mm. the three lock screws present at the model holder were tightened.
- A protractor was placed onto the support shaft followed by its rotation so that the pointer for the model holder was at zero.
- The lock screw of the protractor was then tightened.
- The blanking plug was then removed from the panel that had been removed in step 2.
- Flexible tappings were then carefully fed through the hole in the panel followed by refitting of the side panel.
- The long screws (supplied) were then used to fit the manifold plate to the working section’s bottom and the tapping pipes were connected to the front sockets. This was made easier by adhering to the numbers assigned to each pipe.
- The supplied pipe was then used in the connection of the outlets at the back of the manifold to manometer that was suitably located.
- This set the stage for testing procedure which begun with creation of a blank table of results and the ambient pressure and temperature recorded.
- The wind tunnel was then started at a velocity of 20m/s at -5 to 20 degrees where the angle was changed by 2/3 degree increments. The upstream wall pressure of the working section was recorded.
- The manometer reading for each pressure tapping was recorded.
- The experiment was repeated at velocity of 30 m/s at -5 to 20 degrees where the angle was changed by 2/3 degree increments. The upstream wall pressure of the working section was recorded. Also the manometer reading for each pressure tapping was recorded.
Results
After the experiment the results were represent both in tabular and graphical forms. However, the results were recorded for 20 m/s and 30 m/s respectively and on varying incidence angle degrees ranging from -5 to 20. Therefore, the results are represented in the tables and graphs below as follows:
Table1: 20 m/s L/D and AoA
| 20 m/s | |||
| Lift | Drag | AoA | |
| (N) | (N) | L/D | (Degrees) |
| -6.12493 | 0.522985 | -11.7115 | -4.7 |
| -1.00136 | 0.163879 | -6.11036 | -0.2 |
| 4.577368 | -0.05549 | -82.49 | 5 |
| 9.225824 | 0.169341 | 54.48075 | 9.8 |
| 11.19084 | 0.840279 | 13.31801 | 12.2 |
| 11.84669 | 1.975407 | 5.997088 | 14.2 |
| 12.03313 | 2.820606 | 4.266151 | 16.2 |
| 11.94211 | 3.493816 | 3.418071 | 18 |
| 14.38547 | 5.173793 | 2.780449 | 20 |
Graph 1: 20 m/s L/D against AoA
Table 2: 30 m/s L/D and AoA
| 30 m/s | |||
| Lift | Drag | ||
| (N) | (N) | l/d | (Degrees) |
| -15.1261 | 1.187724 | -12.73536613 | -4.1 |
| -1.9472 | 0.259677 | -7.498546271 | 0.391398 |
| 9.655957 | -0.09915 | -97.38736258 | 5.096809 |
| 21.58859 | 0.048747 | 442.8746679 | 10.18169 |
| 25.25439 | 1.475152 | 17.11985612 | 12.196 |
| 26.88254 | 4.271579 | 6.293349602 | 14.09912 |
| 28.00405 | 6.915119 | 4.049684467 | 16.29762 |
| 28.51154 | 8.931154 | 3.192369094 | 18.00128 |
| 26.39 | 9.515547 | 2.773356067 | 20 |
Graph 2: 30 m/s L/D against AoA
Table 3: 20 m/s Coefficient of Lift
| 20 m/s | ||
| Coefficient of Lift^2, CL^2 | Coefficient of Lift, CL | Coefficient of Drag, CD |
| 0.296785248 | -0.54478 | 0.046269 |
| 0.001641773 | 0.0405188 | 0.015794 |
| 0.164177315 | 0.405188 | -0.00383 |
| 0.662882559 | 0.814176 | 0.016154 |
| 0.991551918 | 0.995767 | 0.074977 |
| 1.083173052 | 1.040756 | 0.173198 |
| 1.206225548 | 1.098283 | 0.257778 |
| 1.220152206 | 1.104605 | 0.323026 |
| 1.214556081 | 1.102069 | 0.394089 |
Graph 3: 20 m/s Coefficient of Lift
Table 4: 30 m/s Coefficient of Lift
| 30 m/s | ||
| Coefficient of Lift^2, C L^2 | Coefficient of Lift, C L | Coefficient of Drag, C D |
| 0.369360063 | -0.60775 | 0.048428 |
| 0.006010901 | -0.07753 | 0.012151 |
| 0.145940044 | 0.382021 | -0.20394 |
| 0.7564094 | 0.869718 | 0.019155 |
| 1.021324487 | 1.010606 | 0.059091 |
| 1.158645571 | 1.076404 | 0.171404 |
| 1.270183097 | 1.127024 | 0.277976 |
| 1.259862574 | 1.122436 | 0.351795 |
| 1.534603142 | 1.238791 | 0.444297 |
Graph 4: 30 m/s Coefficient of Lift
Discussion
Several graphs were plotted both on coefficients and force and there comparison indicates that the coefficients progressively increased with increased force. The force was provided by the flowing wing and the coefficients according to the plotted graphs were directly proportional to the strength of the wind. This was mainly because the Lift, Drag and PM curves were plotted against AoA for both speeds such as 20 m/s and 30 m/s. This was in addition to plotting CL, CD, CM curves against AoA for 20 m/s and 30 m/s speeds.
The CL0, CLalpha, CD0, CDalpha values can be directly taken from the plotted graphs, where,
CL = CLalpha x AoA + CL0
CD = CDalpha x AoA^2 + CD0
Therefore,
CL = 0.405188 x 10.18169 + 1.102069
= 5.22756760772
In addition;
CD = 0.015794 x 10.18169 + 0.394089
= 0.55489861
According to the plots L/D against AoA, plot CL against CD and CL^2 against CD the angle that would be suggested as the cruise angle of attack for this aerofoil, that is, the most aerodynamically efficient condition would be 10.18169. This is due to the fact that this cruise angle presents the highest lift which makes sure that the airplane appropriately balances between the lift and the weight or gravity force. Therefore, according to this wind tunnel testing experiment of an aerofoil this angle would be the most effective for the airplane to cruise as indicated in the plotted graphs. Moreover, the proportionality relationship between CL and CD, indicate that the lift and drag are at appropriate balance. For instance, since for the airplane to cruise the lift in this case represented by CL require overcoming the drag in this case represented by CD, the proportionality relationship indicated in the plotted graphs seems appropriate to facilitate effective cruising of the airplane.
However, when the wind tunnel data obtained from the experiment is compared to information of this aerofoil from the literature, it is clearly seen that there is an obvious difference (Munson, Young & Okiishi, 1994). This is mainly because there are relatively significant differences between your experimental data obtained from the wind tunnel experiment compared to that taken from the literature comprising of secondary sources of information. These differences are mainly attributable to the fact that the conditions for this experiment might have been varied from those of the experiments that were used to collect data presented in the literature been checked. Moreover, the instruments used to collect data presented in the literature might have been different from those used in collecting data presented in this experiment. In addition, the other cause of variation in the data presented might be the difference in the research design adopted. Any slight difference might lead to significant difference in the obtained data.
Table 5: Literature data on wind tunnel parameters
Angle Lift (V) Drag (V) Lift (N) Drag (N) Lift Coeff. Drag Coeff.
-5 -0.000315 0.000001 -9.8005 0.0374 -0.21669 0.00083
-4 -0.000234 0.000002 -7.2804 0.0748 -0.16097 0.00165
-3 -0.000151 0.000005 -4.6980 0.1869 -0.10387 0.00413
-2 -0.000059 0.000008 -1.8357 0.2990 -0.04059 0.00661
-1 0.000028 0.000011 0.8712 0.4112 0.01926 0.00909
0 0.000102 0.000015 3.1735 0.5607 0.07017 0.01240
1 0.000197 0.000019 6.1292 0.7102 0.13552 0.01570
2 0.000272 0.000025 8.4627 0.9345 0.18711 0.02066
3 0.000365 0.000032 11.3562 1.1962 0.25109 0.02645
4 0.000448 0.000040 13.9385 1.4952 0.30818 0.03306
5 0.000532 0.000048 16.5520 1.7942 0.36597 0.03967
6 0.000595 0.000055 18.5121 2.0559 0.40930 0.04546
7 0.000680 0.000065 21.1567 2.4297 0.46778 0.05372
8 0.000734 0.000071 22.8368 2.6540 0.50492 0.05868
9 0.000802 0.000080 24.9525 2.9904 0.55170 0.06612
10 0.000826 0.000089 25.6992 3.3268 0.56821 0.07356
Furthermore, further research into the literature indicate that there are various plotted graphs about the data found in the literature concerning the wind tunnel data as shown below;
Graph 5: A plot of CL against alpha, deg
Graph 6: A plot of CD against CL
Trapezium Rule Integration has extensively been used to find an approximate value for a numerical integral, on the basis of finding the sum of the trapezia areas. Trapezium rule integration method can also be regarded as method of approximate integration. However, this method requires the use of narrower intervals to improve accuracy.
b∫aIn x dx = h/2 [(y0 + y5) + 2(y1 + y2 + y3 + y4)]
Where ‘x’ is an equation ‘a’ refers Upper Limit and ‘b’ refers Lower Limit.
= ∫???2??+ ∫???2???????= −∫?(??2−??2)??
The results are as below:
| Vtheoretical (m/s) | Vtrue (m/s) | Cv |
| 28.76 | 29.591 | 1.028894 |
| 27.246 | 27.774 | 1.019379 |
| 22.562 | 22.83 | 1.011878 |
| 13.662 | 13.075 | 0.957034 |
| 5.449 | 4.566 | 0.837952 |
Application question
Assuming that the body weight is 75kg and the weight of the wing is 25kg and unlimited thrust as well as using CL value obtained from experiment the wing parameters will be calculated as follows:
Conclusion
Reference
Barlow, J.B., Rae, W.H. & Pope, A. (2009). Low speed wind tunnels testing, 3rd ed. Hoboken, NJ: John Wiley & Sons Inc.
Munson, B. R., Young, D. F. & Okiishi, T. H. (1994). Fundamentals of Fluid Mechanics, 2nd ed. Hoboken, NJ: John Wiley and Sons, Inc.
