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Optimization of Landing Gear Fairings Final Report Group7 Daniel Ablog Michael Fuget Seunghyun Ko and Keisuke Tsujita AA 322 Aerospace Laboratory II WILLIAM E BOEING DEPARTMENT OF AERONAUTICS amp ASTRONAUTICS University of Washington Seattle Washington 98195 2400 2014 05 05 Wind tunnel and ow visualization experiments were conducted to deter mine the aerodynamic e ects of landing gear
Re Reynolds Number,S Surface area of test samples in2. t Thickness in,Angle of attack,Density of air slug ft2. absolute viscosity coefficient slug ft s,I Introduction. Small airplanes such as Cessna 150 cover their non retractable landing gear wheels with. the streamlined fairings which is also known as wheel pants while large airplanes have a. retractable landing gear to decrease the drag associated with the landing gear during their. Fig 1 Landing gear fairing of Cessna 150, American Institute of Aeronautics and Astronautics. Fig 2 Fairing design that does not cover the wheel. Currently the fairings of non retractable landing gear cover the top half of the wheel as. shown in Fig 1 1 Previously the drag of several types of existing fairings were measured by. Herrnstein and Biermann 2,Fig 3 The Smallest Drag Fairing Configuration. Should I have this part here or in the theory From the experiment the fairing shown in. Fig 3 with modification 2 had the smallest drag among the six types of fairings tested The. other configurations of fairing tested are shown in Appendix A Herrnstein and Biermann2. concludes that the fairing which covered both sides of the wheel and had minimal frontal. area was most effective in reducing drag Although the fairings reduce the drag they are. the extra weight added to the airplane which should be minimized One way to reduce the. weight of a landing gear fairing is to utilize the new design of the fairing which does not. American Institute of Aeronautics and Astronautics. cover the wheel as shown in Fig 2 The new design of the fairing does not increase the. frontal area and may reduce the weight while keeping the drag performance similar to that. of the existing fairing The purpose of this project is to measure the aerodynamic forces. and visualize the flows around the fairings with five different lengths using the configuration. shown in Fig 2 a fairing with the existing configuration of fairings as shown in Fig 1 and a. wheel without a fairing The results are compared to determine the differences in drag and. A Assumptions and Simplifications, Several assumptions and simplifications were made for the test samples First the test. model can consist of only a wheel and fairing behind the wheel Also although a real wheel. is made of several parts and different materials a circular cylinder whose ratio of thickness. and diameter is the same as that of a real wheel can model the wheel Also the test. models can be modeled by any materials as long as they are consistent for all models These. simplifications are valid because the purpose of the project is only to investigate the effect of. the variation of length of the fairings In addition the assumption that the wheel modeled. by a cylinder does not rotate was made because wheels on a Cessna 150 do not rotate during. cruising conditions, Title of the subsection must be changed Aerodynamic forces exerted by the flow of air are. due to the pressure distribution of the surface and shear stress on the surface 3. Fig 4 Aerodynamic forces on the test sample, The sting mount in the 3x3 wind tunnel measures the normal and axial force shown in. Fig 4 Drag at a certain angle of attack can be calculated from normal and axial forces as. D N sin A cos 1, American Institute of Aeronautics and Astronautics. The Reynolds number which is strongly associated with drag can be found by 3. Since drag depends on the dynamic pressure and projected area drag coefficient coeffi. cient are defined as, As explained in Section II A the test sample models the wheel by a circular cylinder. whose pressure drag is higher than the friction drag 4 The pressure drag is produced by the. pressure difference between the front and back side of the cylinder This difference is caused. by the wake formed behind the cylinder 5,Fig 5 Flow over a blunt body and streamlined body. The drag characteristic of the blunt body such as a circular cylinder and the streamlined. body are shown in Fig 5 The blunt body has large pressure drag caused by the wake formed. behind the bod due to the flow separation The new model shown in Fig 2 uses the wheel. as the front portion and itself as the back portion of the streamlined body This may lead. to less pressure drag compared to the case in which only the wheel is employed. American Institute of Aeronautics and Astronautics. C Previous work related to the design and dimensions of the fairings. Fig 6 Pressure distribution over the cylinder with splitter plate behind. Roshko6 states that when a cylinder is placed in front of a flat splitter plate which has a. length of 5 times the diameter of the cylinder as shown in Fig 6 the drag coefficient of the. cylinder is reduced from 1 1 to 0 7 Fig 6 shows the pressure distribution behind the circular. cylinder with splitter plate and that without splitter plate The splitter plate is effective in. reducing the pressure drop behind the circular cylinder Also Roshko states that by using a. splitter plate with a length of 1 14 times the diameter the vortex shedding frequency varies. depending on the location of the plate whereas the drag did not decrease 6 These previous. work indicate that the length of the splitter plate is what affects the reduction in drag The. length of the splitter plate was used as the reference length in the project as explained in. the Section III A,III Designing and Fabrication of the Models. A Designing of the test models and Determining the test conditions. The dimensions of the wheel were determined based on the actual wheel used The. Cessna 150 employs a wheel whose diameter is 15 in and thickness is 6 in 7 Due to the size. of the wind tunnel the wheel s diameter was scaled down to be 5 in and the thickness was. decided to be 2 in in order to match the ratio of diameter and thickness of the actual wheel. The lengths of the fairings to be tested were chosen as 0 7d 1 0d 1 14d 1 5d and 2 0d. Even though the previous work described in Section III shows that the splitter plate whose. length is 5d placed behind the circular cylinder is effective in reducing the drag 6 the largest. length was decided to be 2d This decision was made because the fairing may touch the. ground during the takeoff or landing if the length is too long Other dimensions which are. kept constant among the model fairings were also decided The width of the fairing was. decided to be the same as the thickness of the wheel because the previous experiment by. Herrnstein and Biermann2 described in Section II shows that fairing with the smallest frontal. area had the lowest drag The spacing between the fairing and the wheel was chosen to be. 0 5 in which corresponds to 1 5 in in the real scale This spacing was chosen because certain. clearance between the fairing and the wheel is required in case of landing on muddy ground. The height of the fairing was decided to be 80 of the diameter of the wheel to allow the. smooth transition of the flow from the wheel to the fairing. American Institute of Aeronautics and Astronautics. It is ideal to execute the experiments under the dynamically similar condition represented. by the Reynolds number First the Reynolds number of the flow which the actual landing. gear experiences was calculated The cruising speed of the Cessna 150 at altitude of 7000 ft. is 117 mph 8 The Reynolds number for its wheel at that speed and altitude was calculated. with Eq 2 with the cruising speed and properties of air at 7000 ft listed in Table 1 The. Table 1 Properties of Air at 7000 ft,Properties Values. 9 Temperature R 493 73,Pressure lb f t2 1 6331 103. Density slug f t3 1 9270 103, viscosity of air was found by interpolating the data in Table 6 in Appendix B The Reynolds. number was calculated to be 1 13 106 with Eq 2, However due to the limitation of the testing facilities it was not possible to conduct the. experiment with the same Reynolds number The air speed and water speed of the testing. facilities were limited as well as the maximum size of the models which can be used The. 3x3 wind tunnel can not be operated above 45 psf of dynamic pressure Under the standard. atmosphere at sea level the Reynolds number was calculated to be only 5 0 105 with. Eq 2 The Reynolds number in water tunnel was also calculated to be 6 5 104. The dimensions of the test samples are shown in Fig 7. American Institute of Aeronautics and Astronautics. a Wheel with new model of fairing l 0 7d 1 0d 1 14d. 1 5d or 2 0d Unit is inch, b Current design of the fairing For simplicity the fairing. and the wheel was designed as one body Unit is inch. Fig 7 The design of the new and current fairing test samples. Because the experiments are planned to be executed in the 3x3 wind tunnel and in the. water tunnel a mount must be created for each tunnel The CAD drawing for the sting. mount which measures the aerodynamic forces and moments was obtained to design the. extension mount for the wind tunnel A CAD model of the extension mount for the sting. mount in the 3 x 3 wind tunnel was created as shown in Fig 8. American Institute of Aeronautics and Astronautics. Fig 8 Extension mount design for wind tunnel testing. In addition the mount for the experiments in the water tunnel was designed as shown. in Fig 9 These configurations were chosen to minimize the effect of the mount on the. experiments,Fig 9 Mount design for water tunnel testing. B Fabrication of test samples, The materials purchased for fabrication are listed in Appendix B A model wheel and five. fairings with different lengths were fabricated from wood The test samples were cut with. the band saw in the machine shop in the Mechanical Engineering Department The fabri. cated test samples are shown in Fig Making a perfectly symmetrical test samples was. extremely difficult and the thickness of the fairings became slightly smaller than what was. originally designed as they were sanded to be tapered After sanding sealer was applied the. models were waterproofed for the experiment in the water tunnel tyotto The machine shop. in the Mechanical Engineering Department and Aeronautics and Astronautics Department. were used to fabricate the test samples The 3 3 Wind Tunnel and Water Tunnel were used. to conduct the experiments, The test samples used in the water tunnel and the wind tunnel are shown in Fig. IV Experimental Apparatus, The setup of the experiment in the water tunnel and in the wind tunnel are included in. Fig 11a and Fig 11c respectively The mount shown in Fig 9 was used to hold the test. samples and change angles of attack in the water tunnel. American Institute of Aeronautics and Astronautics. a Wheel b Fairing with l 0 7d,c Fairing with l 1 14d d Fairing with l 1 14d. e Fairing with l 1 5d f Fairing with l 2d,Fig 10 Fabricated Test Samples. Table 2 Apparatus used in the water tunnel,Water Tunnel. Dye Injection Tube,Lucks Color Blue,Lucks Color Green. Gear Pump Controller,Water Tunnel Mount, Table 2 includes the apparatus used for the flow visualization in the water tunnel exper. American Institute of Aeronautics and Astronautics. a Experimental Setup for the Water b Dye Injection System for the Water. Tunnel Tunnel Experiment,c Experimental Setup for the Wind. Fig 11 Experiment Setups,Table 3 Apparatus used in the 3 3 wind tunnel. Equipment Model,3x3 Wind Tunnel,Attachment Mount, Table 3 includes the apparatus used for the drag measurements in the 3 3 wind tunnel. V Procedure,A Water Tunnel Test, The bare wheel and the fairings attached to the wheel were tested in the water tunnel. to visualize the flows around the samples One of the injection port of the Micropump was. connected to the dye container with a dye injection tube The other injection port was. connected with the dye injection tube which was attached to the metal rod for the dye. injection in the water tunnel The gear pump controller was connected to the Micropump to. American Institute of Aeronautics and Astronautics. control the amount of dye to be injected Schematic of the dye injection system is shown in. Fig 11b The test samples were attached to the water tunnel mount The following angles. of attack were tested 10 7 2 0 10 15 20 and 20 While the experiment the. dye injecting tube attached to the metal rod was moved by hand to see the flow over and. under the test samples In order to analyze the flows over and under the test samples green. dye and blue dye were injected from the top surface and bottom surface of the test samples. respectively However there was not enough green dye for all runs thus blue dye was injected. from for both top and bottom of the flows The frequency of the water tunnel was set to 50. Hz which corresponds to the 50cm sec The conversion chart between the frequency and. the water speed is in Appendix Movies were taken to analyze the flows. B Wind Tunnel Test, The atmospheric pressure was recorded The test samples were attached to sting mount. by the attachment mount as shown in Fig 11c with aluminum tape to reinforce the attach. ment Angles of attack was changed from 0 to 20 with increment of 2 For each run the. dynamic pressure was gradually increased from zero to near 45 psf LabView was used to. record dynamic pressure temperature axial forces and normal forces. VI Discussion of Results,A Measurement of drag, Fig 12 Drag Coefficient measured at 0 for each test sample with various Reynolds. Fig 12 is the plot of the drag coefficient measured at 0 with Reynolds number from. Re 3 5 105 to 7 0 105 Except in the lower Reynolds number range all of the samples. with fairing had the lower coefficient of drag than that with the bare wheel The current. model of the fairing and the test sample with l 0 7 d l 1 0 d and l 1 14 d of fairing. American Institute of Aeronautics and Astronautics. showed slight decrease as the Reynolds number increased On the other hand the sample. with l 1 5 d l 2 0 d of fairing showed slight increase in drag coefficient At zero angle of. attack which is typical in cruise the current model of fairing showed the best aerodynamic. performance over the range of Reynolds number tested. Table 4 The Coefficient of drag and percent decrease compared to the bare wheel model. for low and high Re with weight ratio of the models at zero angle of attack. Low Re High Re,Weight Ratio,Cd Decrease Cd Decrease. Bare Wheel 0 245 0 256 0,0 7 d 0 186 24 2 0 177 30 9 23 1. 1 0 d 0 255 3 83 0 226 12 0 35 8,1 14 d 0 308 25 6 0 269 4 76 42 4. 1 5 d 0 165 32 9 0 193 24 9 56 7,2 0 d 0 194 21 0 0 196 23 6 73 4. Current Model 0 157 36 0 0 135 47 4 1, Table 4 includes the Cd and the percent decrease compared to the bare wheel for each. low and high Reynolds number with the weight ratio of the fairing compared to the current. model at zero angle of attack, Among the new design of the fairings at lower Reynolds number the fairing with the. length of 1 50 d had the best aerodynamic performance as shown in Fig On the other. hand the at higher Reynolds number the fairing with the length of 0 7 d was the lowest. coefficient of drag Overall length The all the models with new fairing design had better. performance except the sample with l 1 14 d of fairing in the high Reynolds number range. For most of the, a Low Reynodls Number Re 4 0 b High Reynodls Number Re 7 0. 105 to 4 5 105 105 to 7 5 105, Fig 13 Avarage drag Coefficient measured at lower and higher Reynolds number for each. test sample with various angles of attack, The drag coefficients for each fairing were plotted against varying angle of attack Low. American Institute of Aeronautics and Astronautics. Reynold s number testing data is shown in Fig 13a and High Reynold s number data in. Fig 13b The general trend shown in the plot is the increase in the coefficient of drag of all. models except the bare wheel with increasing angle of attack In addition from these plots. it can be seen that among the new design of the fairings the 2d length fairing generally had. the lowest drag coefficient compared to those of the individual wheel alone over the range of. the angle of attack However the covered fairing still had the best aerodynamic performance. with the lowest drag coefficients This was expected as the covered fairing has no separation. between the wheel and after body allowing for less chance of flow separation Regardless. of this fact it is still good to note that the weight difference between these fairings and the. covered fairing is substantial despite having a slight difference in drag performance. I am not sure about this Also evident in Fig 13 is the decrease in drag coefficient going. from the low Reynold s number data to the high Reynold s number data A comparison. between these two plots shows that in general the drag coefficient reduces as the Reynold s. number is increased This is due to turbulent flow which occurs in higher Reynold s number. flight Turbulent flows remain attached to the fairing surface longer than flow during low. Reynold s number This increases pressure acting on the rear of the model and therefore. reduces the drag component of the landing gear, This may not be correct The outlier in the wind tunnel experiment was the 1 14d fairing. Drag coefficient for this fairing came out to be much higher than the 1d fairing Some data. points even exceeded those of the individual wheel This may be attributed to manufacturing. error such as having substantial roughness or improper shaping of the model. B Flow Visualization, Flow over the test samples for each run is shown in Fig XX The flow over the bare wheel. is shown in Fig XX, As indicated by Fig XX when the test sample has a positive angle of attack the flow. bottom of the wheel goes in the gap between the wheel and fairing moving up to the top. side of the test model When the flow from the bottom side mixes with the flow on top of. the sample the separation occurs When the test sample has a negative angle of attack. the flow on top side of the sample moves in the gap and causes the flow to separate on the. bottom side of the model, In the case of model having zero and small angles of attack the flow stays attached on. the wheel and fairing as shown in Fig XX, When the wheel has large angles of attack the pressure difference between the top and. bottom sides of the models is large causing the flow suction into the gap. VII Conclusion, As of May 5th the design of test samples and mounts is completed Also fabrication of. the test models has begun However experiments have not been executed and therefore. data cannot be presented, American Institute of Aeronautics and Astronautics. References, Federal Aviation Administration Aircraft Landing Gear Systems Aviation Maintenance Technician. Handbook Airframe Vol 2 2012 Chapter 13, Herrnstein H and Biermann D The drag of airplane wheels wheel fairings and landing gears I. NACA Report 485 1935, Anderson J D Basic Aerodynamics Introduction to Flight 7th ed McGraw Hill New York NW. 2012 pp 134 288, Anderson J D Introduction to the Fundamental Principles and Equations of Viscous Flow Fun. damentals of Aerodynamic 3rd ed McGraw Hill New York NW 2001 pp 713 745. Anderson J D Fundamentals of Invisid Incompressible Flow Fundamentals of Aerodynamic 3rd. ed McGraw Hill New York NW 2001 pp 177 275, Roshko A On the Wake and Drag of Bluff Bodies AIAA Journal Vol 22 No 2 1955 pp 124 132. The Goodyear Tire Rubber Company Flight Special II GOODYEAR Aviation Tires Ohio. 2013 http www goodyearaviation com cfmx web aviattiresel details cfm sortorder 35 FB ACCESSED. Cessna Sales and Service Description and Operating Details Model 150 Owner s Manual Wichita. Kansas 1968, Anderson J D Standard Atmosphere English Engineering Units Introduction to Flight 7th ed. McGraw Hill New York NW 2010 pp 770 773, Cengel Y A Conversion Factors Thermodynamics Engineering Approach 7th ed McGraw Hill. New York NW 2012 pp 655 687, The Engineering ToolBox Absolute and kinematic viscosity of air at temperatures rang. ing 40 1000 C 40 1500 oF at standard atmospheric pressure Imperial and SI Units. http www engineeringtoolbox com air absolute kinematic viscosity d 601 html ACCESSED 5 2 2014. American Institute of Aeronautics and Astronautics.