JAVAFOIL User's Guide Contents
Contents
1. 2ss20 0 een ann anna nungen ann ces cetdstausieesunecectede tuned cteeatececunuereturestereduetesteedee 30 Multi Element Airfoils unsnreenneennnenneennnnnnnnnnnannennnnnnnennnnennennnnnnnennnnenneennnennernnnen nennen nnmnnn naa 31 Automating JavaFoil with a Script uuuuuesusnneansnnnnnnnnnnnnnnnnnnnnnnnnannnnnnnnnnnnnnnnannnnnnnnnnnnnnnnnnnnnannnnnannnn nn 33 FETE 61 ale Saan neien e a ARE a AEN E EER Ra a ENE Eaa a E RA 33 JAVAFOIL JAVAFOIL is a relatively simple program which uses several traditional methods for the analysis of airfoils in subsonic flow The main purpose of JAVAFOIL is to determine the lift drag and moment characteristics of airfoils The program will first calculate the distribution of the velocity on the surface of the airfoil For this purpose it uses a potential flow analysis module which is based on a higher order panel method linear varying vorticity distribution This local velocity and the local pressure are related by the Bernoulli equation In order to find the lift and the pitching moment coefficient the distribution of the pressure can be integrated along the surface 1 Next JAVAFOIL will calculate the behavior of the flow layer close to the airfoil surface the boundary layer The boundary layer analysis module a so called integral method steps along the upper and the lower surfaces of the airfoil starting at the stagnation point It solves a set of differential equations to find the va2. txt multi element airfoil geometry in form of simple two columnar x y coordinate sets arranged in two columns Multi element airfoils must be separated by a pair of x and y values larger than 999 xml multi element airfoil geometry in JAVAFOIL s hierarchically structured xm1 format A OSE multi element airfoil geometry in AutoCad drawing exchange format Many CAD programs can read this file format but the interpretation is not always perfect igs or iges multi element airfoil geometry in Initial Graphics Exchange Standard format Many CAD programs can read this file format Note that JAVAFOIL selects the output file format according to the file name extension Importing airfoil geometry JAVAFOIL can read airfoil geometry from the following file types 4 Oe EXT multi element airfoil geometry in form of simple two columnar x y coordinate sets arranged in two columns Multi element airfoils must be separated by a pair ofx and y values larger than 999 xml multi element airfoil geometry in hierarchically structured xml format me ong g1f bmp Jpg single element airfoil geometry from an image file Note that JAVAFOIL selects the input file format according to the file name extension Importing scanned images You can also load an airfoil from a bitmap image in GIF PNG BMP or JPG format JAVAFOIL then tries to find an airfoil shape in this image by comparing the image points with the color found in the up
3. JAVAFOIL User s Guide Martin Hepperle 5 December 2011 Contents Contents casa ana aaaea arae aaae aaao ae a a a aaa Taa Eaua aaae SaaS Esra apada A Saidia Sot 1 AV A o E EE EE E E E E E A E E A NPT 1 imitations Stea Re ee a a a ER 2 JAVAFOIL S Card S Eee else 2 Geometry Card na ee pa tee ee hes eee 3 JAVAFOIL S Geometry Generators ursnseensnnnnsnennnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnnsnnnnnnnnnnenn nn 5 Modify Gardinen leerer einer lend T A A rennen 14 The Panel Method uusnnssrsnnnnesnnnnnnnnennnnnnennnnnnnennnnnnnnnnnnnnnnennnnnnnnnnnnnnnnnernnnnnernnnnenernnnnnnernnnneernnnnnnenn nn 17 Boundary Layer Analysis nunsnnssnnnnannnennnnnnennnnnnnennnnnnnnnnnnnnnennnnannenennnennennnnennernnnennenn nennen nn 18 Transition Eritena rn aan a ER lana als BAN ea 18 Effect of RROUQMNeSS isoin iiaii niii i e A evi de ee a aiis 20 Sstall G rteeti ns au RE aie eal ee a Gaeta A eas 21 Compressible FlOW 2 2 2 araa are ara aaa a raa se ccetaceweetectenccshscaventuceseucertaeerselssenstestadaes 23 Eritical Pressure Goefficienti uanun nennen ARNEE ESAR AEREA ih 24 Gompressibility Gorrections 2 2 2er HER BEER ih 24 Finite Wings in JAVAFOIL ccseccececeeeeeeeeneeeeeeeneeeeeeeneeseeeeneee seen eneeseeennee see nneeseeesneeseeesneesesessneesesesnenseeeenaes 25 The Aerodynamic Genter 2 cnn essen ann cecsasesseunedcadees seisseupecenenssoces 29 Gro nd Effett
4. and Aircraft cards contains a column with the position of the aerodynamic center A C The aerodynamic center is a point on the airfoil at which the pitching moment is constant not necessarily zero for all angles of attack It can be calculated from the gradient of the pitching moment over lift coefficient curve OC 025 X c 0 25 According to thin airfoil theory the aerodynamic center is located at 25 of the chord length and does not move when the angle of attack is changed In real life airfoils are thick and the location typically can vary about 2 around this location 29 The aerodynamic center is not to be confused with the center of pressure C P which is the point at which the total aerodynamic force acts This total force produces the same effect as the lift and pitching moment The location of the center of pressure changes with angle of attack and can even move in front or behind the airfoil shape The center of pressure can be calculated from lift and pitching moment coefficients C025 C Xop 0 25 Note that both center of pressure as well as the aerodynamic center are for the airfoil only not for the complete aircraft with tailplanes Ground Effect When a wing is brought close to the ground its characteristics are changed considerably First the pressure distribution around the two dimensional airfoil shape a wing of infinite span is affected by the presence of the ground
5. t c Example NACA 2412 2 camber at 40 chord 12 thickness The thickness distribution for the 10 thick section is given by the polynomial y 0 29690 Vx 0 12600 x 0 35160 x 0 28530 x 0 10150 x The coefficients of this thickness distribution had been chosen according to the following constraints 4 for a 10 thick section e maximun thickness at x c 0 3 amp 0 3 0 e finite thickness at trailing edge 0 004 e finite trailing edge angle at x c 1 0 gt amp 1 0 0 234 e nose shape defined by y c 0 078 at x c 0 1 Modified NACA 4 digit airfoils The modification adds the position of the maximum thickness as well as the nose radius to the parameter set of the 4 digit series see 3 Parameters e Free t c f c 7G 8 Naming Scheme The name consists of a 4 digit prefix which is identical to the 4 digit series designation followed by a dash and two additional digits 1 digit maximum ordinate of camber 100 f c 2 digit position of the maximum camber 10 x c 3 and 4 digit maximum thickness 100 t c dash 5 digit indicates the leading edge radius and is usually one of 0 3 6 or 9 o 0 sharp leading edge o 3 the normal radius of the 4 digit series o 6 normal radius o 9 3 times the normal radius e 6 digit position of the maximum thickness 10 x c Example NACA 1410 35 1 camber at 40 chord 10 thickness reduced leading ed
6. Local Criteria Many methods predict transition by applying a criterion based on local boundary layer parameters These criteria are based on relations which can be evaluated at any station along the surface They do not need an extra integration of some instability parameter but of course are affected by the history of the flow Most of these criteria are relating Re to the shape of the boundary layer profile Eppler sa 18 4 H 21 74 0 36 r Transition is assumed to occur when Re gt e TR Eppler enhanced Rid oe 18 4 Hyy 21 74 125 Hyy 1 573 0 36 Transition is assumed to occur when Re gt e Hin 2 Michel 1 This simple criterion assumes transition to occur when Re gt 1 535 Re 18 Michel 2 Transition is assumed to occur when Re gt 1 174 1 22400 Re Re See 22 H12 Res Transition is assumed to occur when 2 1 lt H lt 2 8 and log Re gt 40 4557 64 8066 H 26 7538 H 3 3819 H See 23 Criteria based on a region of instability n Re envelopes These methods first determine a local point of instability and then begin at this point to integrate a measure for the amplification of instability Drela approximates the envelopes of the amplification rate n versus Re by straight lines of the form n f Re H Two versions of this approximation were used in his codes of the XFOIL and MSES ISES family The approximation is
7. Second the lift and the induced drag of a wing of finite span are affected also JAVAFOIL simulates the ground effect on the flow around the two dimensional airfoil by using a mirror image of the airfoil section The mirror plan is always located at y 0 Note that for a proper simulation the baseline airfoil must be translated into the positive y direction so that it does not intersect the horizontal line y 0 This translation can be performed using the Modify card In contrast to the flow around a free airfoil where the flow field can be constructed from a superposition of the solutions for zero and 90 degrees angle of attack the ground effect case cannot be created by superposition Any change of angle of attack also changes the geometry of the airfoil and mirror airfoil pair Therefore a new panel solution is required for each angle of attack which slows down the calculation of a polar somewhat y pivot y point 25 25 a 10 100 X 100 X mirror image mirror image Changing the angle of attack in ground effect rotates around the pivot point The angle of attack of the airfoil is always changed by rotating the section around the pivot point specified on the Modify card If you want to analyze an airfoil at a height of 25 of the chord length and want to maintain the trailing edge point you would first translate the airfoil in y direction by 25 and then set the pivot point to x 100 y 25 Then any subsequent change of angle o
8. Sth digit maximum thickness 100 t c Example NACA 23112 like NACA 23012 design lift coefficient 0 3 maximum camber at 15 chord 12 thickness but with a reflexed aft camber line NACA 1 series airfoils The development of these airfoils was aiming at high subsonic speed applications like propellers see 5 Their shape was designed with the help of the new that is in the 1930s numerical design methods JAVAFOIL can create airfoils of the NACA 16 type which are the only members of the 1 series published by NACA The maximum thickness and the maximum camber are located at 50 chord whereas the minimum pressure is reached at 60 of the chord length Parameters e Free t c C design e Fixed x c 0 5 x c 0 5 Naming Scheme Ist digit 1 series designation 2nd digit position of minimum pressure of the thickness distribution 10 x c a dash 3rd digit 10 C design e 4rd and 5th digit maximum thickness 100 t c Example 16 212 1 series minimum pressure at 60 chord design lift coefficient 0 2 12 thickness While these airfoil shapes are not based on analytical expressions the published coordinates have been approximated to produce an accurate representation of these airfoils The camber lines used are of the uniform load type a 1 0 NACA 6 and 6A series airfoils These airfoils were the first NACA airfoils which had been systematically developed with the inverse design method by Theodorsen The c
9. expressed by on Re Re Res ai rm Transition can occur when Re gt Re et and gt Nar In JAVAFOIL transition is assumed crit to occur when the value n 9 r is exceeded crit Drela XFOIL 1 1 and 5 4 si 2 4 H 2 5 tanh 1 5 H 4 65 3 7 0 25 Re logy Reg u 1215 gol ianh 12 9 225 40 44 i H u 18 7 127 These approximations can be found in 1 and 2 Drela XFOIL 5 7 Modification in 1991 ep Re a 252 2 Hy l e 14 1 0 43 log Rey air 0 7 tanh 9 24 2 492 0 66 B 19 Drela XFOIL 6 8 only a tiny modification term 0 66 0 62 19 034 I Era rte eee ORe Ser 25 2 Hy l 0 43 14 1 log Re a 0 7 tanh 9 24 2 492 0 62 12 7 12 Method of Arnal A set of tables produced by D Arnal has been approximated by W W rz with polynomials On Re log Rey en b b H b Hi Here the envelope is not a straight line as in Drela s method For details see 19 In JAVAFOIL transition is assumed to occur when the value n 9 r is exceeded a a H a Hi crit Method of Granville This method is not described here It also works by integrating a stability parameter starting from a point of instability Abbreviations approximation of n roughness factor 0 smooth r displace
10. is generated even for single element airfoils therefore a Cl min in the order of 5 to 6 can be seen Polar Analyze 500000 500000 500000 15 10 1 100 100 0 0 IH finally export coordinates in XML format Geometry Save Z groundforce example xml 32 Automating JAVAFOIL with a Script Sometimes it is useful to have JAVAFOIL execute a command sequence in a script automatically and then terminate This allows running JAVAFOIL inside a parameter sweep or as part of an optimization loop For this purpose you prepare the script file and start JavaFoil with the name of the script file Then JAVAFOIL runs invisible without opening a window and executes the commands in the script file The name of the script file can be transferred to JAVAFOIL in two ways First you can define a system property using the D command line option of the java command like so java exe DScript Path Script cp Path mhclasses jar jar Path javafoil jar secondly you can specify the script file using as a command line argument to JAVAFOIL like this java exe cp Path mhclasses jar jar Path javafoil jar Script Path Script J Both ways methods are equivalent and produce the same result Note that this example in Windows style uses the backslash as file separator for other Unix like systems you have to use the appropriate separator usually a forward slash Note As JAVAFOIL is running withou
11. perform the following calculation sequence inital value Re 10 for a a to a step Aa iterate Re Rel Re while gt Note that the result still is an airfoil polar even if wing loading and chord length are involved Only when you additionally specify an aspect ratio on the options card the polars include the induced drag and approximate a finite wing A precaution must be undertaken to handle cases where C 0 Here JAVAFOIL limits the Reynolds number to a value corresponding to a small lift coefficient e g C 0 02 Note One can also derive the Reynolds number for a constant ratio 4 eliminating the chord length L This has not been implemented in JAVAFOIL as it was considered more abstract to think in terms of instead of the aircraft design parameters and But as the relation is 4 l it would be sufficient to use 4 instead of in JAVAFOIL while setting l 1 27 Correction of Lift for given Aspect Ratio and Mach number For a given angle of attack a 3D wing of finite aspect ratio produces less lift than the 2D airfoil section which corresponds to an infinite aspect ratio Another correction has to be applied when the Mach number is larger than zero In subsonic flight more lift is produced when the Mach numbers is increased The 3D wing correction is applied only if you specify a value for the aspect ratio of the wing A b S span bb and wi
12. system consists therefore of a of panels 1 sized matrix and two right hand sides These are for 0 and 90 angle of attack and can be solved efficiently at the same time for the two corresponding vorticity distributions The vorticity distribution for any arbitrary angle of attack is then derived from these two solutions remember that potential theory is linear and allows for superposition There is no interaction with the boundary layer as in XFOIL though For a shape discretization by N panels the equation system of this classical panel method consists of the matrix of influence coefficients the unknown circulation strength at each panel corner point and the two right hand side vectors These represent the no flow through the surface conditions for 0 and 90 angle of attack Each coefficient C reflects the effect the influence of the triangular vorticity distribution due to the vortex strength at each corner point on the center point of each panel j The last row contains the tangential flow condition at the trailing edge AKA Kutta condition 17 C ae On Yo Y 90 RAS o RAS o f Y2 0 Y2 90 RHS o RHS go Ciy Cyan i 1 82605 1 Ynat0 N 1 90 RES y 41 0 RAS x 41 90 Boundary Layer Analysis The boundary layer analysis module implements an integral boundary layer integration scheme following publications by Prof Richard Eppler Such integral methods are based on differe
13. this first airfoil element to be used later Modify Select 1 Modify Duplicate now select the first element again Modify Select 1 ee pete Up sadesdown Modify Flip 25 0 Vie ss seele ale vo WSR Modify Scale 75 rotate it 5 degrees trailing edge up around its nose 0 0 Modify Rotate 0 0 5 bd now select the second element Modify Select 2 eee lOmin OS Clem down Modify Flip 25 0 Il soo Beale this GOD co 30 Modify Scale 30 move it back so that there is 5 overlap Modify Move 70 0 12 rotate it by 30 degrees around a point at 70 12 Modify Rotate 70 12 30 finally THIS IS OMPORTANT select both elements again for all further analyses if only one element is selected on the modify card only this element will be considered during the calculation of polars etc Modify Select 1 2 and move both elements up by 25 for ground clearance note that the airfoil may not cross the y 0 line which is the ground plane Modify Move 0 0 25 2 switch ground effect simulation ON Options GroundEffect 1 prepare for analysis Options MachNumber 0 Options StallModel 0 Options TransitionModel 1 Options AspectRatio 0 velocity versus x c should show no strong suction peaks in the nose region of 2nd element Velocity Analyze 0 0 1 0 A polar for Re 500 000 alfa 15 to 10 degrees with ground present a string suction force
14. 65 9 A C Kravets Characteristics of Aircraft Profiles Moscow 1939 33 10 11 12 13 14 15 16 17 18 19 20 21 22 23 Anonymous Measurement of Force Coefficients on the Aerofoils EC 1240 and EC 1240 0640 in the High Speed Tunnel at the National Physical Laboratory British ARC R amp M 2246 S Goldstein Approximate Two Dimensional Aerofoil Theory Part I Velocity Distributions for Symmetrical Aerofoils British A R C Technical Report C P No 68 1952 British A R C R amp M 4726 British A R C R amp M 4978 Katz Plotkin Low Speed Aerodynamics McGraw Hill 1991 H B Helmbold F Keune Beitr ge zur Profilforschung II Geometrie der Profilsystematik Luftfahrt Forschung Band XX 1943 pp 81 ff G Ro ner ber eine Klasse von theoretischen Profilen mit vier frei w hlbaren geometrischen Parametern Jahrbuch 1942 der Deutschen Luftfahrtforschung p I 141 1 159 1942 H H Chun R H Chang Turbulence Flow Simulation for Wings in Ground Effect with two Ground Conditions fixed and moving Ground International Journal of Maritime Engineering Royal Institution of Naval Architects 2003 D Steinbach Comment on Aerodynamic Characteristics of a Two Dimensional Airfoil with Ground Effect Journal of Aircraft Col 34 No 3 pp 455 456 1996 W W rz Hitzdrahtmessungen zum laminar turbulenten Str mungsumschl
15. ag in anliegenden Grenzschichten und Abl seblasen sowie Vergleich mit der linearen Stabilit tstheorie und empirischen Umschlagskriterien Dissertation Institut f r Aero und Gasdynamik Universit t Stuttgart 1995 R Eppler D Somers A Computer Program fort he Design and Analysis of Low Speed Airfoils NASA TM 80210 1980 R Eppler Airfoil Design and Data Springer Verlag 1990 T Cebeci P Bradshaw Momentum Transfer in Boundary Layers Hemisphere Publishing Corporation Washington London 1977 A R Wazzan C Gazley A M O Smith H Rx method for Predicting Transition AIAA Journal Vol 19 No 6 June 1981 pp 810 811 34
16. ation of the induced drag is added to the airfoil drag for the same angle of attack a This correction is also only applied if you specify a value for the aspect ratio of the wing A b S span b and wing area S on the Options card As no information about the real wing shape is available the assumption of having a good wing planform is assumed Therefore the induced drag component is calculated by using the classical formula derived by lifting line theory Prandtl ms ol T A In JAVAFOIL the k Factor is assumed to be 1 0 planar wing with elliptical lift distribution Note that the idea of these simple corrections is to give you a feeling for the relative importance of the induced drag in relation to the airfoil drag only For real wing design you should use a more appropriate 3D aerodynamic analysis tool e g a vortex lattice or panel method Implementation in JAVAFOIL public final static double DragForAspectRatio double dCd double dcl double dAspectRatio double dMachNumber double dReturn dCd if dAspectRatio gt 0 1 add the induced drag of finite wing according to Prandtl dReturn dCl dCl Math PI dAspectRatio return dReturn Note that all finite wing results are only approximations If you need more accurate results you must use a 3D wing analysis code which ideally can also handle friction effects The Aerodynamic Center The output of the Polars
17. ble Flow JAVAFOIL analyzes airfoils in incompressible flow which means low Mach numbers as they are common in model aircraft of general aviation airplanes In practical application this means Mach numbers below M 0 25 It is possible however to extend the Mach number range somewhat by applying compressibility corrections to the incompressible results This is only possible as long as the flow speed is subsonic all over the surface of the airfoil and compressibility effects are small 23 Critical Pressure Coefficient The character of the flow changes dramatically when sonic speed is exceeded anywhere on the surface The pressure coefficient associated with sonic speed is called critical pressure coefficient C rir In most cases pressure recovery from supersonic to subsonic speeds from Cp lt Cycrit to Cp gt Cp crit is leading to an abrupt recompression with a shock The analysis of such flows requires more complex methods than implemented in JAVAFOIL Such methods must be capable of handling compressible flows for example by solving the full compressible potential equations or by solving the Euler equations In order to indicate how close the local flow is to supersonic speeds JAVAFOIL calculates the critical pressure coefficient if a Mach number is specified on the Options card The critical limit is drawn as a wavy line in the graph on the Velocity card Additionally a compressibility correction is applied to the incompre
18. element airfoils is limited by onset of flow separation The achievable limit for single element airfoils seems to be at lift coefficients between 2 and 3 For maximun lift it can be beneficial to split an airfoil unto several elements arranged to form a slotted cascade Each element then develops its own fresh boundary layer and positive interference effects between the elements allow for a higher lift loading per element JAVAFOIL can handle such multi element airfoils to a limited extent Limitations are imposed by the fact that boundary layer effects are not modeled Therefore inaccurate results must be expected when slots are very narrow less than twice the displacement thickness of the boundary layer and when the wake of a leading element interacts with a following element Nevertheless JAVAFOIL should be useful to produce a reasonable first design for a slotted airfoil with appropriate gap overlap and element angle settings Also one can design such sections so that suction peaks and too steep pressure gradients are avoided 31 The following script shows how a two element downforce wing section can be generated starting with a basic NACA 4 digit section A simple JavaFoil example which creates a two element airfoil for downforce generation ea switch to US country settings Options Country 0 create a cambered NACA airfoil for starting Geometry CreateAirfoil 0 61 15 30 6 40 000 0 0 1 fee create a copy of
19. f attack would maintain the trailing edge point and elevate the nose of the airfoil above the y 25 line Note that the airfoil is rotated and thus its projection on the x c axis becomes shorter but pressure velocity or Mach number distributions on the Velocity card are still plotted over x c 30 nn Y 5 92 h c 0 1 Cp Calculation JavaFoil Experiment Steinbach Calculation Chun Chang 1 0 0 0 p Ta PR Eu N We abies ee RI w 0 4 0 6 0 8 1 x c Distributions of the pressure coefficient on a Clark Y airfoil in ground proximity Results of the numerical solution of the Navier Stokes equations have been taken from 17 the experimental results have been reproduced from 18 The experiments were carried out with a fixed ground board equipped with a suction system The results of JAVAFOIL match the experimental results quite well The Navier Stokes solutions should model boundary layer displacement effects more accurately The ground effect on a wing of a finite span is approximated by applying a modified calculation of the induced drag If you specify the aspect ratio of the wing A b S span b and wing area S and the height of the wing above ground h b height h over wing span b on the Options card these values are used to calculate an approximation of the induced drag using 33 h b GG l 1 33 h b Cp i Multi Element Airfoils The maximun lift of single
20. f the 3D effects These effects can applied to the polars produced by JAVAFOIL and make it possible to get a first impression of the relations between induced drag and airfoil drag For example the importance of the airfoil drag is diminishing for higher lift coefficients and lower aspect ratios The three dimensional corrections can be applied to the results for constant Reynolds number Polar card as well as more realistically for the results associated with a constant wing loading Aircraft card G Am S 50kg m A 00 NACA 0012 8 9 m s 50kg m A 5 Ci t m S 50kg m A 10 m S 50kg m A 20 1 0 0 0 amp i i i i 0 000 0 010 0 020 0 030 0 040 0 050 0 060 0 070 0 080 0 090 0 100 Cp Lift versus drag coefficient polars for a NACA 0012 airfoil and wings of different aspect ratio The graph above shows the effect of the wing aspect ratio on lift over drag coefficient Starting with infinite aspect ratio aspect ratio 0 on the Options card three wings with increasing aspect ratio have been analyzed For each curve the maximum of the lift over drag L D ratio is indicated by a filled circle It can be clearly seen that depending on the aspect 25 ratio the additional induced drag distorts the polar so that the optimum L D ratio is shifted to lower lift coefficients While the two dimensional airfoil achieves its maximum of L D at slightly above C 1 0 the low aspect ratio win
21. g of A 5 requires to operate the airfoil at C 0 5 because this is the optimum C of the whole wing If we compare with another airfoil we would better compare the airfoils at the lift coefficients corresponding to the wing aspect ratios Note that the results as shown above are accurate for a wing having an elliptical lift distribution and an elliptical untwisted planform Due to the spanwise lift distribution on a generic wing the airfoils along the span of the wing will operate somewhat above and below the total lift coefficient of the wing To study this effect requires using a more sophisticated three dimensional wing analysis code e g for subsonic flow lifting line vortex lattice or panel methods Also no additional wing effects like Reynolds number variation due to taper are taken into account Polars for Constant Wing Loading Airfoil data has traditionally been presented in form of graphs and tables for constant Reynolds numbers This form results from the typical way wind tunnel experiments and numerical analyses are conducted In a wind tunnel it is relatively easy to maintain a constant wind speed and Reynolds number Now the lift coefficient of a real airplane depends on the speed because the wing loading is usually fixed during flight flying at low lift coefficients results in high speeds and high Reynolds numbers and vice versa Therefore the operating points during flight would slice through a set of polars having c
22. ge radius maximum thickness at 50 x c NACA 5 digit airfoils These sections use the same thickness distributions as the 4 series but have new camber lines leading to lower pitching moments The camber line is composed of a cubic in the forward part to which a straight line is attached towards the rear Instead of the camber f c a design lift coefficient C design 18 now used to define the maximum height of the camber line In practical applications these airfoils are often used with a maximum camber at x c 0 15 i e relatively far forward Parameters e Free t c x c C design e Fixed x c 0 3 Naming Scheme e Ist digit design 10 2 3 C accion e 2nd and 3rd digit 2 100 x c Note that the 3 digit is usually a zero i e the position of the maximum camber is a multiple of 5 e 4th and 5th digit maximum thickness 100 t c Example NACA 23012 design lift coefficient 0 3 maximum camber at 15 chord 12 thickness Modified NACA 5 digit airfoils The rear part of the camber line of these sections has been modified to a cubic curve which provides some reflex Therefore the pitching moments are reduced further or may become even positive Parameters e Free t C design e Fixed x c 0 3 Naming Scheme e Ist digit design 10 2 3 C accion e 2nd and 3rd digit 2 100 x c plus 1 Assuming that the position of the maximum camber is a multiple of 5 the 3 digit is always a 1 e 4th and
23. he official description of the airfoil geometry of the EQ and EQH aerofoils especially how the quartic curve was defined and how the hyperbolic closure was attached to the quartic curve It seems to be that the procedure to generate these shapes was not published Biconvex airfoils These are symmetrical airfoils formed by two arcs They can be represented by the following formula y a x 2 The exponent b can be found from the location of the maximum thickness i e the point where Oy Ox 0 1 1 F X max len b while the factor a depends on the value of the maximum thickness tie a x t max t max If the maximum thickness is placed at x c 0 5 the airfoil is composed of two equal circular arcs These airfoils are intended for supersonic flow Parameters e Free t c x c Double Wedge airfoils These are symmetrical airfoils composed of straight lines They are intended for supersonic flow 11 Parameters e Free t c x c Plate airfoils These sections are generated to represent flat plates with rounded noses and sharp trailing edges The shape can be superimposed over a NACA 4 series camber line to produce a cambered plate Parameters e Free t c f c x c e Fixed leading edge radius r 1 2 t c trailing edge closure begins at x c 08 Newman airfoils These sections consist of a circular nose to which straight tapered tail is attached It can be manufactured easi
24. he boundary layer method does not include any feedback to the potential flow solution which means that it is limited to mostly attached flows Flow separation as it occurs at stall is modeled to some extent by empirical corrections so that maximum lift can be estimated for conventional airfoils If you analyze an airfoil beyond stall the results will be quite inaccurate On the other hand it is somewhat questionable whether any two dimensional analysis method can be used at all in this regime as the flow field beyond stall becomes fully three dimensional with spanwise flow and strong vortices developing In the case of multi element airfoils one must be aware that in the real world very complex flows can develop due to interaction of trailing wakes and the boundary layers of the individual elements or if the boundary layers separate locally An accurate analysis would require a more sophisticated solver for the Navier Stokes equations which would also imply an increase in computer time in the order of 1000 Nevertheless a simple tool like JAVAFOIL can be helpful to estimate the main effects and to improve a design to avoid suction peaks and flow separation JAVAFOIL s Cards The user interface of JAVAFOIL is divided into a stack of cards Each card contains interface elements for a specific task The content of some cards is also relevant for actions executed on other cards for example the Mach number specified on the Options card affect
25. ign Velocity Flow field Boundary Layer Polar Aircraft Options Modify Airfoil o Neme acao TraingEoge Gap 2 Di Element Number of Points jr Pivot x jso Thickness tic jessa Pivot y eo ay re Sealey a Tension x 5 be js 3 0 Cl Camber fic Flap Chord xfic Translation y 13 1 Flap Deflection 5 Duplicate Delete Fipy te 5 865 64 83 1 0 Thickness and Camber work in y direction only Und NACA 0012 Copy Text View of the Modify card showing a two element airfoil with element 2 selected The following modifications can be performed NAME Changes the name of the airfoil NUMBER OF POINTS Changes the number of coordinate points of the selected element s THICKNESS Scales the thickness of the selected element s by decomposing the shape into a thickness distribution and a camber line Only the thickness distribution is scaled so that the camber line is maintained Note that small changes to the camber may occur due to numerical errors 15 CAMBER Scales the camber line to a new height This works only ifthe airfoil is already cambered Scaling the camber line of a symmetrical airfoil accomplishes nothing SCALE BY Scales the airfoil shape by multiplying the coordinates with the given scaling factor FLAP DEFLECTION Modifies the coordinates by deflecting a plain flap of the given chord length The axis of rotation is always the middle betwee
26. inclined e g close to the leading edge or close to the trailing edge of airfoils with a high amount of aft camber The construction method may lead to points extending slightly into the negative x range when a large amount of camber is located close to the leading edge This is a correct behavior and an expected result Note also that most NACA sections have a thick trailing edge by definition In order to produce a thin sharp trailing edge JAVAFOIL has an option to close the airfoil shape by bending the upper und lower surfaces to close the trailing edge NACA 4 digit airfoils The calculation of these classical airfoils is easy because their shape and the associated camber lines are defined by rather simple formulas The maximum thickness is located at x c 30 whereas the maximum camber is typically located at x c 40 See 3 and 4 for more details The camber lines are composed of two parabolic arcs which are joined with equal tangents but a kink in the curvature This kink can be seen in the velocity distributions especially when the position of the maximum camber is different from the common 40 chord station Parameters e Free t c f c x c e Fixed x c 0 3 Naming Scheme The first two integers define the camber line while the last two integers define the thickness e Ist digit maximum ordinate of camber 100 f c e 2nd digit location of maximum camber 10 x c e 3rd and 4th digit maximum thickness 100
27. ion of Aerodynamic Flows p 32 vV Co 24 Finite Wings in JAVAFOIL In the 1920s it has been found by Prandtl and also by Lanchester that the finite span of wings affects their aerodynamic performance They found that the effects could be expressed as a function of the aspect ratio a k a slenderness or finesse of the wing Prandtl s Lifting Line theory was developed and successfully applied to design wings up to the 1940s and even today it is useful for unswept wings of relatively high aspect ratio A gt 5 The aspect ratio can be determined from A b b S span b divided by the mean chord length or span squared divided by wing area S The main result of this theory is that the airfoil drag is increased by an additional drag force induced drag a k a vortex drag which is caused by the finite wing span The vortex drag coefficient of a wing can be expressed by Cy induced K Cr T A where C is the lift coefficient of the whole wing and k is a factor to account for the shape of the lift force distribution along the span for good wing designs k 1 Now JAVAFOIL is a program for the analysis of two dimensional airfoils Nevertheless it supports a very simple model of finite wings to allow for a more realistic comparison of airfoils When the user supplies a value for the aspect ratio on the Options card classical wing theory formulas are used to determine an approximation o
28. le on a composite material sailplane wing r 1 smooth but slightly rough surface as for example a painted cloth surface r 2 similar to the NACA standard roughness r 3 dirty surface with spots of dirt bugs and flies Note that the NACA standard roughness is usually applied to the leading edge only It consists of a sparse 5 10 of the area leading edge coating up to 8 x c The grain size is about 0 45 bo of the chord length Thus for a wing chord length of Im the grain size would be 0 45mm Stall Corrections Empirical Stall Correction 1 CalcFoil if a gt 0 handle separation on upper surface drag increment Cy upper Cy upper F 2 sep upper 2 0 025 cosa x0 e sep upper sin a X x lift multiplier reduces lift linearly with length of separated length En 1 0 2 Ku X se else if a lt 0 handle separation on lower surface drag increment 2 2 Cy lower Cy lower ag sin Oe x0 X sop ie ER xi X sep ies lift multiplier reduces lift linearly with length of separated length C C a 1 my 0 2 i a X sop lower moment multiplier 0 9 x x sep lower sep upper m corrected Cn panel method lift multiplier due to suction peak criterion 21 CC I 5 where AC AC max Prams is the difference between the minimum pressure P max 20 coefficient close to the nose of the ai
29. ly but has a curvature jump at the junction between the nose and the trailing wedge leading to suction peaks and a risk of flow separation Parameters e Free t c Joukowsky airfoils These classical airfoil sections are generated by applying a conformal mapping procedure They were the first practical airfoils developed on a theoretical model Besides producing the airfoil shape the mapping procedure was also used to find the flow field around the airfoil as well as the force and the moment acting on the wing section The airfoils have very thin cusped trailing edges and are therefore difficult to analyze with panel methods and difficult to manufacture The conformal mapping is performed using the Joukowsky transformation of the complex points Zae ON a unit circle with is center at x y circle 12 Z airfoil Z circle Ar where A Xo T 1 Z vi 2 circle In order to match the prescribed airfoil thickness and camber JAVAFOIL performs an iterative search for the center of the circle As usual the resulting coordinates are scaled to unit length Parameters u Free t c f c Van de Vooren airfoils In contrast to the classical Joukowsky airfoils these airfoils have a finite trailing edge angle The transformation function is of the type 1 zu RR Z airfoil kl e m Zani They can be used to create sections with thick trailing edge regions e g for fairings A description of this shape can be f
30. ment thickness 6 momentum thickness Ol shape factor displacement thickness momentum thickness H F 2 Reynolds number based on local momentum thickness Re Re Reynolds number based on local arc length Re Effect of Roughness The effect of roughness on transition and drag is complex and cannot be simulated accurately Even modern resource hungry direct numerical simulation methods have difficulties to simulate the effect In JAVAFOIL two effects of surface roughness are modeled m laminar flow on a rough surface will be destabilized leading to premature transition m laminar as well as turbulent flow on rough surfaces produce a higher skin friction drag The effect on toughness is modeled in the following transition models Eppler IR 4H 21 74 0 36 ppe Transition is assumed to occur when Re gt e 11170307 Standard Eppler ea 18 4 Hyy 21 74 4125 Hy49 1 573 0 36 PP Transition is assumed to occur when Re 2e a cae enhanced 2 Drela Transition is assumed to occur when the value n 9 r is exceeded 20 e approx Arnal W rz Transition is assumed to occur when the value n 9 r is exceeded rz The global effect on drag is taken into account by a simple scaling of the total drag coefficient C C 1 r 10 The roughness factor r is meant to represent the following surface conditions r 0 perfect smooth surface as for examp
31. moothing factor is 0 1 the y coordinate of the smoothed point is 90 of its initial value and 10 of the linear interpolation between the two neighboring points according to oa 1 f yi f Yia t Yaa a eee iH T Sia This filter can be applied several times but subsequent application will also smooth out details like a pointed airfoil nose This option is only available via the scripting language 16 Jame NACA 0012 5 11 928 2 View of the controls on the Modify card You can also modify individual points by dragging them up or down with the left mouse button depressed This modification method is restricted to movements in the y direction If you need more freedom you have to modify the numerical coordinate values on the Geometry card Note concerning multi element airfoils Modifications are applied only to the airfoil element s selected in the Elements list box The selection is also used by other cards Only selected elements are taken into account when total force moment and drag coefficients are determined The Panel Method JAVAFOIL implements a classical panel method to determine the linear potential flow field around single and multi element airfoils In JAVAFOIL the airfoil surfaces carry a linearly varying vorticity distribution This is the same type of distribution as used in XFOIL but simpler than the higher order parabolic distribution as used in Eppler s PROFIL code The resulting equation
32. n upper and lower surface TRAILING EDGE GAP Modifies the shape so that the prescribed trailing edge gap is produced Generally it is recommended to use closed trailing edges for analysis except if the airfoil is extremely thin towards the trailing edge This function can also be applied before exporting airfoil shapes suitable for manufacturing ROTATE Rotates the selected airfoil element s around the specified Pivot point TRANSLATE X Moves the selected airfoil element s by the given distance horizontally TRANSLATE Y Moves the selected airfoil element s by the given distance vertically DUPLICATE Creates a copy of the currently selected element s Note that you have to move the new element from its initial location so that it is not overlapping with other elements DELETE Deletes the selected element s FLIP Y Reflects the selected elements across a horizontal line passing through the pivot point SMOOTH Y This command currently supports two smoothing variants If the smoothing factor is positive the coordinates are approximated by a smoothing spline curve Only this first variant is available via the graphical user interface There it uses a hard wired smoothing factor of 0 1 Other factors can be used when the scripting interface is used If the smoothing factor is negative a filter is applied to the y coordinates to reduce waviness This filter applies a weighted average to each point and its two neighbor points If for example the s
33. ng area S on the Options card The following correction is applied to the lift coefficient of a 2D airfoil C in order to approximate the lift coefficient C of the 3D wing in compressible flow The correction is divided into two regimes of aspect ratios For small aspect ratios A lt 4 the following formula is used C 2 gt 2 7 2 7 1 M gt AT AT If the aspect ratio is larger A gt 4 the simplified approximation is applied C d 2 7 J1 M2 u Implementation in JAVAFOIL public final static double LiftForAspectRatio double dCl double dAspectRatio double dMachNumber double dReturn dCl correction for finite wings if dAspectRatio gt 0 1 Source Anderson Aircraft Performance and Design lift gradient reduction factor a_0 pi AR double dGradientRatio 2 0 Math PI Math PI dAspectRatio if dAspectRatio lt 4 0 low aspect ratio compressible Anderson 2 18b dReturn Math sqrt 1 0 Math pow dMachNumber 2 0 Math pow dGradientRatio 2 0 dGradientRatio else high aspect ratio compressible Anderson 2 16 dReturn Math sqrt 1 0 Math pow dMachNumber 2 0 dGradientRatio return dReturn 28 Determination of Drag for given Aspect Ratio and Mach number After the lift coefficient of the 2D airfoil for a given angle of attack a has been corrected to the effect of the 3D wing an approxim
34. nimum negative camber at x c 0 9204 e The maximum camber is linked to the thickness by the expression f c 0 168 t c I am still looking for more information about Russian airfoil developments NPL EC ECH and EQH airfoils These British symmetrical airfoil sections are composed of an elliptical forward portion E and a cubic C or quartic Q rear end The tail closure is built from a hyperbolic curve H series The location of the maximum thickness can be varied between 30 and 70 of the chord length A limited description is contained in 10 13 Parameters 10 Parameters e Free t c x c After some reverse engineering I have used the following assumptions for these airfoils e the trailing edge thickness is 2 of the airfoil thickness e incase of the C and Q series the rear end is attached with C0 C1 C2 continuity position tangent curvature to the elliptic front part e incase of the Q series the second derivative at the trailing edge is set to 0 2 this gave the best approximations for 1240 to 1260 airfoils e the H modification closes the thick trailing edge by a hyperbolic curve which is attached with CO C1 continuity position tangent to the thickness distribution at x c 0 965 Camber lines are 3 order polynomials which allow to place the location of the maximum camber approximately between 30 and 60 of the chord length Note I am still looking for t
35. ntial equations describing the growth of boundary layer parameters depending on the external local flow velocity These equations are then integrated starting at the stagnation point While accurate analytical formulations are available for laminar boundary layers some empirical correlations are needed to model the turbulent part Note the local skin friction coefficient as given on the Boundary Layer card is twice the value as used by Eppler to follow the more common convention C T v In JAVAFOIL there is no interaction between the boundary layer and the external flow as in XFOIL though Therefore largely separated flows cannot be analyzed a short flow separation Separated lt 10 at the trailing edge does not affect the results very much Also laminar separation bubbles are not modeled when laminar separation is detected the code switches to turbulent flow Transition Criteria Methods to predict transition from laminar to turbulent flow have been developed by many authors since the early days of Prandtl s boundary layer theory While it is possible to analyze the stability of a boundary layer numerically all methods which are practical and fast are more or less approximate and rely on empirical relations usually derived from experiments Because the local boundary layer parameters at a station s are the result of an integration process starting at the stagnation point they contain a history of the flow
36. onformal mapping algorithm was able to deliver a shape for a given pressure distribution This means that no closed form equations describing the thickness distributions exist Earlier JAVAFOIL versions used a very approximate algorithm which had been lifted from the Digital Datcom programs but this produced rather inaccurate representations of the 6 series airfoils Therefore since version 2 09 August 2009 JAVAFOIL uses a more elaborate algorithm which is based on the work of Ladson 6 This new method is using quite accurate tables of the stream function for most of the 6 series airfoils JAVAFOIL can generate individual members of the 63 64 65 66 and 67 as well as the 63A 64A and 65A families The A modification leads to a less cusped trailing edge region The 63 64 65 66 and 67 families can be combined with camber lines of the a 0 toa 1 type The 63A 64A and 65A sections use a modified a 0 8 camber line which is straight aft of x c 0 8 The thickness distribution of these airfoils has also been modified to yield straight lines from x c 0 8 to the trailing edge The a camber line shapes are specified in terms of the design lift coefficient and the position x c where the constant loading ends This is indicated with an additional a x y label in the airfoil name If you specify a gt 1 in JAVAFOIL s input the camber line has a constant loading from leading edge to trailing edge The resulting airfoils do n
37. onstant Reynolds numbers It is possible to create polars more closely related to the conditions during flight This would require adjusting the wind speed to each lift coefficient which is cumbersome and expensive in a wind tunnel but feasible in a numerical tool like JAVAFOIL Abbreviations mass of aircraft m kg gravity constant g density of medium Px m s kinematic viscosity v m s flight speed Vo m s wing area S m chord length m Reynolds number Re Basic Equations The definition of the lift coefficient is C a Solving the definition of the a Soe Vo f Re v A j Reynolds number Re for the velocity v yields v a Inserting this result v into the definition of the lift coefficient produces m g l Cus 5 Rem s gt e V S Solving for the Reynolds number yields 26 N 2 Note that this equation can also be written Re C a VV Poo can also calculate polars of constant Re C to match a given aircraft m g 3 which means that we Using these results one can derive an aircraft oriented airfoil polar for a given wing loading and given mean chord length Due to the dependency between lift coefficient and Reynolds number an iterative calculation procedure is used e prescribe the environmental condition density p and kinematic viscosity v e prescribe a wing loading z and a reference chord length e
38. ot carry the a label Note that officially no intermediate airfoils e g a NACA64 5 012 exist Naming Scheme e st digit 6 series designation e 2nd digit chordwise position of minimum pressure of the thickness distribution 10 x c e single digit suffix following a comma which is 10 AC It represents the range AC above and below C design Where favorable accelerating pressure gradients for laminar flow exist therefore AC is approximately the semi width of laminar bucket e adash e 3rd digit design 10 C gesien e 4rd and Sth digit maximum thickness 100 t c A camber line shape different from a 1 0 is indicated by the additional designation a x y where x y is replaced by the location x c where the constant part of the loading ends and the linear drop towards the trailing edge starts TsAGI B airfoils The TsAGI also ZAGI CAGI was and is Russia s leading aeronautical research organization Not much is known about early airfoil development but the available literature 6 9 shows that similar to other nations Russia has developed airfoil families based on analytical shape descriptions The TsAGI series B is just one such airfoil family The very simple shape description is using just the maximum thickness The resulting sections have a reflexed camber line and hence low pitching moment Parameters e Free t c e Fixed x c 0 3388 maximum positive camber at x c 0 3018 mi
39. ound in 13 Note that not all thicknesses can be achieved for all trailing edge angles therefore the final maximum thickness may not be what was desired Also only symmetrical sections are generated in JAVAFOIL Parameters Free t c Opp Helmbold Keune airfoils In the 1940s many attempts were made to extend the then classical NACA airfoil section methodology to more general airfoil shapes Helmbold and Keune 15 developed elaborate methods to characterize and parameterize airfoil sections While the mathematical approach allowed for representation of a wide range of shapes the methodology was not really successful in these years of manual calculation Later in the age of numerical shape optimization similar methods have been developed e g the Parsec shape functions The parameters of the symmetrical airfoil must be carefully chosen to generate a realistic airfoil shape The center curvature must be large enough to avoid self crossing of the outline 13 curvature Parameters Free t c x c trailing edge angle curvature radius at middle nose radius Ro ner airfoils Another algorithm to generate analytical airfoil shapes based on conformal mapping was published by Ro ner 16 Like all methods using conformal mapping his solution also allowed for the exact analytical determination of the corresponding pressure distributions Parameters Free t c x c trailing edge angle nose radius Pa
40. per left corner of the image Therefore the image should have no border and a monochrome background Before scanning the image a smoothing filter is applied to remove spurious points from the image To achieve acceptable results an image width of 1000 or more pixels is recommended The interior of the airfoil shape can be empty or arbitrarily filled because the algorithm searches from the top and bottom edges of the image and stops when it detects the border of the shape The resulting points are filtered again to improve the smoothness of the shape Nevertheless the results will not be perfect but this method can be considered as a last resort to quickly determine airfoil coordinates if only a scanned image is available It is recommended to inspect the resulting velocity distribution and to use the inverse design method for smoothing the airfoil shape Airfoil image top and comparison of the original dashed and the reconstructed airfoil shape solid using JAVAFOIL s bitmap import capability on the Geometry card JAVAFOIL S Geometry Generators General Remarks on NACA airfoils The construction of the cambered NACA airfoil sections requires that the thickness distribution is erected at right angles to the camber line Some computer programs do not follow this construction principle and add the thickness just to the y coordinates of the camber line This leads to larger deviations from the true airfoil section when the camber line is
41. rfoil and the pressure close to the trailing edge Empirical Stall Correction 2 Eppler if a gt 0 handle separation on upper surface if X sop upper lt Xp trailing edge angle of upper surface Y sep upper YTE Org arctan X sop upper X else O O drag increment 2 Ca upper za Cy upper oF 0 2 g sin a ag Orp ie X sop ius AC Ze C max fudge F a Ore T re X sep upper if AC gt 0 lift reduction C C 4C else lift multiplier C C 1 sina kr X gep upper moment increment Ca Cy Z Sin O X15 Kuppe 0 5 I x 0 25 sep upper else if a lt 0 handle separation on lower surface if x Rn sep lower 22 trailing edge angle of lower surface Y sep lower YTE Ore arctan ____ X sop lower Xp drag increment 2 Cy lower Cy lower 0 2 sm a a Org ee z X sop en AG gt C max fudge a Orp Te e E X sep if AC lt 0 lift reduction C C AC else lift multiplier eb f sina Kr e ja moment increment Ca Ca sina e hey ee 0 5 1 FI 0 25 j lift multiplier due to modified suction peak criterion Cy Cj where AC is the difference between the minimum pressure P max coefficient close to the nose of the airfoil and the pressure close to the trailing edge Compressi
42. rious boundary layer parameters The boundary layer data is then be used to calculate the drag of the airfoil from its properties at the trailing edge Both analysis steps are repeated for each angle of attack which yields a complete polar of the airfoil for one fixed Reynolds number Additional tools for the creation and modification of airfoils have been added to fill the toolbox These tools are wrapped in a Graphical User Interface GUI which was designed to be easy to use and not overly complicated The GUI is organized into a stack of cards which will be described later All calculations are performed by a computer code of my own JAVAFOIL is neither a rewrite of Eppler PROFIL nor of Drela s XFOIL program The boundary layer module is based on the same equations which are also used in the initial version of the Eppler program Additions include new stall and transition models The panel method was developed with the help of the extensive survey of panel methods found in 14 Compared with similar programs JAVAFOIL can also handle multi element airfoils and also simulate ground effect Limitations As already noted JAVAFOIL is a relatively simple program with some limitations Like with all engineering computer codes it is up to the user to judge and to decide how far he wants to trust a program As JAVAFOIL does not model laminar separation bubbles and flow separation its results will become inaccurate if such effects occur T
43. rsec airfoils The Parsec geometry parameterization was developed by H Sobietzky in the 1990s It tries to model airfoil shapes by superposition of selected polynomial terms The parameters resemble the Helmbold Keune approach and are mainly intended to be used for numerical shape optimization The parameters of the symmetrical airfoil must be carefully chosen to generate a realistic airfoil shape The center radius the nose radius as well as the trailing edge wedge angle must be carefully adjusted to avoid self crossing of the outline Due to the limited number of text entry fields in JAVAFOIL user interface the parameters of a Parsec 11 formulation have reduced so that they model symmetrical sections only Parameters Free t c x c trailing edge angle curvature at x c nose radius Modify Card This card can be used to perform various modifications to the airfoil geometry It consists of an input and action area and a geometry view below The modification of parameters is performed by entering new values into a text field and then pressing the button at the left of the text field Thus it is easy to apply certain operations 14 several times A modification will be applied to the airfoil elements which are currently selected in the Element list box The geometry view is automatically scaled to fit all airfoil elements The currently selected elements are highlighted in red JavaFoil E 10 x Geometry Modify Des
44. s the analyses on all other cards Geometry Card The Geometry card is used to store and prepare the geometry of your airfoil It contains the current or working airfoil The geometry is described by a set of coordinate points each having an x and a y value The working airfoil is used and modified by the actions you perform in JAVAFOIL The Geometry card shows a list of x and y coordinate pairs and a plot of the resulting airfoil shape You can enter or paste arbitrary coordinates into this field and press the Update View button to copy the coordinates into the working airfoil The coordinates must be ordered so that they describe the shape in a continuous sequence The order must be trailing edge upper surface nose lower surface trailing edge JAVAFOIL comes with a set of shape generators for a variety of airfoils which is accessible from this card These airfoils represent classical airfoil sections for which analytical descriptions exist e g NACA sections or which can be constructed from geometrical constraints e g wedge sections Despite their age many classical airfoil sections are still applicable to many problems or form a good starting point for new developments Today modern airfoil sections are usually developed for specific purposes and their shapes are usually not published More recent developments lead towards the direct design of three dimensional
45. ssible solution to model first order compressibility effects Note however that the theory becomes invalid when flow reaches or exceeds sonic speed In JAVAFOIL the critical pressure coefficient is calculated from the relation 2 2 Jt 1 k 1 2 p crit kK M K chemann The Aerodynamic Design of Aircraft p 115 In terms of the velocity ratio the critical limit is found from 2 SARET E eee crit k 1 M Co K chemann The Aerodynamic Design of Aircraft p 114 Compressibility Corrections There are different ways to correct incompressible flow results for compressibility effects One should keep in mind that these are only corrections they can never produce the correct physical effects Therefore the applicability of all compressibility corrections is limited to cases where the local flow velocity which can be much higher than the onset flow velocity is well beyond the speed of sound In practical application one can use such corrections well up to about Mach 0 5 the error grows very rapidly when Mach exceeds 0 7 In JAVAFOIL the panel analysis is always running on the given airfoil the shape is never geometrically distorted Compressibility corrections are applied later to the local surface velocities according to the Karman Tsien approximation usually written for C M l 5 W AE a Vo comp Vo inc M V if 2 Mo inc Cebeci An Engineering Approach to the Calculat
46. t showing a window you must make sure that the script ends with an Exit command to terminate the JAVAFOIL run properly Otherwise the JAVAFOIL process will continue to run In Windows you can check for running JAVAFOIL processes using the Task Manager window In Unix like operating systems you can use the ps command to list all processes running under your user account References 1 Mark Drela Michael B Giles Viscous Inviscid Analysis of Transonic and Low Reynolds Number Airfoils AIAA 86 1786 CP 1986 2 Xiao liang Wang Xue xiong Shan Shape Optimization of Stratosphere Airship Journal of Aircraft V43N1 2006 3 Ira Abbott and Albert Von Doenhoff Theory of Wing Sections 4 Eastman N Jacobs Kenneth E Ward Robert M Pinkerton The Characteristics of 78 Related Airfoil Sections from Tests in the Variable Density Wind Tunnel NACA Rep No 460 1933 5 John Stack Tests of Airfoils Designed to Delay the Compressibility Burble NACA Rep No 763 1943 6 Charles L Ladson and Cuyler W Brooks Jr Development of a Computer Program To Obtain Ordinates for NACA 6 and 6A Series Airfoils NASA Technical Memorandum TM X 3069 September 1974 7 Charles L Ladson Cuyler W Brooks Jr and Acquila S Hill Computer Program To Obtain Ordinates for NACA Airfoils NASA Technical Memorandum 4741 December 1996 8 A K Martynov Practical Aerodynamics Pergamon Press 19
47. wing shapes eliminating the classical steps of two dimensional airfoil design and three dimensional wing lofting In most cases modern airfoil sections are not described anymore by analytical formulas just by a set of points The row of buttons at the bottom allows for copying saving loading and printing of airfoil coordinate sets 10 x Geometry Modity Design Velocity Flowtield Boundary Layer Polar Aircraft Options Airfoil Geometry Name Coordinates Clear Naca 0012 Create an Airfoil 1 00000000 0 00000000 aj Family NACA 4 digit e g 2412 0 99726095 0 00038712 Bei 0 98907380 0 00153653 Number of Points jet 0 97552826 0 00341331 0 95677273 0 00596209 Thickness tic 12 0 93301270 0 00911073 0 90450850 0 01277464 Thickness Location xtic 30 0 87157241 0 01686084 0 83456530 0 02127128 Camber fic 0 000 1 en ces amier leesten ste 0000 ma 0 70336832 0 03542434 BR 56 0 65450850 0 04009273 O 60395585 0 04454642 IV Modify NACA section to have closed trailing edge 0 55226423 0 04866201 0 50000000 0 05231025 This is a general purpose airfoil series 0 44773577 0 05535862 0 34549150 0 05913940 gt Airfoil Shape For later analysis the trailing edge should be closed Update View Copy Text Paste Text Open Print View of JAVAFOIL s Geometry card Exporting airfoil geometry JAVAFOIL can write airfoil geometry from the following file types
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