AIRFOIL SECTION

Abstract: The invention refers to an airfoil section that can be used in the areospace and naval industries and also to construct propellers, fans, air-conditioning equipment, hydraulic turbines, missiles, and torpedoes. The principle requires to meet the Kutta-Jukovsky condition at both the leading and trailing edges, i.e. a zero jump of the pressure in these points. This condition is satisfied if the equalities g(0)=g'(0)=0, and g(1)=g'(1)=0 hold where g(x) is the one half thickness distribution along the chord, and g'(x) is the first order derivative of g, both being taken in the specified points. The above-mentioned conditions can be used either to modify the leading and trailing edges of a given classical profile, or to generate new airfoil sections by using a definite skeleton line and a definite thickness distribution. Using this airfoil section, new lifting surfaces with higher performances as to the present-day configurations can be constructed.

Description: AIRFOIL SECTION TECHNICAL FIELD The invention refers to an airfoil section, especially used in the aerospace industry for the components of airplanes, rockets, space - shuttles, propellers, in the naval industry for the hull of ships, propellers, submarines also for fans, air- conditioning plants, hydraulic turbines, missiles and iorpedoes. BACKGROUND ART The class of airfoil sections having a round leading edge is well- known each member of this class being characterized by a leading edge in form of a circle with a given radius. The main disadvantage associated with the above-mentioned class consists in the existence of a stagnation point, which causes great loss of mechanical energy .that is a great pressure drag coefficient C D . Increasing flight speed worsens the phenomenon that takes dramatic proportions in supersonic and hypersonic regimes . Well-known are also the airfoil sections having an acute leading edge defined by a dihedral angle, that is the conical profile, or those having a half moon shape. The disadvantages of these airfoil sections are the early flow separations that release many negative effects, among which the abrupt increase of drag. DISCLOSURE OF INVENTION According to this invention the airfoil section is obtained by modifying the geometry of the leading edge of a given profile so that for every skeleton line shape, the half thickness distribution g(x) satisfies the conditions g(0)=g'(0)=0. It means that thickness and its first order derivative cancel in the leading edge. The modified profile consists on the upper side of a concave arc tangent to the skeleton line s(x) in point O. which passes through the first inflection point on the upper side, and a convex arc which passes through this inflection point, and also through the first junction point (with the old profile) on the upper side having a defined abscissa. On the lower side the leading edge outline consists of a convex arc tangent to s(x) in the point O, which passes through the first inflection point on the lower side, and a concave arc which passes through this point and the first junction point on the lower side having also a defined abscissa . The airfoil section, according to the invention, is obtained by modifying also the trailing edge geometry , so that the one half thickness distribution satisfies in the trailing edge point of an airfoil section the conditions g( 1 )=g'( 1 )=0. i.e. the thickness and its first order derivative cancel in this point. On the upper side the modified profile consists of a convex arc which passes through the second junction point on the upper side, also through the second inflection point on the upper side, and a concave arc which passes through the above mentioned second inflection point and is tangent to the skeleton line in the trailing edge F. On the lower side the modified profile consists of a concave arc which passes through the second junction point on the lower side, also through the second inflection point on the lower side, and a convex arc which passes through the second inflection point on the lower side, being also tangent to the skeleton line in the trailing edge F. The airfoil section, according to the invention, can be also obtained by using a defined skeleton line s(x) and constructing an one half thickness distribution g(x) which satisfies in the leading and trailing edges the above-mentioned conditions, the profile geometry being determined by convex and concave arcs which pass through the mentioned points, i.e. the defined ones. The airfoil section, according to the invention, can be also used for the construction of aerodynamic axial symmetric bodies by turning the bi-directional profile around the symmetry axis perpendicular on the middle of the chord. The main advantages of the proposed technical solution are: - leads to obtain a very rich class of airfoil sections which permits an adequate solution for every concrete design problem; - a reduced pressure drag, the efficiency of the solution is improving while increasing flight speed; - the possibility to obtain motions without shock waves on the body surface: - high lift coefficients at low speeds: - high lift/drag ratios at all speeds; - a relatively uniform loading distribution along the chord, that is a reduction of pressure gradients in the leading edge zone; - an extension of the useful incidence angles range, and the framing of this range between convenable limits according to the imposed design criteria; - a significant delay of the stall; - the possibility of moving the aerodynamic center forwards, to the leading edge, or downwards to 50 percent of the chord .in order to reduce the jump of the pitching moment, when the body passes from subsonic to supersonic regime; - small gradients of pitching moments with the useful incidence angles at low and high speeds: - the possibility to obtain the bi-directional profile, i.e. the airfoil sections for which the leading and trailing edges have the same statute with respect to the aerodynamic performances: - the axisymmetric bodies obtained by turning a bi-directional profile around the symmetry axis which is perpendicular on the middle of its chord has the same aerodynamic performances in all flight directions, and the aerodynamic center lies in a small neighbourhood of the symmetry axis at all flight regimes, i.e. subsonic, transonic. supersonic and hypersonic. BRIEF DESCRIPTION OF DRAWINGS Further, four examples of realizing the invention are given, also by using Figs.1-6 which represent: Fig.1 - a modified rounded nose airfoil section; Fig.2 - the pressure coefficients on the airfoil section NACA 23012. and a modified version of this airfoil section in the leading edge zone, both at α = 3.3°, incompressible regime; Fig.3 - an airfoil section with modified leading and trailing edges zones; Fig.4 - the one half thickness distribution along the chord for an arbitrary imposed skeleton line; Fig.5 - some modified airfoil sections having the thickness distribution defined by using the skeleton line which is supposed to be a positive function; Fig.6 - a bi-directional airfoil section which by turning around the perpendicular axis passing through the middle of the chord produces an axial symmetric body with equal aerodynamic performances in all flight directions. BEST MODE FOR CARYING OUT Let us consider an airfoil section having a rounded leading edge, and which is placed in the Oxz plane, its unity chord being along the Ox axis, and the origin O is that unique point on the contour where the Oz tangent line is perpendicular on the chord. There are used the well known notations: s(x) - the skeleton line distribution along the unity chord; g(x) - the one half thickness distribution along the chord; s'(x). g'(x) - the first order derivatives of functions s. and g respectively in the point having abscissa x; z e (x) - the upper side of the airfoil section. z,.(x)=s(x)+g(x); z,(x) - the lower side. z,(x)=s(x)-g(x); α - the incidence angle; V - the flight speed: C p - the pressure coefficient: C D - the pressure drag coefficient: C L - the lift coefficient. Because no matter the magnitude V of the flight speed, in the stagnation point (which in our case coincides with O) the local speed cancels that means that in a small neighbourhood of O the speed has great negative gradients and the first losses of kinetic energy occur. On the other hand, in another small neighbourhood of the stagnation point O. on both the upper and lower side the local speed abruptly increases unto values of V order, having great positive gradients, and again losses of kinetic energy. Therefore, the stagnation point produces two times losses of mechanical energy and this means increased pressure drag coefficient C D .This negative phenomenon is more and more significant with the increasing V. Having in mind that the C D coefficient is a very important performance criterium for every airfoil section, it means that the stagnation point must be removed, i.e. we must give up the class of rounded nose airfoil sections. This giving up is possible if we impose at the leading edge the Kutta-Jukovsky condition, i.e. the equality ΔC p (0)=0. where Δ C p the jump of pressure coefficient C p in this point is. The Kutta-Jukovsky condition is fulfilled if the thickness distribution satisfies in this point the conditions g(0)=g'(0)=0. i.e. the leading edge is sharp, and both the upper side Z j (x), also lower side z,(x) are tangent to the skeleton line s(x) in O. As a first example of realization, having in mind the imposed conditions, and using an airfoil section with rounded leading edge which has the known geometric characteristics, one obtains an airfoil section with modified leading edge. On the upper side the airfoil section consists of a concave arc 1 tangent to s(x) in the point O and passes through the point A t having a defined abscissa x,, also a convex arc 2 passing through A e and a point B e with defined abscissa x 2 . On the lower side the contour consists of a convex arc 3 tangent to s(x) in O, and passes through the point A, having a defined abscissa x,, also a concave arc 4 passing through A„ and through the point B, with defined abscissa x 4 . The abscissas x,, x,, x, and x 4 are defined by using the imposed design criteria, also the compatibility conditions as obtained from the continuity of contour and its gradient in these points. The A e . B e , A, and B, represent: A e - the first inflection point on the upper side; B e - the first junction point (with the old airfoil section) on the upper side; A, - the first inflection point on the lower side; B, - the first junction point on the lower side. The backwards limit position of junction points B e and B, is restricted by the imposed design criteria for the airfoil section. For illustration one considers a modified alternative of the airfoil section NACA 23012 in the leading edge zone which is defined by taking Xι=0.06, x 2 =0.17. x,=0.08, x 4 =0.2. The four arcs which define the modified airfoil section are approximated by conic sections. The pressure coefficients C p are theoretically obtained in ideal incompressible flow for the incidence angle α=3.3°, and observes that both the initial airfoil section with rounded leading edge and the modified airfoil section have a lift coefficient C L of about 0.5. For the modified airfoil section one finds a spectacular diminution of pressure gradients in the leading edge zone as in Fig.2. In a second example of realization the Kutta-Jukovsky condition must be imposed in the trailing edge too. and for this it is sufficient for the thickness distribution to satisfy at x= 1 the conditions g(l)=g'(l)=0, i.e. the trailing edge must be sharp, and both the upper and lower sides of contour be tangent to the skeleton line in this point. The modified airfoil section in the trailing edge zone consists on the upper side of a convex arc 5 passing through the points C e and D e having the defined abscissas x,. and x 6 respectively, also the concave arc 6 passing through D e and is tangent to s(x) in the trailing edge F having abscissa x=l . On the lower side the modified airfoil section composes of a concave arc 7 passing through the points C, and D, having the defined abscissas x 7 . and x 8 respectively. also a convex arc 8 passing through D, and is tangent to s(x) in the trailing edge F. The abscissas x 5 , x«, x 7 and x g are defined by using the imposed design criteria, and the compatibility conditions as obtained from continuity of contour and its gradient in these points. The points C e , D e , C„ D, represent: C e - the second junction point (with the old airfoil section) on the upper side; D e - the second inflection point on the upper side; C, - the second junction point on the lower side; D, - the second inflection point on the lower side. The forwards limit position of junction points C e and C, is restricted by the imposed design criteria for the modified airfoil section, and always the inequalities x 2 X|, x 4 >x-,, 6 X and x 8 >x 7 must be satisfied. In a third example of realization one considers an airfoil section which has the skeleton line s(x) as an arbitrary smooth function, and the one half thickness ditribution g(x) as a positive function which satisfies at the leading/trailing edges the conditions g(0)=g'(0)=0, and g(l) = g'(l)=0 respectively. In the leading edge zone g(x) composes of a concave arc 9 tangent to the Ox axis in the origin O, and passes through a point G having a defined abscissa x 9 , also a convex arc 10 passing through G and a point H which has the defined abscissa x, 0 .In the trailing edge zone g(x) composes of a convex arc 11 which passes through the points I and J having the defined abscissas x, and x, 2 respectively, also the concave arc 12 passing through J and is tangent to the Ox axis in the trailing edge F. The abscissas x 9 , x, 0 , x n , and x, 2 are defined by using the design criteria for the thickness distribution 2g(x) in the interval [0,1], also the compatibility conditions as obtained from the continuity conditions of g(x) and g'(x) in these points. If g(x) has two inflection points only along the whole chord then the arcs 10 and 1 1 coincide, i.e. we have x,ι=x 9 , and X| 2 =X| 0 - The points G,H,I, and J represent: G-the first inflection point of thickness distribution; H-the second inflection point of thickness distribution (if it exists); I-the last but one inflection point of thickness distribution (if it exists); J-the last inflection point of thickness distribution. Let us consider an airfoil section having a positive skeleton line s(x), tangent to the chord in both the leading and trailing edges, that means the conditions s'(0)=s'( 1 )=0 are satisfied. Moreover, let us admit s(x) as having a single extremum point (maximum) along the chord [0,1], and this maximum being ζ =0.05. attained at 30 percent of the chord as in Fig.5a). Using this skeleton line we can define the one half thickness distribution g(x) of the airfoil section with sharp leading and trailing edges by putting g(x)=ks(x). where k is a real positive number. Because the conditions g(0)=g'(0)=0 and g( 1 )=g'( 1 )=0 are satisfied, for every k we obtain an airfoil section which has the maximum thickness 0.1 k placed at 30 percent of the chord. In this case the function g(x) has a single maximum along the chord, also the arcs 10 and 1 1 coincide, i.e. we have x π =x<,<0.3, x, 2 =X| 0 >0.3. To illustrate the construction manner of an aero and/or hydrodynamic profile we consider in the relation g(x)=ks(x) three particular values of k as follows: i) k=l . In this case the upper surface is z e (x)=2s(x), and the lower surface is z,(x)=0 which coincides with the chord. The airfoil section has a flat lower side as in Fig.5b) and the maximum thickness is 10 percent of the chord: ii) k=l/2. Now we obtain a thi 5percent of the chord, the upper side is as in Fig.5c); iii) k=2. The obtained airfoil section is very thick, having the maximum thickness 20percent of the chord. The upper side is z c (x)=3s(x). the lower side is Zj(x)=-s(x). the airfoil section being sketched in Fig.5d). An important subclass of airfoil sections with sharp leading edge is that for which the skeleton line s(x) and the one half thickness distribution g(x) satisfy the equalities s(x)=s(l-x), g(x)=g(l-x), x[0,l], where g(x) satisfies also the conditions g(0)=g'(0)=0. These conditions tell that every airfoil section is symmetric about the axis (β) perpendicular to the middle of the chord as in Fig. 6, and also we have the relations X|,=l-x, 0 , x 12 =l-Xq, s'(0.5)*=g'(0.5)=0. Because from the aerodynamic point of view the leading and trailing edges can be permuted, we call this airfoil section as bi-directional. For given V and α the bi-directional airfoil section has the same aerodynamic performances when the flow attack it from the leading, or the trailing edges. In a fourth example of realization we turn a bi-directional airfoil section around the symmetry axis (β) and obtain an axisymmetric body which for given V and α has equal aerodynamic performances in all flight directions. REFERENCES 1. Abbot I.H., Doenhoff A.E.. Stivers L.S. Summary of airfoil data. NACA Rep.824. 1945, pp.258-522. 2. Liebeck R.H. A class of airfoils designed for high lift in incompressible flow. Journal of Aircraft, vol.10. Oct.1973. pp.610-617. 3. Schlichting H.. Truckenbrodt E. Aerodynamik des Flugzeuges, I. Springer Verlag, 1959. pp.371. 4. Landau L.. Lifchitz E. Mecanique des fluides. Editions Mir. Moscou 1971. pp.68-70. 5. Carafoli E.. Constantinescu V.N. Dinamica fluidelor incompresibile. Ed. Academiei. Bucuresti. 1981, pp.405-406.

Claims: CLAIMS 1. The airfoil section having a defined one half thickness distribution g(x) along the chord, and a skeleton line function s(x) corresponding to a known airfoil section, characterized by the fact that no matter the skeleton line or thickness shapes both the one half thickness and its first derivative cancels in the leading edge, i.e. we have g(0)=g'(0)=0, so that on the upper side the airfoil section composes of a concave arc (1) tangent to the skeleton line s(x) in the point (0), and which passes through a point (A e ) called the first inflexion point on the upper side and having the defined abscisa (x,). also the convex arc (2) passing through (A e ) and through a point (B e ) called the first junction point on the upper side having the defined abscissa (x : ). on the lower side the contour of the leading edge zone being composed of a convex arc (3) tangent to s(x) in O and passing through the point (A,) called the first inflexion point on the lower side and having the defined abscissa (x,), also passing through the point (B,) called the first junction point on the lower side having the defined abscissa (x 4 ), in the trailing edge zone both the function g(x) which define the one half thickness distribution, also its first derivative must cancel in the trailing edge. i.e. g( 1 )=g'( 1 )=0. so that on the upper side the airfoil section composes of a convex arc (5) which passes through the points (C e ) called the second junction point on the upper side and (D e ) called the second inflexion point on the upper side having the defined abscissas (x 5 ). and (x respectively, also the concave arc (6) which passes through (D e ) and is tangent to the skeleton line in the leading edge (F). on the lower side the airfoil section composes of a concave arc (7) which passes through the points (C,) called the second junction point on the lower side and (D,) called the second inflexion point on the lower side, having the defined abscisas (x 7 ) and (x 8 ) respectively, also a convex arc (8) passing through (D,) and is tangent to the skeleton line in the trailing edge (F). the abscissas (x,), (x 2 ), (x 3 ), (x 4 ), (x 5 ), (x , (x 7 ) and (x 8 ) being defined by using the imposed design criteria for the airfoil section and the compatibility conditions as obtained by asking for the continuity of contour and its gradient, the limit backwards positions of points (B e ) and (B,), and limit forwards positions of points (C e ) and (C,) being also conditioned by the imposed design criteria for the modified airfoil section, and always the restrictions (x 2 )>(x,), (x 4 )>(x,), (x 6 )>(x 5 ), (x 8 )>(x,) must by satisfied. 2. The airfoil section, according to claim 1 , characterized by a defined skeleton line s(x) that constructs a positive one half thickness distribution g(x) which satisfies in the leading and trailing edges the conditions g(0)=g'(0)=0, respectively g(l)=g'(l)=0. and which is formed in die leading edge zone of a concave arc (9) tangent to the (Ox) axis in (O), that passes through a point (G) called the first inflexion point of thickness distribution and having the defined abscissa (x 9 ), also a convex arc ( 10) passing through (G) and a point (H) called the second inflexion point of thickness distribution, having the defined abscissa (x, 0 ), in the trailing edge zone the function g(x) composes of a convex arc ( 11 ) passing through the points (I) called the last but one inflexion point of thickness distribution, and (J) called the last inflexion point of thickness distribution, having the defined abscissa (x, ,), and (x 12 ) respectively, also the concave arc (12) passing through the point (J) and is tangent to the (Ox) axis in the trailing edge (F). the abscissas (x 9 ), (x 10 ), (x ] t ) and (x, 2 ) being defined by using the design criteria imposed for the one half thickness distribution g(x) along the chord, also the continuity conditions for positions and slopes of component arcs. 3. The airfoil section according to claims 1 and 2, characterized by the fact that in the bi-directional configuration it can generate lifting axisymmetric bodies when rotating the airfoil section around the symmetry axis (β) perpendicular on the middle of the chord.

Inventors: TAPOSU IOSIF, TAPOSU, IOSIF
Applicants: ROMANIAN AIR TRANSPORT TAROM, TAPOSU IOSIF, ROMANIAN AIR TRANSPORT - TAROM -, TAPOSU, IOSIF
Application year: 1995
Main classification: F01D5
IPC codes: B64C3, F03D1