hapter 4 Thin airfoil theory 4. ompressible potential flow 4.. The full potential equation In compressible flow, both the lift and drag of a thin airfoil can be determined to a reasonable level of accuracy from an inviscid, irrotational model of the flow. Recall the equations developed in hapter 6 governing steady, irrotational, homentropic (rs = ) flow in the absence of body forces. r Ū = Ū Ū r P = P + rp = (4.) The gradient of the isentropic relation is Recall from the development in hapter 6 that rp = a r. (4.) P r = rp. (4.3) 4-
HAPTER 4. THIN AIRFOIL THEORY 4- Using (4.3) the momentum equation becomes r P + Ū Ū =. (4.4) Substitute (4.) into the continuity equation and use (4.3). becomes The continuity equation Ū ra +( ) a r Ū =. (4.5) Equate the Bernoulli integral to free stream conditions. a + Ū Ū = a + U = a + M = p T t (4.6) Note that the momentum equation is essentially equivalent to the statement that the stagnation temperature T t is constant throughout the flow. Using (4.6) we can write a = h t Ū Ū. (4.7) The continuity equation finally becomes ( ) h t Ū Ū Ū Ū r Ū Ū r =. (4.8) The equations governing compressible, steady, inviscid, irrotational motion reduce to a single equation for the velocity vector U. The irrotationality condition r Ū =permits the introduction of a velocity potential. and (4.8) becomes Ū = r (4.9) r r ( ) h t r r r r r =. (4.) For complex body shapes numerical methods are normally used to solve for. However the equation is of relatively limited applicability. If the flow is over a thick airfoil or a blu
HAPTER 4. THIN AIRFOIL THEORY 4-3 body for instance then the equation only applies to the subsonic Mach number regime at Mach numbers below the range where shocks begin to appear on the body. At high subsonic and supersonic Mach numbers where there are shocks then the homentropic assumption (4.) breaks down. Equation (4.8) also applies to internal flows without shocks such as fully expanded nozzle flow. 4.. The nonlinear small disturbance approximation In the case of a thin airfoil that only slightly disturbs the flow, equation (4.8) can be simplified using small disturbance theory. onsider the flow past a thin 3-D airfoil shown in Figure 4.. Figure 4.: Flow past a thin 3-D airfoil The velocity field consists of a free-stream flow plus a small disturbance Ū = U + u V = v (4.) W = w where u U << v U << w U <<. (4.) Similarly the state variables deviate only slightly from free-stream values
HAPTER 4. THIN AIRFOIL THEORY 4-4 P = P + P T = T + T (4.3) = + and a = a + a. (4.4) This decomposition of variables is substituted into equation (4.8). Various terms are and Ū Ū + uu + u + v + w (4.5) r Ū = u x + v y + w z = U Ū Ū r =(u x U + uu x + vv x + ww x,u y U + uu y + vv y + ww y,u z U + uu z + vv z + ww z ) (4.6) as well as ( ) h t Ū Ū r Ū = U ( ) h t + uu + u + v + w u x + U ( ) h t + uu + u + v + w u y + U ( ) h t + uu + u + v + w u z (4.7) and, finally
HAPTER 4. THIN AIRFOIL THEORY 4-5 ( ) h t Ū Ū Ū Ū r Ū Ū r = U ( ) h t + uu u x + U ( ) h t + uu v y + (4.8) U ( ) h t + uu w z u x U + uu x U + vv x U + ww x U uu x U vu y U wu z U. Neglect terms in (4.7) and (4.8) that are of third order in the disturbance velocities. Now ( ) h t Ū Ū Ū Ū r Ū Ū r = U ( ) h t + uu u x + U ( ) h t + uu v y + (4.9) U ( ) h t + uu w z u x U + uu x U + vv x U + ww x U uu x U vu y U wu z U or
HAPTER 4. THIN AIRFOIL THEORY 4-6 ( ) h t Ū Ū Ū Ū r Ū Ū r = ( ) (h uu )(u x + v y + w z ) (4.) u x U + uu x U + vv x U + ww x U uu x U vu y U wu z U. Recall that ( ) h = a. Equation (4.) can be rearranged to read ( ) h t Ū Ū Ū Ū r Ū Ū r a (u x + v y + w z ) u x U = ( + ) uu x U (vv x U + ww x U ) (4.) ( ) (uv y U + uw z U ) vu y U wu z U. Divide through by a. h a t Ū Ū r Ū Ū Ū Ū r a = M ( + ) M u x + v y + w z uu x a (4.) M a (( ) (uv y + uw z )+vu y + wu z + vv x + ww x ) Equation (4.) contains both linear and quadratic terms in the velocity disturbances and one might expect to be able to neglect the quadratic terms. But note that the first term becomes very small near M =. Thus in order to maintain the small disturbance approximation at transonic Mach numbers the uu x term must be retained. The remaining quadratic terms are small at all Mach numbers and can be dropped. Finally the small disturbance equation is M u x + v y + w z ( + ) M a uu x =. (4.3)
HAPTER 4. THIN AIRFOIL THEORY 4-7 The velocity potential is written in terms of a free-stream potential and a disturbance potential. =U x + (x, y, z) (4.4) The small disturbance equation in terms of the disturbance potential is M xx + yy + zz = ( + ) M a x xx. (4.5) Equation (4.5) is valid over the whole range of subsonic, transonic and supersonic Mach numbers. 4..3 Linearized potential flow If we restrict our attention to subsonic and supersonic flow, staying away from Mach numbers close to one, the nonlinear term on the right side of (4.5) can be dropped and the small disturbance potential equation reduces to the linear wave equation xx ( yy + zz ) = (4.6) where = q M. In two dimensions xx yy =. (4.7) The general solution of (4.7) can be expressed as a sum of two arbitrary functions (x, y) =F (x y)+g (x + y). (4.8) Note that if M < the -D linearized potential equation (4.7) is an elliptic equation that can be rescaled to form Laplace s equation and (4.8) expresses the solution in terms of conjugate complex variables. In this case the subsonic flow can be analyzed using the methods of complex analysis. Presently we will restrict our attention to the supersonic case. The subsonic case is treated later in the chapter. If M > then (4.7) is the -D wave equation and has solutions of hyperbolic type. Supersonic flow is analyzed using the fact that the properties of the flow are constant along the characteristic lines x ± y = constant. Figure 4. illustrates supersonic flow past a
HAPTER 4. THIN AIRFOIL THEORY 4-8 thin airfoil with several characteristics shown. Notice that in the linear approximation the characteristics are all parallel to one another and lie at the Mach angle µ of the free stream. Information about the flow is carried in the value of the potential assigned to a given characteristic and in the spacing between characteristics for a given flow change. Right-leaning characteristics carry the information about the flow on the upper surface of the wing and left-leaning characteristics carry information about the flow on the lower surface. Figure 4.: Right and left leaning characteristics on a thin -D airfoil. All properties of the flow, velocity, pressure, temperature, etc. are constant along the characteristics. Since disturbances only propagate along downstream running characteristics we can write the velocity potential for the upper and lower surfaces as (x, y) =F (x y) y > (x, y) =G (x + y) y <. (4.9) Let y = f (x) define the coordinates of the upper surface of the wing and y = g (x) define the lower surface. The full nonlinear boundary condition on the upper surface is v = df U y=f dx. (4.3) In the spirit of the thin airfoil approximation this boundary condition can be approximated by the linearized form v = df U y= dx (4.3) which we can write as
HAPTER 4. THIN AIRFOIL THEORY 4-9 @ (x, y) df = U @y y= dx (4.3) or F (x) = U df dx. (4.33) On the lower surface the boundary condition is G (x) = U dg dx. (4.34) In the thin airfoil approximation the airfoil itself is, in e ect, collapsed to a line along the x-axis, the velocity potential is extended to the line y = and the surface boundary condition is applied at y = instead of at the physical airfoil surface. The entire e ect of the airfoil on the flow is accounted for by the vertical velocity perturbation generated by the local slope of the wing. The linearized boundary condition is valid on -D thin wings and on 3-D wings of that are of thin planar form. Recall that (4.6) is only valid for subsonic and supersonic flow and not for transonic flow where M <<. 4..4 The pressure coe cient Let s work out the linearized pressure coe cient. The pressure coe cient is P = P P U = P. (4.35) M P The stagnation enthalpy is constant throughout the flow. Thus T =+ U T p T U + v + w. (4.36) Similarly the entropy is constant and so the pressure and temperature are related by P = + U P p T U + v + w. (4.37)
HAPTER 4. THIN AIRFOIL THEORY 4- The pressure coe cient is P = M + U p T U + v + w!. (4.38) The velocity term in (4.37) is small. U U + v + w = uu + u + v + w (4.39) Now use the binomial expansion ( ") n = n" + n (n ) " / to expand the term in parentheses in (4.39). Note that the expansion has to be carried out to second order. The pressure coe cient is approximately P = u u + M U U + v + w. (4.4) U Equation (4.4) is a valid approximation for small perturbations in subsonic or supersonic flow. For -D flows over planar bodies it is su we use the expression cient to retain only the first term in (4.4) and P = u U. (4.4) For 3-D flows over slender approximately axisymmetric, bodies we must retain the last term so P = u + v + w U.. (4.4) U As was discussed above, if the airfoil is a -D shape defined by the function y = f (x) the boundary condition at the surface is df dx = v = tan () (4.43) U + u where is the angle between the airfoil surface and the horizontal. For a thin airfoil this is accurately approximated by
HAPTER 4. THIN AIRFOIL THEORY 4- df dx = y U =. (4.44) For a thin airfoil in supersonic flow the wall pressure coe cient is where du /U = Pwall = M / df dx q / M d, which is valid for M (4.45) has been used. In the thin airfoil approximation in supersonic flow the local pressure coe cient is determined by the local slope of the wing. Figure 4.3 schematically shows the wall pressure coe cient on a thin, symmetric biconvex wing. Figure 4.3: Pressure distribution on a thin symmetric airfoil. This airfoil will have shock waves at the leading and trailing edges and at first sight this would seem to violate the isentropic assumption. But for small disturbances the shocks are weak and the entropy changes are negligible. 4..5 Drag coe cient of a thin symmetric airfoil A thin, -D, symmetric airfoil is situated in a supersonic stream at Mach number M and zero angle of attack. The y-coordinate of the upper surface of the airfoil is given by the function y (x) =ASin x (4.46) where is the airfoil chord. The airfoil thickness to chord ratio is small, A/ <<. Determine the drag coe cient of the airfoil. Solution
HAPTER 4. THIN AIRFOIL THEORY 4- The drag integral is Z D = (P P )Sin( ) dx. (4.47) where the factor of accounts for the drag of both the upper and lower surfaces and is the local angle formed by the upper surface tangent to the airfoil and the x-axis. Since the airfoil is thin the angle is small and we can write the drag coe cient as D D = U = Z P P U! x ( ) d. (4.48) The local tangent is determined by the local slope of the airfoil therefore Tan( ) = dy/dx and for small angles = dy/dx. Now the drag coe cient is D D = U = Z! P P dy x d. (4.49) U dx The pressure coe cient on the airfoil is given by thin airfoil theory (4.45) as P = P P U = p M dy dx (4.5) and so the drag coe cient becomes D = 4 p M The drag coe Z dy x d = dx 4A p M Z os x d x = cient is proportional to the square of the wing thickness-to-chord ratio. A p M. (4.5) 4..6 Thin airfoil with lift and camber at a small angle of attack A thin, cambered, -D, airfoil is situated in a supersonic stream at Mach number M and a small angle of attack as shown in Figure 4.4. The y-coordinate of the upper surface of the airfoil is given by the function
HAPTER 4. THIN AIRFOIL THEORY 4-3 Figure 4.4: Trajectory of a piston and an expansion wave in the x-t plane. x f = A x + B x x (4.5) and the y-coordinate of the lower surface is x x g = A + B x x (4.53) where A/ << and / <<. The airfoil surface is defined by a dimensionless thickness function x () = () = (4.54) a dimensionless camber function x () = () = (4.55) and the angle of attack Tan( ) =. (4.56) Determine the lift and drag coe cients of the airfoil. Let = x. (4.57) Solution
HAPTER 4. THIN AIRFOIL THEORY 4-4 The lift integral is L = Z (P lower P )os( lower ) dx Z (P upper P )os( upper ) dx. (4.58) where is the local angle formed by the tangent to the surface of the airfoil and the x-axis. Since the angle is everywhere small we can use os( ) =. The lift coe cient is D L = U = Z Plower d Z Pupper d. (4.59) The pressure coe cient on the upper surface is given by thin airfoil theory as Pupper = p M df = d p M A d d + B d d. (4.6) On the lower surface the pressure coe cient is Plower = p M dg = d p M A d d + B d d. (4.6) Substitute into the lift integral Z Z L = p df M d d + dg d d = p M Z Z df + In the thin airfoil approximation, the lift is independent of the airfoil thickness. L = 4 p M dg. (4.6) (4.63) The drag integral is D = Z Z (P upper P )Sin( upper ) dx + (P lower P )Sin( lower ) dx. (4.64)
HAPTER 4. THIN AIRFOIL THEORY 4-5 Since the airfoil is thin the angle is small, and we can write the drag coe cient as D D = = U Z P upper P U! Z ( upper ) d + P lower P U! ( lower ) d. (4.65) The local tangent is determined by the local slope of the airfoil therefore Tan( ) = dy/dx and for small angles = dy/dx. Now the drag coe cient is D D = = U Z dyupper Pupper d + d Z Plower dylower d. (4.66) d The drag coe cient becomes D = p M Z A d d + B d d Note that most of the cross terms cancel. The drag coe Z d + A d d + B d d d! cient breaks into several terms.. (4.67) D 4 = p M Z A d d d + The third term in (4.68) is B Z d d d B Z d Z! d + d d (4.68) B Finally Z d d = d Z B d = B ( () ()) =. (4.69) D = 4 p M A Z d d + d B Z d d d + Z d!. (4.7)
HAPTER 4. THIN AIRFOIL THEORY 4-6 In the thin airfoil approximation, the drag is a sum of the drag due to thickness, drag due to camber, and drag due to lift. 4. Similarity rules for high speed flight The Figure below shows the flow past a thin symmetric airfoil at zero angle of attack. The fluid is assumed to be inviscid and the flow Mach number, U /a where a = P / is the speed of sound, is assumed to be much less than one. The airfoil chord is c and the maximum thickness is t. The subscript one is applied in anticipation of the fact that we will shortly scale the airfoil to a new shape with subscript and the same chord. Figure 4.5: Pressure variation over a thin symmetric airfoil in low speed flow. The surface pressure distribution is shown below the wing, expressed in terms of the pressure coe cient. P = P s P U (4.7) The pressure and flow speed throughout the flow satisfy the Bernoulli relation. Near the airfoil surface P + U = P s + U s. (4.7) The pressure is high at the leading edge where the flow stagnates, then as the flow accelerates about the body, the pressure falls rapidly at first, then more slowly reaching a minimum at the point of maximum thickness. From there the surface velocity decreases
HAPTER 4. THIN AIRFOIL THEORY 4-7 and the pressure increases continuously to the trailing edge. In the absence of viscosity, the flow is irrotational. r u = (4.73) This permits the velocity to be described by a potential function. u = r (4.74) When this is combined with the condition of incompressibility, r ū =, the result is Laplace s equation. Let the shape of the airfoil surface in (x,y ) be given by @ @x + @ @y = (4.75) y c = g x. (4.76) c where = t /c is the thickness to chord ratio of the airfoil. The boundary conditions that the velocity potential must satisfy are @ dy = U @y y =c g( x c ) dx @! U. @x body = U dg (x /c) d (x /c) (4.77) Any number of methods of solving for the velocity potential are available including the use of complex variables. In the following we are going to restrict the airfoil to be thin, <<. In this context we will take the velocity potential to be a perturbation potential so that u = U + u v = v (4.78) or
HAPTER 4. THIN AIRFOIL THEORY 4-8 u = U + @ @x v = @ @y. (4.79) The boundary conditions on the perturbation potential in the thin airfoil approximation are @ @y @ @x The surface pressure coe y = dy dg (x /c) = U = U dx body d (x /c)!. cient in the thin airfoil approximation is (4.8) P = U @. (4.8) @x y = Note that the boundary condition on the vertical velocity is now applied on the line y =. In e ect the airfoil has been replaced with a line of volume sources whose strength is proportional to the local slope of the actual airfoil. This sort of approximation is really unnecessary in the low Mach number limit but it is essential when the Mach number is increased and compressibility e ects come in to play. Equally, it is essential in this example where we will map a compressible flow to the incompressible case. 4.. Subsonic flow M < Now imagine a second flow at a free-stream velocity, U in a new space (x,y )overanew airfoil of the same shape (defined by the function g (x/c) but with a new thickness ratio = t /c <<. Part of what we need to do is to determine how and are related to one another. The boundary conditions that the new perturbation velocity potential must satisfy are @ @y @ @x y = dy dg (x /c) = U = U dx body d (x /c)!. (4.8)
HAPTER 4. THIN AIRFOIL THEORY 4-9 In this second flow the Mach number has been increased to the point where compressibility e ects begin to occur: the density begins to vary significantly and the pressure distribution begins to deviate from the incompressible case. As long as the Mach number is not too large and shock waves do not form, the flow will be nearly isentropic. In this instance the -D steady compressible flow equations are @u @u u + v + @x @y @P = @x @v @v u + v + @x @y @P = @y u @ @x + v @ @y + @u @x + @v @y = P = P. (4.83) Let u = U + u v = v = + (4.84) P = P + P where the primed quantities are assumed to be small compared to the free stream conditions. When quadratic terms in the equations of motion are neglected the equations (4.83) reduce to U @u + @v =. (4.85) @x @y P Introduce the perturbation velocity potential. u = @ @x v = @ @y (4.86)
HAPTER 4. THIN AIRFOIL THEORY 4- The equation governing the disturbance flow becomes @ M @x + @ @y =. (4.87) Notice that (4.87) is valid for both sub and supersonic flow in the thin airfoil approximation. Since the flow is isentropic, the pressure and velocity disturbances are related to lowest order by P + U u = (4.88) and the surface pressure coe case. cient retains the same basic form as in the incompressible P = @ @x U y = (4.89) Since we are at a finite Mach number, this last relation is valid only within the thin airfoil, small disturbance approximation and therefore may be expected to be invalid near the leading edge of the airfoil where the velocity change is of the order of the free stream velocity. For example for the thin airfoil depicted in Figure 4.5 which is actually not all that thin, the pressure coe cient is within. < p <. except over a very narrow portion of the chord near the leading edge. Equation (4.87) can be transformed to Laplace s equation, (4.75) using the following change of variables x = x y = q M U y (4.9) = A U where, at the moment, A is an arbitrary constant. by (x,y )= A U U The velocity potentials are related (x,y ) (4.9)
HAPTER 4. THIN AIRFOIL THEORY 4- or (x,y )=A U U @x, q M y A (4.9) and the boundary conditions transform as @ @y @ @x @ @x y = = U @ A q A dg (x /c) M d (x /c)!!. (4.93) The transformation between flows one and two is completed by the correspondence = q A (4.94) M or t c = Finally the transformed pressure coe A t q. (4.95) M c cient is P = A P. (4.96) These results may be stated as follows. The solution for incompressible flow over a thin airfoil with shape g (x /c) and thickness ratio, t /c at velocity U is identical to the subsonic compressible flow at velocity, U and Mach number M over an airfoil with a similar shape but with the thickness ratio t c = q M A t. (4.97) c
HAPTER 4. THIN AIRFOIL THEORY 4- The pressure coe cient for the compressible case is derived by adjusting the incompressible coe cient using P = P /A. This result comprises several di erent similarity rules that can be found in the aeronautical literature depending on the choice of the free constant A. Perhaps the one of greatest interest is the Prandtl-Glauert rule that describes the variation of pressure coe cient with Mach number for a body of a given shape and thickness ratio. In this case we select A = q M (4.98) so that the two bodies being compared in (4.97) have the same shape and thickness ratio. The pressure coe cient for the compressible flow is P = q P. (4.99) M Several scaled profiles are shown in Figure 4.6. Keep in mind the lack of validity of (4.99) near the leading edge where the pressure coe cient is scaled to inaccurate values. Figure 4.6: Pressure coe cient over the airfoil in Figure 4.5 at several Mach numbers as estimated using the Prandtl-Glauert rule (4.99). 4.. Supersonic similarity M > All the theory developed in the previous section can be extended to the supersonic case by simply replacing M with M. In this instance the mapping is between the equation M @ @x @ @y = (4.)
HAPTER 4. THIN AIRFOIL THEORY 4-3 and the simple wave equation @ @x @ @y =. (4.) A generalized form of the pressure coe cient valid for subsonic and supersonic flow is P A = F where A is taken to be a function of M. @ q A (4.) A M 4..3 Transonic similarity M = When the Mach number is close to one, the simple linearization used to obtain (4.87) from (4.83) loses accuracy. In this case the equations (4.83) reduce to the nonlinear equation @u M + @v ( + ) M u @u =. (4.3) @x @y U @x In terms of the perturbation potential @ M @x + @ @y ( + ) M @ @ U @x @x =. (4.4) This equation is invariant under the change of variables x = x q M y = q y M U (4.5) = A U where
HAPTER 4. THIN AIRFOIL THEORY 4-4 + M M A = +. (4.6) M M Notice that, due to the nonlinearity of the transonic equation (4.4) the constant A is no longer arbitrary. The pressure coe cient becomes and the thickness ratios are related by + M M P = + P (4.7) M M t + c = M + M 3 M M t c. (4.8) In the transonic case, it is not possible to compare the same body at di erent Mach numbers or bodies with di erent thickness ratios at the same Mach number except by selecting gases with di erent. For a given gas it is only possible to map the pressure distribution for one airfoil to an airfoil with a di erent thickness ratio at a di erent Mach number. A generalized form of (4.) valid from subsonic to sonic to supersonic Mach numbers is P ( + ) M 3 3 = F @ M A. (4.9) ( + ) M 3 Prior to the advent of supercomputers capable of solving the equations of high speed flow, similarity methods and wind tunnel correlations were the only tools available to the aircraft designer and these methods played a key role in the early development of transonic and supersonic flight. 4.3 Problems Problem - A thin, -D, airfoil is situated in a supersonic stream at Mach number M and a small angle of attack as shown in Figure 4.7. The y-coordinate of the upper surface of the airfoil is given by the function f (x) =A x x x (4.)
HAPTER 4. THIN AIRFOIL THEORY 5 Figure 4.7: Bi-convex airfoil at angle of attack and the y-coordinate of the lower surface is g (x) = A x x x (4.) where A/ << and / <<. Determine the lift and drag coe cients of the airfoil.