Category Fundamentals of Aerodynamics

. Lifting Flows over Arbitrary Bodies: The Vortex Panel Numerical Method

The thin airfoil theory described in Sections 4.7 and 4.8 is just what it says—it ap­plies only to thin airfoils at small angles of attack. (Make certain that you understand exactly where in the development of thin airfoil theory these assumptions are made and the reasons for making them.) The advantage of thin airfoil theory is that closed – form expressions are obtained for the aerodynamic coefficients. Moreover, the results compare favorably with experimental data for airfoils of about 12 percent thickness or less. However, the airfoils on many low-speed airplanes are thicker than 12 percent. Moreover, we are frequently interested in high angles of attack, such as occur during takeoff and landing. Finally, we are sometimes concerned with the generation of aerodynamic lift on other body shapes, such as automobiles or submarines. Hence, thin airfoil theory is quite restrictive when we consider the whole spectrum of aero­dynamic applications. We need a method that allows us to calculate the aerodynamic characteristics of bodies of arbitrary shape, thickness, and orientation. Such a method is described in this section. Specifically, we treat the vortex panel method, which is a numerical technique that has come into widespread use since the early 1970s. In reference to our road map in Figure 4.2, we now move to the left-hand branch. Also, since this chapter deals with airfoils, we limit our attention to two-dimensional bodies.

The vortex panel method is directly analogous to the source panel method de­scribed in Section 3.17. However, because a source has zero circulation, source panels are useful only for nonlifting cases. In contrast, vortices have circulation, and hence vortex panels can be used for lifting cases. (Because of the similarities between source and vortex panel methods, return to Section 3.17 and review the basic philosophy of the source panel method before proceeding further.)

The philosophy of covering a body surface with a vortex sheet of such a strength to make the surface a streamline of the flow was discussed in Section 4.4 We then went on to simplify this idea by placing the vortex sheet on the camber line of the airfoil as shown in Figure 4.11, thus establishing the basis for thin airfoil theory. We

now return to the original idea of wrapping the vortex sheet over the complete surface of the body, as shown in Figure 4.10. We wish to find у (s) such that the body surface becomes a streamline of the flow. There exists no closed-form analytical solution for y(s); rather, the solution must be obtained numerically. This is the purpose of the vortex panel method.

Let us approximate the vortex sheet shown in Figure 4.10 by a series of straight panels, as shown earlier in Figure 3.40. (In Chapter 3, Figure 3.40 was used to discuss source panels; here, we use the same sketch for discussion of vortex panels.) Let the vortex strength у (s) per unit length be constant over a given panel, but allow it to vary from one panel to the next. That is, for the n panels shown in Figure 3.40, the vortex panel strengths per unit length are y, y2,…, yj,…, y„. These panel strengths are unknowns; the main thrust of the panel technique is to solve for yj, j = 1 to n, such that the body surface becomes a streamline of the flow and such that the Kutta condition is satisfied. As explained in Section 3.17, the midpoint of each panel is a control point at which the boundary condition is applied; that is, at each control point, the normal component of the flow velocity is zero.

image347

Let P be a point located at (x, y) in the flow, and let rpj be the distance from any point on the yth panel to P, as shown in Figure 3.40. The radius rpl makes the angle 0Pj with respect to the. r axis. The velocity potential induced at P due to the y’th panel, Дг/),, is, from Equation (4.3),

In Equation (4.72), yj is constant over the y’th panel, and the integral is taken over the jth panel only. The angle 6PJ is given by

image348[4.73]

In turn, the potential at P due to all the panels is Equation (4.72) summed over all the panels:

image349[4.74]

Since point P is just an arbitrary point in the flow, let us put P at the control point of the ith panel shown in Figure 3.40. The coordinates of this control point are (*,-, уi). Then Equations (4.73) and (4.74) become

Подпись: and Подпись: [4.75]
Подпись: 6ij = tan

-і Уі – Уі Xi-Xj

Equation (4.75) is physically the contribution of all the panels to the potential at the control point of the ith panel.

At the control points, the normal component of the velocity is zero; this velocity is the superposition of the uniform flow velocity and the velocity induced by all the vortex panels. The component of VA normal to the / th panel is given by Equation (3.148):

V0o. n = V0C cos p, [3.148]

The normal component of velocity induced at (a, , v, ) by the vortex panels is

v„ = [4.76]

an.

Подпись: V„ Подпись: [4.77]
image351

Combining Equations (4.75) and (4.76), we have

where the summation is over all the panels. The normal component of the flow velocity at the ith control point is the sum of that due to the freestream [Equation (3.148)] and that due to the vortex panels [Equation (4.77)]. The boundary condition states that this sum must be zero:

Подпись: V 4_ у — о y 00.П і v П — Подпись: Voo cos p,image352[4.78]

Substituting Equations (3.148) and (4.77) into (4.78), we obtain

[4.79]

Equation (4.79) is the crux of the vortex panel method. The values of the integrals in Equation (4.79) depend simply on the panel geometry; they are not properties of the flow. Let Jij be the value of this integral when the control point is on the ith panel. Then Equation (4.79) can be written as

П

Езо cos p, – ~ Ji. j = 0 [4.80]

i=і 171

Equation (4.80) is a linear algebraic equation with n unknowns, jq, Y2- ■ ■ ■■ Yn• It represents the flow boundary condition evaluated at the control point of the ith panel. If Equation (4.80) is applied to the control points of all the panels, we obtain a system of n linear equations with n unknowns.

To this point, we have been deliberately paralleling the discussion of the source panel method given in Section 3.17; however, the similarity stops here. For the source panel method, the n equations for the n unknown source strengths are routinely solved, giving the flow over a nonlifting body. In contrast, for the lifting case with vortex panels, in addition to the n equations given by Equation (4.80) applied at all the panels, we must also satisfy the Kutta condition. This can be done in several ways. For example, consider Figure 4.26, which illustrates a detail of the vortex panel distribution at the trailing edge. Note that the length of each panel can be different; their length and distribution over the body are up to your discretion. Let the two panels at the trailing edge (panels і and і — 1 in Figure 4.26) be very small. The

image353

Kutta condition is applied precisely at the trailing edge and is given by у (ТЕ) = 0. To approximate this numerically, if points і and г — 1 are close enough to the trailing edge, we can write

Yi = ~Yi- [4.81]

such that the strengths of the two vortex panels і and і — 1 exactly cancel at the point where they touch at the trailing edge. Thus, in order to impose the Kutta condition on the solution of the flow, Equation (4.81) (or an equivalent expression) must be included. Note that Equation (4.80) evaluated at all the panels and Equation (4.81) constitute an overdetermined system of n unknowns with n + 1 equations. Therefore, to obtain a determined system, Equation (4.80) is not evaluated at one of the control points on the body. That is, we choose to ignore one of the control points, and we evaluate Equation (4.80) at the other n — 1 control points. This, in combination with Equation (4.81), now gives a system of n linear algebraic equations with n unknowns, which can be solved by standard techniques.

At this stage, we have conceptually obtained the values of y, y2,…, yn which make the body surface a streamline of the flow and which also satisfy the Kutta condition. In turn, the flow velocity tangent to the surface can be obtained directly from y. To see this more clearly, consider the airfoil shown in Figure 4.27. We are concerned only with the flow outside the airfoil and on its surface. Therefore, let the velocity be zero at every point inside the body, as shown in Figure 4.27. In particular, the velocity just inside the vortex sheet on the surface is zero. This corresponds to U2 = 0 in Equation (4.8). Hence, the velocity just outside the vortex sheet is, from Equation (4.8),

у = U — U2 = U — 0 = Ml

In Equation (4.8), и denotes the velocity tangential to the vortex sheet. In terms of the picture shown in Figure 4.27, we obtain Va = ya at point а, V/, = yh at point b, etc. Therefore, the local velocities tangential to the airfoil surface are equal to the local values of у. In turn, the local pressure distribution can be obtained from Bernoulli’s equation.

The total circulation and the resulting lift are obtained as follows. Let sj be the length of the yth panel. Then the circulation due to the у th panel is y, s r In turn, the

image354

Figure 4.27 Airfoil as a solid body, with zero velocity inside the profile.

total circulation due to all the panels is

n

г = &л [4-821

./=1

Hence, the lift per unit span is obtained from

П

L’ = Poo Too y. isJ [4.83]

./=>

The presentation in this section is intended to give only the general flavor of the vortex panel method. There are many variations of the method in use today, and you are encouraged to read the modem literature, especially as it appears in the A/A4 Journal and the Journal of Aircraft since 1970. The vortex panel method as described in this section is termed a “first-order” method because it assumes a constant value of у over a given panel. Although the method may appear to be straightforward, its numerical implementation can sometimes be frustrating. For example, the results for a given body are sensitive to the number of panels used, their various sizes, and the way they are distributed over the body surface (i. e., it is usually advantageous to place a large number of small panels near the leading and trailing edges of an airfoil and a smaller number of larger panels in the middle). The need to ignore one of the control points in order to have a determined system in n equations for n unknowns also introduces some arbitrariness in the numerical solution. Which control point do you ignore? Different choices sometimes yield different numerical answers for the distribution of у over the surface. Moreover, the resulting numerical distributions for у are not always smooth, but rather, they have oscillations from one panel to the next as a result of numerical inaccuracies. The problems mentioned above are usually overcome in different ways by different groups who have developed relatively sophisticated panel programs for practical use. For example, what is more common today is to use a combination of both source and vortex panels (source panels to basically simulate the airfoil thickness and vortex panels to introduce circulation) in a panel solution. This combination helps to mitigate some of the practical numerical problems just discussed. Again, you are encouraged to consult the literature for more information.

Such accuracy problems have also encouraged the development of higher-order panel techniques. For example, a “second-order” panel method assumes a linear variation of у over a given panel, as sketched in Figure 4.28. Here, the value of у at the edges of each panel is matched to its neighbors, and the values y, y2, уз, etc. at the boundary points become the unknowns to be solved. The flow-tangency boundary condition is still applied at the control point of each panel, as before. Some results using a second-order vortex panel technique are given in Figure 4.29, which shows

image355

Figure 4.28 Linear distribution of у over each panel—a second-order panel method.

image356

Figure 4.29 Pressure coefficient distribution over an NACA 0012 airfoil; comparison

between second-order vortex panel method and NACA theoretical results from Reference 1 1. The numerical panel results were obtained by one of the author’s graduate students, Mr. Tae-Hwan Cho.

the distribution of pressure coefficients over the upper and lower surfaces of an NACA 0012 airfoil at a 9° angle of attack. The circles and squares are numerical results from a second-order vortex panel technique developed at the University of Maryland, and the solid lines are from NACA results given in Reference 11. Excellent agreement is obtained.

Again, you are encouraged to consult the literature before embarking on any serious panel solutions of your own. For example, Reference 14 is a classic paper on panel methods, and Reference 15 highlights many of the basic concepts of panel methods along with actual computer program statement listings for simple applica­tions. Reference 66 is a modern compilation of papers, several of which deal with current panel techniques. Finally, Katz and Plotkin (Reference 67) give perhaps the most thorough discussion of panel techniques and their foundations to date.

Normal Shock Waves. and Related Topics

Shock wave: A large-amplitude compression wave, such as that produced by an explosion, caused by supersonic motion of a body in a medium.

From the American Heritage Dictionary of the English Language, 1969

8.1 Introduction

The purpose of this chapter and Chapter 9 is to develop shock-wave theory, thus giving us the means to calculate the changes in the flow properties across a wave. These changes were discussed qualitatively in Section 7.6; make certain that you are familiar with these changes before continuing.

The focus of this chapter is on normal shock waves, as sketched in Figure lAb. At first thought, a shock wave that is normal to the upstream flow may seem to be a very special case—and therefore a case of little practical interest—but nothing could be further from the truth. Normal shocks occur frequently in nature. Two such ex­amples are sketched in Figure 8.1; there are many more. The supersonic flow over a blunt body is shown at the left of Figure 8.1. Here, a strong bow shock wave exists in front of the body. (We study such bow shocks in Chapter 9.) Although this wave is curved, the region of the shock closest to the nose is essentially nor­mal to the flow. Moreover, the streamline that passes through this normal portion of the bow shock later impinges on the nose of the body and controls the values

image522

Figure 8.1 Two examples where normal shock waves are of interest.

 

Normal shock inside the nozzle

image523

Overexpanded flow through a nozzle

 

of stagnation (total) pressure and temperature at the nose. Since the nose region of high-speed blunt bodies is of practical interest in the calculation of drag and aerodynamic heating, the properties of the flow behind the normal portion of the shock wave take on some importance. In another example, shown at the right of Figure 8.1, supersonic flow is established inside a nozzle (which can be a super­sonic wind tunnel, a rocket engine, etc.) where the back pressure is high enough to cause a normal shock wave to stand inside the nozzle. (We discuss such “over­expanded” nozzle flows in Chapter 10.) The conditions under which this shock wave will occur and the determination of flow properties at the nozzle exit down­stream of the normal shock are both important questions to be answered. In sum­mary, for these and many other applications, the study of normal shock waves is important.

Finally, we will find that many of the normal shock relations derived in this chapter carry over directly to the analysis of oblique shock waves, as discussed in Chapter 9. So once again, time spent on normal shock waves is time well spent.

The road map for this chapter is given in Figure 8.2. As you can see, our objectives are fairly short and straightforward. We start with a derivation of the basic continuity, momentum, and energy equations for normal shock waves, and then we employ these basic relations to obtain detailed equations for the calculation of flow properties across the shock wave. In addition, we emphasize the physical trends indicated by the equations. On the way toward this objective, we take three side streets having to do with (1) the speed of sound, (2) special forms of the energy equation, and (3) a further discussion of the criteria used to judge when a flow must be treated as compressible. Finally, we apply the results of this chapter to the measurement of airspeed in a compressible flow using a Pitot tube. Keep the road map in Figure 8.2 in mind as you progress through the chapter.

image524

Figure 8.2 Road map for Chapter 8.

Specification of the Flow Field

In Section 2.2.3 we defined both scalar and vector fields. We now apply this concept of a field more directly to an aerodynamic flow. One of the most straightforward ways of describing the details of an aerodynamic flow is simply to visualize the flow in three-dimensional space, and to write the variation of the aerodynamic properties as a function of space and time. For example, in cartesian coordinates the equations

P = p(x, y, z, t) [2.33a]

p = p(x, y,z, t) [2.33b]

T = T(x, y, z, t) [2.33c]

and V = ui + vj + iuk [2.34a]

where и = u(x, y,z, t) [2.34b]

v = v(x, y,z, t) [2.34c]

w = w(x, y,z, t) [2.34<f]

represent the flow field. Equations (2.33a-c) give the variation of the scalar flow field variables pressure, density, and temperature, respectively. (In equilibrium thermody­namics, the specification of two state variables, such as p and p, uniquely defines the values of all other state variables, such as T. In this case, one of Equations (2.33) can be considered redundant.) Equations (ІЗАа-d) give the variation of the vector flow field variable velocity V, where the scalar components of V in the x, y. and г directions are u, v, and w, respectively.

Figure 2.14 illustrates a given fluid element moving in a flow field specified by Equations (2.33) and (2.34). At the time t, the fluid element is at point 1, located at (xi, yi, zi) as shown in Figure 2.14.

At this instant, its velocity is V i and its pressure is given by

p = p(xi, y,z,t)

and similarly for its other flow variables.

By definition, an unsteady flow is one where the flow field variables at any given point are changing with time. For example, if you lock your eyes on point 1 in

Figure 2.14, and keep them fixed on point 1, if the flow is unsteady you will observe p, p, etc. fluctuating with time. Equations (2.33) and (2.34) describe an unsteady flow field because time t is included as one of the independent variables. In contrast, a steady flow is one where the flow field variables at any given point are invariant with time, that is, if you lock your eyes on point 1 you will continuously observe the same constant values for p, p, V etc. for all time. A steady flow field is specified by the relations

p = p(x, y,z) p = p(x, y,z) etc.

The concept of the flow field, and a specified fluid element moving through it as illustrated in Figure 2.14, will be revisited in Section 2.9 where we define and discuss the concept of the substantial derivative.

Подпись: Example 2.1

Подпись: Figure 2.14 A fluid element passing through point 1 in a flow field.

The subsonic compressible flow over a cosine-shaped (wavy) wall is illustrated in Figure 2.15. The wavelength and amplitude of the wall are l and h, respectively, as shown in Figure 2.15. The streamlines exhibit the same qualitative shape as the wall, but with diminishing amplitude as distance above the wall increases. Finally, as у —у oo, the streamline becomes straight. Along this straight streamline, the freestream velocity and Mach number are and M^,

Подпись: Figure 2.1 5 Subsonic compressible flow over a wavy wall; the streamline pattern.

Streamline at <»

Подпись: and Specification of the Flow Field Specification of the Flow Field

respectively. The velocity field in cartesian coordinates is given by

where p = У1 —

Consider the particular flow that exists for the case where і = 1.0 m, h = 0.01 m, = 240 m/s, and = 0.7. Also, consider a fluid element of fixed mass moving along a streamline in the flow field. The fluid element passes through the point (x/i, y/t) = (J, 1). At this point, calculate the time rate of change of the volume of the fluid element, per unit volume.

Solution

From Section 2.3.4, we know that the time rate of change of the volume of a moving fluid element of fixed mass, per unit volume, is given by the divergence of the velocity V • V. In

12,411

Equation (2.41) gives the time rate of change of the volume of the fluid element, per unit volume, as it passes through the point (x/i, у ft) = (j, 1). Note that it is a finite (nonzero) value; the volume of the fluid element is changing as it moves along the streamline. This is consistent with the definition of a compressible flow, where the density is a variable and hence the volume of a fixed mass must also be variable. Note from Equation (2.40) that V • V = 0 only along vertical lines denoted by x/l = 0, j, 1, 1 ^,…, where the sin(2jrx/£) goes to zero. This is a peculiarity associated with the cyclical nature of the flow field over the cosine-shaped wall. For the particular flow considered here, where і = 1.0 m, h = 0.01 m, Vx = 240 m/s, and Мы = 0.7, where

P = 7l – Ml = Vl – (0.7)2 = 0.714

Equation (2.41) yields

Подпись: -0.7327 s'1V • V = ^0.714 – (240)(0.01) e-2*<0’714) =

The physical significance of this result is that, as the fluid element is passing through the point (f, 1) in the flow, it is experiencing a 73 percent rate of decrease of volume per second (the negative quantity denotes a decrease in volume). That is, the density of the fluid element is increasing. Hence, the point (і, 1) is in a compression region of the flow, where the fluid element will experience a decrease in density. Expansion regions are defined by values of x 11 which yield negative values of the sine function in Equation (2.40), which in turn yields a positive value for V • V, which gives an increase in volume of the fluid element, hence a de­crease in density. Clearly, as the fluid element continues its path through this flow field, it experiences cyclical increases and decreases in density, as well as the other flow field properties.

Combination of a Uniform Flow with a Source and Sink

Consider a polar coordinate system with a source of strength Л located at the origin. Superimpose on this flow a uniform stream with velocity Vx moving from left to right, as sketched in Figure 3.22. The stream function for the resulting flow is the sum of Equations (3.57) and (3.72):

i/r = Loor sin в H—- 0 [3.74]

2тт

Since both Equations (3.57) and (3.72) are solutions of Laplace’s equation, we know that Equation (3.74) also satisfies Laplace’s equation; that is, Equation (3.74) de­scribes a viable irrotational, incompressible flow. The streamlines of the combined flow are obtained from Equation (3.74) as

jf = Vxr sin0 H—– в = const [3.75]

2jt

The resulting streamline shapes from Equation (3.75) are sketched at the right of Figure 3.22. The source is located at point D. The velocity field is obtained by

image237

Ф = looT sin в + в

 

+

 

Uniform stream

 

Source

image236

 

ф = Voor sin в

 

Combination of a Uniform Flow with a Source and Sink

Figure 3.22 Superposition of a uniform flow and a source; flow over a semi-infinite body.

differentiating Equation (3.75):

Подпись: A 2nr Подпись: [3.76]1 дф

Vr =——– = Voo cos в +

Подпись: and Подпись: Ve = = -VooSin0 dr Подпись: [3.77]

r dO

Note from Section 3.10 that the radial velocity from a source is А/2л r, and from Section 3.9 the component of the freestream velocity in the radial direction is Voo cos в. Hence, Equation (3.76) is simply the direct sum of the two velocity fields—a result which is consistent with the linear nature of Laplace’s equation. Therefore, not only can we add the values of ф or r/r to obtain more complex solutions, we can add their derivatives, that is, the velocities, as well.

The stagnation points in the flow can be obtained by setting Equations (3.76) and

(3.77)

Подпись: and Combination of a Uniform Flow with a Source and Sink Combination of a Uniform Flow with a Source and Sink

equal to zero:

Solving for r and в, we find that one stagnation point exists, located at (г, в) = (A/2jrV00, tt), which is labeled as point В in Figure 3.22. That is, the stagnation point is a distance (A/2jrV0O) directly upstream of the source. From this result, the distance DB clearly grows smaller if is increased and larger if A is increased— trends that also make sense based on intuition. For example, looking at Figure 3.22, you would expect that as the source strength is increased, keeping the same, the stagnation point В will be blown further upstream. Conversely, if V-y_ is increased, keeping the source strength the same, the stagnation point will be blown further downstream.

If the coordinates of the stagnation point at В are substituted into Equation (3.75), we obtain

Л

sin ті H—— tt = const

2 tt

 

image238

image239

Hence, the streamline that goes through the stagnation point is described by ijr = A/2. This streamline is shown as curve ABC in Figure 3.22.

Examining Figure 3.22, we now come to an important conclusion. Since we are dealing with inviscid flow, where the velocity at the surface of a solid body is tangent to the body, then any streamline of the combined flow at the right of Figure 3.22 could be replaced by a solid surface of the same shape. In particular, consider the streamline ABC. Because it contains the stagnation point at B, the streamline ABC is a dividing streamline; that is, it separates the fluid coming from the freestream and the fluid emanating from the source. All the fluid outside ABC is from the freestream, and all the fluid inside ABC is from the source. Therefore, as far as the freestream is concerned, the entire region inside ABC could be replaced with a solid body of the same shape, and the external flow, that is, the flow from the freestream, would not feel the difference. The streamline xfr — A/2 extends downstream to infinity, forming a semi-infinite body. Therefore, we are led to the following important interpretation. If we want to construct the flow over a solid semi-infinite body described by the curve ABC as shown in Figure 3.22, then all we need to do is take a uniform stream with velocity Voc and add to it a source of strength Л at point D. The resulting superposition will then represent the flow over the prescribed solid semi-infinite body of shape ABC. This illustrates the practicality of adding elementary flows to obtain a more complex flow over a body of interest.

The superposition illustrated in Figure 3.22 results in the flow over the semi­infinite body ABC. This is a half-body that stretches to infinity in the downstream direction (i. e., the body is not closed). However, if we take a sink of equal strength as the source and add it to the flow downstream of point D, then the resulting body shape will be closed. Let us examine this flow in more detail.

Consider a polar coordinate system with a source and sink placed a distance b to the left and right of the origin, respectively, as sketched in Figure 3.23. The strengths of the source and sink are A and —A, respectively (equal and opposite). In addition, superimpose a uniform stream with velocity Ex,, as shown in Figure 3.23. The stream function for the combined flow at any point P with coordinates (г. 0) is obtained from Equations (3.57) and (3.72):

A A

if = ExT Sin0 + —01 – —02 2л 2л

or if = Vxr sin 0 + — (01 – 02) [3.80]

The velocity field is obtained by differentiating Equation (3.80) according to Equa­tions (2.151a and b). Note from the geometry of Figure 3.23 that 0] and 02 in Equation (3.80) are functions of r, 0, and b. In turn, by setting V = 0, two stagnation points are found, namely, points A and В in Figure 3.23. These stagnation points are located

image240

Figure 3.23 Superposition of a uniform flow and a

source-sink pair; flow over a Rankine oval.

such that (see Problem 3.13)

J.b

О A — OB = b2 [3.81]

V ^Voo

The equation of the streamlines is given by Equation (3.80) as

Л r,

ф = V^r sin 0 H——- (0i — 02) = const [3.82]

2ix

The equation of the specific streamline going through the stagnation points is obtained from Equation (3.82) by noting that 0 = 0; = 02 = 7Г at point A and 0 = 0i = 02 = 0 at point B. Hence, for the stagnation streamline, Equation (3.82) yields a value of zero for the constant. Thus, the stagnation streamline is given by = 0, that is,

Л r,

Voor sin0 H——- (0i — 02) = 0 [3.83]

2tt

the equation of an oval, as sketched in Figure 3.23. Equation (3.83) is also the dividing streamline; all the flow from the source is consumed by the sink and is contained entirely inside the oval, whereas the flow outside the oval has originated with the uniform stream only. Therefore, in Figure 3.23, the region inside the oval can be replaced by a solid body with the shape given by т/r = 0, and the region outside the oval can be interpreted as the inviscid, potential (irrotational), incompressible flow over the solid body. This problem was first solved in the nineteenth century by the famous Scottish engineer W. J. M. Rankine; hence, the shape given by Equation (3.83) and sketched in Figure 3.23 is called a Rankine oval.

A Numerical Nonlinear Lifting-Line Method

The classical Prandtl lifting-line theory described in Section 5.3 assumes a linear vari­ation of с/ versus aeff. This is clearly seen in Equation (5.19). However, as the angle of attack approaches and exceeds the stall angle, the lift curve becomes nonlinear, as shown in Figure 4.4. This high-angle-of-attack regime is of interest to modern aero – dynamicists. For example, when an airplane is in a spin, the angle of attack can range from 40 to 90°; an understanding of high-angle-of-attack aerodynamics is essential to the prevention of such spins. In addition, modern fighter airplanes achieve optimum maneuverability by pulling high angles of attack at subsonic speeds. Therefore, there are practical reasons for extending Prandtl’s classical theory to account for a nonlinear lift curve. One simple extension is described in this section.

The classical theory developed in Section 5.4 is essentially closed form; that is, the results are analytical equations as opposed to a purely numerical solution. Of course, in the end, the Fourier coefficients A„ for a given wing must come from a solution of a system of simultaneous linear algebraic equations. Until the advent of the modern digital computer, these coefficients were calculated by hand. Today, they are readily solved on a computer using standard matrix methods. However, the elements of the lifting-line theory lend themselves to a straightforward purely numerical solution which allows the treatment of nonlinear effects. Moreover, this
numerical solution emphasizes the fundamental aspects of lifting-line theory. For these reasons, such a numerical solution is outlined in this section.

Consider the most general case of a finite wing of given planform and geometric twist, with different airfoil sections at different spanwise stations. Assume that we have experimental data for the lift curves of the airfoil sections, including the nonlinear regime (i. e., assume we have the conditions of Figure 4.4 for all the given airfoil sections). A numerical iterative solution for the finite-wing properties can be obtained as follows:

1. Divide the wing into a number of spanwise stations, as shown in Figure 5.26. Here к + 1 stations are shown, with n designating any specific station.

2. For the given wing at a given a, assume the lift distribution along the span; that is, assume values for Г at all the stations Г і, Г2,…, Г„,…, Г*+1. An elliptical lift distribution is satisfactory for such an assumed distribution.

3.

Подпись: Оіі(Уп) Подпись: 1 ґ2 (dT/dy)dy 4л- Foo J_b/2 Уп-У Подпись: [5.75]

With this assumed variation of Г, calculate the induced angle of attack a, from Equation (5.18) at each of the stations:

The integral is evaluated numerically. If Simpson’s rule is used, Equation (5.75) becomes

Подпись: ОіііУп)1 Ay ул (dr/dy)j_i | ^ (dr/dy)j | {dr/dy)j+l [5 y6] 4nVoc 3 j^6(yn ~ У]-) Уп-Уі Уп-yj+i ‘

where Ay is the distance between stations. In Equation (5.76), when yn = yj-1, у,, or yj+1, a singularity occurs (a denominator goes to zero). When this singularity occurs, it can be avoided by replacing the given term by its average value based on the two adjacent sections.

4. Using a, from step 3, obtain the effective angle of attack o^ff at each station from

aeff(y«) = a – a,(y„)

5.

image439

With the distribution of o^ff calculated from step 4, obtain the section lift coeffi­cient (c/)„ at each station. These values are read from the known lift curve for the airfoil.

Figure 5.26 Stations along the span for a numerical solution.

6. From (c;)„ obtained in step 5, a new circulation distribution is calculated from the Kutta-Joukowski theorem and the definition of lift coefficient:

^ (У«) = Рэо^эсГСУл) = 2 Poo (u)«

Hence, nj„) = iy00c„(c,)„

where c„ is the local section chord. Keep in mind that in all the above steps, n ranges from 1 to к + 1.

7. The new distribution of Г obtained in step 6 is compared with the values that were initially fed into step 3. If the results from step 6 do not agree with the input to step 3, then a new input is generated. If the previous input to step 3 is designated as Told and the result of step 6 is designated as rnew, then the new input to step 3 is determined from

Подпись: input= Told + £(Tnew — Told)

where D is a damping factor for the iterations. Experience has found that the iterative procedure requires heavy damping, with typical values of D on the order of 0.05.

8. Steps 3 to 7 are repeated a sufficient number of cycles until Tnew and Told agree at each spanwise station to within acceptable accuracy. If this accuracy is stipulated to be within 0.01 percent for a stretch of five previous iterations, then a minimum of 50 and sometimes as many as 150 iterations may be required for convergence.

9. From the converged Г (у), the lift and induced drag coefficients are obtained from Equations (5.26) and (5.30), respectively. The integrations in these equations can again be carried out by Simpson’s rule.

The procedure outlined above generally works smoothly and quickly on a high­speed digital computer. Typical results are shown in Figure 5.27, which shows the circulation distributions for rectangular wings with three different aspect ratios. The solid lines are from the classical calculations of Prandtl (Section 5.3), and the symbols are from the numerical method described above. Excellent agreement is obtained, thus verifying the integrity and accuracy of the numerical method. Also, Figure 5.27 should be studied as an example of typical circulation distributions over general finite wings, with Г reasonably high over the center section of the wing but rapidly dropping to zero at the tips.

An example of the use of the numerical method for the nonlinear regime is shown in Figure 5.28. Here, Cl versus a is given for a rectangular wing up to an angle of attack of 50°—well beyond stall. The numerical results are compared with existing experimental data obtained at the University of Maryland (Reference 19). The numerical lifting-line solution at high angle of attack agrees with the experiment to within 20 percent, and much closer for many cases. Therefore, such solutions given reasonable preliminary engineering results for the high-angle-of-attack poststall region. However, it is wise not to stretch the applicability of lifting-line theory too far. At high angles of attack, the flow is highly three-dimensional. This is clearly seen in the surface oil pattern on a rectangular wing at high angle of attack shown in Figure

image440

Figure 5.27 Lift distribution for a rectangular wing;

comparison between Prandtl’s classical theory and the numerical lifting-line method of Reference 20.

image441

Figure 5.28 Lift coefficient versus angle of attack; comparison between experimental and numerical results.

5.29. At high a, there is a strong spanwise flow, in combination with mushroom­shaped flow separation regions. Clearly, the basic assumptions of lifting-line theory, classical or numerical, cannot properly account for such three-dimensional flows.

For more details and results on the numerical lifting-line method, please see Reference 20.

image442

Figure 5.29 Surface oil flow pattern on a stalled, finite rectangular wing with a Clark Y-14 airfoil section. AR = 3.5, a = 22.8°, Re = 245,000 (based on chord length). This pattern was established by coating the wing surface with pigmented mineral oil and inserting the model in a low-speed subsonic wind tunnel. In the photograph shown, flow is from top to bottom. Note the highly three-dimensional flow pattern. (Courte$y of Allen E. Winkelmann, University of Maryland.}

Center of Pressure

From Equations (1.7) and (1.8), we see that the normal and axial forces on the body are due to the distributed loads imposed by the pressure and shear stress distributions. Moreover, these distributed loads generate a moment about the leading edge, as given by Equation (1.11). Question: If the aerodynamic force on a body is specified in terms of a resultant single force R, or its components such as N and A, where on the body should this resultant be placed? The answer is that the resultant force should be located on the body such that it produces the same effect as the distributed loads. For example, the distributed load on a two-dimensional body such as an airfoil produces a moment about the leading edge given by Equation (1.11); therefore, N’ and A! must be placed on the airfoil at such a location to generate the same moment about the leading edge. If A! is placed on the chord line as shown in Figure 1.18, then N’ must be located a distance xcp downstream of the leading edge such that

MLE — (xcp)N

image32[1.20]

In Figure 1.18, the direction of the curled arrow illustrating M[E is drawn in the positive (pitch-up) sense. (From Section 1.5, recall the standard convention that aerodynamic moments are positive if they tend to increase the angle of attack.) Ex­amining Figure 1.18, we see that a positive N’ creates a negative (pitch-down) moment about the leading edge. This is consistent with the negative sign in Equation (1.20). Therefore, in Figure 1.18, the actual moment about the leading edge is negative, and hence is in a direction opposite to the curled arrow shown.

N’

image33

Figure 1.18 Center of pressure for an ai rfoil.

 

image34

Resultant force at leading edge

 

Center of Pressure

image35

Подпись: Figure 1.19Equivalent ways of specifying the force-and-moment system on an airfoil.

In Figure 1.18 and Equation (1.20), xcp is defined as the center of pressure. It is the location where the resultant of a distributed load effectively acts on the body. If moments were taken about the center of pressure, the integrated effect of the distributed loads would be zero. Hence, an alternate definition of the center of pressure is that point on the body about which the aerodynamic moment is zero.

Center of Pressure Подпись: ^LE L' Подпись: [1.21]

In cases where the angle of attack of the body is small, sin a ~ 0 and cos a ~ 1: hence, from Equation (1.1), L’ ~ N’. Thus, Equation (1.20) becomes

Examine Equations (1.20) and (1.21). As N’ and L’ decrease, xcp increases. As the forces approach zero, the center of pressure moves to infinity. For this reason, the center of pressure is not always a convenient concept in aerodynamics. However, this is no problem. To define the force-and-moment system due to a distributed load on a body, the resultant force can be placed at any point on the body, as long as the value of the moment about that point is also given. For example, Figure 1.19 illustrates three equivalent ways of specifying the force-and-moment system on an airfoil. In the left figure, the resultant is placed at the leading edge, with a finite value of { . In the middle figure, the resultant is placed at the quarter-chord point, with a finite value of M’y4. In the right figure, the resultant is placed at the center of pressure, with a zero moment about that point. By inspection of Figure 1.19, the quantitative relation between these cases is

MhE = — — L + Мсц = —xcpL [1.22]

Подпись: Example 1.3In low-speed, incompressible flow, the following experimental data are obtained for an NACA 4412 airfoil section at an angle of attack of 4°: q = 0.85 and cmx/4 = —0.09. Calculate the location of the center of pressure.

Solution

From Equation (1.22),

Center of Pressure

Center of Pressure

C

4

 

cp

 

-^cp 1 (^c/4/^00^ ) 1 Off, г/4

c 4 (L’/q^c) 4 с,

Подпись: 0.356_ 1 (-0.09)

“4 0.85

(Note: In Chapter 4, we will learn that, for a thin, symmetrical airfoil, the center of pressure is at the quarter-chord location. However, for the NACA 4412 airfoil, which is not symmetric, the center-of-pressure location is behind the quarter-chord point.)

Pathlines, Streamlines, and Streaklines

of a Flow

In addition to knowing the density, pressure, temperature, and velocity fields, in aerodynamics we like to draw pictures of “where the flow is going.” To accomplish this, we construct diagrams of pathlines and/or streamlines of the flow. The distinction between pathlines and streamlines is described in this section.

Consider an unsteady flow with a velocity field given by V = V(x, y, z, t). Also, consider an infinitesimal fluid element moving through the flow field, say, element A as shown in Figure 2.25a. Element A passes through point 1. Let us trace the path

of element A as it moves downstream from point 1, as given by the dashed line in Figure 2.25a. Such a path is defined as the pathline for element A. Now, trace the path of another fluid element, say, element В as shown in Figure 2.25b. Assume that element В also passes through point 1, but at some different time from element A. The pathline of element В is given by the dashed line in Figure 2.25b. Because the flow is unsteady, the velocity at point 1 (and at all other points of the flow) changes with time. Hence, the pathlines of elements A and В are different curves in Figure 2.25a and b. In general, for unsteady flow, the pathlines for different fluid elements passing through the same point are not the same.

In Section 1.4, the concept of a streamline was introduced in a somewhat heuristic manner. Let us be more precise here. By definition, a streamline is a curve whose tangent at any point is in the direction of the velocity vector at that point. Streamlines are illustrated in Figure 2.26. The streamlines are drawn such that their tangents at every point along the streamline are in the same direction as the velocity vectors at those points. If the flow is unsteady, the streamline pattern is different at different times because the velocity vectors are fluctuating with time in both magnitude and direction.

In general, streamlines are different from pathlines. You can visualize a pathline as a time-exposure photograph of a given fluid element, whereas a streamline pattern is like a single frame of a motion picture of the flow. In an unsteady flow, the streamline pattern changes; hence, each “frame” of the motion picture is different.

However, for the case of steady flow (which applies to most of the applications in this book), the magnitude and direction of the velocity vectors at all points are fixed, invariant with time. Hence, the pathlines for different fluid elements going through the same point are the same. Moreover, the pathlines and streamlines are identical. Therefore, in steady flow, there is no distinction between pathlines and streamlines;

Velocity vector

image150

Figure 3.36 Streamlines.

image151

through point 1

Figure 3.37 For steady flow, streamlines and pathlines are the same.

they are the same curves in space. This fact is reinforced in Figure 2.27, which illus­trates the fixed, time-invariant streamline (pathline) through point 1. In Figure 2.27, a given fluid element passing through point 1 traces a pathline downstream. All sub­sequent fluid elements passing through point 1 at later times trace the same pathline. Since the velocity vector is tangent to the pathline at all points on the pathline for all times, the pathline is also a streamline. For the remainder of this book, we deal mainly with the concept of streamlines rather than pathlines; however, always keep in mind the distinction described above.

Question: Given the velocity field of a flow, how can we obtain the mathematical equation for a streamline? Obviously, the streamline illustrated in Figure 2.27 is a curve in space, and hence it can be described by the equation f(x, y,z) =0. How can we obtain this equation? To answer this question, let ds be a directed element of the streamline, such as shown at point 2 in Figure 2.27. The velocity at point 2 is V, and by definition of a streamline, V is parallel to ds. Hence, from the definition of the vector cross product [see Equation (2.4)],

ds x V = 0 I [3.115]

Equation (2.115) is a valid equation for a streamline. To put it in a more recognizable form, expand Equation (2.115) in cartesian coordinates:

Equations (2.117a to c) are differential equations for the streamline. Knowing u, v, and w as functions of x, y, and z, Equations (2.117a to c) can be integrated to yield the equation for the streamline: fix, y, z.) = 0.

To reinforce the physical meaning of Equations (2.117a to c), consider a stream­line in two dimensions, as sketched in Figure 2.28a. The equation of this streamline is у = fix). Hence, at point 1 on the streamline, the slope is dyjdx. However, V with x and у components и and v, respectively, is tangent to the streamline at point 1. Thus, the slope of the streamline is also given by v/u, as shown in Figure 2.26. Therefore,

Подпись: ds = dx і + dy'] + dz к Y = ui + pj + wk

[2.1 18]

image152

(a)

Figure 2.28 (a) Equation of a stream in two-dimensional

cartesian space, (b) Sketch of a streamtube in three-dimensional space.

 

image153

Equation (2.118) is a differential equation for a streamline in two dimensions. From Equation (2.118),

v dx — и dy = 0

which is precisely Equation (2.117c). Therefore, Equations (2.117a to c) and (2.118) simply state mathematically that the velocity vector is tangent to the streamline.

A concept related to streamlines is that of a streamtube. Consider an arbitrary closed curve C in three-dimensional space, as shown in Figure 2.28b. Consider the streamlines which pass through all points on C. These streamlines form a tube in space as sketched in Figure 2.28b; such a tube is called a streamtube. For example, the walls of an ordinary garden hose form a streamtube for the water flowing through the hose. For a steady flow, a direct application of the integral form of the continuity equation [Equation (2.53)] proves that the mass flow across all cross sections of a streamtube is constant. (Prove this yourself.)

Подпись: Example 2.3Consider the velocity field given by и = y/(x2 + y2) and v = —x/(x2 + y2). Calculate the equation of the streamline passing through the point (0, 5).

Solution

From Equation (2.118), dy/dx = v/u = —x/y, and

ydy = —x dx

Integrating, we obtain

+ c

where c is a constant of integration.

For the streamline through (0, 5), we have

52 = 0 + c or c — 25

Подпись: x2 + y2 = 25

Thus, the equation of the streamline is

Note that the streamline is a circle with its center at the origin and a radius of 5 units.

Streamlines are by far the most common method used to visualize a fluid flow. In an unsteady flow it is also useful to track the path of a given fluid element as it moves through the flow field, i. e., to trace out the pathline of the fluid element. However, separate from the ideas of a streamline and a pathline is the concept of a streakline. Consider a fixed point in a flow field, such as point 1 in Figure 2.29. Consider all the individual fluid elements that have passed through point 1 over a given time interval t2 — q. These fluid elements, shown in Figure 2.29, are connected with each other, like a string of elephants connected trunk-to-tail. Element A is the fluid element that passed through point 1 at time q. Element В is the next element that passed through point 1, just behind element A. Element C is the element that passed through point 1 just behind element B, and so forth. Figure 2.29 is an illustration, made at time t2, which shows all the fluid elements that have earlier passed through point 1 over the time interval (t2 — q). The line that connects all these fluid elements is, by definition, a

image154

Figure 2.29 Illustration of a streakline through point I.

streakline. We can more concisely define a streakline as the locus of fluid elements which have earlier passed through a prescribed point. To help further visualize the concept of a streakline, imagine that we are constantly injecting dye into the flow field at point 1. The dye will flow downstream from point 1, forming a curve in the x, y, z space in Figure 2.29. This curve is the streakline shown in Figure 2.29. A photograph of a streakline in the flow of water over a circular cylinder is shown in Figure 3.48. The white streakline is made visible by white particles that are constantly formed by electrolysis near a small anode fixed on the cylinder surface. These white particles subsequently flow downstream forming a streakline.

For a steady flow, pathlines, streamlines, and streaklines are all the same curves. Only in an unsteady flow are they different. So for steady flow, which is the type of flow mainly considered in this book, the concepts of a pathline, streamline, and streakline are redundant.

Airfoil Nomenclature

The first patented airfoil shapes were developed by Horatio F. Phillips in 1884. Phillips was an Englishman who carried out the first serious wind-tunnel experiments on airfoils. In 1902, the Wright brothers conducted their own airfoil tests in a wind tunnel, developing relatively efficient shapes which contributed to their successful first flight on December 17, 1903 (see Section 1.1). Clearly, in the early days of powered flight, airfoil design was basically customized and personalized. However, in the early 1930s, the National Advisory Committee for Aeronautics (NACA)—the

image301

Figure 4.2 Road map for Chapter 4.

forerunner of NASA—embarked on a series of definitive airfoil experiments using airfoil shapes that were constructed rationally and systematically. Many of these NACA airfoils are in common use today. Therefore, in this chapter we follow the nomenclature established by the NACA; such nomenclature is now a well-known standard.

Consider the airfoil sketched in Figure 4.3. The mean camber line is the locus of points halfway between the upper and lower surfaces as measured perpendicular to the mean camber line itself. The most forward and rearward points of the mean camber line are the leading and trailing edges, respectively. The straight line connecting the leading and trailing edges is the chord line of the airfoil, and the precise distance from

image302

Figure 4.3 Airfoil nomenclature.

the leading to the trailing edge measured along the chord line is simply designated the chord c of the airfoil. The camber is the maximum distance between the mean camber line and the chord line, measured perpendicular to the chord line. The thickness is the distance between the upper and lower surfaces, also measured perpendicular to the chord line. The shape of the airfoil at the leading edge is usually circular, with a leading-edge radius of approximately 0.02c. The shapes of all standard NACA airfoils are generated by specifying the shape of the mean camber line and then wrapping a specified symmetrical thickness distribution around the mean camber line.

The force-and-moment system on an airfoil was discussed in Section 1.5, and the relative wind, angle of attack, lift, and drag were defined in Figure 1.10. You should review these considerations before proceeding further.

The NACA identified different airfoil shapes with a logical numbering system. For example, the first family of NACA airfoils, developed in the 1930s, was the “four­digit” series, such as the NACA 2412 airfoil. Here, the first digit is the maximum camber in hundredths of chord, the second digit is the location of maximum camber along the chord from the leading edge in tenths of chord, and the last two digits give the maximum thickness in hundredths of chord. For the NACA 2412 airfoil, the maximum camber is 0.02c located at 0.4c from the leading edge, and the maximum thickness is 0.12c. It is common practice to state these numbers in percent of chord, that is, 2 percent camber at 40 percent chord, with 12 percent thickness. An airfoil with no camber, that is, with the camber line and chord line coincident, is called a symmetric airfoil. Clearly, the shape of a symmetric airfoil is the same above and below the chord line. For example, the NACA 0012 airfoil is a symmetric airfoil with a maximum thickness of 12 percent.

The second family of NACA airfoils was the “five-digit” series, such as the NACA 23012 airfoil. Here, the first digit when multiplied by | gives the design lift coefficient[13] in tenths, the next two digits when divided by 2 give the location of maximum camber along the chord from the leading edge in hundredths of chord, and the final two digits give the maximum thickness in hundredths of chord. For the NACA 23012 airfoil, the design lift coefficient is 0.3, the location of maximum camber is at 0.15c, and the airfoil has 12 percent maximum thickness.

One of the most widely used family of NACA airfoils is the “6-series” laminar flow airfoils, developed during World War II. An example is the NACA 65-218. Here,

the first digit simply identifies the series, the second gives the location of minimum pressure in tenths of chord from the leading edge (for the basic symmetric thickness distribution at zero lift), the third digit is the design lift coefficient in tenths, and the last two digits give the maximum thickness in hundredths of chord. For the NACA 65-218 airfoil, the 6 is the series designation, the minimum pressure occurs at 0.5c for the basic symmetric thickness distribution at zero lift, the design lift coefficient is

0. 2, and the airfoil is 18 percent thick.

The complete NACA airfoil numbering system is given in Reference 11. Indeed, Reference 11 is a definitive presentation of the classic NACA airfoil work up to 1949. It contains a discussion of airfoil theory, its application, coordinates for the shape of NACA airfoils, and a huge bulk of experimental data for these airfoils. This author strongly encourages you to read Reference 11 for a thorough presentation of airfoil characteristics.

As a matter of interest, the following is a short partial listing of airplanes currently in service which use standard NACA airfoils.

Airplane Airfoil

Beechcraft Sundowner Beechcraft Bonanza

Cessna 150 Fairchild A-10

Gates Learjet 24D General Dynamics F-16 Lockheed C-5 Galaxy

NACA 63A415 NACA 23016.5 (at root) NACA 23012 (at tip) NACA 2412 NACA 6716 (at root) NACA 6713 (at tip) NACA 64A109 NACA 64A204 NACA 0012 (modified)

In addition, many of the large aircraft companies today design their own special- purpose airfoils; for example, the Boeing 727, 737, 747, 757, and 767 all have spe­cially designed Boeing airfoils. Such capability is made possible by modern airfoil design computer programs utilizing either panel techniques or direct numerical finite – difference solutions of the governing partial differential equations for the flow field. (Such equations are developed in Chapter 2.)

Compressible Flow: Some Preliminary Aspects

With the realization of aeroplane and missile speeds equal to or even surpassing many times the speed of sound, thermodynamics has entered the scene and will never again leave our considerations.

Jakob Ackeret, 1962

7.1 Introduction

On September 30, 1935, the leading aerodynamicists from all comers of the world converged on Rome, Italy. Some of them arrived in airplanes which, in those days, lumbered along at speeds of 130 mi/h. Ironically, these people were gathering to discuss airplane aerodynamics not at 130 mi/h but rather at the unbelievable speeds of 500 mi/h and faster. By invitation only, such aerodynamic giants as Theodore von Karman and Eastman Jacobs from the United States, Ludwig Prandtl and Adolf Busemann from Germany, Jakob Ackeret from Switzerland, G. I. Taylor from Eng­land, Arturo Crocco and Enrico Pistolesi from Italy, and others assembled for the fifth Volta Conference, which had as its topic “High Velocities in Aviation.” Although the jet engine had not yet been developed, these men were convinced that the future of aviation was “faster and higher.” At that time, some aeronautical engineers felt that airplanes would never fly faster than the speed of sound—the myth of the “sound barrier” was propagating through the ranks of aviation. However, the people who attended the fifth Volta Conference knew better. For 6 days, inside an impressive Re­naissance building that served as the city hall during the Holy Roman Empire, these

individuals presented papers that discussed flight at high subsonic, supersonic, and even hypersonic speeds. Among these presentations was the first public revelation of the concept of a swept wing for high-speed flight; Adolf Busemann, who originated the concept, discussed the technical reasons why swept wings would have less drag at high speeds than conventional straight wings. (One year later, the swept-wing con­cept was classified by the German Luftwaffe as a military secret. The Germans went on to produce a large bulk of swept-wing research during World War II, resulting in the design of the first operational jet airplane—the Me 262—which had a moderate degree of sweep.) Many of the discussions at the Volta Conference centered on the effects of “compressibility” at high subsonic speeds, that is, the effects of variable density, because this was clearly going to be the first problem to be encountered by future high-speed airplanes. For example, Eastman Jacobs presented wind-tunnel test results for compressibility effects on standard NACA four – and five-digit airfoils at high subsonic speeds and noted extraordinarily large increases in drag beyond certain freestream Mach numbers. In regard to supersonic flows, Ludwig Prandtl presented a series of photographs showing shock waves inside nozzles and on various bodies— with some of the photographs dating as far back as 1907, when Prandtl started serious work in supersonic aerodynamics. (Clearly, Ludwig Prandtl was busy with much more than just the development of his incompressible airfoil and finite-wing theory discussed in Chapters 4 and 5.) Jakob Ackeret gave a paper on the design of su­personic wind tunnels, which, under his direction, were being established in Italy, Switzerland, and Germany. There were also presentations on propulsion techniques for high-speed flight, including rockets and ramjets. The atmosphere surrounding the participants in the Volta Conference was exciting and heady; the conference launched the world aerodynamic community into the area of high-speed subsonic and super­sonic flight—an area which today is as commonplace as the 130-mi/h flight speeds of 1935. Indeed, the purpose of the next eight chapters of this book is to present the fundamentals of such high-speed flight.

In contrast to the low-speed, incompressible flows discussed in Chapters 3 to 6, the pivotal aspect of high-speed flow is that the density is a variable. Such flows are called compressible flows and are the subject of Chapters 7 to 14. Return to Figure 1.38, which gives a block diagram categorizing types of aerodynamic flows. In Chapters 7 to 14, we discuss flows which fall into blocks D and F that is, we will deal with inviscid compressible flow. In the process, we touch all the flow regimes itemized in blocks G through J. These flow regimes are illustrated in Figure 1.37; study Figures 1.37 and 1.38 carefully, and review the surrounding discussion in Section 1.10 before proceeding further.

In addition to variable density, another pivotal aspect of high-speed compressible flow is energy. A high-speed flow is a high-energy flow. For example, consider the flow of air at standard sea level conditions moving at twice the speed of sound. The internal energy of 1 kg of this air is 2.07 x 105 J, whereas the kinetic energy is larger, namely, 2.31 x 105 J. When the flow velocity is decreased, some of this kinetic energy is lost and reappears as an increase in internal energy, hence increasing the temperature of the gas. Therefore, in a high-speed flow, energy transformations and temperature changes are important considerations. Such considerations come under

image487

Figure 7.1 Road map for Chapter 7.

the science of thermodynamics. For this reason, thermodynamics is a vital ingredient in the study of compressible flow. One purpose of the present chapter is to review briefly the particular aspects of thermodynamics which are essential to our subsequent discussions of compressible flow.

The road map for this chapter is given in Figure 7.1. As our discussion proceeds, refer to this road map in order to provide an orientation for our ideas.

Summary

Refer again to the road map for Chapter 1 given in Figure 1.6. Read again each block in this diagram as a reminder of the material we have covered. If you feel uncomfortable about some of the concepts, or if your memory is slightly “foggy” on certain points, go back and reread the pertinent sections until you have mastered the material.

This chapter has been primarily qualitative, emphasizing definitions and basic concepts. However, some of the more important quantitative relations are summarized below:

 

Summary

The center of pressure is obtained from

 

Mle

~T~

 

Mle

N

 

r’cp

 

[1.80] and [1.91]

 

The criteria for two or more flows to be dynamically similar are:

1. The bodies and any other solid boundaries must be geometrically similar.

2. The similarity parameters must be the same. Two important similarity parameters are Mach number M = V/a and Reynolds number Re = p V с/ц..

If two or more flows are dynamically similar, then the force coefficients Cl, Cd, etc., are the same.

 

image81image82

In fluid statics, the governing equation is the hydrostatic equation:

dp = ~gp dy

[1.52]

For a constant density medium, this integrates to

p + pgh = constant

[1.54]

or p і + pgh = P2 + pgh2

Such equations govern, among other things, the operation of a manometer, and also lead to Archimedes’ principle that the buoyancy force on a body immersed in a fluid is equal to the weight of the fluid displaced by the body.