Category AERODYNAMICS, AERONAUTICS, AND FLIGHT MECHANICS

During training, a pilot soon learns that the airspeed that appears on the airspeed indicator is not the true airspeed. Instead, in order to determine the true airspeed, the pilot must also read the altimeter and outside air tem­perature. The pilot then resorts to a small hand calculator or, in some instances, adjusts the dial on the airspeed indicator accordingly to allow for the atmospheric properties. ^

The airspeed indicator is nothing more than an accurate differential pres­sure gage calibrated according to Equation 2.31. This equation can be put in the form

The airspeed indicator measures the difference, p0-p (sometimes called the compressible dynamic pressure), but then is calibrated to obtain V by assuming standard sea level values for the acoustic velocity and the free-

/)>t f?’

stream static pressure. Thus the calibrated airspeed is defined by

where a subscript SL is used to denote standard sea level values. As a ratio to Vcai, the true airspeed can be written

_V_ { [[/(VcaQ/S + l^-‘^-ni1’2 Vca, Г L [f(VCil)+l](1’-Wl’-l JJ

^ = /(Vca,)

SL

and can be obtained from Equation 2.40. 0 and S are the temperature and pressure ratios, respectively.

If compressibility can be neglected, the airspeed will be given by Equa-

tion 2.36. In this case, using the standard sea level value of p to calibrate the airspeed indicator dial leads to the equivalent airspeed.

p(Po-p)T/2

L Psl J

where a is the density ratio.

Finally, the indicated airspeed, Viv is defined simply as the airspeed that the pilot reads from the airspeed indicator. If the indicator has no mechanical error (instrument error) and if the static source is located so that it samples the true free-stream static pressure (otherwise a position error incurs), Vf = Vcai – Furthermore, if the Mach number is net too high, V = Vcai = Ve.

As an example in determining true airspeed, suppose a pilot reads an indicated airspeed of 180 m/s for an OAT of 239 °K and an altimeter reading of 6000 m. This altitude, according to Equations 2.8 and 2.10, corresponds to a pressure ratio 8 of 0.466. The measured temperature ratio, в, is equal to 0.826. Hence <7 = 0.564. According to Equation 2.42, the true airspeed will be 239.7 m/s. From Equation 2.41, the true airspeed is calculated to be 231.6 m/s. Thus, using the incompressible relationship to estimate the true airspeed from the indicated airspeed results in a speed a few percent higher than that obtained from the calibrated airspeed relationship.

To be precise one calculates the true airspeed from the calibrated airspeed and then determines the equivalent airspeed from its definition, Equation 2.42. In the previous example, this results in a Ve of 173.9 m/s, a value 3.4% less than the calibrated airspeed.

Bernoulli’s equation is well known in fluid mechanics and relates the pressure to the velocity along a streamline in an inviscid, incompressible flow. It was first formulated by Euler in the middle 1700s. The derivation of this equation follows from Euler’s equations using the fact that along a streamline the velocity vector is tangential to the streamline.

dx _ dy dz

U V w

First, multiply Equation 2.22a through by dx and then substitute Equation 2.26 for v dx and w dx. Also, the first term of the equation will be set equal to zero; that is, at this time only steady flow will be considered.

Similarly, multiply Equation 2.22b by dy, Equation 2.22c by dz, and substitute Equation 2.26 for и dy, w dy and и dz, v dz, respectively. Adding the three equations results in perfect differentials for p and V2, V being the magnitude of the resultant velocity along the streamline. This last term results from the fact that

du 1 du2

dx 2 dx

and

V2 = u2 + v2+ w

Thus, along a streamline, Euler’s equations become

V dV + —= 0

If p is not a function of p (i. e., the flow is incompressible), Equation 2.27 can be integrated immediately to give

P + 2pV2 = constant (2.28)

If the flow is uniform at infinity, Equation 2.28 becomes

P + 2pV2 = constant = p« + 2PVJ (2.29)

Here V is the magnitude of the local velocity and p is the local static pressure. V. and p„ are the corresponding free-stream values. Equation 2.29 is known as Bernoulli’s equation.

The counterpart to Equation 2.29 for compressible flow is obtained by assuming pressure and density changes to follow an isentropic process. For such a process,

plpy = constant (2.30)

у is the ratio of the specific heat at constant pressure to the specific heat at constant volume and is equal approximately to 1.4 for air. Substituting Equation 2,30 into Equation 2.27 and integrating leads to an equation some­times referred to as the compressible Bernoulli’s equation.

Xr H—2_ — = constant (2.31)

2 у 1 p

This equation can be written in terms of the acoustic velocity. First it is necessary to derive the acoustic velocity, which can be done by the use of the momentum theorem and continuity. Figure 2.9 assumes the possibility of a stationary disturbance in a steady flow across which the pressure, density, and velocity change by small increments. In the absence of body forces and viscosity, the momentum theorem gives

But, from continuity, or

Thus

и + du ————

P + dp, p + dp

Figure 2.9 A stationary small disturbance in a steady compressible flow.

If the small disturbance is stationary in the steady flow having a velocity of u, then obviously и is the velocity of the disturbance relative to the fluid. By definition,- it follows that u, given by Equation 2.32, is the acoustic velocity. By the use of Equation 2.30, the acoustic velocity is obtained as

(2.33)

An alternate form, using the equation of state (Equation 2.1), is

a = (yRT)m (2.34)

Thus Equation 2.31 can be written

V2 a2

~2 + = constant (2.35)

The acoustic velocity is also included in Figure 2.3 for the standard atmos­phere.

Determination of Free-Stream Velocity

At low speeds (compared to the acoustic velocity) a gas flow is essen­tially incompressible. In this case, and for that of a liquid, Equation 2.29 applies. If the fluid is brought to rest so that the local velocity is zero then, from Equation 2.29, the local pressure, referred to in this case as the stagnation or total pressure, is equal to> the sum of the free-stream static pressure, p„, and pVj/2. This latter term is called the dynamic pressure and is frequently denoted by the symbol q. Thus,

-P P

where p0 is the total pressure, also referred to as the stagnation or reservoir pressure. The pitot-static tube shown in Figure 2.2 measures (po — p«) and is probably the most common means used to determine airspeed. However, notice that Equation 2.36 contains the mass density that must be determined before the airspeed can be calculated. This is most readily achieved by measuring, in addition to the difference between the stagnation pressure and the static pressure, the static pressure itself and the temperature. The density is then determined from the equation of state (Equation 2.1).

At higher speeds (and we will now examine what is high) Equation 2.29 no longer holds, so that V» must be determined from Equation 2.31.

V2 , 7 P _ 7 Po

2 у – 1 p у – 1 p0

At this point the subscript °° has been dropped, so that p, p, and V without a subscript may be local or free-stream values.

Remembering that yplp is the square of the acoustic velocity, the preceding equation becomes

M2 | 1 _ 1 pop

2 у – 1 у – 1 p po

Using Equation 2.30, this can be written as

The dynamic pressure, q, is defined as

q = PV2 (2.38)

which can be written in terms of the Mach number as

q = |урЛ#

Combining this with Equation 2.37 gives

The square root of Equation 2.39 is presented graphically in Figure 2.10. The departure of this function from unity is a measure of the error to be incurred in calculating the airspeed from the incompressible Bernoulli equa­tion. Below a Mach number of 0.5 the error is seen to be less than 3%.

The principle of conservation of mass, applied to an elemental control surface, led to Equation 2.17, which must be satisfied everywhere in the flow. Similarly, the momentum theorem applied to the same element leads to another set of equations that must hold everywhere.

Referring again to Figure 2.7, if p is the static pressure at the center of the element then, on the center of the right face, the static pressure will be

| dp Ax + dx 2

This pressure and a similar pressure on the left face produce a net force in the x direction equal to

dp *„

Since there are no body forces present and the fluid is assumed inviscid, the above force must equal the net momentum flux out plus the instantaneous change of fluid momentum contained within the element.

The momentum flux out of the right face in the x direction will be

Out of the upper face the corresponding momentum flux will be

Similar expressions can be written for the momentum flux in through the left and bottom faces.

The instantaneous change of the fluid momentum contained within the element in the x direction is simply %

~(pu Ax Ay)

Thus, equating the net forces in the x direction to the change in momentum and momentum flux and using Equation 2.17 leads to

du 8u du _ 1 dp

dt U dx V dy p dx

Generalizing this to three dimensions results in a set of equations known as Euler’s equations of motion.

du , dU, dU, dU 1 dp

dv, dv , dv, dv 1 dp

—+И— +U— + w — =—— r~

dt dx dy dz p dy

dw, dw, dw, dw 1 dp

——+ U — + V ——+ w——=———————— r-

dt dx dy dz p dz

Notice that if и is written as u{x, y, z, t), the left side of Equation 2.22 is the total derivative of u. The operator, d( )ldt, is the local acceleration and exists only if the flow is unsteady.

In vector notation Euler’s equation can be written

§ + (V – V)V = –Vp (2.23)

ol p

If the vector product of the operator V is taken with each term in Equation 2.23, Equation 2.24 results.

<0 is the curl of the velocity vector, V x V, and is known as the vorticity.

(2-25)

One can conclude from Equation 2.24 that, for an inviscid fluid, the vorticity is constant along a streamline. Since, far removed from a body, the flow is usually taken to be uniform, the vorticity at that location is zero; hence, it is zero everywhere.

Fluid passing through an area at a velocity of V has a mass flow rate equal to pAV. This is easily seen by reference to Figure 2.6. Here flow is

Reference

plane

Figure 2.6 Mass flow through a surface.

pictured along a streamtube of cross-sectional area A. The fluid velocity is equal to V. At time t = 0, picture a small slug of fluid of length, 1, about to cross a reference plane. At time //V, this entire slug will have passed through the reference plane. The volume of the slug is Al, so that a mass of pAl was transported across the reference plane during the time 1/V. Hence the mass rate of flow, m, is given by

pAl m (1/V)

= pAV

Along a streamtube (which may be a conduit with solid walls) the quantity pAV must be a constant if mass is not to accumulate in the system. For incompressible flow, p is a constant, so that the conservation of mass leads to the continuity principle

AV = constant

AV is the volume flow rate and is sometimes referred to as the flux. Similarly, pAV is the mass flux. The mass flux through a surface multiplied by the velocity vector at the surface is defined as the momentum flux. Generally, if the velocity vector is not normal to the surface, the mass flux will be

pAV • n

with the momentum flux written as

(pAV • n)V

here n is the unit vector normal to the surface and in the direction in which the flux is defined to be positive. For example, if the surface encloses a volume and the net mass flux out of the volume is to be calculated, n would

be directed outward from the volume, and the following integral would be evaluated over the entire surface.

Consider the conservation of mass applied to a differential control surface. For simplicity, a two-dimensional flow will be treated. A rectangular contour is shown in Figure 2.7. The flow passing through this element has velocity components of и and v in the center of the element in the x and у directions, respectively. The corresponding components on the right face of the element are found by expanding them in a Taylor series in x and у and dropping second-order and higher terms in Ax. Hence the mass flux out through the right face will be

Г, д(ри)Дл:1 A

Г+-Гт]А’

Writing similar expressions for the other three faces leads to the net mass flux

*

Га^+8£»)1

L to dy J

The net mass flux out of the differential element must equal the rate at which the mass of the fluid contained within the element is decreasing, given by

– (p Ax Ay)

у

м pv + dx/2

pi >

pu

Ax

Figure 2.7 A rectangular-differential control surface.

Since Ax and Ay are arbitrary, it follows that, in general,

dp d(pu) djpv) _ – dt dx dy

In three dimensions the preceding equation can be written in vector notation

as

!f + V-(pV) = 0 (2.17)

where V is the vector operator, del, defined by

Any physically possible flow must satisfy Equation 2.17 at every point in the flow.

For an incompressible flow, the mass density is a constant, so Equation 2.17 reduces to

VV = 0 (2.18)

The above is known as the divergence of the velocity vector, div V.

The Momentum Theorem

The momentum theorem in fluid mechanics is the counterpart of New­ton’s second law of motion in solid mechanics, which states that a force imposed on a system produces a rate of change in the momentum of the system. The theorem can be easily derived by treating the fluid as a collection of fluid particles and applying the second law. The details of the derivation can be found in several texts (e. g., Ref. 2.1) and will not be repeated here.

Defining a control surface as an imaginary closed surface through which a flow is passing, the momentum theorem states:

“The sum of external forces (or moments) acting on a control surface and internal forces (or moments) acting on the fluid within the control surface produces a change in the flux of momentum (or angular momentum) through the surface and an instantaneous rate of change of momentum (or angular momentum) of the fluid particles within the control surface.”

Mathematically, for linear motion of an inviscid fluid, the theorem can be expressed in vector notation by

– j pn dS + B = fsf PV(V ‘ n) dS + Yt f jy f PV dr (2.19)

In Equation 2.19, n is the unit normal directed outward from the surface,

S, enclosing the volume, V. V is the velocity vector, which generally depends on position and time. В represents the vector sum of all body forces within the control surface acting on the fluid, p is the mass density of the fluid defined as the mass per unit volume.

For the angular momentum,

Here, Q is the vector sum of all moments, both internal and external, acting on the control surface or the fluid within the surface, г is the radius vector to a fluid particle.

As an example of the use of the momentum theorem, consider the force on the burning building produced by the firehose mentioned at the beginning of this chapter. Figure 2.8 illustrates a possible flow pattern, admittedly simplified. Suppose the nozzle has a diameter of 10 cm and water is issuing from the nozzle with a velocity of 60m/s. The mass density of water is approximately 1000 kg/m3. The control surface is shown dotted. Equation 2.19 will now be written for this system in the x direction. Since the flow is steady, the partial*derivative with respect to time of the volume integral given by the last term on the right side of the equation vanishes. Also, В is zero, since the control surface does not enclose any

bodies. Thus Equation 2.19 becomes

-JsfpndS = Jsf pV(V ■ n) dS

Measuring p relative to the atmospheric static pressure, p is zero everywhere along the control surface except at the wall. Here n is directed to the right so that the surface integral on the left becomes the total force exerted on the fluid by the pressure on the wall. К F represents the magnitude of the total force on the wall,

-IF = J / pV(V • n) dS

For the fluid entering the control surface on the left,

V = 60i n = — і

For the fluid leaving the control surface, the unit normal to this cylindrical surface has no component in the x direction. Hence,

– iF = – J j (1000)60i(-60) dS

= — 36 x 105i J J dS

The surface integral reduces to the nozzle area of 7.85 x 10“3 m2. Thus, without actually determining the pressure distribution on the wall, the total force on the wall is found from the momentum theorem to equal 28.3 kN.

We will now treat a fluid that is moving so that, in addition to gravita­tional forces, inertial and shearing forces must be considered.

A typical flow around a streamlined shape is pictured in Figure 2.4. Note that this figure is labled “two-dimensional flow”; this means simply that the flow field is a function only of two coordinates (x and y, in the case of Figure 2.4) and does not depend on the third coordinate. For example, the flow of wind around a tall, cylindrical smokestack is essentially two-dimensional except near the top. Here the wind goes over as well as around the stack, and the flow is three-dimensional. As another example, Figure 2.4 might represent the flow around a long, streamlined strut such as the one that supports the wing of a high-wing airplane. The three-dimensional counterpart of this shape might be the blimp.

Several features of flow around a body in general are noted in Figure 2.4. First, observe that the flow is illustrated by means of streamlines. A stream­line is an imaginary line characterizing the flow such that, at every point along the line, the velocity vector is tangent to the line. Thus, in two-dimensional

Figure 2.4 Two-dimensional, flow around a streamlined shape.

flow, if y(x) defines the position of a streamline, y(x) is related to the x and у components of the velocity, u(x) and v(x), by

dy _ v(x) dx u(x)

Note that the body surface itself is a streamline.

In three-dimensional flow a surface swept by streamlines is known as a stream surface. If such a surface is closed, it is known as a stream tube.

The mass flow accelerates around the body as the result of a continuous distribution of pressure exerted on the fluid by the body. An equal and opposite reaction must occur on the body. This static pressure distribution, acting everywhere normal to the body’s surface, is pictured on the lower half of the body in Figure 2.4. The small arrows represent the local static pressure, p, relative to the static pressure, po, in the fluid far removed from the body. Near the nose p is greater than p0; further aft the pressure becomes negative relative to po- If this static pressure distribution, acting normal to the surface, is known, forces on the body can be determined by integrating this pressure over its surface.

In addition to the local static pressure, shearing stresses resulting from the fluid’s viscosity also give rise to body forces. As fluid passes over a solid surface, the fluid particles immediately in contact with the surface are brought to rest. Moving away from the surface, successive layers of fluid are slowed by the shearing stresses produced by the inner layers. (The term “layers” is used only as a convenience in describing the fluid behavior. The fluid shears in a continuous manner and not in discrete layers.) The result is a thin layer of slower moving fluid, known as the bbundary layer, adjacent to the surface. Near the front of the body this layer is very thin, and the flow within it is smooth without any random or turbulent fluctuations. Here the fluid particles might be described as moving along in the layer on parallel planes, or laminae; hence the flow is referred to as laminar.

At some distance back from the nose of the body, disturbances to the flow (e. g., from surface roughnesses) are no longer damped out. These disturbances suddenly amplify, and the laminar boundary layer undergoes transition to a turbulent boundary layer. This layer is considerably thicker than the laminar one and is characterized by a mean velocity profile on which small, randomly fluctuating velocity components are superimposed. These flow regions are shown in Figure 2.4. The boundary layers are pictured considerably thicker than they actually are for purposes of illustration. For example, on the wing of an airplane flying at 1(X) m/s at low altitude, the turbulent boundary 1.0 m back from the leading edge would be only approximately 1.6 cm thick. If the layer were still laminar at this point, its thickness would be approximately 0.2 cm.

Returning to Figure 2.4, the turbulent boundary layer continues to thicken toward the rear of the body. Over this portion of the surface the fluid

is moving into a region of increasing static pressure that is tending to oppose the flow. The slower moving fluid in the boundary layer may be unable to overcome this adverse pressure gradient, so that at some point the flow actually separates from the body surface. Downstream of this separation point, reverse flow will be found along the surface with the static pressure nearly constant and equal to that at the point of separation.

At some distance downstream of the body the separated flow closes, and a wake is formed. Here, a velocity deficiency representing a momentum loss by the fluid is found near the center of the wake. This decrement of momentum (more precisely, momentum flux) is a direct measure of the body drag (i. e., the force on the body in the direction of the free-stream velocity).

The general flow pattern described thus far can vary, depending on the size and shape of the body, the magnitude of the free-stream velocity, and the properties of the fluid. Variations in these parameters can eliminate transition or separation or both.

One might reasonably assume that the forces on a body moving through a fluid depend in some way on the mass density of the fluid, p, the size of the body, l, and the body’s velocity, V. If we assume that any one force, F, is proportional to the product of these parameters each raised to an unknown power, then

F oc p°vblc

In order for the basic units of mass, length, and time to be consistent, it follows that

Considering M, L, and T in order leads to three equations for the unknown exponents a, b, and c from which it is found that a = 1, b =2, and c = 2. Hence,

F oc pv42 (2.12)

Fqr a particular force the constant of prdfc>ortionality in Equation 2.12 is referred to as a coefficient and is modified by the name of the force, for example, the lift coefficient. Thus the lift and drag forces, L and D, can be expressed as

L = l2pV2SCL (2.13 a)

D = 12pV2SCD (2.13b)

Note that the square of the characteristic length, l2, has been replaced by a reference area, S. Also, a factor of 1/2 has been introduced. This can be done, since the lift and drag coefficients, CL and CD, are arbitrary at this point. The quantity pV2/2 is referred to as the dynamic pressure, the significance of which will be – made clear shortly.

For many applications, the coefficients CL and CD remain constant for a given geometric shape over a wide range of operating conditions or body size. For example, a two-dimensional airfoil at а Г angle of attack will have a lift coefficient of approximately 0.1 for velocities from a few meters per second up to 100 m/s or more. In addition, CL will be almost independent of the size of the airfoil. However, a more rigorous application of dimensional analysis [see Buckingham’s – n theorem (Ref. 2.1)] will result in the constant of proportionality in Equation 2.12 possibly being dependent on a number of dimensionless parameters. Two of the most important of these are known as the Reynolds number, R, and the Mach number, M, defined by,

where l is a characteristic length, V is the free-stream velocity, ju, is the coefficient of viscosity, and a is the velocity of sound. The velocity of sound is the speed at which a small pressure disturbance is propagated through the fluid; at this point, it requires no further explanation. The coefficient of viscosity, however, is not as well known and will be elaborated on by reference to Figure 2.5. Here, the velocity profile is pictured in the boundary layer of a laminar, viscous flow over a surface. The viscous shearing produces a shearing stress of rw on the wall. This force per unit area is related to the gradient of the velocity u(y) at the wall by

(2.15)

Actually, Equation 2.15 is applicable to calculating the shear stresses between fluid elements and is not restricted simply to the wall. Generally, the viscous shearing stress in the fluid in any plane parallel to the flow and away

у

ІШШШІжшшшіш Figure 2.5 Viscous flow adjacent to a surface.

from the wall is given by the product of ц and, the velocity gradient normal to the direction of flow.

The kinematic viscosity, v, is defined as the ratio of ц to p.

it

P v is defined as a matter of convenience, since it is the ratio of p. to p that governs the Reynolds number. The kinematic viscosity for the standard atmosphere is included in Figure 2.3 as an inverse fraction of the standard sea level value.

A physical significance can be given to the Reynolds number by multi­plying numerator and denominator by V and dividing by l.

QiVfl)

In the following material (see Equation,2.28) the normal pressure will be shown to be proportional to pV2 whereas, from Equation 2.15, pV/l is proportional to the shearing stress. Hence for a given flow the Reynolds number is proportional to the ratio of normal pressures (inertia forces) to viscous shearing stresses. Thus, relatively speaking, a flow is less viscous than another flow if its Reynolds number is higher than that of the second flow.

The Mach number determines to what extent fluid compressibility can be neglected’ (i. e., the variation of mass density with pressure). Current jet transports, for example, can cruise at Mach numbers up to approximately 0.8 before significant compressibility effects are encountered.

At lower Mach numbers, two flows are geometrically and dynamically similar if the Reynolds numbers are the same for both flows. Hence, for example, for a given shape, CD for a body 10 m long at 100 m/s will be the same as CD for a 100-m long body at 10 m/s. As another example, suppose transition occurs 2 m back from the leading edge of a flat plate aligned with a flow having a velocity of 50 m/s. Then, at 25 m/s transition would occur at a distance of 4 m from the leading edge. Obviously the effects of R and M on dimensionless aerodynamic coefficients must be considered when interpreting test results obtained with the use of small models.

For many cases of interest to aerodynamics the pressure field around a shape can be calculated assuming the air to be inviscid and incompressible. Small corrections can then be made to the resulting solutions to account for these “real fluid” effects. Corrections for viscosity or compressibility will be considered as needed in the following chapters.

Before treating the more difficult case of a fluid in motion, let us consider a fluid at rest in static equilibrium. The mass per unit volume of a fluid is defined as the mass density, usually denoted by p. The mass density is a

16

constant for liquids, but it is a function of temperature, T, and pressure, p, for gases. Indeed, for a gas, p, p, and T are related by the equation of state

P = pKT (2.1)

R is referred to as the gas constant and has a value of 287.3 m2/°K-sec2 for air at normal temperatures. In Equation 2.1, T is the thermodynamic or absolute temperature in degrees Kelvin. T and the Celsius temperature, t, are related by

T = t+ 273.15 (2.2)

A container filled with a liquid is pictured in Figure 2.1a. A free-body diagram of a small slug of the fluid is shown in Figure 2.1b. This slug has a

Figure 2.1 The variation of pressure with depth in a liquid.

unit cross-sectional area and a differential length of dh. Acting downward over the upper surface is the static pressure, p, while acting upward over the lower face is this same pressure plus the rate of increase of p with depth multiplied by the change in depth, dh. The static pressure also acts inward around the sides of the element, but this contributes nothing to the balance of forces in the vertical direction. In addition to the pressure forces, the weight of the fluid element, pgdh, acts vertically downward; g is the gravitational constant.

Summing forces on the element in the vertical direction leads to

Integrating Equation 2.3 from h = 0 at the surface to any depth, h, results in the static pressure as a function of the depth.

P = Pa + pgh (2.4)

where pa is the atmospheric pressure at the free surface.

A manometer is a device frequently used to measure pressures. It is based on Equation 2.4. Consider the experimental setup pictured in Figure

2.2. Here, a device known as a pitot-static tube is immersed in and aligned with a gas flow. The impact of the gas being brought to rest at the nose of the tube produces a pressure higher than that along the sides of the tube. This pressure, known as the total pressure, is transmitted through a tube to one side of a U-shaped glass tube partially filled with a liquid. Some distance back from the nose of the pitot-static tube the pressure is sampled through a small opening that is flush with the sides of the tube. This opening, if it is far enough back from the nose, does not disturb the flow so that the pressure sampled by it is the same as the static pressure of the undisturbed flow. This static pressure is transmitted to the right side of the glass U-tube manometer. The total pressure, being higher than the static pressure, causes the liquid in the right side of the U-tube to drop while the level on the left side rises.

If we denote p as the static pressure and p + Ap as the total pressure, the pressure at the bottom of the U-tube can be calculated by Equation 2.4 using either the right or left side of the tube. Equating the results from the two sides gives

p + Ap + pgh0 = p + pg(Ah + h0)

or

Ap = pg Ah (2.5)

Hence, the difference of the liquid levels in the two sides of the manometer is a direct measure of the pressure difference applied across the manometer. In this case we could then determine the difference between the total pressure

Gas flow

and static pressure in the gas flow from which, as we will see later, the velocity of the gas can be calculated. ; Now consider the variation of static pressure through the atmosphere. Again the forces acting on a differential mass of gas will be treated in a manner similar to the development of Equation 2.3 for a liquid. However, h will be taken to be the altitude above the ground and, since the gravitational attrac – tion is now opposite to the direction of increasing h, the sign of Equation 2.3 Ї changes. For the atmosphere,

(2.6)

The mass density, p, is not a constant in this case, so that Equation 2.6 cannot be integrated immediately. In order to perform the integration the equation of state, Equation 2.1, is substituted for p, which leads to

dp _ gdh p RT

(2.7)

From experimental observation, the variation of temperature with al­titude is known or, at least, a standard variation has been agreed on. Up to an altitude of 11 km, the temperature is taken to decrease linearly with altitude at a rate, known as the lapse rate, of 6.51 °C/km. Thus, Equation 2.7 becomes

dp _ dT g p T R(dTldh) or

8 = 05 2561 (2.8)

where 5 is the ratio of the static pressure at altitude to the pressure at sea

level and в is the corresponding absolute temperature ratio.

Using the equation of state, the corresponding density ratio, a, is obtained immediately from Equation 2.8.

8

а = в or

c = 042561 (2.9)

Using the standard lapse rate and a sea level temperature of 288.15 °K, в as a function of altitude is given by

0 = 1— 0.02256 h (2.10)

where h is the altitude in kilometers.

The lower region of the atmosphere up to an altitude for which Equations 2.8 to 2.10 hold is referred to as the troposphere. This is the region in which most of today’s flying is done. Above 11 km and up to an altitude of approximately 23 km, the temperature is nearly constant. This region forms the lower part of the stratosphere. Through the remainder of the stratosphere, the temperature increases, reaching approximately 270 °K at an altitude of around 50 km.

Figure 2.3 presents graphs (taken from Ref. 2.3) of the various properties of the standard atmosphere as a function of altitude. Each property is presented as a ratio to its standard sea level value denoted by the subscript “0.” In addition to p, p, and T, the acoustic velocity and kinematic viscosity are presented. These two properties will be defined later.

One normally thinks of altitude as the vertical distance of an airplane above the earth’s surface. However, the operation of an airplane depends on the properties of the air through which it is flying, not on the geometric height. Thus the altitude is frequently specified in terms of the standard atmosphere. Specifically, one refers to the pressure altitude or the density altitude as the height in the standard atmosphere corresponding tv. the pressure or density, respectively, of the atmosphere in which the airplane is operating. An air­plane’s altimeter is simply an absolute pressure gage calibrated according to

the standard atmosphere. It has a manual adjustment to allow for variations in sea level barometric pressure. When set to standard sea level pressure (760 mm Hg, 29.92 in. Hg), assuming the instrument and static pressure source to be free of errors, the altimeter will read the pressure altitude. When set to the local sea level barometric pressure (which the pilot can obtain over the

radio while in flight), the altimeter will read closely the true altitude above sea level. A pilot must refer to a chart prescribing the ground elevation above sea level in order to determine the height above the ground.

This chapter will stress the principles in fluid mechanics that are especi­ally important to the study of aerodynamics. For the reader whose pre­paration does not include fluid mechanics, the material in this chapter should be sufficient to understand the developments in succeeding chapters. For a more complete treatment, see any of the many available texts on fluid mechanics (e. g., Refs. 2.1 and 2.2).

Unlike solid mechanics, one normally deals with a continuous medium in the study of fluid mechanics. An airplane in flight does not experience a sudden change in the properties of the air surrounding it. The stream of water from a firehose exerts a steady force on the side of a burning building, unlike the impulse on a swinging bat as it connects with the discrete mass of the baseball.

In solid mechanics, one is concerned with the behavior of a given, finite system of solid masses under the influence of force and moment vectors acting on the system. In fluid mechanics one generally deals not with a finite system, but with the flow of a continuous fluid mass under the influence of distributed pressures and shear stresses.

The term fluid should not be confused with the term liquid, since the former includes not only the latter, but gases as well. Generally a fluid is defined as any substance that will readily deform under the influence of shearing forces. Thus a fluid is the antonym of a solid. Since both liquids and gases satisfy this definition, they are both known as fluids. A liquid is distinguished from a gas by the fact that the former is nearly incompressible. Unlike a gas, the volume of a given mass of liquid remains nearly constant, independent of the pressure imposed on the mass.

Consider the airplane in steady, climbing flight shown in Figure 1.3. The term steady means that the airplane is not accelerating; hence, all forces and moments on the aircraft must be in balance. To be more precise, one states that the vector sum of all forces and moments on the airplane must be zero. To depict the angles more clearly, all forces are shown acting through the center of gravity (eg). Although the resultant of all the forces must pass

through the center of gravity, it is not generally true that any one of the forces, with the exception of W, must satisfy this condition.

In this figure V represents the velocity vector of the airplane’s center of gravity. This vector is shown inclined upward from the horizontal through the angle of climb, 0C. The angle between the horizontal and the thrust line is denoted as 0. If this line is taken to be the reference line of the airplane, then one states that the airplane is pitched up through this angle. The angle between the reference line and the velocity vector, в – вс, is referred to as the angle of attack of the airplane. Later we will use other angles of attack referenced to the wing geometry; thus, one must be careful in interpreting lift and drag data presented as a function of the angle of attack.

The thrust, T, is the propelling force that balances mainly the aerody­namic drag on the airplane. T can be produced by a propeller, a turbojet, or a rocket engine.

The lift, L, is defined as the component of all aerodynamic forces generated by the aircraft in the direction normal to the velocity vector, V. In level flight this means principally the upward vertical force produced by the wing. Generally, however, it includes the tail and fuselage forces. For exam­ple, in landing many aircraft require a downward force on the horizontal tail in order to trim out the nose-down pitching moment produced by the wing flaps. This trimming force can be significant, requiring a wing lift noticeably in excess of the airplane’s weight.

Similar to the lift, the drag, D, is defined as the component of all aerodynamic forces generated by the airplane in the direction opposite to the

velocity vector, V. This force is composed of two principal parts, the parasite drag and the induced drag. The induced drag is generated as a result of producing lift; the parasite drag is the drag of the fuselage, landing gear, struts, and other surfaces exposed to the air. There is a fine point concerning the drag of the wing to be mentioned here that will be elaborated on later. Part of the wing drag contributes to the parasite drag and is sometimes referred to as profile drag. The profile drag is closely equal to the drag of the wing at zero lift; however, it does increase with increasing lift. This increase is therefore usually included as part of the induced drag. In a strict sense this is incorrect, as will become clearer later on.

W is the gross weight of the airplane and, by definition, acts at the center of gravity of the airplane and is directed vertically downward. It is composed of the empty weight of the airplane and its useful load. This latter weight includes the payload (passengers and cargo) and the fuel weight.

The pitching moment, M, is defined as positive in the nose-up direction (clockwise in Figure 1.3) and results from the distribution of aerodynamic forces on the wing, tail, fuselage, engine nacelles, and other surfaces exposed to the flow. Obviously, if the airplane is in trim, the sum of these moments about the center of gravity must be zero.

We know today that the aerodynamic forces on an airplane are the same whether we move the airplane through still air or fix the airplane and move the air past it. In other words, it is the relative motion between the air and airplane and not the absolute motion of either that determines the aerody­namic forces. This statement was not always so obvious. When he learned of the Wright Brothers’ wind tunnel tests, Octave Chanute wrote to them on October 12, 1901 (Ref. 1.1) and referred to “natural wind.” Chanute con­jectured in his letter:

“It seems to me that there may be a difference in the result whether the air is impinged upon by a moving body or whether the wind impinges upon the same body at rest. In the latter case each molecule, being driven from behind, tends to transfer more of its energy to the body than in the former case when the body meets each molecule successively before it has time to react on its neighbors. ”

Fortunately, Wilbur and Orville Wright chose to believe their own wind tunnel results.

Returning to Figure 1.3, we may equate the vector sum of all forces to zero, since the airplane is in equilibrium. Hence, in the direction of flight,

T cos (0 — 0C) — D — W sin 0C = 0 (1.1)

Normal to this direction,

I W cos 6C – L-T sin (0 – вс) = 0

These equations can be solved for the angle of climb to give

In this form, ec appears on both sides of the equation. However, let us assume a priori that вс and (в – вс) are small angles. Also, except for very high performance and V/STOL (vertical or short takeoff and landing) airplanes, the thrust for most airplanes is only a fraction of the weight. Thus, Equation 1.3 becomes

For airplanes propelled by turbojets or rockets, Equation 1.4 is in the form that one would normally use for calculating the angle of climb. However, in the case of airplanes with shaft engines, this equation is modified so that we can deal with power instead of thrust.

First, consider a thrusting propeller that moves a distance S in time t at a constant velocity, V. The work that the propeller performs during this time is, obviously,

work = TS

Power is the rate at which work is performed; hence,

power = T —

But Sit is equal to the velocity of advance of the propeller. Hence the power available from the propeller is given by

P avail = TV (1.5)

Similarly, the power required by a body traveling through the air with a velocity of V and having a drag of D will be

Preqd = DV

Thus, returning to Equation 1.4, by multiplying through by WV, we get

W(Vec) = P a vail — Preq’d (1-6)

The quantity V0C is the vertical rate of climb, Vc. The difference between the power that is required and that available is referred to as the excess power, Pxs – Thus Equation 1.6 shows that the vertical rate of climb can be obtained by equating the excess power to the power required to lift the airplane’s weight at the rate Vc. In operating an airplane this means the following. A pilot is flying at a given speed in steady, level flight with the engine throttle only partially open. If the pilot advances the throttle while maintaining a

constant airspeed, the power from the engine will then be in excess of that required for level flight, and the airplane will climb.

Suppose, instead of keeping it constant, the pilot, while opening the throttle, allows the airspeed to increase in such a manner as to maintain a constant altitude. When a wide open throttle (WOT) condition is reached the maximum power available is equal to the power required. This is the con­dition for maximum airspeed, “straight and level.”

From this brief introduction into airplane performance, it is obvious that we must be able to estimate the aerodynamic forces on the airplane before we can predict its performance. Also, a knowledge of the characteristics of its power plant-propulsor combination is essential.

In addition to performance, the area of “flying qualities” is very im­portant to the acceptance of an airplane by the customer. Flying qualities refers primarily to stability and control, but it also encompasses airplane response to atmospheric disturbances.

Let us briefly consider the pitching moment M, shown in Figure 1.3. This moment, which must be zero for steady, trimmed flight, results mainly from the lift on the wing and tail. In addition, contributions arise from the fuselage, nacelles, propulsor, and distribution of pressure over the wing. Suppose now that the airplane is trimmed in steady, level flight when it is suddenly disturbed (possibly by a gust or an input from the pilot) such that it pitches up by some amount. Before it can respond, the airplane’s path is still essentially horizontal, so that the angle between the velocity vector and the plane’s axis is momentarily increased. It will be shown later that, at a given airspeed, the moment, M, is dependent on this angle, defined previously as the angle of attack. Since the moment was initially zero before the airplane was disturbed, it follows that, in general, it will have some value other than zero due to the increase in angle of attack. Suppose this increment in M is positive. In this case the tendency would then be for the angle of attack to increase even further. This is an unstable situation where the airplane, when disturbed, tends to move even further from its steady-state condition. Thus, for the airplane to exhibit a more favorable, stable response, we desire that the increment in M caused by an angle of attack change be negative.

This is about as far as we can go without considering in detail the generation of aerodynamic forces and moments on an airplane and its components. The preceding discussion has shown the importance of being able to predict these quantities from both performance and flying qualities viewpoints. The following chapters will present detailed analytical and experimental material sufficient to determine the performance and stability and control characteristics of an airplane.

As you study the material to follow, keep in mind that it took the early aviation pioneers a lifetime to accumulate only a fraction of the knowledge that is yours to gain with a few months of study.

The primary system of units to be used in this text is the SI (Systems Internationale) system. Since this system is just now being adopted in the United States, a comparison to the English system is presented in Appendix A. l. Also, to assure familiarity with both systems, a limited number of exercises are given in the English system. For a more complete explanation of the SI system, see Reference 1.4.

PROBLEMS

1.1 Calculate the rate of climb of an airplane having a thrust-to-weight ratio of 0.25 and a lift-to-drag ratio of 15.0 at a forward velocity of 70m/s (230 fps). Express Vc in meters per second. Current practice is to express rate of climb in feet per minute. What would be your answer in these units?

1.2

Which of the systems (ball and track) pictured below are in equilibrium? Which are stable?

1.3 An aircraft weighs 45,000 N (10,117 lb) and requires 597 kW (800 hp) to fly straight and level at a speed of 80 m/s (179 mph). If the available power is 895 kW (1200 hp), how fast will the airplane climb when the throttle is advanced to the wide open position?

1.4 For an aircraft with a high thrust-to-weight ratio, the angle of climb is not necessarily small. In addition, for certain V/STOL aircraft, the thrust vector can be inclined upward significantly with respect to the direction of flight. If this angle is denoted as 0T, show that

T cos 6t — D L + T sin вт

1.5 A student pushes against the side of a building with a force of 6N for a period of 4 hr. How much work was done?

1.6 An aircraft has a lift-to-drag ratio of 15. It is at an altitude of 1500 m (4921 ft) when the engine fails. An airport is 16 km (9.94 miles) ahead. Will the pilot be able to reach it?

REFERENCES

1.1 McFarland, Marvin W., editor, The Papers of Wilbur and Orville Wright, Including the Chanute-Wright Letters, McGraw-Hill, New York, 1953.

1.2 Harris, Sherwood, The First to Fly, Aviation’s Pioneer Days, Simon and Schuster, New York, 1970.

1.3 Robertson, James M., Hydrodynamics in Theory and Application, Pren­tice-Hall, Englewood Cliffs, N. J., 1965.

1.4 Mechtly, E. A., The International System of Units, Physical Constants and Conversion Factors, NASA SP-7012, U. S. Government Printing Office, Washington, D. C., 1969.

1.5 Cleveland, F. A., “Size Effects in Conventional Aircraft Design,” J. of Aircraft, 7(6), November-December 1970 (33rd Wright Brothers Lecture).

TWO

Aeronautics is defined as “the science that treats of the operation of aircraft; also, the ^art or science of operating aircraft.” Basically, with aeronautics, one is concerned with predicting and controlling the forces and moments on an aircraft that is traveling through the atmosphere.

A BRIEF HISTORY

Thursday, December 17, 1903

“When we got up a wind of between 20 and 25 miles was blowing from the north. We got the machine out early and put out the signal for the men at the station. Before we were quite ready, John T. Daniels, W. S. Dough, A. D. Etheridge, W. C. Brinkly of Manteo, and Johnny Moore of Nags Head arrived. After running the engine and propellers a few minutes to get them in working order, I got on the machine at 10:35 for the first trial. The wind, according to our anemometers at this time, was blowing a little over 20 miles (corrected) 27 miles according to the government anemometer at Kitty Hawk. On slipping the rope the machine started off increasing in speed to probably 7 or 8 miles. The machine lifted [гот*Щр truck just as it was entering the fourth rail. Mr. Daniels took a picture just as it left the tracks. I found the control of the front rudder quite difficult on account of its being balanced too near the center and thus had a tendency to turn itself when started so that the rudder was turned too far on one side and then too far on the other. As a result the machine would rise suddenly to about 10ft. and then as suddenly, on turning the rudder, dart for the ground. A sudden dart when out about 100 feet from the end of the tracks ended the flight. Time about 12 seconds (not known exactly as watch was not promptly stopped). The level for throwing off the engine was broken, and the skid under the rudder cracked. After repairs, at 20 min. after 11 o’clock Will made the second trial.”

The above, taken from Orville Wright’s diary, as reported in Reference

1.1, describes mankind’s first sustained, controlled, powered flight in a

heavier-than-air machine. The photograph, mentioned by Orville Wright, is shown here as Figure 1.1. Three more flights were made that morning. The last one, by Wilbur Wright, began just at 12 o’clock and covered 260 m in 59 s. Shortly after fhis flight, a strong gust of wind struck the airplane, turning it over and over. Although the machine was severely damaged and never flew again, the Wright Brothers achieved their goal, begun approximately 4yr earlier.

Their success was no stroke of luck. The Wright Brothers were painstak­ing in their research and confident of their own results. They built their own wind tunnel and tested, in a methodical manner, hundreds of different airfoil and wing planform shapes. They were anything but a “couple of bicycle mechanics.” Their letters to Octave Chanute, a respected civil engineer and aviation enthusiast of the day, reveal the Wright Brothers to have been learned men well versed in basic concepts such as work, energy, statics, and dynamics. A three-view drawing of their first airplane is presented in Figure 1.2.

On September 18, 1901, Wilbur Wright was invited to deliver a lecture before the Western Society of Engineers at a meeting in Chicago, Illinois. Among the conclusions reached by him in that paper were:

1. “That the ratio of drift to lift in well-shaped surfaces is less at angles of incidence of five degrees to 12 degrees than at an angle of three degrees. ” (“Drift” is what we now call “drag.”)

2. “That in arched surfaces the center of pressure at 90 degrees is near the center of the surface, but moves slowly forward as the angle becomes less, till a critical angle varying with the shape and depth of the curve is reached, after which it moves rapidly toward the rear till the angle of no lift is found.”

3. “That a pair of superposed, or tandem surfaces, has less lift in proportion to drift than either surface separately, even after making allowance for weight and head resistance of the connections. ”

These statements and other remarks (see Ref. 1.1) show that the Wright Brothers had a good understanding of wing and airfoil behavior well beyond that of other experimenters of the time.

Following their first successful flights at Kitty Hawk, North Carolina, in 1903, the Wright Brothers returned to their home in Dayton, Ohio. Two years later they were making flights there, almost routinely, in excess of 30 km and 30 min while others were still trying to get off the ground.

Most of the success of the Wright Brothers must be attributed to their own research, which utilized their wind tunnel and numerous experiments with controlled kites and gliders. However, their work was built, to some degree, on the gliding experiments of Otto Lilienthal and Octave Chanute. Beginning in 1891, Lilienthal, working near Berlin, Germany, made ap­proximately 2000 gliding flights over a 5-yr period. Based on measurements obtained from these experiments, he published tables of lift and drag measurements on which the Wright Brothers based their early designs. Unfortunately, Lilienthal had no means of providing direct aerodynamic control to his gliders and relied instead on kinesthetic control, whereby he shifted his weight fore and aft and side to side. On August 9, 18%, as the result of a gust, Otto Lilienthal lost control and crashed from an altitude of approximately 15 m. He died the next day. During 1896 and 1897, Octave Chanute, inspired by Lilienthal’s work, designed and built several gliders that were flown by others near Miller, Indiana. Chanute recognized Lilienthal’s control problems and was attempting to achieve an “automatic” stability in his designs. Chanute’s principal contribution was the addition of both vertical and horizontal stabilizing tail surfaces. In addition, he went to the “box,” or biplane, configuration for added strength. Unfortunately, he also relied on kinesthetic control.

When the Wright Brothers began their gliding experiments in the fall of 1900, they realized that adequate control about all three axes was one of the major prerequisites to successful flight. To provide pitch control (i. e., nose up or down), they resorted to an all-movable horizontal tail mounted in front of

the wing. Yaw control (i. e., turning to the left or right) was accomplished by means of an all-movable vertical tail mounted behind the wing. Their method of roll control (i. e., lowering one side of the wing and raising the other) was not as obvious from photographs as the controls about the other two axes. Here, the Wright Brothers devised a means of warping their “box” wing so that the angle of incidence was increased on one side and decreased on the other. The vertical tail, or rudder, was connected to the wing-warping wires so as to produce what pilots refer to today as a coordinated turn.

The Wright Brothers were well ahead of all other aviation enthusiasts of their era. In fact, it was not until 3 yr after their first flight that a similar capability was demonstrated, this by Charles and Gabriel Voisin in Paris, France (Ref. 1.2). On March 30, 1907, Charles Voisin made a controlled flight of approximately 100 m in an airplane similar in appearance to the Wright flyer. A second machine built by the Voisin Brothers for Henri Farman, a bicycle and automobile racer, was flown by Farman later that year on flights that exceeded 2000 m. By the end of that year at least five others succeeded in following the Wright Brothers’ lead, and aviation was on its way.

Today we are able to explain the results of the early experimenters in a very rational way by applying well-established aerodynamic principles that have evolved over the years from both analysis and experimentation. These developments have their beginnings with Sir Isaac Newton, who has been called the first real fluid mechanician (Ref. 1.3). In 1687 Newton, who is probably best known for his work in solid mechanics, reasoned that the resistance of a body moving through a fluid is proportional to the fluid density, the velocity squared, and the area of the body.

Newton also postulated the shear force in a viscous fluid to be propor­tional to the velocity gradient. Today, any fluid obeying this relationship is referred to as a Newtonian fluid.

In 1738, Daniel Bernoulli, a Swiss mathematician, published his treatise, “Hydrodynamics,” which was follov^l in 1743 by a similar work produced by his father, John Bernoulli. The Bemoullis made important contributions to understanding the behavior of fluids. In particular, John introduced the concept of internal pressure, and he was probably the first to apply momen – ( turn principles to infinitesimal fluid elements.

I Leonhard Euler, another Swiss mathematician, first put the science of hydrodynamics on a firm mathematical base. Around 1755, Euler properly ^ formulated the equations of motion based on Newtonian mechanics and the f works of John and Daniel Bernoulli. It was he who first derived along a } streamline the relationship that we refer to today as “Bernoulli’s equation.”

The aerodynamic theories of the 1800s and early 1900s developed from f’, the early works of these mathematicians. In 1894 the English engineer, Frederick William Lanchester, developed a theory to predict the aerodynamic I; behavior of wings. Unfortunately, this work was not made generally known

Table 1.1 (continued)

Source. From F. A. Cleveland, “Size Effects in Conventional Aircraft,” J. of Aircraft, 7(6), November-December 1970 (33rd Wright Brothers Lecture). Reproduced with permission.

“ Counting original(s), subsequent series production, and derivatives—if any. k Counting pilot (Orville Wright) and 5 lb of fuel. c Destroyed in taxi-test which resulted in unintended liftoff. d Set world record 21 Oct. 1929 with 169 onboard.

‘ One ANT-20 is built with six 1100-hp engines.

1 Turbine energy expressed in terms of gas-hp with 0.8 efficiency.

* SP = in series production.

k Used mainly as freighter: 724-seat stretched version projected.

1 Triple deck version.

’ Two in Germany; six in Japan.

* Flew only once on high-speed taxi test.

until 1907 in a book published by Lanchester. By then the Wright Brothers had been flying for 3 yr. Much of the knowledge that they had laboriously deduced from experiment could have been reasoned from Lanchester’s theory. In 1894, Lanchester completed an analysis of airplane stability that could also have been of value to the Wrights. Again, this work was not published until 1908.

Lanchester’s wing theory was somewhat intuitive in its development. In 1918 Ludwig Prandtl, a German professor of mechanics, presented a mathe­matical formulation of three-dimensional wing theory; today both men are credited with this accomplishment. Prandtl also made another important contribution to the science with his formalized boundary layer concept.

Around 1917 Nikolai Ergorovich Joukowski (the spelling has been angli­cized), a Russian professor of rational mechanics and aerodynamics in Moscow, published a series of lectures on hydrodynamics in which the behavior of a family of airfoils was investigated analytically.

The work of these early hydro – and aerodynamicists contributed little, if any, to the progress and ultimate success of those struggling to fly. However, it was the analytical base laid by Euler and those who followed him on which the rapid progress in aviation was built.

After 1908, the list of aviators, engineers, and scientists contributing to the development of aviation grew rapidly. Quantum improvements were accomplished with the use of flaps, retractable gear, the cantilevered wing, all-metal construction, and the turbojet engine. This impressive growth is documented in Table 1.1. Note that in less than 10 yr from the Wright Brothers’ first flight, the useful load increased from 667 N (150 lb) to more than 13,300 N (3000 lb). In the next 10 yr, the useful load increased by a factor of 10 and today is more than 1.78 x 106 N (400,000 lb) for the Lockheed C-5A.

Our state of knowledge is now such that one can predict with some certainty the performance of an airplane before it is ever flown. Where analytical or numerical techniques are insufficient, sophisticated experimental facilities are utilized to investigate areas such as high-lift devices, complicated three-dimensional flows in turbomachinery, and aerothermodynamics.

An airplane, whether it is a sleek, modern jet or a Piper Cub, is a thing of beauty, supported almost miraculously by the invisible medium through which it travels. As an aeronautical engineer and a pilot, the beauty of flight is a pleasure I hope always to enjoy. One of my favourite poems is “High Flight,” by John Gillespie Magee, Jr., a young RAF pilot killed during World War II. Although it is unusual to include a poem in a preface, I must break convention because this poem expresses my thoughts better than I can.

Oh, I have slipped the surly bonds of earth And danced the skies on laughter-silvered wings;

Sunward I’ve climbed, and joined the tumbling mirth _

Of sun-split cloudffland done a hundred things

You have not dreamed of—wheeled and soared and swung

High in the sunlit silence. Hov’ring there,

I’ve chased the shouting wind along, and flung My eager craft through footless halls of air.

Up, up the long, delirious, burning blue I’ve topped the windswept heights with easy grace

Where never lark, or even eagle flew.

And, while with silent, lifting mind I’ve trod

The high untrespassed sanctity of space,

Put out my hand, and touched the face of God.

^ This book explains the many technical aspects of flight; the title reflects $8 contents. Beginning with the fundamentals of incompressible and com – gpssible flows, aerodynamic principles relating to lift, drag, and thrust are |p&veloped. Two-dimensional airfoils and three-dimensional wings with high low aspect ratios are treated for subsonic and supersonic flows.

The operating characteristics of different types of aircraft power plants, including normally aspirated and supercharged-piston engines, turboprops, turbojets, and turbofans, are presented in some detail. Typical operating curves are given for specific engines within these classes. A fairly complete treatment of propellers is also included.

This information, which is necessary for estimating the aerodynamic characteristics of an airplane, is followed by a presentation of methods for calculating airplane performance; items such as takeoff distance, climb rates and times, range payload curves, and landing distances, are discussed.

Finally, the subjects of longitudinal and lateral-directional stability and control, both static and dynamic, are introduced. Because this book is intended for use in a first course in these subjects, it covers only open-loop control.

Features of this textbook include material on the prop-fan, winglets, high lift devices, cooling and trim drag, the latest NASA airfoils, criteria for providing satisfactory open-loop flying qualities, and the use of the SI and English systems of units. Practical examples are given for most developments and, in many cases, are applied to currently operating airplane-engine com­binations. Some numerical treatments of aerodynamic problems and the use of the analog computer for examining longitudinal and lateral dynamic behavior are also introduced. Considerable data are provided relating to lift, drag, and thrust predictions. Stability and control data and performance data on a number of presently operating aircraft are found throughout the book.

Since each subject is developed from first principles, it can be used as a text for a first course in aerodynamics at the junior level. It also can be used in successive courses in compressible flow, airplane performance, and stabil­ity and control as either the primary text or as a reference. At least half of the material in the book is at the senior level of most aerospace engineering programs.

Obviously a book of this nature could not have been written without the help and cooperation of many individuals and organizations. I especially thank Pratt & Whitney and Pratt & Whitney of Canada for the performance curves on their engines; Avco-Lycoming for information on cooling drag and reciprocating engines; Cessna for the data pertaining to the performance of the Citation I; and Piper for their support of this effort in many ways. My friends and acquaintances with these fine companies who helped personally are too numerous to mention. They know who they are, and I am grateful to them.

I am indebted to my secretaries, Charlotte Weldon, and Sharon Symanovich, who typed the manuscript, for a job well done. Their correction of mistakes and their patience with my penmanship and the other trying aspects of preparing the manuscript are sincerely appreciated.

I prepared the manuscript on evenings, weekends and, when I could find the time, during the day at The Pennsylvania State University. This meant many lonely evenings and weekends for my wife, Emily, so I offer my appreciation for her indulgence, patience, and understanding.

Barnes W. McCormick, Jr.