C H A PTE R 7 Confi g u ration Layout a n d Loft 1 71

Component Su rface
Fuselage 70344 . 8
Ve rt tail 26 1 65 . 3
Wing 1 02 6 3 6 . 7
Circular arc canopy 907 1 . 4
Nacelle 25462.9
Total 2 3 3 6 8 1 .0

I \
I \
750. I
'

\
/ '
/
'

I I
p '
\

I'°\ \
e "
500.

I "
r

m
I
['\. h
c
I '/ '' ' ./
....__
l
/ /
/
250. �
'
/
---- -
/
e
/
I "
l£.. \
v ,....-- -1-� r-- _ "- I
0. I/ / ....__ - ' I
0. 200. 400. 600.

Fuselage statjons

Fig. 7.4 Wetted area plot.

Component Vol ume

Fuselage 847 1 24.4
Vert tail 42903.5
Wing 287005.5
Circ u l ar arc canopy 460 1 4.0
Nacelle 95 1 49 . 8
Tota l 1 3 1 8 1 96 . 8
c
/ �
0 4000. "
I I\
f-- / \
I\
/ \
·-� /
s
3000. --
-- \
e
c
'
t '
' -..__
2000.
-�
i /
- --
0
/ f '
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/ / I
n / I ' '-
1 000. ,r
I/ / '
a / '
/ ,.... "- -- -
/ ,.
_,.
/
/ ,... -
,
a
0.
0. 200. 400. 600.

Fuselage stat.ions

Fig. 7 .5 Volume distribution plot.

1 72 A i r c raft Desi g n : A C o n c e p t u a l A p p r o a c h
 C HAPT E R 7 Confi g u ration Layout a n d Loft 1 73

- -®
,.....-t---i�i-'+-f---rr-""11"1.--- ---- - - - (�

-®
�'*"Jl!l'=t--- --- -·J

�
-eo._
0
-e
0
..Cl
c:

·�
0
c
·-- ---®

l!)
cl..
 CHAPT E R 7 Config u ration Layout a n d Loft 1 75

layout work, but the designers still must obtain and model the detailed
geo metry of the aircraft's internal components and subsystems. Such
CAD systems do not require a separate lines control drawing. Instead, at
th e appropriate phase during design, a production-quality "solid model"
geo metry that defines the aircraft surfaces in great detail and accuracy will
be prepared. This, too, is a far-from-trivial task.
After the inboard profile drawing has been prepared, an "inboard
isometric" drawing (Fig. 7.8) can be prepared. These are usually prepared
by the art group for illustration only, used in briefings and proposals.
Isometrics are often published by aviation magazines, and theirs are
usually better than those prepared by the aircraft companies!

Con ic Lofting
"Lofting" is the process of defining the external geometry of the aircraft.
The word itself apparently comes from old shipyards, where the drawings
would be made in a loft over the worl�shop. "Those drawings made in the
loft" became the "loft" of the ship.
"Production lofting," the most detailed form of lofting, provides an exact,
mathematical definition of the entire aircraft including such minor details
as the intake and exhaust ducts for the air conditioning. A production-loft
definition is expected to be accurate to within a few hundredths of an inch
(or less) over the entire aircraft. This allows the different parts of the aircraft
to be designed and fabricated at different plant sites yet fit together perfectly
during final assembly.
Most aircraft companies now use computer-aided design systems that
incorporate methods discussed in [22l . These systems are so accurate that
different parts of the aircraft can be designed and built in different locations,
yet will fit together perfectly.
For an initial layout it is not necessary to go into as much detail. However,
the overall lofting of the fuselage, wing, tails, and nacelles must be defined
sufficiently to show that these major components will properly enclose the
required internal components and fuel tanks while providing a smooth
aerodynamic contour.
Lofting for ship hulls was done using enormous drawings. To provide a
smooth longitudinal contour, points taken from the desired cross sections
were connected longitudinally on the drawing by flexible "splines," long,
thin wood or plastic rulers held down at certain points by lead "ducks"
(pointed weights-see Fig. 7.9).
This technique was used for early aircraft lofting but suffers from two
disadvantages. First, it requires a lot of trial and error to achieve a smooth
surface both in cross section and longitudinally. Second, and perhaps
more important, this method does not provide a unique mathematical
definition of the surface. To create a new cross section requires a
tremendous amount of drafting effort, especially for a canted cross section
1 76 A i rc raft Desig n : A C o n c e pt u a l A p p ro a c h

Lead "d ucks"

Fig. 7 . 9 Spline lofting.

(i.e., a cross-sectional cut at some angle other than perpendicular to the

centerline of the aircraft). In addition to the time involved, this method is
prone to mismatch errors. Quite simply, there is too much "art" involved.
A new method of lofting was developed and used for the first time on the
North American Aviation P-51 Mustang. l23l This method, now considered
traditional, is based upon a mathematical curve form known as the "conic."
The great advantage of the conic method is that it is in fact a mathematically
defined curve, so that it can be plotted with great accuracy for production
lofting, but it is also easy to construct on the drafting table.
Up until roughly the 1 980s, conic lofting was used almost exclusively
for aircraft design. Today, modern three-dimensional CAD systems use
more sophisticated mathematical curves that are plotted on the screen
by laborious point-at-a-time calculations-computers are good at that.
However, this author firmly believes that a good understanding of traditional
conic lofting is a necessary foundation for understanding the process of
aircraft surface design and will help the designer learn to properly use even
the best of modern CAD systems.
A conic is a second-degree curve whose family includes the circle, ellipse,
parabola, and hyperbola. The conic is best visualized as a slanted cut through
a right circular cone (Fig. 7.10). The shape of the conic depends upon the
angle of the cut through the cone. If the cut is flat (i.e., perpendicular to
the axis of the cone), then the resulting curve will be a circle; if somewhat
slanted, an ellipse; if exactly parallel to the opposite side, a parabola. A
greater cut angle yields a hyperbola.
The generalized mathematical form of the conic is given in Eq. (7. 1). This
form of the equation is never used directly. A number of specialized conic
equations are provided in [22l .
 C HA P T E R 7 Configu ration Layout a n d loft 1 77

On a drafting table, the conic curve is constructed from the desired start
an end points (A and B) and the desired tangent angles at those points.
d
These tangent angles intersect at point C. The shape of the conic between
the points A and B is defined by some shoulder point S. (The points
labeled E in Fig. 7.10 are a special type of shoulder point, discussed later.)
figure 7. 1 1 illustrates the rapid graphical layout of a conic curve.
The first illustration in Fig. 7. 1 1 shows the given points A, B, C, and S.
In the second illustration, lines have been drawn from A and B, passing
through S.
The remaining illustrations show the generation of one point on the
conic. In the third illustration a line is drawn from point C at an arbitrary
angle. Note the points where this line intersects the A-S and B-S lines.
Lines are now drawn from A and B through the points found in the last
step. The intersection of these lines is a point P, which is on the desired
conic curve.

E l l i pse Pa ra bola Hyperbola

b \
I

- - -I -�1t,
I
£ 1
u
\

Fig. 7 . 1 0 Conic geometry definition.

1 78 A i rc raft D e s i g n : A C o n c e p t u a l A p p roa c h

A c A c

s
•

G iven A, B, C, S
S is any s h o u l d e r point
B B

A c A c

I I
', " p
I I - - --.. _
I I

B B
I I

Fig. 7 . 1 1 Conic loyout.

B
3

Fig. 7 . 1 2 Conic layout example.

 C H A PT E R 7 Confi g u ration Layout and Loft 1 79

To generate additional points, the last two steps are repeated. Another
C
line is drawn from point at another arbitrary angle, and then the lines
from A and B are drawn and their intersection is found. When enough
points have been generated, a French curve is used to draw the conic.
While this procedure seems complicated at first, with a little practice a
good designer can construct an accurate conic in less than a minute.
Figure 7. 12 illustrates a conic curve generated in this manner. Note that
it is not necessary to completely draw the various lines, as it is only their
intersections that are of interest.

Conic Fuselage Development

To create a smoothly lofted fuselage using conics, the points A, B, C,

and S
in each of the various cross sections are connected longitudinally by smooth
lines. Figure 7.13 shows the upper half of a simple fuselage, in which the A, B,
c, and s points in three cross sections are connected by smooth longitudinal
lines. These are called "longitudinal control lines" because they control the
shapes of the conic cross sections.
Figure 7.14 shows the side and top views of these longitudinal control
lines. Because the cross sections are tangent to horizontal at the top of the
fuselage, the A C
and lines are identical in side view. Similarly, the cross

Fig. 7 . 1 3 Longitudinal control lines.

1 80 A i rc raft Des i g n : A C o n c e p t u a l A p proa c h

B, C
s

Fig. 7 . 1 4 Cross-section development from longitudinal control lines.

sections are tangent to vertical at the side of the fuselage, so that the B and C
lines are identical in top view. This is common, but not required.
In Fig. 7. 14, the longitudinal control lines are used to create a new cross
section, in between the second and third cross sections already defined. This
new cross section is created by measuring, from the longitudinal control
lines, the positions of the A,
B, C, and S points at the desired location of
the new cross section.
A,
As is shown for point each point is defined by two measurements, one
from side view and one from top view. From these points the new cross
section can be drawn using the conic layout procedure illustrated in Fig. 7. 1 1 .
The original cross sections that are used to develop the longitudinal
control lines are called the "control cross sections" or "control stations."
These cross sections are drawn to enclose the various internal components,
such as the cockpit or engine.
Control stations can also be drawn to match some required shape. For
example, the last cross section of a single-engine jet fighter with a conven­
tional round nozzle would have to be a circle of the diameter of the nozzle.
Typically, some 5 - 10 control stations will be required to develop a
fuselage that meets all geometric requirements. The remaining cross sec­
tions of the fuselage can then be drawn from the longitudinal control lines
developed from these control stations.
 CHAPTE R 7 Confi g u ration Layout a n d Loft 181

-8 Fuselage Lofting Example

Figure 7.15 illustrates a common application of conic lofting to define
a fighter fuselage for an initial layout. Five control stations are required
fo r this example. Station 0 is the nose, which is a single point. All of the
longitudinal control lines must originate there.
Station 120 is established for this example by the requirements for the
cockpit (Chapter 9). This station is approximately circular in shape and is
defined using two conics (upper and lower) . Each conic has its own A, B,
C, B
and S points. Note that the (end) point of the upper conic is identical
A
to the (start) point of the lower conic.
Station 240 has a flat side to provide for a side-mounted inlet as can be
seen on the F-4, the MiG-23, the SAAB Gripen, and many other aircraft.
At this station, the end points of the upper and lower conics are moved
apart vertically, with the area between them defined as a straight line. Note
in side view that the longitudinal control lines separate smoothly, not
suddenly. This is to ensure a smooth l c:mgitudinal contour.
Station 370 is similar to station 240, with a relatively square cross­
sectional shape. This could allow room for the landing gear or perhaps to
attach a low wing to the side of the fuselage, without a drag-producing
acute angle.

0 1 20 240 370 500

:;�;''"'l
Fuselage

� 290

�-[--------[-�-1
I

Fig. 7 . 1 5 Typical fuselage lofting.

1 82 A i rc raft Des i g n : A C o n c e p t u a l A p p roa c h

Station 500 is a circular cross section, to allow for a connection with a

round exhaust nozzle. The longitudinal control lines come back together
in a smooth fashion, as shown.
These five control stations are then used to create the longitudinal
control lines. From those lines, additional cross sections can be created as
desired. Section 290 was created in such a fashion, by measuring the conic
control points from the longitudinal control lines and then drawing the
conics as described earlier.
Figure 7. 1 5 shows only the fuselage lofting. The canopy, inlet duct, and
inlet duct fairing would be lofted in a similar fashion, using longitudinal
control lines through a few control stations.

MIJ Conic Shape Parameter

One problem arises with this method of initial lofting. The locations of
the shoulder points S can be difficult to control, creating conics either too
square (shoulder point too close to point C ) or too flat (shoulder point too
far away from point C ). An alternate technique using conics involves a par­
ameter that directly controls the shoulder point's distance from the point C.
The points labeled E in Fig. 7.10 are conic shoulder points that happen
to lie upon the line D-C. D is the point exactly midway between A and B.
Such a shoulder point E determines the conic shape parameter p, as
defined in the following equation:
p= I DE l / I D C I (7.2)
where

IAD I = I BD I (7.3)
Referring to Fig. 7.10, the shoulder points labeled E are based upon the p
values required to obtain the ellipse, parabola, or hyperbola forms of the
conic. These are given below, along with the p value that defines a circle
(a special form of the ellipse):
Hyperbola:
p > 0.5
Parabola:
p = 0.5
Ellipse:
p < 0.5
Circle:

p = 0.4142 and IA C I = I B C I (7.4)

 C HAP T E R 7 Config u ration Layout a n d Loft 1 83

The conic shape parameter allows the designer to specify the conic
c urve's C.
distance from the point A conic with a large p value (approaching
1.0) will be nearly square, with the shoulder point almost touching the
p C.
oint A conic with a small p value (approaching 0.0) will nearly resemble
the straight line from A-B. The parameter p can be used to control the
longitudinal fairing of a fuselage more easily.
Figure 7.16 shows the use of the conic shape parameter p to lay out a
conic. Points A, B, and C are known, but the shoulder point S is not
known. However, the value of p is given.
In the illustration on the right side of Fig. 7.16, the line has been A-B
drawn and bisected to find the point D.
The shoulder point S is found by
measuring along line D- C, starting at D,
by a distance equal to p times the
total length of line D- C. Once the shoulder point is found, the conic can
be drawn as illustrated in Fig. 7. 1 1 .
By using this approach, a fuselage can b e lofted without the use o f a
longitudinal control line to control the location of the shoulder points. If p
is specified to be some constant value (or all of the cross sections, then the
designer need only control the conic endpoints and tangent intersection
points. To permit the fuselage ends to be circular in shape, the value of p
would be fixed at 0.4142.
Greater flexibility can be attained by allowing p to vary longitudinally. For
example, the fuselage of Fig. 7.15 requires a p value of 0.4142 at both ends to
allow a circular shape, but the values of p at the middle of the fuselage are
higher, perhaps around 0.7.
An "auxiliary control line" can be used to coi;itrol the value of p graphi­
cally, as shown in Fig. 7.17. Note the auxiliary control line for p at the

Given control
c
poi nts a n d p

B
A

A I DS I = p l D C I
I AD I = I DB I

Fig. 7 . 1 6 Conic layout using p.

1 84 Ai rc raft D e s i g n : A C o n c e p t u a l A p p ro a c h

bottom. If the value of p varies smoothly from nose to tail, and the conic
endpoints and tangent intersection point are controlled with smooth
longitudinal lines, then the resulting fuselage surface will be smooth.
In Fig. 7.17 the upper conic has a constant p value of 0.4142, while the
lower conic has a p value varying from 0.4142 at the nose and tail to about
0.6 at the middle of the fuselage. This has the effect of "squaring" the
lower fuselage to provide more room for the landing gear.
Figure 7.18 shows the use of p to develop the cross sections labeled A
and B. Observe the development of the upper and lower conics by the
method shown in Fig. 7.16 and the use of different p values for the upper
and lower conics.
Thus far, no mention has been made of the method for developing the
longitudinal control lines and auxiliary control lines. During production
lofting, these control lines would be defined mathematically, using conics
or some form of polynomial.
For initial layouts, sufficient accuracy can be obtained graphically through
the use of the flexible splines discussed earlier. Points are taken from the

Top view

(Radome) Section A Section B (Nozzle)

Side view

-1 I

---\ - 1
- Lower conic
p _
/
0.5

r
p 0 0.4142

-� p-Control l i nes
- - - --
- -

U pper conic

Fig. 7 . 1 7 Conic fuselage development using p.

 CHAPTER 7 Configu ration Layout a n d Loft 1 85

p = 0.4142 p = 0.4142

p = 0.595 p = 0.610

Section A Section B

Fig. 7 . 1 8 Cross-section development using p.

control cross sections and plotted in side and top view and then connected
longitudinally using a spline to draft a smooth line. In fact, a designer with
a "good eye" can obtain sufficient smoothness using a French curve if
spline and ducks are not available.
Figure 7.19 shows an illustrative example of the conic-developed loft
lines for an exotically shaped aircraft, the sup�rsonic SAAB J-35 Draken
(Dragon). In this isometric view you can see the longitudinal control
scheme for fuselage, nacelle, canopy, and inlet duct, and you can also see
the lines definition for wing and tail. Such a detailed loft definition is not
normally done until sometime in preliminary design. But, a good designer
will consider the overall loft definition even from the earliest conceptual
design layout.

Flat-Wrap Fuselage Lofting

An important cost driver for aircraft fabrication is the amount of
compound curvature used in lofting the aircraft. Compound curvature
implies the existence of surface curvature in all directions for some point
on the surface.
For example, a ball is entirely composed of compound-curvature surfaces.
A flat sheet has no curvature, compound or otherwise. A cylinder is curved,
b ut only in one direction, so it does not have any compound curvature.
Instead, a cylinder or any other surface with curvature in only one direction
is said to be "flat-wrapped."
If a surface is flat-wrapped, it can be constructed by "wrapping" a flat
sheet around its cross sections. This is mathematically known as a
1 86 A i rc raft Des ig n : A C o n c e p t u a l A p p r o a c h

;1ii
!flt! !ft! 1t1 11r 1r:n:
. I • 1 /Jll!ilf!f!!H!1

>,
... > ,.,,.,. -.,_>l.;:..
� .. ....

Fig. 7 . 1 9 Isometric view of SAAB Draken major loft lines (courtesy SAAB Aircraft).

"developable surface" and is not necessarily the same as the "ruled surface"
available on most CAD systems.
For aircraft fabrication, flat-wrap lofting allows the skins to be cut from
flat sheets and bent to the desired skin contours. This is far cheaper than
the construction technique for a surface with compound curvature.
Compound curvature requires that the skins be shaped by a stretching or
stamping operation, which entails expensive tools and extra fabrication
steps.
 C H A PT E R 7 Confi g u ration Layout a n d Loft 1 87

The advantage of flat wrap was seen during the design and fabrication of
th e X-31 Enhanced Fighter Maneuver demonstrator. Rockwell's manufactur­
ing personnel pointed out a problem: the compound curves of the aft fuselage
would require hot die forming. Because the material around the engine was
titanium, the die itself would cost about $400,000 (1999 dollars) and be the
pacing item in the fabrication schedule. By changing the last 30 in. {76 cm}
of the aft fuselage to a flat-wrap loft, titanium sheet could be bent to shape
with no forming required.
Aircraft applications of flat-wrap lofting must be defined in the initial loft
definition used for the conceptual layout. There are several ways of lofting a
surface so that it is flat-wrapped. The simplest technique uses a constant
cross section. For example, a commercial airliner usually has the identical
circular-cross-sectional shape over most of its length. In fact, any cross­
section shape will produce a flat-wrap surface if it is held constant in the
longitudinal direction.
If the same cross-sectional shape is maintained but linearly scaled in size,
a flat-wrap contour is produced. For example, a cone is a flat-wrap surface
produced by linearly scaling a circular cross section.
Many aircraft have a tailcone that, although not circular in cross section,
is linearly scaled to produce a flat-wrap surface. This can be accomplished
with conics by maintaining identical tangent angles and p value, using
straight longitudinal control lines, and maintaining the lengths AC and BC
in constant proportion.
Sometimes it is necessary to vary the shape of the cross sections other
than by scaling. Flat wrap cannot be exactly maintained in such cases using
conics. A more sophisticated technique (beyond the scope of this book)
must be used.
However, flat wrap can be closely approximated in most such cases on
two conditions. First, the longitudinal control lines must be straight. This
includes the line controlling the shoulder point S. If the conic shape
parameter p is used instead of a shoulder-point control line, then the p
value must be either constant or linearly varied. Second, the tangent angles
of the conics must not change longitudinally. If the tangent angles are all
either horizontal or vertical, as in Figs. 7.15 and 7.17, this condition can
easily be met.
Figure 7.20 shows such a complex flat-wrapped surface. The fuselage is
defined by five conics plus a straight-line, flat underside. The "bump" on
top could represent the back of the canopy and grows smaller toward the
rear of the fuselage. While the conics change shape and size, their endpoints
hold the same tangent angles.
The use of flat-wrap lofting for a fuselage represents a compromise.
While flat-wrap surfaces are easier and cheaper to fabricate, they are less
desirable from an aerodynamic viewpoint. For example, a smoothly con­
toured teardrop shape will have less drag than a flat-wrap cylinder with a
nosecone and tailcone.
1 88 A i r c ra ft Desig n : A C o n ceptua l A p p ro a c h

Fig. 7 .20 Complex flat-wrapped surface.

Circle-to-Square Adapter
A common problem in lofting is the "circle-to-square adapter." For
example, the inlet duct of many supersonic j et aircraft is approximately
square at the air inlet, yet must attain a circular shape at the engine front-face.
Modern, two-dimensional nozzles also require a circle-to-square adapter.
Flat-wrap can be attained for a circle-to-square adapter by constructing
the adapter of interlocking, V-shaped segments, each of which is itself
flat-wrapped (Fig. 7.21). The flat sides of the square section taper to points

Section A-A

Fig. 7 . 2 1 Circle-to-square adopter.

 CHAPTER 7 Confi g u ration Layout and Loft 1 89

that just touch the circular section. Similarly, the cone-shaped sides of the
circular section taper to points that touch the corners of the square
section. Note the "rounded-off square" shape of the intermediate sections.
The connecting surfaces must be straight longitudinally for a flat-wrap
surface to be maintained.

Loft Verification via Buttock-Plane Cuts

If conic lofting is properly done with smooth longitudinal control
lines, the resulting shape should be smooth. Sometimes, though, it is wise
to check the surface contours. Perhaps part of the fuselage is flat-wrapped
or oddly shaped for some other reason, or perhaps a designer is asked to
evaluate a design created by someone else.
Surface contours can be visualized and evaluated using an ancient
technique borrowed from shipbuilding. Hull contours are checked for
smoothness by laying out the "waterlines." If a ship is floating in the water,
the line around the hull where the surface of the water intersects the hull
is a waterline. For good ship performance, this waterline should be smooth
in the longitudinal direction.
If the hull is raised partly out of the water some arbitrary distance, a new
waterline is formed. Hull designers check for hull smoothness by laying
out a large number of these waterlines, each separated in height by some
arbitrary distance. If all of the waterlines have smooth contours, then the
hull is smooth.
Such horizontal waterline cuts can be used for evaluation of the smooth­
ness of an aircraft fuselage, but it is more common to use vertically oriented
cuts known as "buttock-plane cuts" (Fig. 7.22).
Buttock-plane ("butt-plane") cuts form the intersection of the aircraft
with vertical planes defined by their distance from the aircraft centerline.

Airfoil is a butt-plane cut of the wing

Fig. 7 .22 Buttock-plane cut.

1 90 Ai rcraft D e s i g n : A Conceptu a l Approach

Cut fuselage with vertical planes

p a ra l l e l to centerl i n e

cl

I
I I I
I
Top view
I I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I I
Cross- I
I I section I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I I I
I

FRP

Fig. 7 .23 Buttock-plane cut layout.

For example, "butt-plane 30" is the contour created by intersecting a vertical

plane with the fuselage at a distance of 30 in. from the centerline. Note
in Fig. 7.22 that the butt-plane cuts are oriented such that the airfoil is a
butt-plane cut of the wing. It is for this reason that butt-plane cuts are
more commonly used for aircraft than waterlines.
Figure 7.23 illustrates the development of butt-plane cuts. Vertical lines
are drawn on each cross section, indicating the locations of the arbitrarily
selected butt-planes. The points where these vertical lines intersect the
cross sections are transferred to the side-view drawing and connected
longitudinally. If the fuselage surface is smooth, then these longitudinal
lines for the different butt-planes will all be smooth.
Buttock-plane cuts can also be used to generate new cross sections. Once
the butt-plane cuts are developed as in Fig. 7.23, a new cross section can be
developed by transferring the vertical locations of the butt-plane cuts to the
cross section desired and then drawing a smooth cross-sectional contour
using those points.
 C HAPTE R 7 Confi g u ration Layout a n d Loft 191

Sometimes this archaic method is actually easier than developing the

longitudinal control lines for conic fuselage lofting. This is most likely
when the surface is highly irregular, such as the forebody of a blended wing­
b ody aircraft like the B- lB.

Wi ng/Ta i l Layout and Loft

� Reference Wing /Tai l Layout
Chapter 4 described the selection of the basic geometric parameters for
the wing and tails. These parameters include the aspect ratio taper ratio A,
A, sweep, dihedral, and thickness. Also, the selection of an appropriate
airfo il was considered. In Chapter 6, the actual sizes for the wing, tails, and
fuselage were defined, based upon an initial estimate for the takeoff
gross weight.
From these parameters, the geometric dimensions necessary for layout of
the reference (trapezoidal) wing or tail �an be obtained, as shown in Fig. 7.24
and defined by the following equations:

b = v'AS (7.5)
Croot =
2S (7.6)
b ( l + A)
Ctip = ACroot (7.7)

.._-- Croat ---io--

t
y

i c hord ( C )
Mean aerodyna m i c

b/2

j
Fig. 7 . 24 Reference (trapezoidal) wing/tail.
 ( ) Croot 1
1 92 A i rc raft Desi g n : A C o n c e p t u a l A p p r o a c h

- 2 + A + A2
C= 3 (7.8)

(�) (11: :)
l+A

y= 2
(7.9)

The mean aerodynamic chord (MAC or C, pronounced "C-bar") is the

wing chord that acts like all of the area of the wing is compressed there.
Recall that the wing subsonic pitching moments do not change with angle
of attack if measured about a point 25% of chord back from the leading
edge of the MAC. Y (" Y-bar") is the spanwise location of C-bar.
For a vertical tail, the Y-bar value calculated in Eq. (7.9) must be doubled.
This occurs because the total area is half the value obtained if the vertical
tail were to be laid flat and converted to a symmetrical wing. All other
calculations are the same as for a wing or horizontal tail.
Figure 7.24 also shows a quick method of determining C-bar and Y-bar
from a trapezoidal wing. Y-bar is found as the intersection of the
50%-chord line and a line drawn from a point located at the tip chord
length behind the root chord, to a point at the root chord length ahead of
the tip chord. C-bar is found simply by drawing it at that location.
As mentioned in Chapter 4, the elliptically shaped wing has less drag than
the trapezoidal wing already assumed, and if modern composite construction
is used, it might no longer have the historically expected cost penalty. When
designing an elliptical wing the span is found as usual from Eq. (7.5). The root
chord Croot is found from Eq. (7. lOa). Then the chord length is found as a
function of distance Y from the centerline as in Eq. (7. lOb). The MAC is
84.9% of Croot • and Y-bar is equal to 52.9% of the semispan.

Croot = 7Tb
45
(7 . l Oa)

(7 . lOb)

Note that total area is 'TT/ 4 times the product of span and root chord. Also,
it is common for the chords of elliptical wings to be "slid" in the chordwise X
direction so that the 25% of chord line is straight and unswept. This has no
effect on the above calculations, but does move the 25% of MAC location a
bit forward .

.a Wing Location with Respect to the Fuselage

The location and length of the MAC are important because the designer
locates the wing on the aircraft so that some selected percent of the MAC
is aligned with the aircraft center of gravity. This provides a first estimate
of the wing position to attain the required stability characteristics.
 C HAPTE R 7 Confi g u ration Layout a n d Loft 1 93

If the airplane is a pure flying wing, with no other components than a

wing of trapezoidal shape, it will be neutrally stable if its center of gravity
is at that 25% of C-bar location. The pitching moment doesn't change
around that point so that if the angle of attack changes, the moments do
not change. This is the very definition of neutral stability.
An aft tail adds to the stability. For a stable aircraft with an aft tail, the
wing should be initially located such that the aircraft center of gravity is at
about 30% of the mean aerodynamic chord. When the combined effects of
the fuselage and tail are considered, this gives a reasonable level of stability.
When designing an unstable aircraft with an aft tail such as the F-22,
the wing must be farther forward. A good first approximation is to locate
the wing such that the center of gravity is at about 40% of the mean
aerodynamic chord.
For a canard aircraft, such rules of thumb are less reliable due to the
canard downwash and its influence upon the wing. For a control-type
canard with a computerized flight control system (i.e., an unstable aircraft),
the wing can initially be placed such t�at the aircraft center of gravity is at
about 15-20% of the wing's mean aerodynamic chord.
For a lifting-type canard, the mean aerodynamic chords of the wing and
canard should both be determined, and a point at about the 15% MAC
for each should be identified (20-25% for an unstable aircraft). Then
the combined MAC location can be determined as the average of these
percentage MAC locations for the wing and canard, weighted by their
respective areas. Note that this is a very crude estimate!
After the initial layout is completed and analyzed using the methods of
Chapters 12- 19, the wing will probably need to be moved and the tails
resized to meet all required stability and control characteristics. Hopefully
the initial estimates will be close enough so that major changes will not
be needed.

8IJ Wing/Ta i l Lofting

After positioning and drawing the trapezoidal wing and tails, the actual
exposed wing and tails are designed. As changes are made, those original plan­
forms should be retained on the drawing or as components in the CAD file,
to facilitate their use in aerodynamic calculations. Recall that the trapezoidal
wing is used as the reference wing for aerodynamic coefficients-don't forget
what it is!
The trapezoidal wing is defined to the aircraft centerline and is based
upon the projected area (i.e., dihedral does not affect the top view of the
reference wing). The actual, exposed wing begins at the side of the fuselage
and includes the effect of the dihedral upon the true-view area. The dihedral
angle increases the actual wing area equivalent to dividing by the cosine of
the dihedral angle.
1 94 A i rc raft D e s i g n : A Conceptu a l A p proach

a) Rounded wingtip b) Tra i l i ng-edge "kick" or "bat"

d) Cu rved

Fig. 7.25 Nontrapezoidal wings.

Also, the actual wing planform might not be trapezoidal. Figure 7.25
illustrates several of the many nontrapezoidal wing variations. A typical
rounded wing tip is shown in Fig. 7.25a. This and other wing-tip shapes
have already been discussed. The straightened-out trailing edge shown in
Fig. 7.25b increases the flap chord and provides increased wing thickness
for the landing gear.
Figure 7.25c illustrates a "leading-edge extension" (LEX), which increases
lift for combat maneuvering (see Chapter 12). A highly blended wing/body
is shown in Fig. 7.25d, in which the actual wing looks very little like the
reference wing.* This type of wing is used to minimize the transonic and
supersonic shocks.
Once the designer has settled upon the actual wing and tail planforms,
their surfaces must be lofted to provide accurate cross sections. These are
required to verify that there is sufficient room for the fuel tanks, landing
gear, spars, and other internal components. During production design, this
lofting would be done using conics or some other mathematical surface
definition in a modern CAD system.
For initial design, simpler methods of wing and tail lofting can be used.
These rely upon the assumption that the airfoil coordinates themselves are

* Be careful: if the actual wing looks almost nothing like the original trapezoidal wing, classical
analysis methods based on the original wing parameters may give a poor result. Computational
aerodynamics analysis methods are not so affected.
 CHAPTER 7 Confi g u ratio n Layout a n d Loft 1 95

smoothly lofted. This is an excellent assumption-otherwise the airfoil

performance would be poor.
If the wing or tail uses the identical airfoil section and thickness ratio at all
span stations and is without twist, the airfoils can be drawn simply by scaling
the airfoil coordinates to fit the chord lengths of the selected spanwise
locations.
It is customary to lightly draw the airfoils on the top view of the wing,
superimposing them on their chordline (Fig. 7.26). This layout procedure
simplifies the generation of cross sections, as will be discussed later. For
initial design purposes the airfoils can be quickly drawn using only a few
scaled coordinate points for the top and bottom surfaces.
If twist is incorporated, the incidence at each span station must be deter­
mined and the chord line rotated accordingly before the airfoil is drawn.
Because the chord length is defined in top view, the chord length at each
spanwise station must be increased equivalent to dividing by the cosine of
the appropriate incidence angle.
For a complicated twist distribution, an auxiliary twist control line can be
constructed behind the wing. The airfoil incidence at each span station can
then be read from the control line (Fig. 7.27).
A wing with a complicated aerodynamic design might have the twist,
camber, and thickness all varying from root to tip. These spanwise variations
can be lofted by using a separate auxiliary control line for each, as shown
in Fig. 7.28. The airfoil coordinate points must be calculated by separating
the airfoil into its camber line and thickness distribution, scaling them as
indicated by the auxiliary control lines, and recombining them. Such a
complicated wing design is not normally accomplished until later in the
design process. In fact, the latest computational wing design methods directly
reshape the wing surfaces in the CAD model so that these methods may not
b e needed. Still, it is useful to understand the classical methods.

'
' A i rcraft top view
'
'

Fig. 7 .26 Airfoil layout on wing planform .

1 96 Aircraft Des i g n : A Conceptua l Approach

A i rfoi l incidence (deg)

+ 2 +l 0 -1 -2

_f_ _
-� - -
1 .0 deg

I
O deg�

{
- 1 .0 deg

- - - - �::=;;�
-�
- - - - - -

-2.0 deg
Ang les a re exaggerated for i l l u stration

Fig. 7 . 2 7 Airfoil layout with twist.

Auxiliary control l i nes

( +J ___,-+---4-'--.-+-�-- H

L
J
Ca m ber _} Twist

Th ickness (tic)

Fig. 7 .28 Wing airfoi l layout-nonlinear variations.

 C HA PT E R 7 Confi g u ration Layout and Loft 1 97

U n rigged Rigged

I
This section moved down
- Note T.E.

Fig. 7 . 29 Wing airfoil rigging.

For a wing such as shown in Fig. 7.28, the complex curvatures of the wing
su rface can present difficulties. A spar running from root to tip might be
so curved that it is structurally undesirable. Even worse, the hinge lines for
the ailerons and flaps might not lie in a straight line. Curved hinge lines
are impossible, so the ailerons and flaps might have to be broken into a
shorter segments unless the wing surface can be modified to straighten the
hinge line.
This is done by "wing rigging" (not to be confused with the rigging of a
b iplane wing)-the process of vertically shifting the airfoil sections until
some desired spanwise line is straight.
Figure 7.29 illustrates a complex wing in which the aileron hinge line,
Section A-A, is curved. On the right side of the figure is the same wing
with the midspan airfoil moved downward a few inches. This provides a
straight hinge line shown as Section B-B.

aJ)JJ Airfoil Linear Interpolation

Wings are often initially defined by a root airfoil and a tip airfoil, plus
their incidence angles or relative twist. Frequently the tip airfoil will be
selected for gentle stall characteristics whereas the root airfoil is selected
for best performance. The resulting wing has good overall performance
with good stall characteristics because the tip will stall after the root. The
airfoils between the root and tip can be quickly developed by one of
two methods.
Linear interpolation, the easiest method, is depicted in Fig. 7.30. Here
the new airfoils are created as "weighted averages" of the root and tip airfoils.
Linearly interpolated cross sections are also known as "ruled surfaces" and
are so called in many CAD programs.
1 98 A i rc raft Desi g n : A C o n c e p t u a l A p p r o a c h

\ (3)

\
\

\
\ (5) \
\ (6)
\ • •
• •
• \ •
\

\
\4)
\
(2)

(3)

Graphic i nterpolation between d iffering root and tip a i rfoils

1 -S u perim pose root a n d tip a i rfoils o n pla nform
2-Draw line at some constant percent of chord
3-Swi ng a i rfoi l point down onto chord reference l i n e
4-Con nect root a n d tip poi nts from 3
5-Swi ng point u p to new a i rfoi l location
6-Repeat for other perce nt chord l i n es

Fig. 7 .30 Wing airfoil layout-linear interpolation .

The intermediate airfoils are linearly interpolated by a five-step process.

The root and tip airfoils are drawn (step 1). A constant percent-chord line
is drawn connecting the root and tip airfoil, and vertical lines are drawn
from the intersection of that line with the chord lines (step 2). The airfoil
points found at those vertical lines are "swung down" to the chord line,
using an arc centered at the intersection of the chord line and the vertical
line (step 3). These swung-down points for the root and tip airfoils are
then connected by a straight line (step 4).
At the desired location of an interpolated airfoil, a chord line is drawn.
The intersection of that chord line with the line drawn in step 4 defines
the chordwise location of a point on the interpolated airfoil. In step 5 this
point is "swung up" to its thickness location by an arc centered at the inter­
section of the chord line and the spanwise percent-chord line from step 2.
 CHAPTER 7 Confi g u ration Layout a n d Loft 1 99

This process is repeated for as many points as are needed to draw

th e new airfoil. Then the process is repeated to draw other airfoils.
While it seems complicated, a wing can be developed using this method
in about 15 min by an experienced designer. A CAD system does this
instantly.
Linearly interpolated airfoils have section properties that are approxi­
mately the interpolation of the section properties of the root and tip airfoils.
Note that interpolated section properties may not be correct for modern
laminar airfoils.

al:J Airfoil Flat-Wrap Interpolation

The linear-interpolation method (ruled surface) doesn't necessarily
provide a flat-wrap surface, instead possibly yielding a slight amount of
compound curvature. This is especially true if the wing is twisted or the
airfoil shapes are dissimilar. This requires a modification to the method
just described.
In laying out a fuselage for flat-wrap, it was necessary to hold the same
tangent angle for the conics in the different cross sections. The same is
true for wings. To provide a proper flat-wrap lofted wing, linear interpolation
must be done between airfoil coordinates with the same slope (i.e., tangent
angle) rather than the equal-percent-of-chord method above.
Figure 7.31 illustrates this modification. The only difference is in step 2.
Previously a spanwise line was drawn connecting constant percent chord
locations on the chord line. To obtain a flat-wrap surface this spanwise
line must be drawn connecting locations on the chord line that have
the same surface slope. Note in the figure how the tip chord has the indi­
cated slope at a more-aft percent location of the chord than does the
root chord.
Is this really important? Many composite homebuilts are fabricated by
a method long used for model airplanes. A large block of foam is cut to
the desired wing shape using a hot-wire cutter that is guided by root and
tip airfoil templates attached to the foam block. The templates have tic-marks
that are numbered. The wire is guided around the templates by two home­
builders, one of whom calls out the numbers of the tic-marks.
If the tic-marks are at constant percent-chord locations, and the wing has
dissimilar airfoils or appreciable twist, this method will produce a linearly
interpolated instead of flat-wrap surface. For a wing covered by fiberglass,
this will pose no problem as the fiberglass cloth will easily conform to the
slight amount of compound curvature present.
However, if the wing is to be covered by sheet metal or plywood, the
linearly interpolated foam surface will be depressed relative to the flat­
wrapped skin. This could reduce the strength of the skin bonding. It is
conceivable that such a wing could fail in flight for this simple reason.
Who said lofting is not important?
200 A i rcraft Des i g n : A Conceptu a l A p p ro a c h

l2f'..)

� (3) (1 )

-?,0" " '\\ '

'\ '

'\
'\
\
'\ (5) (6)
'\
. '\ ,, . .
'\ '\\ . .

(1 )

Flat-wra p i n terpolation between d iffering root a n d tip a i rfoi l s

1 -S u perimpose root a n d tip a i rfoils on planform
2A-For a point on the root a i rfoil, fi n d the slope
2 8 - F i n d the point o n the tip a i rfoi l with the same slope
2C-Con nect the percent chord points from (2A) a n d (28)
3-At root a n d tip, swing points down onto chord reference l i n e
4-Con n ect the points from ( 3 )
5-Swing p o i n t u p t o n e w a i rfoi l location
6-Repeat for other points

Fig. 7 . 3 1 Wing airfoi l layout-flat-wrap.

MlJ Wing/Ta i l Cross-Section Layout

One of the important tasks during configuration design is to make sure
that the fuel tanks, landing gear, and other internal components all fit
within the wing contours. The wing shape is defined by airfoils, but this fit
verification is better done using wing and tail cross sections, which are
oriented perpendicular to the aircraft centerline. Such cross sections can
be easily developed once the airfoils are drawn onto the top view of the
wing. Figure 7.32 illustrates the development of one such cross section.
To develop a wing or tail cross-section cut, vertical lines are drawn on the
cross section at the spanwise location of the airfoils shown on the wing top
view. Also, the wing reference plane is shown at the appropriate wing
dihedral angle. Then the airfoil upper and lower points are measured relative
 CHAPTER 7 Configuration Layout a nd Loft 20 1

to th e plane of the wing and drawn accordingly on the cross section. The
cross -section shape can then be drawn using French curves.
The same procedure can be used to develop section cuts at angles other
th perpendicular to the aircraft centerline. The sections of Fig. 7.29
an
labeled A-A and B-B were developed in this manner.
A modern CAD system can easily create these cross-section cuts. Ideally,
those cuts are readily superimposed upon the internal components allowing
either them to be redesigned or relocated as appropriate.

MJll Wing Fillets

For improved aerodynamic efficiency, the wing-fuselage intersection of
many aircraft is smoothly blended using a "wing fillet" (Fig. 7.33). A wing
fillet is usually defined by a circular arc of varying radius, tangent to both
the wing and fuselage. Typically, a wing fillet has a radius of about 10% of
the root-chord length.
The fillet arc radius can be constarrt or can be varied using an auxiliary
radius control line as shown in Fig. 7.33. The fillet radius usually increases
towards the rear of the aircraft to minimize airflow separation. Some aircraft
have a fillet only on the rear part of the wing. In this case the fillet starts, with
zero radius, at or near the wing's maximum thickness point.

'
'
'
'

//
"1
/I
/
/ I
Superim pose a i rfoils on pla nform I I
I I
transfer points to cross-section I I
I I
I I
I I
I I
I I
I I
I I
I I
I I
I I
\
-- Wing reference plane

Fig. 7 . 3 2 Wing/tail cross-section layout.

202 A i rc raft Desi g n : A C o n c e pt u a l A p p r o a c h

Wing fi l let
Lea d i n g-edge/
fi l let rad i u s

�
Arc Auxi l i a ry fi l let control l i n e
ra d I. U S
+
----1
i,
I

Fig. 7 .33 Wing fil let layout.

The fillet circular arc is defined perpendicular to the wing surface, so that
the arc is in a purely vertical plane only at the maximum thickness point of
the wing. At the leading edge, the arc is in a horizontal plane, that is, it is seen
in top view.
For initial layout purposes the fillet is frequently "eyeballed." Only a few of
the 10 or 15 aircraft cross sections developed for an initial layout will show
the wing fillet, so a fillet radius that "looks good" can be used.
Some airplanes have a fillet that is basically a straight and nearly vertical
line running from the maximum width point of the fuselage, down to the top
of the wing and extending towards the rear. While not as beautiful as the
circular fillet, it can work just as well.

MJjJ Winglet Design

Winglets were presented in Chapter 4 as devices to reduce induced drag,
especially for a wing with a fairly high span loading. They are now widely used
and are especially beneficial when an existing design is being recertified to a
higher takeoff weight. This increases span loading, which increases induced
drag unless the wing span is also increased. That is difficult and expensive.
Instead, winglets can be added to counteract the induced drag increase
without the need to extend the wing span.
Fundamentally, the winglet works by producing a side force (inward­
pointing "lift") that has a slight forward component because of the rotation
of the vortex over the top of the wing tip. If there is no side force, then
there is no winglet effect. Thus, the winglet must be wing-like (hence the
name), with both camber and angle of attack to the local flow.
 C HAPTER 7 Confi g u ration Layout and Loft 203

Another way to understand how a winglet works is this: it is a vertical

wing, lifting inwards. Like all lifting surfaces, it makes a "downwash"
behind itself, which in this case is an "outwash." This "blows" the wing's
tip vortices farther outward. In the far field, the effective span of a wing
depends on the separation between its tip vortices, so that the wing, in
effect, has greater span and hence less drag due to lift.
The winglet also acts as an endplate, resisting the tendency of the air to
flow arou nd the tip and therefore allowing the wing to generate more lift
near the tip.
There are many types, shapes, sizes, and geometries of winglets. Almost
every year, a new variation on the basic theme is proposed. Figure 7.34
illustrates what some call the "classic" winglet as defined by R. Whitco mb,
the original developer. The upper winglet should begin at the place where
the wing-tip airfoil has its maximum thickness, and it should be swept
about the same as the wing. It should be at least as tall as the tip chord of
the wing, and even taller is better because the drag reduction is roughly pro­
portional to the winglet height. The Caf!lber of the winglet should be greater
than that of the wing to ensure sufficient side force, and it should have a
4-deg leading-edge-out incidence angle. Typically, the winglet t/ c is about 8%.
The bottom winglet panel, seen o n the original winglet concept,
contributes less to the drag reduction. Because it sticks below the wing, it
threatens to scrape on the ground if the aircraft is rolled too much, so it is
not included on many winglet-equipped aircraft. If included, it should be
twisted with its root at a 7-deg incidence and its tip at 1 1 deg.

11 5 deg·
I

� ·'
I I

I
I
I
Height = c1 I

j
I

Dihedra l

Fig. 7 .34 Wing let design guidelines (after NASA N76-26 l 63, R. Whitcomb).
204 A i r c raft Desig n : A Conceptual A p p ro a c h

A further drag reduction can be obtained by smoothly curving the wing

tip upward to the winglet (when seen from the front) rather than having
the winglet be a separate piece attached to the top of the wing.
One danger with the winglet is that it is adding mass behind the elastic
axis of the wing. Flutter tendencies must always be considered, and a detailed
aeroelastic analysis should be performed to determine if structure stiffening
will be required. Because that will add weight, it can reduce the benefit of
the winglet.

Wetted-Area Determ i nation

Aircraft wetted area Swet , the total exposed surface area, can be visualized
as the area of the external parts of the aircraft that would get wet if it were
dipped into water. The wetted area must be calculated for drag estimation,
as it is the major contributor to friction drag.
The wing and tail wetted areas can be approximated from their plan ­
forms, as shown in Fig. 7.35. The wetted area is estimated by multiplying
the true-view exposed planform area Sex posed times a factor based upon
the wing or tail thickness ratio.
If a wing or tail were paper thin, then the wetted area would be exactly
twice the true planform area (i.e., top and bottom). The effect of finite
thickness is to increase the wetted area, as approximated by Eqs. (7.11) or
(7.12). Note that the true exposed planform area is the projected
(top-view) area divided by the cosine of the dihedral angle.
If t/ c < 0.05,
Swet = 2.003Sexposed (7. 1 1)
If t/ c > 0.05,
Swet = Sexposed [ l.977 + 0.52 (t/c)] (7.12)

---..,-��������-..,��������--..�����.-� cl
\
\
\
\

Fig. 7 . 35 Wing/tail wetted-area estimate.

 C HAPTE R 7 Config u ration Layout and Loft 205

A top

Fig. 7 .36 Quick fuselage wetted-area estimate.

The exposed area shown in Fig. 7.35 ·can be measured from the drawing in
several ways. A professional designer will have access to a "planimeter," a
mechanical device for measuring areas. Use of the planimeter is a dying art
as the computer replaces the drafting board. Alternatively, the area can be
measured by tracing onto graph paper and "counting squares."
The wetted area of the fuselage can be initially estimated using just the
side and top views of the aircraft by the method shown in Fig. 7.36. The
side- and top-view projected areas of the fuselage are measured from
the drawing, and the values are averaged.
For a long, thin body circular in cross section, this average projected area
times 1T will yield the surface wetted area. If the body is rectangular in cross
section, the wetted area will be four times the average projected area. For

)
typical aircraft, Eq. (7.13) provides a reasonable approximation.

Swet � 3.4 top : side

(
A A (7. 13)
A more accurate estimation of wetted area can be obtained by graphical
integration using a number of fuselage cross sections. If the perimeters of
the cross sections are measured and plotted vs longitudinal location, using
the same units on the graph, then the integrated area under the resulting
curve gives the wetted area (Fig. 7.37).
Perimeters can be measured using a professional's "map-measure" or
approximated using a piece of scrap paper. Simply follow around the
perimeter of the cross section making tic marks on the paper, and then
measure the total length using a ruler.
Note that the cross-sectional perimeter measurements should not
include the portions where components join, such as at the wing-fuselage
intersection. These areas are not "wetted."
206 A i rc raft Des i g n : A Conceptu a l A p p ro a c h

Cross-section
peri meter

Wetted a rea = a rea u n d e r cu rve

Fig. 7 .37 Fuselage wetted-area plot.

Volume Determ i nation

The aircraft internal volume can be used as a measure of the reason­
ableness of a new design. A conceptual design layout can't show all of the
internal components that will be packed inside by the time the airplane
flies. Many of them won't be designed until much later in the development
process. A statistical approach can be used to determine if there is enough
room in the design to accommodate all components. This should be done
by the designer as a check after the design layout is completed. It is also
commonly done by customer engineering groups to see if a design is
acceptable or if the designers have "cheated," making a layout that looks
good now due to its small size but will have to grow as the design
is matured.
This is done using statistical plots of total internal volume vs aircraft
takeoff gross weight for different classes of aircraft. An aircraft with
less-than-typical internal volume for its weight will probably experience
problems in development and will likely have poor maintainability in
service due to tight packaging. A more sophisticated density checking
method called net design volume is given in Chapter 19.
Aircraft internal volume can be quickly estimated in a similar fashion to
the wetted-area. Equation (7.14) uses the side- and top-view projected
areas as used in Eq. (7.13) to estimate volume. The 3.4 term assumes that
 CHAPTER 7 Confi g u ration Layout and Loft 207

the cross-section shape is intermediate between a square and a circle. L is

the fuselage length.

Vol � 3.4
(A top ) (A side ) (7. 14)
4L
An accurate estimate of internal volume can be found by a graphical
integration process much like that used for wetted-area determination.
The cross-section areas of a number of cross sections are measured and
plotted vs longitudinal location. The area under the resulting curve is the
volume, as shown in Fig. 7.38.
This "volume distribution plot" is also used predict and minimize
supersonic wave drag and transonic drag rise. In fact, its very shape
determines the supersonic drag. This will be discussed in Chapter 12.

Use of Computer-Aided Design (CAD)

in Conceptua l Design
Today, the previous discussion of drafting table techniques sounds
almost quaint. Everyone, from students to grizzled industry veterans, now
uses a CAD system of some sort for most design work.
Modern CAD systems are amazingly powerful and offer excellent graphi­
cal user interfaces, accurate surface definitions, realistic photo-like rendering
capabilities, and sophisticated data management systems, even on a personal

Cross-section
area

Vol u m e = a rea under curve

Fig. 7 .38 Aircraft volume plot.

208 Aircraft D e s i g n : A Con ceptu a l Approach

computer. Design capabilities allow creation of every imaginable type of

geometry, and various CAD systems have specific geometry creation tools
to simplify development of certain design components and features. The
best modern CAD systems have virtually automated design of certain
parts, such as hydraulic tubing and access doors. In one CAD program the
hydraulic system designer can simply indicate, in three dimensions, the
desired path of a hydraulic line, and the system will create the tubing at
the proper diameter, construct bends with diameters that can be fabricated
without cracking, and include the proper fittings, couplings, and brackets­
all automatically. The future will see more and more such automation of
the design of common parts and systems.
Furthermore, through the industry usage of modern CAD systems the
entire aircraft is being designed digitally, allowing the use of virtual rather
than actual mock-ups. T his saves time and money and does a better j ob of
identifying and fixing component interference problems and potential
difficulties in fabrication and maintenance of the aircraft. The digital
product definition also improves prototype fabrication and aircraft produc­
tion. Transference of the design data to computer-aided manufacturing
(CAM) becomes almost trivially easy, and the resulting parts fit together
perfectly. Altogether, the integrated use of CAD and CAM has been, in
this author's opinion, the single greatest improvement in cost and quality
that the aircraft industry has ever seen.
However, there can be problems with too great of a willingness to "let the
CAD system do it." First of all, with a CAD system there is a tendency to let
the computer lead you in the "easy" direction. If it is easy to retract the
landing gear directly inward with your CAD system, you may do so even if
a better design would result from having it retract inward and forward at
a difficult-to-construct oblique angle. If you can easily calculate the
volume of a square fuel tank, but don't know how to get the volume of a
complicated tank wrapped around the inlet duct, guess which one you are
likely to design!
Another problem is the actual calculation of the volumes, wetted areas,
and other dimensions critical to your analysis of your design. Sometimes a
CAD system may confidently display an incorrect answer! For example, we
might model the wing as a collection of airfoils connected by a mathematical
surface and might readily calculate the wetted area of the wing itself.
However, where that wing intersects the fuselage we must cut away the
surface of the wing where it penetrates the fuselage and cut away the fuselage
where the wing covers it. It is possible in many CAD systems to forget to
account for, or double-account for, the wing root airfoil "wetted area"
that must be removed from the fuselage and not included with the wing!
This potential problem is minimized if true "solid models" are (properly!)
employed. Other examples include the inlet front and the back end of a
fuselage or nacelle with a jet engine, or the front of a propeller nacelle,
where the exhaust or intake areas must not be included. Even a solid
 C HA PT E R 7 Configu ration Layout a n d Loft 209

model could accidentally give the wrong answer in this case, failing to
understand that the "hole" isn't there!
For this reason it is STRONGLY recommended that all CAD users start
by doing a trivially simple "aircraft design" consisting of a tube-plus-cone fuse­
lage and a simple wing, where the correct wetted areas and volumes can be
easily calculated by hand and compared with the answer from the CAD system.
Yet another problem for students is that the aircraft design course can
easily become the "learn how to use a certain CAD system" course. There
is not enough time in a semester course to really learn how to do conceptual
design, and ANY time spent learning which button produces which geometry
is time NOT spent learning the philosophy, methods, and techniques of
aircraft conceptual design.
In industry, a real but subtle problem is that, with a CAD system,
everybody's designs look good whether they are or are not! When everybody
was using a drafting table, you could usually tell from drafting technique that
a design was done by a beginner and therefore whether the design needed to
be reviewed extra carefully. Today, it "t�kes one to know one" -you must be a
pretty good designer yourself to know if a design you are looking at was
done properly.
CAD tools used during conceptual design should be tailored toward
the fluid environment and the unique tasks of aircraft conceptual design.
Quite simply, what is done during conceptual design, the things that are
critical, and the tasks that are boring and repetitive (and therefore ideal for
computerization) are different from those in other, later phases of aircraft
design.
A perfect example is the wing trapezoidal geometry. During detail part
design, it is out of the question to change the wing trapezoidal geometry,
no matter how much the design of, say, a certain wing rib would be improved
as a result. During conceptual design though, those parameters are constantly
being changed, almost every week in the early stages. Conceptual designers
need capabilities to change these instantly and to have the computer
automatically revise the wing's nontrapezoidal shaping to match the new
geometry and also revise the geometries of any parts made from the wing,
such as wing fuel tanks, flaps, ailerons, spars, ribs, and possibly even wing
carry-through structure and landing gear attachments. All that the designer
should have to do is to enter the revised geometric parameter (such as aspect
ratio).
Figure 7.39 shows such an automatic revision of the nontrapezoidal
geometry from changes to the geometric trapezoidal parameters, done with
the RDS-Professional program. l24l At the upper left is trapezoidal wing
geometry. To its right is the wing created from it, with a swept-back tip,
leading-edge strake, and trailing-edge kick. Below is the revised trapezoidal
geometry after the aspect ratio, taper ratio, and sweep are changed in
response to some optimization. To its right is the resulting wing geometry
including the same swept-back tip, leading-edge strake, and trailing-edge kick.
210 A i rc raft Des i g n : A C o n ceptu a l A p p roach

, - -
\
\
\
\

Fig. 7 .39 Automated revision of wing geometry.

Notional Design Layout: Advanced Technology Commuter/Cargo Jet (D. Raymer, courtesy
Conceptua l Research Corp.).
 CHAPTER 7 Confi g u ration Layout a n d Loft 21 1

C- 1 7 Globemaster (NASA photo by Jim Ross) .

What We've Learned

Configuration design layout i s the heart o f the design process: you build the
drawing. The fuselage and similar bodies should be designed using a deliberate
longitudinal control scheme, as illustrated by classic conic lofting. Wings and
tails should be designed using spanwise control lines to place and scale the
selected airfoils.
212 Airc raft D e s i g n : A C o n ceptu a l A p p roach
 Special
Considerations
in Configuration
Layout

•
�.£] • •

W h i l e d o i ng t h e " mecha n i ca l " tasks of a i rcraft layout desig n , the designer is thinking
a bout many other things to m a ke a good a i rp l a n e ,
• Al l a re i m portant, a n d a l l m u st be considered ,
• Often "good t h i n g s " i n o n e a rea w i l l conft ict with those i n a nother a rea (aero vs
·
structure') ,
• The confi g u ration designer w i l l never be expert i n a l l of these, but needs to know
them wel l e n o u g h to m a ke the layout a n d tal k to the experts ,

Introduction

T
he previous chapter discussed the mechanics of configuration layout.
Later chapters will focus on the provisions for specific internal
components, such as the crew station and landing gear. This
chapter discusses various intangible considerations that the designer
should consider when making the initial layout. These include aerodynamics,
structures, detectability, vulnerability, producibility, and maintainability.
These are numerically analyzed in later stages of the design process, but
that is possible only when the initial layout is completed. During configur­
ation layout, the designer must consider their impact in a qualitative sense
and try to "do the right thing."

213
214 A i r c raft Desi g n : A C o n c e pt u a l A p p r o a c h

Aerodynamic Considerations

Aerodynamic analysis will be discussed in Chapter 12 where various

estimation methods will be presented. But aerodynamic design doesn't
start with calculations, it starts with the initial design layout. During
concept layout, the designer must consider the requirements for aerody­
namics based upon experience, understanding, and a "good eye."
The overall configuration arrangement and "cleanliness" of an aircraft has
a major effect upon aerodynamic efficiency. A poorly designed fuselage can
have excessive flow separation, unnecessary interference drag, a bad transo­
nic drag rise, and high supersonic wave drag. A clumsily done wing-fuselage
intersection can cause lift losses or disruption of the desired elliptical lift
distribution and can cause bad handling qualities including spin tendencies.
The sum total of the wetted area of the design is the most powerful
aerodynamic consideration for virtually all aircraft. Friction drag is literally
calculated as a coefficient times the wetted area, so an excess of wetted area is
always to be avoided. Fuselage wetted area can be minimized by tight internal
packaging, that is, less internal volume. This is always a temptation but must
be balanced against the needs of subsystems integration and of maintenance.
For a given volume, wetted area is minimized with a low fineness ratio.
Unfortunately, a short, fat fuselage has a lot of separation in the back
causing a huge pressure drag. As discussed in Chapter 6, the designer
should pick a suitable fineness ratio, that is, the ratio between the fuselage
length and its maximum diameter. If there is a specific layout requirement
that forces a certain cross-section area, such as side-by-side seating for two
people, then a fineness ratio of about three is the best answer. For an aircraft
where the internal components can be rearranged
and the cross-section diameter can be reduced as The greatest
needed, the optimum fineness ratio for subsonic air­ compliment a
craft is somewhere between 6 and 8. Supersonic designer will ever
designs will have a fineness ratio of 10 to 15 or more. receive: "He (or
Good lofting also produces good aerodynamics. she) can see air."
The use of smooth longitudinal control lines in
developing the fuselage contours will reduce the drag. Slope discontinuities
(breaks) in the longitudinal direction are very bad. If there needs to be a
longitudinal break, it should be smoothed following a radius roughly equal
to the fuselage diameter.
Even curvature (second derivative) discontinuities should be avoided as
much as possible. The flow tends to separate right at the discontinuity.
This is shown in Fig. 8.1, where a nicely-radiused front suddenly transitions
to a flat area. If this is a train track, the train comes off the track right where
the arrow points. If this is an airfoil or a fuselage, the air "comes off the track"
causing a lot of drag and a loss of lift.
 CHAPTER 8 Spec i a l Considerations i n Confi g u ration Layout 21 5

Di sconti n u ity in seco n d de rivative

Ten d s to sepa rate here

(
. .. . . Secon d d erivative cont i n u o u s

(
N o t a s l i kely t o sepa rate

.·
·
· ..

··········

Fig. 8. 1 Railroad curves reduce drag.

In 1880 the Railroad Gazette published the solution called the "Track
Transition Curve," also known as an "E uler Spiral." This is a curve whose
curvature (I /radius) changes linearly with curve length, reducing to zero
when the straight segment is reached. We airplane designers just call it a
railroad curve, and eyeball it to look like this. It works.
To prevent separation of the airflow, the aft-fuselage deviation from the
freestream direction should not exceed 10 or 12 deg (Fig. 8.2). This can go
up to 15 deg on the bottom because the higher pressure air tends to push
the air around the corner.

�,.._�����-=======�:-�-,--
1 1 0 deg- 1 2 deg
-===T
c:==-- J
maxi m u m

c �� L\ 30 deg maxi m u m

� g m,,; m " m

\__ S m a l l rad i u s corners

Fig. 8.2 Longitudinal contour guidelines.

216 Ai rc raft D e s i g n : A C o n c e pt u a l Approach

The air inflow induced by a pusher-propeller will "pull" the air around the
corner, preventing separation despite contour angles of up to 30 deg or more.
However, when that push propeller stops working, the flow separates causing
a drag increase to compound the thrust loss. It is for this reason that airplanes
with propellers in front and back generally fly better on the back propelle r
than the front.
In general, aft-fuselage upsweep should be minimized as much as poss­
ible, especially for high-speed aircraft. An upsweep of about 25 deg can be
tolerated for a rear-loading transport aircraft provided that the fuselage
lower corners are fairly sharp. This causes a vortex flow pattern that
reduces the drag penalty. Some aircraft use strakes at the rear of the fuselage
for the same reason.
The shape of the fuselage cross section affects the drag. To reduce cost,
some airplanes have been designed with simple square cross-section
shapes. While easy to build, this can increase drag by 30-40% due to separ­
ation when the high-pressure air underneath tries to flow around the sharp
edges to the sides and top.
If an aircraft's forebody has sharp lower corners or even corners that just
aren't rounded enough, a separated vortex can be formed at high angles of
attack. This can be ingested by the inlets, with bad results, and can have an
unpredictable effect upon the wing or tail surfaces.
The importance of well-designed wing fillets has already been discussed.
Fillets are especially important for low-wing, high-speed aircraft such as
j et transports.
"Base area" is any unfaired, rearward-facing blunt surface. Base area
causes extremely high drag due to the low pressure experienced by the
rearward-facing surface (see Chapter 12).
However, a base area between or very near the jet exhausts can be
"filled-in" by the pressure field of the exhaust, partially alleviating the drag
penalty. The T-38 has such a base area between its nozzles. A base area
fill-in effect is difficult to predict.
The aerodynamic interaction between different components should be
visualized in designing the aircraft. For example, a canard should not be
located such that its wake might enter the engine inlets at any possible
angle of attack. Wake ingestion can stall or even destroy a jet engine.

#:fIJ Isobar Tai loring

In Chapter 4 the importance of wing sweep for delaying the formation of
shocks over the wing was discussed. It was explained that the shocks are
formed over the top of the wing due to the increased velocity causing the
air to go supersonic. It was also explained that theoretically this could be
proven to depend not on the actual velocity of the air over the top of the
wing but by the velocity perpendicular to the leading edge. Sweeping the
 C H A PTE R 8 Spec i a l Considerations i n Confi g u ration Layout 21 7

Isobar l i nes of
consta nt p ressu re

Isobars u n sweep at root a n d tips

Restore isobar
sweep with Restore isobar
"peaky" root sweep with pla nform
a i rfoil

Fig. 8.3 Isobar tailoring for shock suppression .

wing causes this velocity to appear to be reduced, so shock formation

is delayed.
Actually, the wing sweep theory is based not just on leading-edge sweep,
but on the sweep of the wing pressure "isobars." Isobars are lines connecting
regions with the same pressure. This is illustrated in Fig. 8.3. At the upper left
there is an airfoil with its pressure contours shown, and four pressures are
depicted with dots. To the right is a top view of part of the wing with
those same four dots shown and lines (isobars) connecting those dots with
other points on the top of the wing having the same pressure.
The complete wing shown in Fig. 8.3 illustrates two common problems
with "real" wings. At the root, the isobars from the left and right sides of
the wing cannot meet in a "V." Instead, they are joined by a rounded-off
corner. As a result, this swept wing has no effective sweep at the root, at
least according to the wing sweep theory. This causes shocks near the wing
root and is a very real problem. Something similar happens at the wing
tips, as shown.
To solve the isobar unsweep problem at the root, two aerodynamic
strategies can be employed. One is to exaggerate the wing sweep near the
root, blending the wing in a smooth fashion into the forebody of the fuselage.
This is seen on the B- lB and was featured on the North American Rockwell
F-X (F- 15 Proposal) design, shown at the right of Fig. 8.3.
218 A i rc raft Desi g n : A C o n c e pt u a l A p p r o a c h

Another approach, commonly used on large airliners, is to "pull" the wing

root isobars forward by using a strange airfoil shape at the root that is
specially designed to have its pressure peak very near its nose. Such an airfoil
tends to have a large nose radius, a fairly flat top, and, oddly enough, negative
camber. This negative camber tends to create negative lift, so that it must be
placed at a high nose-up twist angle to maintain a good spanwise lift
distribution.
Such design tricks are beyond the scope of early conceptual design but
can be approximated on the initial design layout based on similar aircraft.
At a later date, an airfoil/wing optimization code will be run to define the
best airfoil geometries.

#:fD Superson ic Area Rule

For supersonic aircraft, the greatest aerodynamic impact upon the
configuration layout results from the desire to minimize supersonic wave
drag, a pressure drag due to the formation of shocks. This is analytically
related to the longitudinal change in the aircraft's total cross-sectional
area. In fact, wave drag is calculated using the second derivative (i.e., curva­
ture) of the volume-distribution plot as shown in Fig. 7.38.
Thus, a "good" volume distribution from a wave-drag viewpoint has the
required total internal volume distributed longitudinally in a fashion that
minimizes curvature in the volume-distribution plot. Several mathematical
solutions to this problem have been found for simple bodies-of-revolution,
with the Sears - Haack body (Fig. 8.4; see [25l ) having the lowest wave drag.
If an aircraft could be designed with a volume plot shaped like the Sears­
Haack volume distribution, it would have the minimum wave drag at Mach
1.0 for a given length and total internal volume. (What happens at higher

"'
�
c
"' f-;...-'----+--;--'-:-..-+-'--'-1-,_..,...-i--c_;_�-+-�...,-;->-��.-+-"-'-;-+-+--c-''-"'-+-I
0
·.;:::;
u
QJ
Vl
Vl

0
Vl

u H--'-'_,_.,_,._.__H-+-'--'---+-+-,--;--i-,--1-+-+--,---,_,_-'-'"-+--.,--t--->r_;_..,_,.-'-"-+-:-i

Fuselage stations

Fig. 8.4 Sears-Haack volume distribution.

 C H A PT E R 8 Spec i a l Considerations i n Confi g u ration Layout 219

"Su personic a rea rule" (M = 1 .0)

Cross -sect ion Cross-section

a rea S moother a rea prog ression
area
Lower maxi m u m c ross-section

Fuselage

Fig. 8.5 Design for low wove drag .

Mach numbers is discussed in Chapter 1 2, but for initial layout purposes the
minimization of wave drag at Mach 1.0 is a suitable goal in most cases.)
However, it is usually impossible to exactly or even approximately match
the Sears-Haack shape for a real aircraft. Fortunately, major drag reductions
can be obtained simply by smoothing the volume distribution shape.
As shown in Fig. 8.5, the main contributors to the cross-sectional area are
the wing and the fuselage. A typical fuselage with a trapezoidal wing will have
an irregularly shaped volume distribution with tht= maximum cross-sectional
area located near the center of the wing. By "squeezing" the fuselage at that
point, the volume-distribution shape can be smoothed and the maximum
cross-sectional area reduced.
This design technique, developed by R. Whitcomb of the NACA, [26l is
referred to as "area-ruling" or "coke-bottling" and can reduce the wave
drag by as much as 50%. Note that the volume removed at the center of
the fuselage must be provided elsewhere, either by lengthening the fuselage
or by increasing its cross-sectional area in other places.
While area-ruling was developed for minimization of supersonic drag,
there is reason to believe that even low-speed aircraft can benefit from it
to some extent. The airflow over the wing tends to separate toward the
trailing edge. If an aircraft is designed such that the fuselage is increasing
in cross-sectional area toward the wing trailing edge, this can "push" air
onto the wing, thus reducing the tendency to separate. The Wittman
Tailwind, which is remarkably efficient, uses this approach.

M:fJI Compression Lift

A successful yet almost forgotten aerodynamic concept can be used
to imp rove lift-to-drag ratio at supersonic speeds. Compression lift was
220 A i r c raft Desi g n : A C o n c e p t u a l A p p r oa c h

apparently conceived by two researchers at NACA Langley in 1954 and

used by Richard Child and George Owl of North American Aviation to
configure a huge supersonic bomber that literally rode its own shock wave,
the B-70.
Any body shape will create shock waves at supersonic speeds, forming at
the nose and at anyplace else where the cross-section area is increasing.
These shocks trail back at approximately the Mach angle [arcsine (1/M)],
as shown in Fig. 8.6. In the B-70, the inlet duct was faired back into a wide
nacelle, with a steadily widening cross-sectional area until a maximum was
reached (Fig. 8.7). Engines and payload were carried in this nacelle, which
created a strong shock on either side with greatly increased static pressures
behind the shocks. By placing the wing above these shocks, the increased
pressure beneath the wing provided free lift-roughly 30% of the total lift
required!
The B-70 also used fold-down wing tips. As can be seen, these reflected
the shocks from the nacelle creating even more shocks under the wing­
more free lift! Furthermore, they solved the two big stability problems
inherent in supersonic flight. First, the aerodynamic center moves consider­
ably to the rear requiring some way to move it forward at supersonic speeds,
or to move the center of gravity to the rear. Folding down the wing tips does
the former. Also, at supersonic speeds the effectiveness of a vertical tail
usually reduces. Folding down the wing tips helps this problem, too.

Fig. 8.6 Supersonic shocks.

 CHAPTE R 8 Spec i a l Considerations i n Confi g u ration Layout 221

B-70 bottom view

Fig. 8.7 Compression l ift.

In fact, the B-70 is an excellent example of synergistic design. The design

features are all working together, and many of the design components do
more than one task and offer more than one benefit.
The NAAF-X proposal shown in Fig. 8.3 was configured for compression
lift as well as isobar sweep.

M:fJj Design "Fixes"

Real airplanes have many "things" on them that are not seen on a concep­
tual design layout. Some of these things are equipment such as antennas and
lights, and some of these things are minor design details such as fuel drains
and cooling vents that are not normally considered during conceptual design.
Some of these things, though, are fixes to aerodynamic problems discovered
later in design development or flight test. In conceptual design we think we
have no such aerodynamic problems, and if we did, we would revise the
overall arrangement to avoid them. Later on, it is too difficult to change
the overall geometry, so if unexpected problems are found, they must be
fixed in some other way.
Aerodynamic problems are most often attributable to two phenomena:
separation of the flow, or the formation of an unwanted, "bad" vortex.
Typical devices to fix aerodynamic problems are shown in Fig. 8.8. These
mostly work by creating and controlling "good" vortices.
Flow separation over a wing or fuselage often occurs because the air near
the aircraft has been slowed down too much by viscous effects and no longer
has much energy. When this low-energy air is asked to turn a corner, it
simply can't and separates instead. To fix this, a number of small plates are
bent into an "L" shape and attached just before the region of separation,
set at an angle to the flow. These create small vortices that stir up the air
near the surface, bringing high-energy air into the boundary layer. This
action permits the flow to follow a much greater turn. Such "vortex
222 Aircraft De s i g n : A C o n c e ptu a l A p p roa c h

0&
Vortex Fence & notch S n a g or notch
generators

0a,�J0a,9-Y

Nose stra ke S h a r k nose Body o r nacelle strakes

Fig. 8.8 Aerodynamic fixes.

generators" are commonly found on the tops of wings and near the back of a
long fuselage, but can be found almost anywhere on airplanes except right at
the nose!
The best locations for vortex generators to fix some particular problem
are found by trial and error, both in the wind tunnel and in flight test.
Strangely enough, the vortex generators cause almost no increase in parasitic
drag, even on a flat plate. They are so small that they are mostly in the bound­
ary layer, and their own effect on drag is negligible whereas, if they prevent
separation, they can greatly reduce the total drag of the aircraft.
At high angle of attack, the flow experiences a disastrous form of
separation called wing stall. Properly placed vortex generators can delay
this and are commonly found on wings for this purpose, but still don't
allow the wing to reach its maximum lift.
Wing stall tends to start at the wing root and spread outward. By placing a
"fence" just outboard of where the stall has been found to begin, the stall can
be prevented from spreading outward until such a high angle of attack is
reached that the outboard part of the wing stalls on its own.
A fence can also be used to cure a problem common in highly swept
wings. The sweep of the wing tends to push the air outward, especially in
the boundary layer where the air is low in energy. It is not uncommon
for the boundary-layer air from the root of the wing to travel outward, all
of the way to the tip of the wing. This increases boundary-layer thickness,
and that tends to cause flow separation and wing stall. A fence can physically
prevent that occurrence and can improve stall characteristics.
One can create a "virtual fence" by placing a notch or snag at the location
just outboard of where the stall begins. These form a vortex that, like a fence,
 CHAPTER 8 Spec i a l Considerations i n Confi g u ration Layout 223

acts to separate the stalled from the un-stalled flow and stop the stall
from spreading.
The leading edge outboard of the wing notch can be cambered downward
to rther reduce the outboard wing panel's tendency to stall. Properly done,
fu
this can also greatly reduce spin tendencies and promote spin recovery and is
highly recommended for general aviation and training aircraft.
Nose strakes, or the similar sharp-sided "shark nose," are used to force
vortices to form simultaneously on both sides of the forebody at higher
angles of attack. With a rounded forebody, at some high angle of attack
such vortices will form, but the vortex on one side might form sooner than
on the other. Having a vortex on only one side of the forebody creates a
strong suction force that can pull the nose to one side, causing a spin.
Sharp edges on the nose fix this.
Finally, large strakes or fins can be strategically placed to form vortices
and do something good. For example, the vertical tails of the F- 18 were
having structural fatigue problems resulting from an unexpected tendency
of the vortices from the wing strakes .to hit the vertical tails. To fix this,
small upright strakes were added to the top of the aircraft to create vortices
that divert the wing strake vortices. As can be imagined, they were not on the
conceptual design layouts!
Many airliners have similar strakes on the engine nacelles. These can be
used to improve flow over the wing flaps, or to fix a flow problem at the
horizontal tail, or both. The DC- 10, perhaps the first to use such nacelle
strakes, needed them because the nacelle and pylon were causing the flow
to separate resulting in a premature stall. Th{l nacelle strakes fixed the
separation and increased maximum lift.
The growth versions of the DC-9 had flow problems at the vertical tail,
leading to directional stability reduction at moderate sideslip. Strakes
below the cockpit were found to cure this problem, even though they are
located about 100 ft {30 m} ahead of the tail.
Another type of vortex-generating strake called a "vortilon" is placed just
below the wing leading edge and is aligned with the flight direction. (It looks
like a miniature engine pylon that lost its engine!) At high angle of attack, the
local flow at the leading edge is diverted outward toward the wing tip so that
the vortilon finds itself at an angle to the local flow and produces a vortex.
This vortex wraps over the top of the wing and energizes the boundary
layer while acting like a stall fence.

Structura l Considerations
M:fll Load Paths
Except in the smallest of projects, the configuration designer does not
actually do the detailed structural design of the airplane. That is the respon­
sibility of the structural design group. However, the configuration designer
224 Aircraft Design: A Conceptual Approach

does create the overall structural arrangement as a part of the initial configur ­
ation design, defining-with guidance from the structures staff-the major
fuselage frames, longerons, wing spars, carrythrough structure, and attach­
ment locations for the major load items. Well done, this structural arrange ­
ment will create a design that seems to glide through detailed structural
design and produces a lighter-than-usual structures group weight. Poorly
done, nothing awaits but blood, toil, tears, and sweat.
The main concern in the development of a good structural arrangement
is the provision of efficient "load paths"-the structural elements by whic h
opposing forces are connected. The primary forces to be resolved are the
lift of the wing and the opposing weight of the major parts of the aircraft,
such as the engines and payload. The size and weight of the structural
members will be minimized by locating these opposing forces near to
each other.
Carried to the extreme, this leads to the flying wing concept. In a flying
wing the lift and weight forces can be located at virtually the same place.
In the ideal case, the weight of the aircraft would be distributed along the
span of the wing exactly as the lift is distributed (Fig. 8.9). This is referred
to as "spanloading" and eliminates the need for a heavy wing structure to
carry the weight of the fuselage to the opposing lift force exerted by the
wing. The structure can then be sized by lesser requirements such as the
landing-gear loads.
While ideal span-loading is rarely possible, the span-loading concept can
be applied to more conventional aircraft by spreading some of the heavy
items such as engines out along the wing. This will yield noticeable weight
savings, but must be balanced against the possible drag increase, especially
if it requires a larger vertical tail to handle an engine-out situation.
If the opposing lift and weight forces cannot be located at the same place,
then some structural path will be required to carry the load. The weight of
structural members can be reduced by providing the shortest, straightest
load path possible.
Figure 8.10 illustrates a structural arrangement for a small fighter. The
major fuselage loads are carried to the wing by "longerons," which are
typically I- or H-shaped extrusions running fore and aft and attached to
the skin. Longerons are heavy, and their weight should be minimized by
designing the aircraft so that they are as straight as possible.
For example, the lower longerons in Fig. 8.10 are high enough that they
pass over the wing-carrythrough box. Had the longerons been placed
lower, they would have required a kink to pass over the box.
On the other hand, the purpose of the longeron is to prevent fuselage
bending. This implies that the lightest longeron structure occurs when the
upper and lower longerons are as far apart vertically as possible. In
Fig. 8. 1 1 the longerons are farther apart, but this requires a kink to pass
over the box. Only a trade study can ultimately determine which approach
is lighter for any particular aircraft.
 CHAPTE R 8 Spec i a l Considerations i n Config u ration Layout 225

Idea l ly
s pa n loaded Wing (rea r view)
wing

Weight d i stribution

Center
line

Rea listic @
b
.
wrng Fuse-
Wing
lage - "T �-r---------
l--""""\J
'-.../
and
fuselage

Weight d i stribution

Wingtip
store
Nacelle

Fuselage

Fig. 8.9 Spanloading for weight reduction .

In some designs similar to Fig. 8. 10, the lower longerons are placed near
the bottom of the aircraft. A kink over the wing box is avoided by passing
the longeron under or through the wing box. This minimizes weight but
complicates both fabrication and repair of the aircraft.
For aircraft such as transports, which have fewer cutouts and concen­
trated loads than a fighter, the fuselage will be constructed with a large
number of "stringers," which are distributed around the circumference of
the fuselage ( Fig. 8.12). Weight is minimized when the stringers are all
straight and uninterrupted.
Another major structural element used to carry fuselage bending loads is
the "ke elson." This is like the keel on a boat, and it is a large beam placed at
the bottom of the fuselage as shown in Fig. 8.12. A keelson is frequently used
226 A i rcraft Design- A Conceptu a l A p proach

K e y concept: " l o a d paths"

E n g i n e mou nts

� �� -- �� � �� ..
�....�-- � ...
. ....�-
�...;.__. cl
Ta i l s p i n d l e

Longerons

Wing box

Ta i l atta c h m e nt
fittings

<:�::: l�;;-/!:.;: =::=:'. :J2zdJl:=: ���t:�: -

Eng i n e m ou nts

\ Dcog b cooo '""h

S
ot

Fig. 8. 1 0 Structural arrangement.

to carry the fuselage bending loads through the portion of the lower fuselage,
which is cut up by the wheel wells.
As the wing provides the lift force, load-path distances can be reduced by
locating the heavy weight items as near to the wing as possible. Similarly,
weight can be reduced by locating structural cutouts away from the wing.
Required structural cutouts include the cockpit area and a variety of doors
(passenger, weapons bay, landing gear, engine access, etc.).

K i n ked
lower l o n g e ro n

Fig. 8. 1 1 Kinked lower longeron .

 CHAPTER 8 Spec i a l Considerations i n Confi g u ration Layout 227

Fig. 8. 1 2 Structural concepts for fuselage loads.

An especially poor arrangement (seen on some older fighter aircraft) has

the main landing gear retracting into the wing-box area, which requires a
large cutout where the loads are the greatest.
When possible, structural cutouts should be avoided altogether. For
example, a jet engine that is buried in the fuselage requires a cutout for the
inlet, a cutout for the exhaust, and in most cases another cutout for
removal of the engine. The resulting weight penalty compared to a podded
engine must be balanced against the reduced drag of a buried engine
installation.
Figure 8.10 illustrates another important concept in structural arrange­
ment. Large concentrated loads such as the wing and landing-gear
attachments must be carried by a strong, heavy structural member such as
a major fuselage bulkhead. The number of such heavy bulkheads can be
minimized by arranging the aircraft so that the bulkheads each carry a
number of concentrated loads, rather than requiring a separate bulkhead
for each concentrated load.
In Fig. 8.10 the two bulkheads in the aft fuselage carry the loads for the
engines, tails, and arresting hook. Had the tails and engine been located
without this in mind, the structural designer would have had to provide
four or five heavy bulkheads rather than the two shown.
Note that an aircraft designed using composite materials must consider
the special properties of composites. Properly employed, they can reduce
part-count, simplify manufacture, and reduce both weight and cost.
228 Ai rcraft De s i g n : A Conceptua l A p proach

Composites, even more so than aluminum, add substantial weight penalties

when loads are concentrated such as for wing attachment fittings, so good
design practice for such aircraft will avoid load concentrations. It is for this
reason that many newer composite airplanes (Lancair, Cirrus, Kestrel)
have the vertical tail molded right into the fuselage rather than fabricated
separately then bolted on. For more information, see Chapter 14.

*:ffJ Carrythrough Structure

The lift force on the wing produces a tremendous bending moment
where the wing attaches to the fuselage. The means by which this bending
moment is carried across the fuselage is a key parameter in the structural
arrangement and will greatly influence both the structural weight and the
aerodynamic drag of the aircraft.
Figure 8.13 illustrates the four major types of wing carrythrough struc­
ture. The "box carrythrough" is virtually standard for high-speed transports
and general aviation aircraft. The box carrythrough simply continues the
wing box through the fuselage. The fuselage itself is not subjected to any
of the bending moment of the wing, which minimizes fuselage weight.
However, the box carrythrough occupies a substantial amount of
fuselage volume and tends to add cross-sectional area at the worst possible

Wing box ca rryt h ro u g h

Strut-braced

Fig. 8. 1 3 Wing carrythrough structure.

 C H A PTE R 8 Spec i a l Consideratio n s i n Confi g u ration Layout 229

place for wave drag, as already discussed. Also, the box carrythrough inter­
feres with the longeron load-paths.
The "ring-frame" approach relies upon large, heavy bulkheads to carry
th bending moment through the fuselage. The wing panels are attached
e
to fittings on the side of these fuselage bulkheads. While this approach
is usually heavier from a structural viewpoint, the resulting drag reduction
at high speeds has led to the use of this approach for most modern
fighters.
The "bending beam" carrythrough can be viewed as a compromise
between these two approaches. Like the ring-frame approach, the wing
panels are attached to the side of the fuselage to carry the lift forces.
However, the bending moment is carried through the fuselage by one of
several beams that connect the two wing panels. This approach has less
of a fuselage volume increase than does the box-carrythrough approach.
The bending-beam carrythrough is common in sailplanes and is also seen
on a number of advanced composite general aviation designs. Frequently
there is a separate bending beam for each wing half, which simplifies
manufacture.
Many light aircraft and slower transport aircraft use an external strut to
carry the bending moments. This is typically set at around 40 degrees up
from horizontal such that the lift outboard of the strut attachment point is
nearly balanced by the lift inboard of that point. As a result, there is little
load remaining at the place where the wing is attached to the fuselage. For
a high-wing aircraft, rather than "hanging" from the wing, the fuselage is
"sitting" on the bottom strut attachments. While the strut-braced wing is
probably the lightest of all, it obviously has a substantial drag penalty at
higher speeds. Wing structural analysis with a strut is described in Section
14. 10.6.
Aircraft wings usually have the front spar at about 20-30% of the chord
back from the leading edge. The rear spar is usually at about the 60-75%
chord location. Additional spars can be located between the front and rear
spars forming a "multispar" structure. Multispar structure is typical for
large or high-speed aircraft.
If the wing skin over the spars is an integral part of the wing structure, a
"wing box" is formed that in most cases provides the minimum weight.
Aircraft with the landing gear in the wing will usually have the gear
located aft of the wing box, with a single trailing-edge spar behind the gear
to carry the flap loads, as shown in Fig. 8. 14.
Ribs carry the loads from the control surfaces, store stations, and landing
gear to the spars and skins. A multispar wing box will have comparatively few
ribs, located only where major loads occur.
Another form of wing structure, the "multirib" or "stringer panel" box,
has only two spars, plus a large number of spanwise stringers attached to
the wing skins. Numerous ribs are used to maintain the shape of the box
under ben ding.
230 Aircraft Desig n : A Conceptual Approach

Carrythrough box o r
ring fra mes

Wing
atta c h m ents -+---+---- "Kick s p a r"

M a i n �-----'\"'\�
s p a rs

Fig. 8. 1 4 Typical wing box structure.

Variable sweep and folding capability add considerably to the wing struc­
tural weight. On the other hand, use of a delta wing will reduce the structural
weight. These are further discussed in Chapter 15.

#:fD Clearances and Allowances

First-order structural sizing will be discussed in Chapter 14. For initial
layout purposes the designer must guess at the amount of clearance required
for structure around the internal components. A good designer with a
"calibrated eyeball" can prevent a lot of lost effort, for the aircraft might
require substantial redesign if later structural analysis determines that
more room is required for the structural members.
A large airliner will typically require about 4 in. { IO cm} of clearance from
the inner wall of the passenger compartment to the outer skin ("mold line").
The structure of a conventional fighter fuselage will typically require about
2 in. {5 cm} of offset from the mold line for internal components. For a
small general aviation aircraft, I-in. clearance or less might be acceptable.
The type of internal component will affect the required clearance. A jet
engine contained within an aluminum or composite fuselage will require
perhaps an additional inch of clearance to allow for a heat shield. The heat
shield can be constructed of titanium, steel, or a heat-proof matting. On
the other hand, an "integral" fuel tank in which the existing structure is
 CHAPTE R 8 Spec i a l Consideratio n s i n Confi g u ration Layout 231

simply sealed and filled with fuel will require no clearance other than the
thickness of the skin.
There is no easy formula for the estimation of structural clearance. The
designer must use judgment acquired through experience. The best way to
gain this judgment other than actual design experience is by looking at
existing designs.

#:fJI Flutter
Flutter is an unfortunate dynamic interaction between the aerodynamics
and the structure of an aircraft. It occurs when some structural deflection of
the aircraft such as wing bending causes an aerodynamic load that tends to
amplify the deflection during each oscillation until structural failure is
reached. There are many possible flutter modes. An aileron with its center
of mass well behind its hinge line will tend to lag when accelerated
upwards by oscillating wing bending. This lagging is similar to a flap deflec­
tion, increasing the wing lift and amplifying the wing bending. On the way
back down, the aileron lags upward, driving the wing down even further.
Similar flutter modes occur in elevators and rudders that have center of
masses behind their hinge lines. Early Learjets were crashing because water
was freezing inside the elevators behind the hinge line, causing flutter. This
was difficult to uncover because, of course, the ice melted by the time the
accident investigators got to the scene. Even a trim tab or servo tab can
cause flutter if it has its center of mass behind its hinge line (see Fig. 8.15).
The solution to this control surface flutter is obvious: don't allow the
center of mass to be behind the hinge line! Instead, add mass balancing in
the form of weight ahead of the hinge line, and ruthlessly avoid weight
behind it. A control surface is said to be statically balanced if its chordwise
center of gravity is on its hinge line. Many World War II vintage planes
had fabric-covered control surfaces to keep the center of gravity forward to
avoid flutter.
Complete balancing of a control surface requires the product of inertia
about the hinge axis to be zero. This leads to placing balance weights near
the tips of control surfaces to reduce the product of inertia. Dynamic
balance is obtained when a control surface moves with its wing or tail
without any tendency towards relative rotation between the two, so they
act as if they were welded together.
Control surface flutter is more likely if there is play (looseness) in the
control linkages or play in the trim tab linkage. For this reason, stiff
pushrod linkages are preferred over wire cables. Also, pilots should always
inspect control linkages before flight.
The shaping of the control surfaces has an effect on flutter. They should
never be convex, bulging out into the airflow, because it sets up unstable flow
at the trailing edge. Instead, they should be flat-sided or concave. It is
232 A i rcraft De s i g n : A C o n c e p tu a l A p p r o a c h

I n it i a l nose- u p d istu rba n ce

Pitch sta b i l ity p ro d u ces

nose-down resto r i n g moment
res u l t i n g i n rotationa l a cceleration
Mass of e levator resists
a cceleration, downward
d eflection g i ves more l ift

Fig. 8. 1 5 Elevator lag pitching flutter.

desirable to have a beveled trailing edge, and a control surface that is

"fattened" at the hinge line will tend to reattach the flow, improving flutter
characteristics.
It is bad for flutter if the natural frequency of the vibration of the aileron
about its hinge is nearly the same as the wing natural bending frequency
(analysis of this requires complicated calculations or shake-testing). It is
also bad if the tip of the aileron is in the wing-tip vortex. For this reason, ailer­
ons should not extend all the way to the wing tip. Another potential source of
flutter problems is an excessive amount of aileron aerodynamic balance. If a
deflecting aileron produces almost no restoring moments, flutter can result.
There is less of a tendency for a fuselage torsional flutter if the rudder is
halfway below the fuselage rather than solely above it, mounted on the
vertical tail. To increase their relative torsional stiffness, a rigid torque tube
should connect left-and right-side elevators.
Another type of flutter that has nothing to do with control surface
problems is called wing flexure-torsion binary flutter. In it, a torsional
vibration or oscillation of the wing sets up aerodynamic forces in phase
with an up-and-down wing bending flexural motion. The wing is bending
up and down and twisting at the same time (in-phase oscillation) such that
it has a positive angle of attack when the wing is going up and a negative
angle of attack when it is going down. The resulting change in lift amplifies
 C H A PT E R 8 Special Considerations in Configu ration Layout 233

the up and down bending, possibly leading to flutter and divergence (wing
breaks). This can be avoided by increasing the wing's torsional rigidity and
by keeping the wing's chordwise center of gravity at or in front of the
wing's structural elastic axis. In other words, avoid any weight behind
roughly the middle of the wing, and try to give the wing a strong and rigid
box structure.
Yet another type of flutter is a problem peculiar to high-speed aircraft.
Aerodynamic forces on structural panels can set up an in-and-out
oil-can-like flutter, with the potential to rip the panel right off the aircraft.
This is avoided by making sure that the panels do not have too great an
unsupported length, or by using honeycomb panels or some other stiffened
skin. This panel flutter is not typically addressed in conceptual design.

9 Radar Detectabil ity

Ever since the dawn of military aviation, attempts have been made to
reduce the detectability of aircraft. During World War I, the only "sensor"
in use was the human eyeball. Camouflage paint in mottled patterns was
used on both sides to reduce the chance of detection.
Radar (acronym for radio detection and ranging), the primary sensor used
against aircraft today, consists of a transmitter antenna that broadcasts a
directed beam of electromagnetic radio waves and a receiver antenna that
picks up the faint radio waves that bounce off objects "illuminated" by the
radio beam. Usually the transmitter and receiver antennas are collocated
("monostatic radar"), although some systems have them in different locations
("bistatic radar").
Detectability to radar has been a concern since radar was first used in
World War IL "Chaff' was an early counter-radar "stealth" technology.
Chaff, also called "Window," consists of bits of metal foil or metallized
fibers dropped by an aircraft to create many radar echos that hide its
actual echo return. Chaff is still useful against less sophisticated radars.
Chaff obscures the actual location of the aircraft, but does not allow the
aircraft to pass unnoticed. To avoid detection, the aircraft must return such a
low amount of the transmitted radio beam that the receiver antenna cannot
distinguish between it and the background radio static.
Radar stealth treatments go back further than many people realize. In
World War II, German U-boats had RAM-covered periscopes to avoid detec­
tion. The first jet flying wing in the world, the Horten IX (which flew in 1945),
used a charcoal-and-glue RAM and configuration shaping to reduce its
signature. By 1960 the U.S. Air Force had flown a T-33 covered entirely
with RAM and a B-47 with its inlets covered by screens and its exhausts
obscured. U-2s were flown with RAM and other more exotic RCS-reduction
techniques. The North American Aviation Hound Dog air-launched cruise
missile (operationally carried by the B-52) incorporated RAM, and NAA's
F-X proposal for the F- 1 5 had inlet ducts treated with RAM, in 1969.
234 A i r c ra ft Design : A C o n c eptu a l A p proa c h

By the 1970s most U.S. military aircraft companies had the ability to
design for stealth in terms of overall configuration shaping. This included
sloping the fuselage sidewalls; hiding inlets and engine front and rear faces;
sweeping the edges of the wing, tail, and other edges; and similar fundamental
stealth techniques. However, that is only part of the capability required to be
a qualified "stealth house." Other key capabilities include the ability to
analytically estimate signature, the technologies for stealth treatments of
surfaces, edges, and details such as access doors and running lights, and
the technology for "stealthy" integration of avionics, including radomes.
Two companies were clear leaders in these areas and, as a result, were the
most successful in development of actual stealth aircraft. Lockheed gained
such expertise through its development of spy planes (U-2 and SR-71 ),
while Northrop apparently made a corporate decision in the 1960s that
stealth was a critical emerging technology and invested accordinglyJ2 7l
Development of radar stealth technology capable of making an aircraft
operationally undetectable was accelerated by the Defense Advanced
Research Projects Agency (DARPA) Project Harvey, begun around 1970
and named after the invisible rabbit in the play of the same name. This classi­
fied program led to the Have Blue flight test demonstrator, awarded to Lock­
heed over Northrop, the only other serious competitor. Have Blue led in turn
to the operational F- 1 17, which proved the operational worth of stealth.
The fundamental mathematical relationships governing radar cross
section and other electromagnetic phenomena are Maxwell's equations,
defined over 150 years ago. These, like the Navier-Stokes equations for aero­
dynamics (see Chapter 1 2), are complete governing equations and, if solved,
would tell us everything we want to know! However, like the Navier-Stokes
equations, they are currently impossible to solve exactly in their complete
form for any complicated geometry, so we solve simplified versions of the
equations to attempt to predict RCS. These simplified versions of Maxwell's
equations can only consider a limited version of the physical phenomena that
cause radar energy to return.
The extent to which an object returns electromagnetic energy is the
object's radar cross section (RCS). RCS is usually measured in square
meters or in decibel square meters, with zero dBsm equal to 10 to the zero
power, or 1 m 2 . Twenty dBsm equals 10 to the second power, or 100 m2 .
Because radar signal strength is an inverse function of the fourth power of
the distance to the target, it takes a very substantial reduction in RCS to
obtain a meaningful operational benefit.
The RCS of an aircraft is not a single number. The RCS is different for
each "look-angle" (i.e., direction from the threat radar) . When graphed in
polar coordinates, RCS appears as shown in Fig. 8.16. These are actual data
for the B-70 supersonic bomber. As can be seen, RCS varies widely from
different directions, by almost four orders of magnitude for this design.
We use the expression "spikes" to describe directions from which the RCS
of an aircraft is very high. These are typically perpendicular to the leading and
 CHAPTER 8 Spec i a l Consideration s i n Confi g u ration Layout 235

/
,,....---- -

S q u a re

/ /
meters
100,000
Red uction with
10,000
RAM
\

\
I \
I I 1000

I I
\
\

/
/
\
I
\ I

\
/
/ I

Fig. 8. 1 6 Radar crass section i n polar coordinates.

trailing edges of the wing, perpendicular to the flat side of the aircraft unless
it is properly shaped and treated, and directly off the nose and tail due to
the inlets, nozzles, radome, and other features. For the B-70, huge spikes
are evident to the sides perpendicular to the big flat sides of the nacelle.
However, with a cruising speed of Mach 3.0 at almost 80,000 ft
{24,300 mg}, it would be difficult to intercept a B- 70 that was already flying
past. Unfortunately, there are also substantial spikes just off the nose, perpen­
dicular to the leading edges of the canards, which have fairly low sweep.
These spikes would have warned defenders that a B-70 was coming. As
shown, treatment with RAM was under serious consideration for operational
B-70s. This would have reduced the nose-on signature by several orders
of magnitude.
Actual signature levels are, for obvious reasons, highly classified numbers.
Reference [28] gives the signature of the B-52 as 100 m2 , or 20 dBsm, and the
stealth-treated B- lB as 1 m2 , or 0 dBsm. The Lockheed A- 12, similar to the
SR-71 and highly treated for stealth using the technology of the early 1960s,
is quoted in [28l as having an RCS of 0.014 m2 or -8 dBsm. Nonstealth fighters
typically have nose-on signatures on the order of 10 sqm, or 10 dBsm. The
stealthy MiG 1 .42 fighter technology demonstrator is quoted by MiG as
having an RCS of 0.1 sqm, or - 1 0 dBsm. Reference [29] suggests that
"where stealth is a primary design objective, RCS will probably be in the
region of 0.01 to 0.1 square meter" {-20 to -10 dBsm}.
236 A i rcraft Desig n : A C o n c eptu a l A p p r o a c h

RCS varies depending upon the frequency and polarization of the threat
radar (see [30 - 32l ). The following comments relate to typical threat radars
seen by military aircraft.
There are many electromagnetic phenomena that contribute to the RCS
of an aircraft. These require different design approaches for RCS reduction
and can produce conflicting design requirements. Figure 8.17 illustrates
the major RCS contributors for a typical, untreated fighter aircraft.
One of the largest contributions to airframe RCS occurs any time a
relatively flat surface of the aircraft is perpendicular to the incoming radar
beam. Imagine shining a flashlight at a shiny aircraft in a dark hanger. Any
spots where the beam is reflected directly back at you will have an enormous
RCS contribution.
Typically this "specular return" occurs on the flat sides of the aircraft
fuselage and along an upright vertical tail (when the radar is abeam the
aircraft). To prevent these RCS "spikes," the designer can slope the fuselage
sides, angle the vertical tails, and so on, so that there are no flat surfaces
presented toward the radar (Fig. 8.18) .
Note that this RCS reduction approach assumes that the designer knows
where the threat radar will be located relative to the aircraft. This information
is usually provided by the operations-analysis department or by the customer
as a design driver. Also, this assumes a monostatic radar.
Another area of the aircraft that can present a perpendicular bounce for
the radar is the round leading edge of the wing and tail surfaces. If the aircraft

F lat side
of ta i l

Exhaust
cavity

Cockpit

Lea d i n g
sides Missile
edges
of i n stal lation
cavity fuse lage gaps and
Radome i rreg u l a rities

Fig. 8. 1 7 Major RCS contributors.

 C H A PT E R 8 Spec i a l Consideratio n s i n Confi g u ration Layout 237

� Radar

High RCS

� Radar
Lower RCS

Fig. 8. 1 8 Flat-side RCS reduction.

is primarily designed for low detectability by a nose-on threat radar, the wings
and tails can be highly swept to reduce their contribution to RCS. Note that
this and many other approaches to reducing the RCS will produce a penalty
in aerodynamic efficiency.
Aircraft cavities such as inlet front faces and engine exhausts create a
radar return perpendicular to the plane of the opening. All around the
opening there will be small perpendicular bounces. When the threat radar
is at a direction perpendicular to the opening, those small bounces will be
"in phase" and so will sum to a single large ,return. This is avoided by
sweeping the plane of the opening well away from the expected directions
of threat radars, as can be seen on the F-22, B- lB, F / A- 18£, and other
designs. To further reduce this RCS contribution, the inlet lips are often
treated with radar absorbers.
It is also important to avoid any "corner reflectors," that is, intersecting
surfaces that form approximately a right angle, as shown in Fig. 8.17 at the
wing-fuselage junction.
Another contributor to airframe RCS occurs due to the electromagnetic
currents that build up on the skin when illuminated by a radar. These cur­
rents flow across the skin until they hit a discontinuity such as at a sharp
trailing edge, a wing tip, a control surface, or a crack around a removable
panel or door. At a discontinuity, the currents "scatter," or radiate electro­
magnetic energy, some of which is transmitted back to the radar (Fig. 8.19) .
This effect i s much lower i n intensity than the specular return, but i s still
sufficient for detection. The effect is strongest when the discontinuity is
straight and perpendicular to the radar beam. Thus, the discontinuities
such as at the wing and tail trailing edges are usually swept to minimize
the detectability from the front. Carried to the extreme, this leads to
diamond- or sawtooth-shaped edges on every door, access plate, and other
discontinuity on the aircraft, as seen on the B-2 and F- 1 17.
238 A i rcraft D es i g n : A C o n c e ptu a l A p p ro a c h

Edge
scattering

Fig. 8. 1 9 Surface current scatterings.

This scattering of surface currents actually represents three different

types of radar returns. If the surface is being directly illuminated by the
radar, a surface discontinuity causes diffraction, which is the same phenom­
enon that causes a rainbow. Diffraction will occur not just at a physical edge,
such as a wing trailing edge, but any place that the surface has a sharp corner.
There is even an apparent shadow edge anyplace where the surface is no
longer illuminated by the radar energy, such as at the transition from front
to back of a wing airfoil when illuminated from the front.
Two other radar returns, from traveling waves and creeping waves, are
caused by the flow of electromagnetic energy from the front to the back,
nonilluminated side of the body. Traveling waves occur when a sharp discon­
tinuity is reached and the energy (which cannot be destroyed) travels back to
the front where it reradiates. Traveling waves and edge diffraction both call
for the avoidance of trailing edges perpendicular to the threat radar.
Creeping waves occur when the backside of the body is smoothly curved,
so the energy "creeps" all the way around the body, slightly radiating as it
goes around. Radar absorbers, as discussed next, are useful for suppressing
creeping waves. References (33] and (34] are suggested for discussions of
RCS theory.
First-generation stealth designs such as the Lockheed F- 1 17 and the never
constructed North American Rockwell "Surprise Fighter" relied upon faceted
shaping in which the aircraft shape is constructed of interlocking flat
triangles and trapezoids. This has advantages in ease of construction and
 CHAPTER 8 Spec i a l Consideratio n s i n Confi g u ration Layout 239

signature analysis, but offers a large number of sharp edges to create

diffraction returns, and so is no longer in favor. [35]
Current stealth design begins with the aiming of the RCS spikes. For
good stealth design, we "aim them where the bad guys ain't," based on
requirements as to what signatures are required from what directions.
This starts with a decision as to what directions pose a severe threat, and
what directions pose a lesser threat. Typically, we assume that the forward
direction poses a severe threat because we are flying toward someone that
we plan to attack. Toward the rear we have a severe threat because we
have just attacked them and are running away, and they are really angry!
Around the sides of the aircraft at similar altitudes we assume a lesser,
but still significant threat, based on the supposition that opponents at our
altitude can most readily attack us. Directly above and below us are not as
likely to pose a threat. The design requirements of a stealth aircraft will
include desired levels of RCS from different directions (azimuth and
elevation), such as, "no greater than (classified) decibel square meters
within (classified) degrees off the . nose at plus-or-minus (classified)
degrees elevation."
Given the definition of the directions of likely threat, we can select where
to aim the spikes by appropriately aligning the spike-producing features
such as wing leading and trailing edges. With current technology, all aircraft
will have at least four spikes, namely, the perpendicular bounces and edge
diffractions from the leading and trailing edges of the wing. From the direc­
tions of those spikes, a threat radar will be able to see the aircraft anyway,
so we align any additional spikes in the same direction rather than allowing
a spike at another direction. To paraphrase an old song, "one big spike is
better than two little spikes" when it comes to stealth design.
For example, the edge diffractions off the wing trailing edge when seen
from the front will create a spike perpendicular to that trailing edge. By
setting the wing trailing edge to the same angle as a wing leading edge, the
spikes align. If we align the left trailing edge with the right leading edge,
we get a diamond planform as seen to the left of Fig. 8.20. Aligning the left
trailing edge with the left leading edge generates an untapered planform
(highly swept to avoid spikes to the front and rear) .
As was discussed in Chapter 4, an untapered wing (taper ratio of one) or a
wing with a taper ratio of zero as in the diamond wing are the worst possible
wings from an aerodynamic standpoint. The untapered wing has excessive
lift outboard, especially if swept, while the diamond wing has insufficient
lift outboard to form an elliptical lift distribution as desired for minimum
drag due to lift. However, if we combine these aerodynamically bad
planforms as shown at the bottom of Fig. 8.20 and carefully twist and
camber the resulting planform, we can obtain a fairly good aerodynamic
efficiency. This illustration is quite similar to the original B-2 configuration,
but it was later revised to the current configuration to better balance the
design and reduce the length of the external exhaust run.
240 A i r c raft Des i g n : A Conceptual A p p r o a c h

Fig. 8.20 "Aiming the spikes" for RCS reduction.

Smaller, but nontrivial spikes also arise from the edges of an access door,
landing-gear door, or weapons bay door. Where possible, we design such
doors rotated roughly 45 deg so that the edges align with the existing
spikes from the wing leading and trailing edges, creating the characteristic
diamond shape. If this is not feasible, we put sawtooth edges on the doors
to avoid strong spikes forward and to the rear.
This design approach leads to an aircraft planform composed entirely of
straight, highly swept lines, much like the first-generation stealth designs.
However, the desire to eliminate the edge diffractions caused by the facets
of first-generation stealth now produces designs in which cross-sectional
shapes are smooth, not sharp-edged. The steep angles on the fuselage sides
as shown in Fig. 8.18 are employed to prevent broadside perpendicular
bounce returns, but these angled sides flow smoothly over the top and
bottom of the fuselage. Such shaping can be seen on the B-2, F-22, F-23,
and F-35 Joint Strike Fighter (JSF) and is apparent in this notional
fighter design developed for pre-JSF requirements trade studies at RAND
Corporation ( [3 6l , see Fig. 8.21).
RCS can also be reduced simply by eliminating parts of the aircraft. A
horizontal tail that does not exist cannot contribute to the radar return!
Modern computerized flight controls combined with the use of vectored­
thrust engines can solve many of the difficulties of the tailless configuration.
This author expects that eventually fighters will be designed with neither
vertical nor horizontal tails (no canards, either) to
minimize signature, with vectored nozzles and Get rid of things­
forebody vortex control used to control the aircraft. the radar can't see
Similarly, RCS can be reduced if the nacelles can it if it isn't there!
be eliminated through the use of buried engines, or
 C HA PT E R 8 Specia l Considerations in Configuration Layout 24 1

b etter yet, by eliminating the entire fuselage through the use of the
flying-wing concept. This approach is used in the Northrop B-2.
In addition to reshaping the aircraft, detectability can be reduced through
the use of skin materials that absorb radar energy. These are called radar
absorbing materials (RAM) and are typically carbon or ferrite particles
embedded in a "binder," which can be a composite matrix material such as
urethane, or a type of silicone, or certain ceramics for high-temperature
appli cations.
These particles are heated by the radar electromagnetic waves, thus
absorb ing some of the energy. This will reduce (not eliminate!) the radar
return due to perpendicular bounce; it can also reduce the surface currents
and thus reduce the RCS due to scattering at sharp edges. The thickness of
the radar absorbing material should be about one-fourth of the wavelength
of the threat radar.
RAM can be applied parasitically, to the outside of the structure as
attached non-load-bearing panels or even as a paint. RAM can also be
built into the aircraft's structural material, which is then called radar

Fig. 8.2 1 Next-generation attack fighter (D. Raymer 1 994) .

242 A i rc raft Des i g n : A C o n c e ptu a l A p p r o a c h

absorbing structure, or RAS. A typical RAS is a honeycomb panel with

Kevlar-epoxy skins that are transparent to radar, an inner skin of graphite­
epoxy that is reflective of radar, and a Nomex honeycomb core in between
that includes radar absorbers in increasing density from the outside to the
inside so as to gradually trap the radar energy.
Each time the radar energy bounces off RAM, it loses more energy
(attenuates), so the geometry should be designed to force multiple
bounces. This is especially suitable in an inlet duct, where a long and
curved duct will cause the radar energy to bounce of the sidewalls many
times. Eventually, most of the radar energy will be absorbed.
As there are many types of RAM and similar treatments, no quick esti­
mate for the weight impact of their use can be provided here. However,
one can probably assume that such use will reduce or eliminate any weight
savings otherwise assumed for the use of composite materials.
For most existing aircraft, the airframe is not the largest contributor to
RCS, especially nose-on. A conventional radome, covering the aircraft's
own radar, is transparent to radar for obvious reasons. Therefore, it is also
transparent to the threat radar, allowing the threat radar's beam to bounce
off the forward bulkhead and electronic equipment within the radome.
Even worse, the aircraft's own radar antenna, when illuminated by a
threat radar, can produce a radar magnification effect much like a cat's eye.
These effects can be reduced with a "bandpass" radome, which is transparent
to only one radar frequency (that of the aircraft's radar) .
Other huge contributors to the RCS for a conventional aircraft are the
inlet and exhaust cavities. Radar energy gets into these cavities, bounces off
the engine parts, and sprays back out the cavity towards the threat radar.
Also, these cavities represent additional surface discontinuities.
The best solution for reducing these RCS contributions is to hide them
from the expected threat locations. For example, inlets can be hidden from
ground-based radars by locating them on top of the aircraft (Fig. 8.22).
Exhausts can be hidden through the use of two-dimensional nozzles.
The F- 1 1 7 used a mesh screen at the front of the inlet duct to keep out the
radar energy. For this to work, the mesh must be smaller than the radar's
wavelength, leading to a loss of inlet pressure recovery that in turn reduces
thrust and increases fuel consumption. Also, icing becomes a concern.
More recent stealth designs allow the radar energy into the inlet duct, but
uses RAM inside the duct to absorb it as already described. Also, if the
radar energy is allowed inside, some provision must be made for hiding
any direct view of the engine front face from the outside. This can be done
by extreme snaking of the duct, or by putting curved vanes or an onion­
shaped bulb in front of the engine. If such devices are put into the duct,
care must be taken that the mean flowpath does not decrease in cross­
sectional area, and provisions for anti-icing might be required.
Cockpits provide a radar return for a similar reason. The radar energy
enters the cockpit, bounces around off the equipment inside, and then
 CHAPTER 8 Spec i a l Considerations i n Confi g u ration Layout 243

Fig. 8.22 Detectability reduction approaches.

reradiates back outside. One solution for this is to thinly coat the canopy with
some conductive metal such as gold, causing the canopy to reflect the radar
energy away.
Finally, the aircraft's weapons can have a major impact on RCS. Missiles
and bombs have fins that form natural corner reflectors. The carriage and
release mechanisms have numerous corner refle ctors, cavities, and surface
discontinuities. Gun ports present yet another kind of cavity. The only real
solution for these problems is to put all the weapons inside, behind closed
doors. However, the weight, volume, and complexity penalties of this
approach must be carefully considered.
Electronic countermeasures (ECM) -devices to trick the threat radar­
usually consist of some sort of radar receiver that picks up the threat radar
emissions and some sort of transmit antenna to send a deceiving signal
back to the threat radar. The many techniques for tricking radar (and
ECM) go beyond the scope of this book. However, designers should be
aware that there is a tradeoff between the aircraft's RCS level and the required
amount of ECM.

Infrared Detectab i l ity

Infrared (IR) detectability also concerns the military aircraft designer.
Many short-range air-to-air and ground-to-air missiles rely upon IR
seekers. Modern IR sensors are sensitive enough to detect not only the
radiation emitted by the engine exhaust and hot parts, but also that
emitted by the whole aircraft skin due to aerodynamic heating at transonic
244 A i rc raft D e s i g n : A C o n ceptu a l Approach

and supersonic speeds. Also, sensors can even detect the solar IR radiation
that reflects off the skin and cockpit transparencies (windows).
IR detectability can be reduced by reducing the engine exhaust tempera­
tures with a high-bypass-ratio engine. These have large fans up front, who se
cool airflow can be mixed with the hotter turbine exhaust before it exits the
nozzle. This reduces both exhaust and hot-part temperatures. However,
there might be a performance penalty especially at higher speeds.
Emissions from the exposed engine hot parts (primarily the inside of the
nozzle) can be reduced by cooling them with air bled off the engine compres­
sor. This will also increase fuel consumption slightly. Another approach hides
the nozzles from the expected location of the threat IR sensor. For example,
the H-tails of the A- 10 hide the nozzles from some angles. Unfortunately, the
worst-case threat location is from the rear, and it is difficult to shield the
nozzles from that direction!
Plume emissions are reduced by quickly mixing the exhaust with
the outside air. As mentioned, a high-bypass engine is the best way of
accomplishing this. Mixing can also be enhanced by the use of a wide,
thin nozzle rather than a circular one. Another technique is to angle the
exhaust upward or downward relative to the freestream. This will have an
obvious thrust penalty, however.
Sun glint in the IR frequencies can be somewhat reduced by the use of
special paints that have low IR reflectivity. Cockpit transparencies (which
can't be painted!) can be shaped with all flat sides to prevent continuous
tracking by an IR sensor.
Emissions due to aerodynamic heat are best controlled by slowing the
aircraft down.
IR missiles can sometimes be tricked by throwing out a flare that burns to
produce approximately the same IR frequencies as the aircraft. However,
modern IR seekers are getting better at identifying which hot source is the
actual aircraft.
IR fundamentals are more thoroughly discussed in [37l .

Visual Detectab i l ity

The human eyeball is still a potent aircraft detection sensor. On a clear
day, an aircraft or its contrail can be spotted visually before detection by
the onboard radar of a typical fighter. Also, fighter aircraft usually have
radar only in front, which leaves the eyeball as the primary detector for
spotting threat aircraft that are abeam or above.
Visual detection depends upon the size of aircraft and its color and
intensity contrast with the background. In simulated combat, pilots of the
small F-5 can frequently spot the much-larger F-15s well before the F-5s
are seen. However, aircraft size is determined by the mission requirements
and cannot be arbitrarily reduced.
 C H A PT E R 8 Spec i a l Consideratio ns i n Confi g u ratio n Layout 245

Background contrast is reduced primarily with camouflage paints, using

colors and surface textures that cause the aircraft to reflect light at an inten­
sity and color equal to that of the background. This requires assumptions as
to the ap propriate background as well as the lighting conditions.
Fre quently aircraft will have a lighter paint on the bottom, because the
backgro und for look-up angles is the sky. Current camouflage paint
schemes are dirty blue-grey for sky backgrounds and dull, mottled grey­
greens and grey-browns for ground backgrounds.
Diffe rent parts of the aircraft can contrast against each other, which
in crea ses detectability. To counter this, paint colors can be varied to
lighten the dark areas, such as where one part of the aircraft casts a
shadow on another. Also, small lights can sometimes be used to fill in a
shadow spot .
Canopy glint is also a problem for visual detection. The use of flat trans­
parencies can be applied as already discussed, but will tend to detract from
the pilot's outside viewing.
At night, aircraft are visually detectt". d mostly by engine and exhaust glow
and by glint off the transparencies. These can be reduced by techniques
already discussed for IR and glint suppression.
There are also psychological aspects to visual detection. If the aircraft
does not look like an aircraft, the human mind might ignore it. The irregular
mottled patterns used for camouflage paints exploit this tendency.
In air-to-air combat, seconds are precious. If a pilot is confused as to
the opponent's orientation, the opponent can obtain favorable positioning.
To this end, some aircraft have even had fak� canopies painted on the
underside. Forward-swept and oblique wings can also provide momentary
disorientation.

Aura l Signature
Aural signature (noise) is important for civilian as well as military aircraft.
Commercial airports have anti-noise ordinances that might restrict some air­
planes. Aircraft noise is largely caused by airflow shear layers, primarily due
to the engine exhaust.
A small-diameter, high-velocity jet exhaust produces the greatest noise,
whereas a large-diameter propeller with a low tip-speed produces the least
noise. A turbofan falls somewhere in between. Blade shaping and internal
duct shaping can somewhat reduce noise.
Because much of the noise comes from the exhaust shear layers, anything
that promotes rapid mixing between the exhaust stream and the outside air
will reduce noise. Some jet engines have a special nozzle called a "daisy mixer"
that looks, from the rear, like the flower. Rather than a circular exhaust pipe,
the final nozzle shape goes in and out (Fig. 8.23a). The exhaust follows this
contour and continues in that shape as it leaves the nozzle, increasing the
246 A i rc raft Desi g n : A C o n c e pt u a l A p p r oa c h

b)

Fig. 8.23 Two approaches to exhaust noise reduction .

mixing surface between the exhaust and the outside air. Another approach is
seen in the Boeing 787 nacelle where the fan's exit nozzle is cut away in a
wedge-like pattern. The high-pressure fan air blows outward a bit through
the cutout portions, creating a flow pattern downstream that is just like
that of the daisy mixer (Fig. 8.23b).
There is also mechanical noise from jet engines-spinning bearings ,
vibrations, air slapping against compressor blades, and the like. Accessory
drives can also create noise.
Piston exhaust stacks are an obvious source of noise. This noise can be
controlled with mufflers and by aiming the exhaust stacks away from the
ground and possibly over the wings. Mufflers are heavy, though, so many
general aviation airplanes have small mufflers or sometimes, none at all.
Recent research has discovered, surprisingly, that the airflow around the
extended landing gear and flaps has a large contribution to the noise heard
when a big plane flies overhead. Aerodynamic "cleanup" including stream­
lined fairings and better-designed linkages has been proven to reduce noise
substantially.
Within the aircraft, noise is primarily caused by the engines. Well­
designed engine mounts, mufflers, and insulation materials can be used to
reduce the noise. Internal noise will be created if the exhaust from a piston
engine impinges upon any part of the aircraft, especially the cabin.
Wing-mounted propellers can have a tremendous effect on internal noise
if they are too close to the cabin. Propellers should have a minimum clear­
ance to the fuselage of about 1 ft {30 cm} and should preferably have an
even greater clearance of about one-half of the propeller radius.
However, the greater the propeller clearance, the larger the vertical
tail must be to counter the engine-out yaw. Some airplanes have the propel­
lers so close to the fuselage that you can barely slide your fingers between
them!
 C H APTE R 8 Spec i a l Considerations i n Config u ration Layout 247

Jet engines mounted on the aft fuselage (DC-9, B727, etc.) should be
lo ed as far away from the fuselage as structurally permitted to reduce
cat
cabin noise. Also, they should be located as far aft as possible, preferable
aft of the cabin pressure vessel.
The traditional approach to in-cabin sound suppression has been heavy
insul ation blankets, strategically located to block the noise. A newer technol­
ogy called "active sound suppression" uses a microphone to detect noise in
the cabin then employs a speaker to send a noise signal 180 deg out of
phase, cancelling the cabin noise. Although not perfect, this system works
well on aircraft such as the SAAB 2000.

Vu lnerabil ity Considerations

Vulnerability concerns the ability of the aircraft to sustain battle damage,
co nti nue flying, and return to base. An aircraft can be "killed" in many ways.
A single bullet through a nonredundant elevator actuator is as bad as a big
missile up the tailpipe! ,
"Vulnerable area" is a key concept. This refers to the product of the
projected area (square feet or meters) of the aircraft components times the
probability that each component will, if struck, cause the aircraft to be lost.
Vulnerable area is different for each threat direction. Typical components
with a high aircraft kill probability (near 1 .0) are the crew compartment,
engine (if single engine), fuel tanks (unless self-sealing), and weapons.
Figure 8.24 shows a typical vulnerable-area calculation.
When assessing the vulnerability of an aircrqft, the first step is to deter­
mine the ways in which it can be "killed." Referred to as a "failure modes
and effects analysis" (FMEA), or "damage modes and effects analysis"
(DMEA), this step will typically be performed during the later stages of
conceptual design. The FMEA considers both the ways in which battle
damage can affect individual aircraft components and the ways in which
damage to each component will affect the other components. "Failure
mode, effects and criticality analysis" (FMECA) adds study of the probability
of failure modes vs the impact of their occurrence.
During initial configuration layout, the designer should strive to avoid
certain features known to cause vulnerability problems. Fire is the greatest
danger to a battle-damaged aircraft. Not only is the fuel highly flammable,
but so is the hydraulic fluid. The second Have Blue stealth demonstrator
crashed due to a crack in a weld in a hydraulic line, which sprayed fuel on
the engine.
If at all possible, fuel should not be located over or around the engines
and inlet ducts. While tanks can be made self-sealing to a small puncture,
a large hole will allow fuel to ignite on the hot engine. The pylon-mounted
engines on the A-10 insure that leaking fuel cannot ignite on the engines.
Similarly, hydraulic lines and reservoirs should be located away from
the engines.
248 A i rc raft D e s i g n : A C o n ceptua l Approach

S a m p l e calcu lation

Component P resented a rea Pk given hit Vu l nera b l e a rea

P i l ot (a) 5 ft2 1 .0 5 ft2

C o m p u te r (b) 4 ft2 0.5 2 ft2
F u e l (c) 80 ft2 0.3 24 ft2
E n g i n e (d) 50 ft2 0.4 20 ft2
Total v u l n erable a rea 51 ft2

Fig. 8.24 Vulnerable-area calculation .

Firewalls should be used to prevent the spread of flames beyond a burning

engine bay. Engine bays, fuel bays, and weapon bays should have a fire­
suppression system.
When an engine is struck, turbine and compressor blades can fly off at
high speeds. Avoid placing critical components such as hydraulic lines or
weapons anyplace where they could be damaged by an exploding engine.
Also, a twin-engine aircraft should have enough separation between
engines to prevent damage to the good engine. If twin engines are together
in the fuselage, a combined firewall and containment shield should separate
them. This requires at least 1 ft {30 cm} of clearance between engines.
Combat aircraft carry gun ammo, bombs, and missiles. An aircraft can
survive a burst of cannon shells only to explode from a fire in the ammo box.
Propeller blades can fly off either from battle damage or during a
wheels-up landing. Critical components, especially the crew and passenger
compartments, shouldn't be placed within a 5-deg arc of the propeller disk.
Avoid placing guns, bombs, or fuel near the crew compartment. Fuel
should not be placed in the fuselage of a passenger plane.
Redundancy of critical components can be used to allow the survival of
the aircraft when a critical component is hit or fails for any other reason.
Typical components that could be redundant include the hydraulic system,
electrical system, flight control system, and fuel system. Note that while
 C H A PT E R 8 Spec i a l Considerati o ns i n Confi g u ration Layout 249

redundancy improves the survivability and reliability, it worsens the mainten­

ance requirements because there are more components to fail.
While normally considered a topic for military aircraft, the concepts for
re cing vulnerability also apply to civil aircraft. FMEA should be conducted
du
to minimize the possibility that a failure or damage in one system can cause
the aircraft to crash. One concern for commercial airliners is that turbine
engines sometimes lose turbine and compressor blades even without being
shot at! Those should not pierce the passenger compartment, nor fly
across to another engine and damage it.
For more information on vulnerability, l37l is again suggested.

Crashworthiness Considerations
Airplanes crash. Careful design can reduce the probability of injury in a
moderate crash. Several suggestions have already been mentioned, including
positioning the propellers so that the blades will not strike anyone if they fly
off during a crash. Also mentioned was. the desire to avoid placing fuel tanks
in the fuselage of a passenger airplane (although fuel in the wing box carry­
through structure is usually considered acceptable).
To protect the crew and passengers in the event of a crash, the aircraft
should be designed to act like a shock absorber. A shock absorber works by
deflecting in a controlled fashion, spreading the load from a sudden impact
over a specified distance (the "stroke") and over time (see Chapter 1 1). The
aircraft's structure can be designed to work the same way, crushing in a
controlled fashion over distance and time. Helicopters are routinely designed
in this way, with extensive analysis and test of the deflections of the structure
during a crash.
For aircraft, one can see the benefits of collapsing structure very starkly
when studying accidents of low-wing general aviation aircraft. It is tragically
common that the back-seat passengers will survive a crash, while the pilot
and front seat passenger, who are sitting on the hard, noncollapsing wing
box, will not survive. There is some concern that composite structures,
which tend to be very stiff and do not deflect so readily during a crash,
might be less survivable in accidents.
Figure 8.25 shows several other design suggestions that were learned the
hard way. A normal, vertical firewall in a propeller aircraft has a sharp lower
corner that tends to dig into the ground, stopping the aircraft dangerously
fast. Sloping the lower part of the firewall back as shown will prevent
digging in, therefore reducing the deceleration.
For a large passenger aircraft, the floor should not be supported by braces
from the lower part of the fuselage. As shown, these braces can push upward
through the floor in the event of a crash, unless special collapsing braces
are used.
Common sense will avoid many crashworthiness problems. For example,
things will break loose and fly forward during a crash. Therefore, do not put
250 A i rc raft Desi g n : A C o n c e p t u a l A p p roa c h

This:

Sca rfed fi rewa l l p revents scoo p i n g N o fl oor struts

Not t h i s :

F i rewa l l scoop i n g i n creases c r a s h l o a d s F loor struts p u s h

fl oor u pward

Fig. 8.25 Crashworthiness design.

heavy items behind and/ or above people. This sounds obvious, but there are
some aircraft with the engine in a pod above and behind the cockpit.
There are also some military j ets with large fuel tanks directly behind the
cockpit, offering the opportunity to be bathed in j et fuel during a crash.
However, the pilot would probably try to eject rather than ride out a crash
bad enough to rupture the fuel tanks.
One should also consider secondary damage. For example, landing gear
and engine nacelles will frequently be ripped away during a crash. If possible,
they should be located so that they do not rip open fuel tanks in the process.
Some form of protection should be provided in the not-unlikely event
that the aircraft flips over during a crash. This is lacking in several small
homebuilt designs.

Produci b i l ity Considerations

1:11111 Design for Production
It is often said that aircraft are bought "by the pound." While it is true that
aircraft cost is most directly related to weight, there is also a strong cost
impact due to the materials selected, the fabrication processes and tooling
required (forging, stamping, molding, etc.), and the assembly man-hours.
The configuration designer does not usually determine the materials used
or exactly how the aircraft will be fabricated. However, the ease of producing
the aircraft can be greatly facilitated by the overall design layout.
 C H A PT E R 8 Spec i a l Considerations i n Config u ration Layout 251

One impact the configuration designer has upon producibility is the

extent to which flat-wrap structure is incorporated. This has a major
imp act upon the tooling costs and fabrication man-hours, as discussed in
the last chapter.
Part commonality can also reduce production costs. If possible, the left
and right main landing gear should be identical (left-right common) . It
might be desirable to use uncambered horizontal tails to allow left-right
commonality even if a slight aerodynamic penalty results. In some cases
the wing airfoil can be slightly reshaped to allow left-right common ailerons.
Forgings are the most expensive type of structure in common usage and
are also usually the longest-lead-time items for production tooling. Forgings
can be required whenever a high load passes through a small area. Forgings
are used for landing-gear struts, wing-sweep pivots, and all-moving tail pivots
(trunnions). The designer should avoid, if possible, such highly loaded
structure.
Installation of internal components and routing of hydraulic lines, electri­
cal wiring, and cooling ducts comprise another major production cost due to
the large amount of manual labor required. To ease installation of com­
ponents and routing, avoid the tight internal packaging so desirable for
reduced wetted area and wave drag. When evaluating proposed designs, gov­
ernment design boards will compare the overall aircraft density (weight
divided by volume) with historical data for similar aircraft to ensure
packaging realism.
Routing can be simplified through provision of a clearly defined "routing
tunnel." This can be internal or, as shown in -Fig. 8.26, an external and
nonstructural fairing that typically runs along the spine or belly of the
aircraft. However, if all routing is concentrated in one area, the aircraft
vulnerability will be drastically worsened.
Routing can be reduced by careful placement of the internal components.
For example, the avionics and the crew station will both require cooling air
("environmental control") . If the avionics, crew station, and environmental

Routing
t u n nel

Fig. 8.26 External routing tunnel .

252 Aircraft Des i g n : A Conceptual Approac h

control system (ECS) can be located near to each other, the routing distances
will be minimized.
Sometimes clever design can reduce routing. The Rutan Defiant, a "push­
pull" twin-engine design, uses completely separate electrical systems for the
front and rear engines, including separate batteries. This requires an extra
battery, but a trade study determined that the extra battery weighs less
than the otherwise-required electrical cable and eliminates the front-to-rear
routing requirement.
Another factor for producibility concerns manufacturing breaks. Aircraft
are built in subassemblies as shown in Fig. 8.27. Typically, a large aircraft
will be built up from a cockpit, an aft-fuselage, and a number of mid-fuselage
subassemblies. A small aircraft can be built from only two or three
subassemblies.
It is important that the designer consider where the subassembly breaks
will occur and avoid placing components across the convenient break
locations. Figure 8.28 shows a typical fighter with a fuselage production
break located just aft of the cockpit. This is very common because the
cockpit pressure vessel should not be broken for fabrication.
In the upper design, the nose-wheel well is divided by the production
break, which prevents fully assembling the nose-wheel linkages before
the two subassemblies are connected. The lower illustration shows a better
arrangement.

Fig. 8.27 Production subossemblies of SAAB Droken (courtesy of SAAB Aircraft) .

 CHAPTE R 8 Spec i a l Consi deratio n s i n Confi g u ration Layout 253

Fig. 8.28 Production breaks.

l:ll1fJ Review of Aircraft Fabrication

Design for producibility requires experience that no book can provide. A
good understanding of structural design and fabrication and the basic prin­
ciples of operation for the major subsystems provides the background for
developing producible designs. The following material provides a brief intro­
duction to aircraft fabrication.
While there have been tremendous advances in aircraft production in
recent years, much of the modern factory would be recognizable to a manu­
facturing engineer from the Wright Brothers' days. Aircraft production, then
and now, involves the application of the mechanical arts of machining,
forming, finishing, joining, assembly, and testing.
Machining involves the removal of a carefully controlled amount of
material from a part, typically by the application of a cutting tool via relative
motion between the part and the tool. The cutting tool is generally based
upon the inclined wedge and acts to peel away a thin shaving of the part.
(A drill bit can be seen as a set of inclined wedges positioned radially
around an axis.) The relative motion between tool and part can be rotational,
as with the drill, lathe, and mill, or it can be translational, as with the broach
and planer.
Forming refers to the numerous ways in which materials, especially
metals, are changed in shape other than by machining. Forming includes
casting, forging, extruding, stamping, punching, bending, and drawing. In
casting, the metal is brought up to its melting temperature then poured
into a mold. Forging involves forcing nonmolten metal into a mold
through pressure or impact. Extrusion is the process of forcing metal to
flow out a hole with the desired cross-sectional area, creating shaped bar
stock. Stamping and punching are used to cut out shapes and holes in
254 Ai rcraft D e s i g n : A C o nceptu a l Approach

sheet metal. Bending is self-explanatory, and drawing is the process of forcing

sheet metal into a form creating cup-like geometries.
Finishing encompasses a number of processes applied to formed and/ or
machined parts. Some finish processes include further material removal to
create a smoother surface, such as deburring, lapping, and finish grinding.
Other finish processes, such as painting, anodization, and plating, involve
application of a surface coating.
Composite fabrication is sufficiently unlike metal fabrication that it
deserves special mention. In thermoset composite production, a liquid or
pliable semisolid plastic material undergoes a chemical change into a new,
solid material, usually accompanied by the application of heat and/ or
pressure. For aircraft applications the plastic "matrix" material is reinforced
by a fiber, typically of graphite material. Thermoset composite manufacture
is unique in that the material itself is produced at the same time and place as
the part. A second class of composites, the thermoplastics, involves a plastic
matrix that is heated in a mold until it deforms readily, assuming the shape of
the mold. Composite fabrication is further described in Chapter 14.
Joining is simply the attachment of parts together, by processes including
brazing, soldering, welding, bonding, riveting, and bolting. All of these
processes historically have a high manual-labor content, and all are being
automated to various extents in modern factories. For example, modern
car factories have long lines of robotic spot-welders attaching body panels.
Automatic riveting machines, applicable for simple geometries such as
rivets in a row down a wing spar, can be found in the modern aircraft factory.
Assembly is the process of combining parts and subassemblies into the
final product. Assembly usually involves joining operations such as riveting
or bolting, but is distinguished from joining by the greater level of complete­
ness of the subassemblies. For example, when you attach a wing skin to the
wing ribs it is "joining," but when you attach the wing to the fuselage, it is
"assembly."
Testing is a key part of the manufacturing process. In traditional factories,
testing was generally done by random selection of finished product and was
frequently of a destructive nature. While helping to keep average quality up,
such random destructive testing did not insure that any given part was accep­
table because the only parts known by testing to be acceptable were destroyed
in the process!
Today's factories are tending toward nondestructive testing techniques
such as magnaflux, ultrasonic, and nuclear magnetic resonance, and are
also applying advanced statistical techniques to better select samples to
test and to determine the corrective action required.

l:ll1fl CAD /CAM, Automation, and Robotics

CAD/CAM, or computer-aided design/computer-aided manufacture, is
a generic term for the many different ways in which computers are being used
 C H APTE R 8 Spec i a l Consi derations i n Confi g u ration Layout 255

in design and manufacture. Typically, CAD/CAM refers specifically to the

us e of computers for component design and the use of the resulting
CA D database as the input for the programming of numerically controlled
machinery and robots (as described below) . The benefits of CAD/CAM
are well-established and include improved design quality, reduction in
design time and/ or increase in the number of design iterations possible,
earlier discovery of errors, integration of design, analysis, and manufacturing
engineering, and facilitation of training.
Automation refers to almost any use of computerized equipment during
manufacture. However, the generic term "automation" is most frequently
applied to tasks such as riveting, parts retrieval, and process control (such
as autoclave cycling), whereas the more specific terms "numerical control"
and "robotics" are used as described next.
Numerical control (NC) programming refers to the creation of digital
instructions that command a computer-controlled machine tool such as a
mill or lathe. This area is probably one of the highest leverage in terms of
reducing cost and improving quality. While machine tools themselves have
experienced little fundamental change in this century (this author knows
of a company recently making high-tech wind turbines on a 100-year-old
lathe!), the application of numerical control replacing the skilled but bored
machinist has had a tremendous effect on productivity and quality.
The most sophisticated subset of automation is robotics, in which a
computer-controlled machine performs tasks involving highly complex
motions that previously might have been performed by a human. Note that
it is the ability to physically manipulate objects in response to programming
that distinguishes the robot from other forms of automation or mechanism.
Robotics examples include part pickup and positioning, painting, composite­
ply laydown, material handling, simple assembly, and welding, and are usually
limited to "semiskilled" jobs, at least to date.
A key robotics technology for composites is in the labor-intensive tape
lay-up process. Programmable robot arms with tape dispenser end effectors
are widely used to place the prepreg. Computer-controlled filament dispen­
sers are being used to wind approximately round bodies such as tanks and
even entire fuselages. Also, autoclave cycle control is widely automated.

l:ll1JI Add itive Manufacturing

Rapid fabrication of parts without tools is being performed using
techniques broadly known as "additive manufacturing" (AM). The part is
made by "addition" of material rather than by molding or the "subtractive"
processes of machining. Generally, AM works by mathematically slicing
the CAD-based design into thin cross sections that are traced out, one
slic e at a time, onto some material. There is no tooling or other fabrication
process. The part comes complete, right out of the machine like paper
comes out of a printer, so it is sometimes known as "three-dimensional
256 Aircraft Design : A Conceptual Approach

printing." It would be the ultimate fabrication method if cost comes down

and material properties can match those of molding and machining.
In addition to the obvious savings in touch labor, AM offers a huge
improvement in the "buy-to-fly" ratio. To make a 100-lb part by machining,
you may need to start with a 2,000-lb block of aluminum or titanium, so
1,900 lb of material is recycled, or thrown away. With AM, there is almost
no waste. The buy-to-fly ratio approaches unity.
Another benefit of AM is that you can design and build things that
would be literally impossible any other way. Lightening holes and bracing
structures can be incorporated in inaccessible places. It is even possible to
build up mechanisms in ready-to-go fashion, such as gears that are produced
in place, already meshed together!
According to NASA's Karen Taminger, "AM is extremely useful in
prototype development because it can be used to build parts directly fro m
CAD without molds or tooling. This flows into production enabling
changes late in the design cycle and offering design flexibility in the materials
and shapes that can be built."
The first AM method to enter widespread usage was stereolithography
(SLA), which uses an ultraviolet laser beam to trace out the slices on a vat
of photosensitive chemicals that solidify as they are irradiated. After each
layer is completed, an "elevator" holding the part moves down slightly, and
the next layer is solidified on top of it. To date, only relatively fragile plastics
can be used by SLA devices, but the plastic prototypes can then be used to
create molds for casting materials such as epoxy or aluminum.
Another form of AM is selective laser sintering (SLS) in which a laser
melts a powdered material to produce the layers, which can be as little as
20 µ thick. Workable materials include plastic polymers, steel, titanium,
aluminum, Inconel, ceramics, and glass. The final material properties pro­
duced by SLS are not quite as good as a machined block of vendor-supplied
material, but SLS is getting very close.
Electron beam freeform fabrication (EBF 3 ) uses electron beams to build
up a part and works with weldable alloys including aluminum and titanium.
If costs are competitive and the part quality proves to be as good as machined
parts, this could be the best solution for creation of aircraft components
including airframe and mechanisms. l38l
Other AM approaches include fused deposition modeling (FDM) and
laminated object manufacturing, in which thin layers of material are cut to
shape and attached together.
Note that it is also possible with some types of AM to add on to an
existing part. For example, one could fabricate a major titanium fuselage
frame by superplastic forming and then use AM to add special attachment
fittings for production model variations.
When considering the use of AM, it is important to consider the size
of the parts. At present, many types of AM lend themselves only to parts
small enough to hold in your hands. We want to fabricate wing spars for
 CHAPTE R 8 Spec i a l Considerations i n Config u ration Layout 257

co mmercial transports! Eventually we want to make the entire structure of a

large airc raft all at once, perhaps with landing gear and other mechanisms
for med in place. This may take 50 years. In the meantime, part size and
material properties must be carefully considered when choosing an Additive
M anufacturing method.
AM is rapidly gaining acceptance in the aircraft design world. The Lock­
heed Pole cat, an unmanned proof-of-concept vehicle with a 90-ft span {27.4 m}
and weighing 9,000 lb {4,090 kg}, was largely made via AM. This allowed it to
be designed and built in only 1 8 months. Yes it crashed, but not due to AM.
AM methods are improving every year with faster fabrication, larger part
c ap ab ilities, different materials, and better material properties. In fact, they
say th at the material properties of metal AM parts are now virtually identical
to ma chined parts. Mixed-material AM devices are also in development.
A remaining issue for AM is validation and certification, especially for use
in safety-of-flight components. Each AM process must be proven to the full
sati sfaction of the regulating authorities. This is being done.

ifIJ Mai ntainabil ity Considerations

Maintainability means simply the ease with which the aircraft can be
fixed. Reliability and maintainability (R&M) are frequently bundled together
and measured in maintenance man-hours per flighthour (MMH/FH).
MMH/FHs range from less than one for a small private aircraft to well
over a hundred for a sophisticated supersonic bomber or interceptor.
Reliability is usually out of the hands of the conceptual designer.
Reliability depends largely upon the detail design of the avionics, engines,
and other subsystems. The configuration designer can only negatively
impact reliability by placing delicate components, such as avionics, too
near to vibration and heat sources such as the engines.
Anybody who has attempted to repair a car will already know this key
problem for maintainability. Getting at the internal components frequently
takes longer than fixing them! Accessibility depends upon the packaging
density, number and location of doors, and number of components that
must be removed to get at the broken component.
For large aircraft, just getting to the access doors can be a major under­
taking. Many airliners have the APU (auxiliary power unit; see Chapter 1 1)
installed in the tail, 20 ft {6 m} off the ground! This is acceptable for airliners
because they are serviced at major airports where work platforms can be
rolled into position. This can pose a problem for military aircraft that are
expected to operate away from main bases.
Figure 8.29 shows the actual servicing diagram for the B-70 supersonic
bomber, which is so large that a tall man can barely touch its bottom.
Notice the extra access panels near the engines and near the cockpit (for
avionics servicing) . For all its size, though, an engine on the B-70 could be
changed in 25 min-still a good time today!
 Qc;GHT SIDE
258 A i rcraft D e s i g n : A Conceptu a l A p p roa c h

11
0

LEFT SIDE

16 11
19
-

Fig. 8.29 Servicing diagram.

Packaging density has already been discussed. The number and location
of doors on modern fighters have greatly improved over prior-generation
designs. Frequently, the ratio between the total area of the access doors
and the total wetted area of the aircraft's fuselage is used as a measure of
merit, with modern fighters approaching a value of one-half.
A structural weight penalty must be paid for such access. This leads to the
temptation to use "structural doors" that carry skin loads via heavy hinges
and latches. These are always more difficult to open than non-load-bearing
doors because the airframe's deflection from its own weight will bind the
latches and hinges. In extreme cases, the aircraft must be supported on
j acks or a cradle to open these structural doors.
As a general rule, the best access should be provided to the components
that break the most often or require the most routine maintenance. Engine
access doors that allow most of the engine to be exposed should definitely
be provided. Also, large doors should be provided for the avionics compart­
ment, hydraulic pumps, actuators, electrical generators, environmental
control system, auxiliary power unit, and gun bay.
The worst feature an aircraft can have for maintainability is a requirement
for major structural disassembly to access or remove a component. For
example, the V/STOL AV-SB Harrier requires that the entire wing be
 CHAPTE R 8 Spec i a l Consideratio ns i n Confi g u ration Layout 259

removed before removing the engine. Several aircraft require removal of a

part of the longeron to remove the wing.
Similarly, the designer should avoid placing internal components such
that one must be removed to get to another. In the F-4 Phantom, an ejection
seat must be removed to get to the radio (a high-break-rate item) . It is not
uncommon for the ejection seat to be damaged during this process.
"One-deep" design will avoid such problems.

B-70 with wing tips drooped for supersonic flight (photo from U .S. Air Force) .

What We've Lea rned

We've discovered some of the things the designer is thinking a n d doing while
making that first design layout. These include aerodynamics, structures, pro­
ducibility, maintainability, crashworthiness, noise, and for military aircraft,
signature and vulnerability.
260 A i rcraft D e s i g n : A Conceptu a l A p p ro a c h
 Crew Station,
Passengers, and
Payload

Passengers a n d payload a re t h e rea l reason we des i g n a n d b u i l d a i rpla nes, s o do

rig ht!
•

them
• Strict g overn ment reg u l ations m ust be understood and fol lowed .
• The c rew sta t i o n d e s i g n s the front end of s m a l ler a i rp l a nes.
• The cabin d e s i g n s the a i rl i n e r fus e l a g e .
• Safety i s p a ra mo u nt.

Introduction

F
or the initial configuration layout ("Dash-One"), it is not necessary
to go into the details of crew-station design, such as the actual
arrangement and location of controls and instruments, or the details
of passenger and payload provisions. However, the basic geometry of the
crew station and payload/passenger compartment must be considered so
that the subsequent detailed cockpit design and payload integration efforts
will not require revision of the overall aircraft.
If it is a passenger plane, that very first configuration layout must have
allowances for head room, leg room, exits, galleys, and toilets. If it is a military
fighter, the bombs have to fit, and the gun has to be indicated in a workable
location. For manned aircraft, the pilot needs to see out, needs enough room
inside, and needs enough space for an instrument panel even if the actual
arrangement of the instruments won't be done until much later.
This chapter presents dimensions and "rule-of-thumb" design guidance
for conceptual layout of aircraft crew stations, passenger compartments,
payload compartments, and weapons installations. Information for more

261
262 A i rc raft Desig n : A C o n c e p t u a l A p p r o a c h

detailed design efforts is contained in the various civilian and military

specifications and in subsystem vendors' design data packages.

Crew Station
The crew station will affect the conceptual design primarily in the vis ion
requirements. Requirements for unobstructed outside vision for the pilo t
can determine both the location of the cockpit and the fuselage shape in
the vicinity of the cockpit.
For example, the pilot must be able to see the runway while on final
approach, so the nose of the aircraft must slope away from the pilot's eye
at some specified angle. While this can produce greater drag than a more
streamlined nose, the need for safety overrides drag considerations. Similarly,
the need for overside vision might prevent locating the cockpit directly above
the wing.
When laying out an aircraft's cockpit, it is first necessary to decide what
range of pilot sizes to accommodate. For most military aircraft, the design
requirements include accommodation of the 5th to the 95th percentile of
male pilots, that is, a pilot height range of 65.2-73.1 in. {1 .66- 1 .86 m}.
Because of the expense of designing aircraft that will accommodate smaller
or larger pilots, the services exclude such people from pilot training.
Women are now entering the military flying profession in substan­
tial numbers. Future military aircraft will require the accommodation of
approximately the 20th percentile female, about 98 lb and 59 in. tall {44 kg
and 1.5 m}. This can affect the detailed layout of cockpit controls and
displays but should have little impact upon conceptual cockpit layout as
described next.
General aviation cockpits are designed to whatever range of pilot sizes the
marketing department feels is needed for customer appeal, but typically are
comfortable only for those under about 72 in. {1 .83 m}. Commercial-airliner
cockpits are designed to accommodate pilot sizes similar to those of
military aircraft.
Figure 9.1 shows a typical pilot figure useful for conce ptual design
layout. This 95th percentile pilot, based upon dimensions from l 39l , includes
allowances for boots and a helmet. A cockpit designed for this size of pilot
will usually provide sufficient cockpit space for adjustable seats and controls
to accommodate down to the 5th percentile of pilots.
Designers sometimes copy such a figure onto cardboard in a standard
design scale such as 20-to- l, cut out the pieces, and connect them with
pins to produce a movable manikin. This is placed on the drawing, positioned
as desired, and traced onto the layout. A computer-aided aircraft design
system can incorporate a built-in pilot manikin (see [2 1 l ) .
Dimensions fo r a typical cockpit sized t o fit the 95th-percentile pilot are
shown in Fig. 9.2. The two key reference points for cockpit layout are shown.
The seat reference point, where the seat pan meets the back, is the reference
 C HAPTE R 9 Crew Station, Passengers. a n d Payload 263

S h o u l d e r w i d t h - 26 i n .
{66 cm}

A l low 30 i n . (76 cm}

for cleara n ce

Fig. 9. 1 Average 95th percentile pilot.

for the floor height and the leg-room requirement. The pilot's eye point
is used for defining the overnose angle, transparency grazing angle, and
pilot's head clearance (10-in. {25-cm} radius).
This cockpit layout uses a typical 13-deg seatback angle, but seatback
angles of 30 deg are in use (F- 16), and angles of up to 70 deg have been con­
sidered for advanced fighter studies. This entails a substantial penalty in
outside vision for the pilot but can improve his ability to withstand high-g
turns and also can reduce drag because of a reduction in the cockpit height.
When designing a reclined-seat cockpit, rotate both the seat and the
pilot's eye point about the seat reference point and then use the new position
of the pilot's eye to check overnose vision.
Overnose vision is critical for safety especially during landing and is also
important for air-to-air combat. Military specifications typically require
17-deg overnose vision for transports and bombers and 1 1 - 1 5 deg for
fighter and attack aircraft. Military trainer aircraft in which the instructor
pilot sits behind the student require 5-deg vision from the back seat over
the top of the front seat.
Various military specifications and design handbooks provide detailed
requirements for the layout of the cockpit of fighters, transports, bombers,
and other military aircraft.
General aviation aircraft land in a fairly level attitude and so have over­
nose vision angles of only about 5-10 deg. Many of the older designs have
such a small overnose vision angle that the pilot loses sight of the runway
from the time of flare until the aircraft is on the ground and the nose
is lowered.
264 Ai rcraft D e s i g n : A C o nceptu a l A p p roach

C ross section

Head c l e a ra n ce

1------ //
1 0 i n . {25 c m }

/
I
1 5 i n ._J {38 cm)
40 deg

/ ::;_ 1 {8 c m }
\
/,

3 in.

/
-
/

//
I
Fo r l o n g e ro n
I
cl

}
1 7 in

/� , }
, f43 cm
7 3 in
f33 crn
I , .......____. ,
I
I
,'

P i lot's eye

32 i n . {8 1 cm)

8 in. {20 cm}

I ""- -______....

� SO i n .
{ 1 .3 m)

Seat refe rence point

Fig. 9.2 Typical fighter cockpit.

Civilian transports frequently have a much greater overnose v1s10n

angle, such as the Lockheed L-101 1 with an overnose vision angle of
21 deg. Civilian overnose vision angles must be calculated for each aircraft
based upon the ability of the pilot to see and react to the approach lights
at decision height (100 ft {30.5 m}) during minimum weather conditions
(1200-ft {366-m} runway visual range). The higher the approach speed, the
greater the overnose vision angle must be.
 C H A PT E R 9 Crew Station. Passengers. a nd Payl oad 265

Reference [40) details a graphical technique for determining the required

o no se angle, but it can only be applied after the initial aircraft layout
ver
is complete and the exact location of the pilot's eye and the main
landing gear is known. For initial layout, Eq. (9. 1) is a close approximation,
based upon the aircraft angle of attack during approach and the approach
speed.
ll'overnose '.::::'. ll'app roach + 0 .07 Vapp roach ( V in kt)
= ll'app roach + 0 . 04 Vapp roach ( V in k m/hr ) (9 . 1)
Figure 9.2 shows an over-the-side vision requirement of 40 deg, measured
from the pilot's eye location on centerline. This is typical for fighters
and attack aircraft. For bombers and transports, it is desirable that the
pilot be able to look down at a 35-deg angle without head movement and
at a 70-deg angle when the pilot's head is pressed against the cockpit glass.
This would also be reasonable for general aviation aircraft, but many
general aviation aircraft have a low wing blocking the downward view.
The vision angle looking upward is also important. Transport and
bomber aircraft should have unobstructed vision forwards and upwards
to at least 20 deg above the horizon. Fighters should have completely
unobstructed vision above and all of the way to the tail of the aircraft. Any
canopy structure should be no more than 2 in. {5 cm} wide to avoid
blocking vision.
The transparency grazing angle shown in Fig. 9.2 is the smallest angle
between the pilot's line of vision and the cockpit windscreen. If this angle
becomes too small, the transparency of the glass or plexiglass will become
substantially reduced, and under adverse lighting conditions the pilot
might only see a reflection of the top of the instrument panel instead of
whatever is in front of the aircraft! For this reason, a minimum grazing
angle of 30 deg is recommended.
The cockpit of a transport aircraft must contain anywhere from two to
four crew members as well as provisions for radios, instruments, and stowage
of map cases and overnight bags. Reference [40) suggests an overall length
of about 150 in. {3.8 m} for a four-crew-member cockpit, 130 in. {3.3 m} for
three crew members, and 100 in. {2.5 m} for a two-crew-member cockpit.
The cockpit dimensions shown in Fig. 9.2 will provide enough room
for most military ejection seats. An ejection seat is required for safe escape
when flying at a speed that gives a dynamic pressure above about 230 psf
{ 1 1 kN/m2 } (equal to 260 kt {48 1 km/h} at sea level).
At speeds approaching Mach 1 at sea level (dynamic pressure above 1200
{58 kN/m2 }), even an ejection seat is unsafe and an encapsulated seat or
separable crew capsule must be used. These are heavy and complex. A separ­
able crew capsule is seen on the FB- 1 1 1 and the prototype B- lA. The latter,
inc lu ding seats for four crew members, instruments, and some avionics,
weighed about 9000 lb {4,082 kg}.
 266 A i r c ra ft Desig n : A C o n c eptua l A p p ro a c h

..,.. 9. 3 Passenger Compa rtment

The actual cabin arrangement for a commercial aircraft is determine d
more by marketing than by regulations. Figure 9.3 defines the dimensio ns
of interest. "Pitch" of the seats is defined as the distance from the back of
one seat to the back of the next. Pitch includes fore and aft seat length as
well as leg room. "Head room" is the height from the floor to the roof over
the seats. For many smaller aircraft the sidewall of the fuselage cuts off a
portion of the outer seat's head room, as shown. In such a case it is importan t
to ensure that the outer passenger has a 10-in. {25-cm} clearance radius about
the eye position.
Table 9.1 provides typical dimensions and data for passenger compart­
ments with first-class, economy, or high-density seating. This informatio n
(based upon l4o , 4 1 l , and other references) can be used to lay out a cabin
floor plan.
Sad to say, today the typical design values presented in this table are
rarely used in practice. Recent measurements of actual seats indicate that
the airlines are using roughly 3 1 in. pitch and 17 in. width {79 x 43 cm} for
economy seats on commercial jets. Such cramped quarters, in years past,
were only inflicted upon passengers flying short commuter flights.
However, it is probably good to design the aircraft to the larger dimensions
in the table-so the airlines can cram in more rows after they have bought
the plane.
There should be no more than three seats accessed from one aisle, so an
aircraft with more than six seats abreast will require two aisles. Also, doors

r
Aisle
r - - - - - - - - - - ,

height
I I
I I
Head room I I
I I
I
I
I
I
I
, ....___ s eat -----#-T-
.
1
pitch
Seat Aisle
I
- -
width width
I
I
I

Fig. 9.3 Commercial passenger al lowances.

 C H A PT E R 9 Crew Station, Passengers, a n d Payl oad 267

Table 9. 1 Typical Passenger Compartment Data

High-Density/
First Class Economy Small Aircraft
seal pitch, in. {cm} 38-40 {97-1 02} 34-36 {86-9 1 } 30-32 {76-8 1 }
seat width, in. {cm} 20-28 {5 1 -7 1 } 1 7-22 {43-56} 1 6-1 8 {4 1 -46}
Head room, in. {cm} > 65 {1 65} >65 { 1 65}
Aisle width, in. {cm} 20-28 {5 1 -7 1 } 1 8-20 {46-5 1 } :::: 1 2 {30}
Aisle heig ht, in. {cm} > 76 { 1 93} > 76 { 1 93} > 60 { 1 52}
Pa ssengers per cabin staff 1 6-20 3 1 -36 ::; 50
(international-domestic)
Passengers per lavatory 1 0-20 40-60 40-60
(40 x 40 in.) {l x l m}
Galley volume per passenger, 5-8 {0. 1 4-0.23} 1 -2 {0.03-0.06} 0-1 {0-0.03}
tt 3 {m 3 }

and entry aisles are required for approximately every 10-20 rows of seats.
These usually include closet space and occupy 40-60 in. { l - 1 .5 m} of cabin
length each.
Passengers can be assumed to weigh an average of 1 80 lb {82 kg} (dressed
and with carry-on bags) and to bring about 40-60 lb { 1 8-27 kg} of checked
luggage. A current trend toward more carry-on luggage and less checked
luggage has been overflowing the current aircrafts' capacity for overhead
stowage of bags.
The cabin cross section and cargo bay dimensions (see the following) are
used to determine the internal diameter of the fuselage. The fuselage external
diameter is then determined by estimating the required structural thickness.
This ranges from 1 in. {2.5 cm} for a small business or utility transport to
about 4 in. { 1 0 cm} for a jumbo jet.

Cargo Provisions
Cargo must be carried in a secure fashion to prevent shifting while in
flight. Larger civilian transports use standard cargo containers that are pre­
loaded with cargo and luggage and then placed into the belly of the aircraft.
During conceptual design, it is best to attempt to use an existing container
rather than requiring purchase of a large inventory of new containers.
Two of the more widely used cargo containers are shown in Fig. 9.4. Of
the smaller transports, the Boeing 727 is the most widely used, and the 727
container shown is available at virtually every commercial airport.
The "Lower Deck" LD-3 container is used by all of the widebody trans­
ports. The B-747 carries 30 LD-3s plus 1000 ft3 {28.3 m3 } of bulk cargo
volume (non-containerized) . The L- 1 0 1 1 carries 16 LD-3s plus 700 ft3
{19.8 m3 } of bulk cargo volume, and the DC- 10 and Airbus A-300 each
268 A i rc r a ft D e s i g n : A C o n c e pt u a l A p p r oa c h

72 7-200 C conta i n e r LD-3 conta i n e r

78 c u b i c feet 1 58 c u b i c feet
{2.2 m 3 } {4.5 m 3 }

�
44 . 4

41 . 1

Fig. 9.4 Cargo containers.

carry 14 LD-3s plus 805 {22.8} and 565 ft3 {16 m 3 }, respectively, of bulk
cargo volume.
To accommodate these containers, the belly cargo compartments
require doors measuring approximately 70 in. {1.8 m} on a side. As was
discussed in the section on wing vertical placement, low-wing transports
usually have two belly cargo compartments, one forward of the wing box
and one aft.
The cargo volume per passenger of a civilian transport ranges from
about 8.6 - 15.6 ft3 {0.24-0.44 m 3 } per passenger. l4 1 l The smaller number
represents a small short-haul j et (DC-9). The larger number represents a
transcontinental j et (B-747). The DC- 10, L-101 1, Airbus, and B-767 all
have about 1 1 ft 3 {0.3 1 m 3 } per passenger. Note that these volumes provide
room for paid cargo as well as passenger luggage.
Smaller transports do not use cargo containers, but instead rely upon
hand-loading of the cargo compartment. For such aircraft a cargo provision
of 6-8 ft3 {0. 17-0.23 m 3 } per passenger is reasonable.
Military transports use flat pallets to preload cargo. Cargo is placed upon
these pallets, tied down, and covered with a tarp. The most common pallet
measures 88 x 108 in. {2.2 x 2.7 m}.
Military transports must have their cargo compartment floor approxi­
mately 4-5 ft { 1 .4 m} off the ground to allow direct loading and unloading
of cargo from a truck bed at air bases without cargo-handling facilities.
However, the military does use some commercial aircraft for cargo transport
and has pallet loaders capable of raising to a floor height of 13 ft {4 m} at the
major military airlift command bases.
The cross section of the cargo compartment is extremely important for
a military transport aircraft. The C-5 and C- 1 7, largest of the U.S. military
transports, are sized to carry so-called outsized cargo, which includes
 C HAPT E R 9 C rew Station. Passengers, and Payload 269

M-60 tanks, helicopters, and large trucks. The C-5 cargo bay is 19 ft wide,
1 3.5 ft high, and 121 ft long {5.8 x 4.1 x 36.9 m}. It can carry a payload of
263,000 lb { 1 1 9,295 kg}.
The C-130 is used for troop and supply delivery to the front lines and
cannot carry outsized cargo. Its cargo bay measures 10.3 ft wide, 9.2 ft
high, and 41.5 ft long {3. 1 x 2.8 x 12.7 m}.

Wea pons Carriage

Carriage of weapons is the purpose of most military aircraft. Traditional
weapons include guns, bombs, and missiles. Lasers and other exotic technol­
ogies might someday become feasible as airborne weapons, but will not be
discussed here.
The weapons are a substantial portion of the aircraft's total weight.
This requires that the weapons be located near the aircraft's center of
gravity. Otherwise the aircraft would pitch up or down when the weapons
are released.
Missiles differ from bombs primarily in that missiles are powered. Today,
virtually all missiles are also guided in some fashion. Many bombs are
"dumb," or unguided, and are placed upon a target by some bombsight
mechanism or computer that releases them at the proper position and
velocity so that they freefall to the desired target. "Smart bombs" have
some guidance mechanism, typically homing on a laser spot or guiding to
a GPS (global positioning system) coordinate. ,
Missiles are launched from the aircraft in one of two ways. Smaller
missiles such as the AIM-9 and AIM- 120 are usually rail-launched. A rail­
launcher is mounted to the aircraft, usually at the wing tip or on a pylon
under the wing. Attached to the missile are several mounting lugs, which
slide onto the rail as shown on Fig. 9.5. For launch, the missile motor
powers the missile down the rail and free of the aircraft.
Ejection-launch is used mainly for larger missiles. The missile is attached
to the aircraft through hooks that are capable of quick-release, powered by an
explosive charge. This explosive charge also powers two pistons that shove
the missile away from the aircraft at an extremely high acceleration. The
missile motor is lit after it clears the aircraft by some specified distance.
The newer weapons ejectors are pneumatic. Compressed air is stored in a
bottle and quickly released into pistons. This approach reduces maintenance
and avoids the logistics associated with those explosive charges. Pneumatic
ejectors are used on the F-35 and are claimed to save weight and cost.
Bombs can also be ejected using explosive or pneumatic ejectors, or can
simply be released and allowed to fall free of the aircraft.
There are four options for weapons carriage. Each has pros and cons,
depending upon the application. External carriage is the lightest and simplest
and offers the most flexibility for carrying alternate weapon stores.
270 A i rc raft D e s i g n . A Conceptua I Approach
.

Rail Ejector
Pylon Explos ive
_.....- or c h a rge
wingtip

l Release
mecha n is m

Fig.
. 95 · Missile carriage/launch .

C\

S e m i -s u b m e rged

Co nformal

Fig. 9.6 Weapons coma ge options.

·