Note: Descriptions are shown in the official language in which they were submitted.
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LAMINAR FLOW NACELLE
FIELD OF THE INVENTION
This invention relates to a nacelle for housing
an aircraft engine, and, more particularly, to a nacelle
e~fective for reducing aerodynamic drag therefrom.
BACKGROUND OF THE I~v~NllON
In a subsonic aircraft having an externally
mounted engine, for example, a gas turbine engine mounted
below a wing by a pylon, aerodynamic drag due to freestream
airflow over the nacelle of the engine can typically represent
approximately 4~ of the total engine thrust output. Any
reduction in this aerodynamic drag can result in a sig-
nificant saving in the amount of fuel consumed.
Accordingly, a desired function of an engine
nacelle is to provide a lightweight housing for the
aircraft engine which produces relatively low aerodynamic
drag. An example of a prior art low drag nacelle is
disclosed in U~S. Patent No. 3,533,237, issued
October 13, 1970 to G.R. Rabone et al, assigned to the
present assignee.
The aerodynamic drag due to a nacelle is
determined by the pressure distribution and a dimensionless
friction coefficient Cf, known to those skilled in the art,
over the outer surface of the nacelle over which the
freestream air flows during aircraft flight. It is known
to those skilled in the art that reduced aerodynamic
drag exists where the surface pressure distribution
promotes a laminar boundary layer over the nacelle
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outer surface without any boundary layer separation
thereof. The friction coefficient C~, and thus
aerodynamic drag, have reduced values when a laminar
boundary layer exists.
Also known to those skilled in the art is that
where the boundary layer along the nacelle outer surface
transitions from laminar to turbulent, the friction
coefficient Cf~ and thus aerodynamic drag, have increased
values. Accordingly, it is desirable to provide a nacelle
which promotes a surface pressure distribution effective
for increasing the extent of laminar boundary layer flow,
reducing the extent of turbulent flow and avoiding
boundary layer separation.
A nacelle is typically an annular member which
houses an aircraft engine, such as a gas turbine engine.
Unlike a wing which extends longitudinally and has upper
and lower surfaces designed for maximum lift and reduced
drag, a nacelle extends circumferentially and has an
outer surface designed primarily to house an engine and
reduce drag.
However, in both a nacelle and a wing the
pressure distribution over the surfaces thereof is a
significant factor in determining the extent of l~mi n~r
and turbulent airflow thereover. In a wing, for example,
the pressure distribution is dependent on the contours
of the leading and trailing edges and the upper and
]ower surfaces. A change in any contour affects the
entire pressure distribution over the wing.
In a nacelle, in contrast, the pressure
distribution is primarily affected by the contours of
the leading and trailing edge regions and the outer
surface. The inner surface of the nacelle has little
interaction with the freestream airflow, and therefore
has less affect on the pressure distribution.
Furthermore, inasmuch as a nacelle is
typically mounted to an aircraft near a fuselage, pylon
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or wing, the pressure distribution over the nacelle
can also be affected by ~he presence of these adjacent
structuxes. A chang~ in any contour of the elements of
the nacelle and the presence of adjacent structures
affects the entire pressure distribution over the outer
surface of the nacelle.
Past attempts at maintaining and extending
laminar flow on wings and nacelles have involved the
use of ac~ive control devices. An active control device
requires an auxiliary source of energy to cooperate
with the surface for energizing or removing the boundary
layer for maintaining laminar flow and preventing
boundary layer separation.
For example, boundary layer suction or blowing
slots or holes disposed in the surface to be controlled
are known in the art. The slot is connected to a pump
by internal ducting and is effective for reducing or
pre~enting turbulent flow, and thereby maintaining
laminar boundary layer flow. Ho~ever, the additional
2Q weight and energy required to power the active control
device typically offsets advantages due to reduced
aerodynamic drag.
Accordingly, it is one object of this invention
to pro~ide an improved nacelle for housing an aircraft
engine which is effective for reducing aerodynamic drag
during aircraft operation.
Another object of this invention is to provide
an improved nacelle which does not require an active
device for re~ucing aerodynamic drag.
Another object of this invention is to provide
an improved naceIle having increased areas of l~mi n~r
flow and decreased areas of turbulent flow thereover.
Another object of this invention is to provide
an improved nacelle having a profile effective for
controlling surface pressure distribution thereover for
reducing aerodynamic drag.
13DV~8063
SUM~ARY OF TH~ INVENTION
An improved nacelle is provided for use on an
aircraft that reduces aerodynamic drag during aircraft
operation. In one embodiment, the nacelle houses a gas
turbine engine and comprises a leading edge and a
trailing edge, having a reference chord extending
therebetween, and an outer surface which is continuous
from the leading edge to the trailing edge. The outer
surface includes a forward portion, an intermediate
portion and an aft portion and has a profile defined by
a relative thickness measured perpendicularly from the
reference chord to the outer surface. The profile has a
maximum thickness at the intersection of the forward
and intermediate portions, which intersection is
located greater than about 36% of the chord from the
leading edge. The profile of the outer surface is
effective for producing l~m; n~r flow along the forward
portion and pressures which decrease continuously at a
negative gradient from the leading edge to the
intersection, and turbulent flow along the intermediate
and aft portions and pressures which increase continuously
at a positive gradient from the intersection to the
trailing edge.
B~IEF DES~l~lION OF T~IE DRAWINGS
The invention, together with further objects and
advantages thereof, is more particularly described in
the following detailed description taken in conjunction
with the accompanying drawings in which:
Figure 1 is a view partially in cross section
3Q of a turbofan engine mounted to a wing of an aircraft
by a pylon and incorporating a nacelle according to one
form of the present invention.
Figure 2 is an enlarged sectional view of the
nacelle of Figure 1.
Figure 3 is a graph according to one form of
the present invention representing pressure distribution
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along the outer surface of the nacelle of Figure 2
relative to a reference chord extending from a leading
edge to a trailing edge thereof.
Figure 4 is a graph illustrating a profile of
the nacelle of Figure 2 normalized with respect to the
reference chord which is effective for obtaining a
pressure distribution according to Figure 3.
Figure 5 is a graph illustrating the radius of
curvature of the nacelle of Figure 2 normalized with
respect to the reference chord.
Figure 6 is an enlarged view of a leading edge
region of the nacelle of Figure 2.
Figure 7 is an enlarged view of a trailing
edge region of the normalized nacelle profile shown in
Figure 4.
Figure 8 is a sectional view of a single
exhaust gas turbine engine incorporating another embodiment
of the present invention.
DETAI1ED DES~CRI:PTION
2Q Illustrated in Figure 1 is a high bypass gas
turbofan engine 10 mounted to a wing 12 of an aircraft
(not shown) by an aerodynamically shaped pylon 14. The
turbofan engine 10 includes a fan assembly 16 driven by
a core engine 18.
Housing the engine 10 is an annular nacelle 20
including a core cowl 22 surrounding the core engine 18
and a fan cowl 24, according to one form of ~he present
invention, surrounding the fan assembly 16. The fan
cowl 24 also surrounds and is spaced from a forward
portion of the core cowl 22 for defining an annular fan
discharge nozzle 26. The fan cowl 24 includes an inlet
throat 28 for receiving an engine airflow portion 30 of
a freestream airflow 32.
During aircraft operation, such as for example
at cruise, the engine airflow 30 is accelerated by the
fan assembly 16 and is discharged from the fan nozzle 26
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over the core cowl 22 for generating thrust. The
freestream airflow 32 flows downstream over the fan
cowl 24 of the nacelle 20 and interacts with or scrubs
the fan cowl 24 and produces aerodynamic drag, a
significant portion of which is frictional drag acting
in a direction opposite to that of the cruising aircraft.
A primary purpose of the present invention is to
provide a nacelle, such as the fan cowl 24, that is
effective for reducing aerodynamic drag due to freestream
airflow 32 thereover during subsonic aircraft cruise.
Reduced aerodynamic drag at cruise is achieved by providing
the fan cowl 24 with a predetermined aerodynamic surface
profile effective for producing a pressure distribution
to promote a natural laminar boundary layer over an
increased portion of the outer surface of the fan cowl 24
of the nacelle 20 without causing boundary layer separation.
However, inasmuch as engine airflow 30 discharged from the
fan nozzle 26 primarily flows over the core cowl 22, the
profile of the core cowl 22 of the nace~le 20 is
preferably determined according to conventional standards.
Illus~rated in more detail in Figure 2 is the fan
cowl 24 of Figure l. The fan cowl 24 includes an
annular leading edge 34 and an annular trailing edge 36
having a reference chord 38 of length C extending there-
between. The fan cowl 24 also includes an outer surface 40which is continuous from the leading edge 34 to the
trailing edge 36. The outer surface 40 includes a forward
portion 42, an intermediate portion 44 and an aft portion
46. The forward portion 42 extends from the leading edge
34 to a first intersection 48 joining the forward portion
42 and the intermediate portion 44. The aft portion 46
extends from a second intersection 50 to the trailing
edge 36 and joins the intermediate portion 44.
A significant feature of the fan cowl 24 is the
profile of the outer surface 40. The profile is the
outline of the outer surface 40 and can be defined by a
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varying relative thickness T representing the perpendicular
distance of the outer surface 40 from the reference chord
38. The thickness T increases along the chord 38 from
the leading edge 34 to a position of maximum thickness
T at the first intersection 48. The thickness T then
decreases along the chord 38 from the first intersection
48 to the trailing edge 3~.
Another significant feature of the fan cowl 24
is that the m~x; ml~m thickness T is located farther
max
aft along the chord 38 than the maximum thickness TmaX2
of a typical prior art nacelle 52, indicated in broken
line in Figure 2 for comparison purposes. This feature
and features to be described hereinafter promote laminar
flow along the forward portion 42 while limiting turbulent
flow to the intermediate portion 44 and the aft portion 46
without boundary layer separation.
To more fully appreciate the significance of
the present invention, a descxiption of the pressure
distribution over the fan cowl 24 is appropriate. It
is known to those skilled in the art that a pressure
gradient due to freestream airflow imposed on a nacelle
surface, such as outer surface 40 of the fan cowl 24,
affects the location of boundary layer transition from
laminar flow to turbulent flow. Generally a negative
pressure gradient, i.e., pressure decreasing in the
flow direction delays transition from laminar to
turbulent flow.
It is also known that a positive pressure
gradient must follow a negative pressure gradient to return
the pressure back to an ambient, freestream value. It
is in this positive pressure gradient region that the flow
over the nacelle becomes turbulent, resulting in increased
drag.
However, to increase the extent of laminar flow
in a nacelle of finite length, the length in which the
pressure can be returned to ambient must necessarily
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decrease. This has been a limiting factor in prior art
nacelles because the decreased length remaining to return
the pressure to ambient promotes boundary layer
separation. Boundary layer separation initiating in
the turbulent flow region significant increases drag and
is therefore undesirable. Accordingly, nacelles of the
prior art typically include relatively large regions of
turbulent flow for suitably returning pressures ~o ambient
for preventing boundary layer separation.
~owever, according to the present invention, a
significant increase in the extent of laminar flow
without boundary layer separation can be realized by
providing a predeterminedly shaped fan cowl 24, such as
the one shown in Figure 2; that is effective for promoting
a predetermined pressure distribution over the outer
surface 40 of the fan cowl 24.
Illustrated in Figure 3 is a graph according to
the present invention indicating pressure distributions
due to freestream airflow over an outer surface of a
nacelle such as the fan cowl 24 shown in Figure 2. The
abscissa represents a normalized, nondimensional distance
X/C, where C is the length of the chord 38 and X is a
distance measured along the chord 38 from the leading
edge 34 (as shown in Figure 2). For example~ the leading
edge 34 and the trailing edge 36 are located at X/C =
0.0 and X/C = l.C, respectively, which can alternatively
be stated as 0~ C or 100% C, respectively.
The ordinate in Figure 3 represents pressure
over the surface of the fan cowl 24 at each point of
abscissa X/C. The pressure can be, for example, a pressure
coefficient Cp defined as 2 (PS ~ P)/dv ; where P, v and d re-
present the pressure, velocity and density, respectively,
of the freestream airflow 32, and PS represents static
pressure measured at the nacelle outer surface. The
pressure could also be represented, for example~ by
PS/PT, where PT represents freestream airflow total pressure.
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An example of a prior art Cp distribution 54
for a nacelle is represented by the broken line in Figure
3 and substantially corresponds to the prior art nacelle
52 shown in broken line in Figure ~. The prior art
Cp distribution 54 includes a negative pressure gradient
portion 56 which extends from 0% C to about 10% CO The
negative gradient portion 56 produces a small length
of laminer flow having a relatively low value of the
friction coefficient Cf~ indicating relatively low
drag. At about 10% C, the prior art Cp distribution
54 includes a minimum, negative, Cp 58 about which the
Cp distribution changes abruptly from the negative gradient
56 to a positive pressure gradient portion 60. The
positive pressure gradient portion 60 extends from about
10% C to 100% C. The abrupt Cp change at 10% C and the
positive gradient portion 60 produces a relatively large
length of turbulent flow having a relatively high friction
coefficient Cf resulting in increased aerodynamic drag.
It is to be noted that boundary layer separation is reduced
or avoided in the prior art nacelle 52 by increasing the
extent of turbulent flow at the expense of reducing the
extent of l~m' n~r flow, resulting in increased drag.
Also shown in the graph of Figure 3, is a
predetermined l~m; n~r flow Cp distribution 62 according
to one form of the present invention. The laminar flow
Cp distribution 62 provides for an increased extent of
laminar flow over that of the prior art, without boundary
layer separation. The Cp distribution 62 is characterized
by a continuously decreasing pressure coefficient Cp from
0% C to a position of minimum, negative, Cp 64 located
greater than about the 10% C of the prior art. In the
particular embodiment shown in Figure 3, the po~ition
of minimum Cp 64 is between about 50% C and about 60% C,
and is, preferably, at about 56~ C. Furthermore, the
position of the minimum Cp 64 corresponds with the position
of maximum thickness T at the first intersection 48 of
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Figure 2. This is in contrast with the prior art nacelle
52 of Figure 2 f wherein the position of minimum Cp 58 of
Figure 3 is spaced forwardly of the position of the prior
art maximum thickness T a 2 in Figure 2.
In the embodiment of the present invention shown
in Figure 3, the laminar flow Cp distribution 62 includes
a first negative gradient portion 66 which decreases from
a positive value of Cp at 0% C to a negative value of Cp
at approximately 10% C. The Cp distribution 62 includes
a second negative gradient portion 58 which is continuous
with the first portion 66 and extends from approximately
10% C to the minimum Cp 64 at about 56% C. The second
portion 6~ has a negative gradient of lesser magnitude
than the gradient of the first portion 66. Furthermore,
both the first por~ion 66 and the second portion 68 are
substantially convex with respect to the abscissa X/C.
The term convex is intended to indicate that a
curve, such as the second portion 68, has a center of
radius of curvature located between the curve and the
abscissa X/C. Correspondingly, the term concave is
intended to indicate that a curve has a center of
radius of curvature located on that side of the curve
opposite to the abscissa X/C.
A significant feature of the present invention
which permits increased extent of reduced drag laminar
flow along the nacelle surface 40 is a predetermined
positive gradient portion 70. The positive gradient
portion 70 extends from abGut 56% C to 100% C and is
effective for preventing boundary layer separation.
More specifically, at about 56% C, the laminar flow
Cp distribution 62 includes a transition portion about
the minimum Cp 64 wherein the slope or gradient of the
curve changes from negative to positive in value. This
change occurs more gradually than the abrupt change found
in the prior art Cp distribution 54 and is a factor in
preventing boundary layer separation. From approximately
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56% C to 100% C, the positive gradient portion 70 extends
from the minimum Cp 64 to a positive value of Cp,
respectively. In a preferred embodiment, the positive
gradient portion 70 along the aft portion 46 adjacent
the trailing edge 36 (as in Figure 2~ decreases at a
decreasing rate and has a substantially concave profile with
respect to the abscissa X/C, and, for example, can be
parabolic.
When a nacelle, such as the fan cowl 24 shown in
Figure 2, is contoured to provide a pressure distribution
such as represented by the laminar flow Cp distribution 62
shown in Figure 3, laminar flow can be made to e~ist from
0% C to approximately 56% C. The laminar flow and the low
roefficient of friction Cf associated therewith results in
a nacelle surface having si~nificantly reduced aerodynamic
drag during aircraft cruise operation without boundary
layer separation.
Illustrated in Figure 4 is a normalized profile
72 of a nacelle profile according to one embodiment of the
present invention. The abscissa is X/C, as above-described,
and the ordinate represents the thic~ness T divided by the
chord length C. The nacelle profile 72 is effective for
promoting the l~m; n~r flow Cp distribution 62 of Figure 3.
Inasmuch as the nacelle profile 72 is normalized, it is
applicable for defining any nacelle simply by appropriate
scaling. In this regard, the nacelle profile 72 of
Figure 4 is a nondimensional representation of the fan
cowl 24 shown in Figure 2.
Although the desired l~m;n~r flow Cp distribution
62 of Figure 3 according to the invention has been
determined, it is not possible to completely predetermine
a specific profile of the fan cowl 24 suitable for all
aircraft engine applications. This is so inasmuch as the
pressure distribution about the fan cowl 24 is influenced
by many factors as above-described.
Accordingly, the specific profile of the fan cowl
24 of Figure 2 which is effective for promoting the desired
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laminar flow Cp distribution 62 of Figure 3 will vary
according to the particular structural requirements in
any given applicationO To determine the specific
profilel an inverse method of analysis, known to those
skilled in the art, can be used. By this inverse method,
the profile of the fan cowl 24 is systematically varied
and a resulting Cp distribution is analytically or
experimentally determined taking into account any
appropriate factors until the desired Cp distribution
is produced. ~owever, although generally no two
l~m' n~r flow nacelle profiles according to the present
invention will be identical, such nacelles will possess
common features which distinguishthe nacelle over those
of the prior art.
One common feature, as above-described, is the
location of the maximum thickness T along the chord
max
38 at about 50% C to about 60% C and at the position of
minimum Cp 64.
Another feature is shown in the normalized
thickness graph of Figure 4. The m~i mllm thickness TmaX
of the fan cowl 24 is greater than that of the prior
art nacelle 52. Furthermore, the ~agnitude of TmaX
according to the invention ranges between about 6% and about
10% of the chord length C and is preferably about 7%
thereof.
The curvature of the profile of the fan cowl
24 of Figures 2 and 4 according to the invention is also a
significant factor for obtaining the l~ml n~r flow Cp
distribution 62 of Figure 3. Beginning with the region
near the leading edge 34 of the fan cowl 24, as shown
in Figure 2 and in more detail in Figure 6, the leading
edge 34 has a radius of curvature Rl which is less than
about 0.5% of the chord length C. Rl is typically
smaller than those of the prior art nacelle 52 and ranges
between 0.1% and about 0.5% of the chord length C, with
0.1% thereof being preferred.
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adjacent the leading edge 34, shown in Figures 2 and 6,
is suitably aerodynamically blended to the inlet throat
28 according to conventional standards. The curvature
of the outer surface 40 of the fan cowl 24 is defined
in more detail in Figure 5~ which illustrates a graph
of the radius of curvature R of the profiles of Figure
2 normalized with respect to the chord length C and
plotted against X/C. A laminar flow R/C curve 74
according to the invention is indicated in solid line
and, for comparison, a prior art R/C curve 76 for the
prior art nacelle 52 of Figure 2 is indicated in broken
line. The R/C curve 74 is also a significant factor in
defining the profile of the surface 40 for obtaining
reduced aerodynamic drag without boundary layer
separation.
Between 10% C and 56% C, which corresponds
to the intermediate portion 44 of the fan cowl 24 of
Figure 2, the R/C curve 74 is concave with respect
to the abscissa X/C and includes a local minimum R/C
78 at about 65% C.
The R/C curve 74 for both the forward portion
42 and the intermediate portion 44 remains positive in
value indicating that the actual profile of the outer
surface 40 of the fan cowl 24 of Figure 2 is convex with
respect to the chord 38. At about 85% C, corresponding
to the second intersection 50, the R/C curve 74 approaches
an infinite value indicating that the actual profile of
the fan cowl 24 approaches a straight line. Between 85% C
and 100% C, which corresponds to the aft portion 46 of
the fan cowl 24, the actual profile of the fan cowl 24
can remain substantially straight or concave/ with the
R/C curve 74 being negative in value.
In contrast to the prior art R/C curve 76 shown
in broken line in Figure 5, which is continuous and
substantially convex with respect to the abscissa X/C,
the laminar flow R/C curve 74 contains discontinuities
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and both convex and concave portions as above described
which are preferred for increasing the extent of laminar
flow over the fan cowl 24 without causing boundary layer
separation thereover.
Figure 7 illus~rates in further detail the
graph o Figure 4 between 56% C and 100% C. This region
of the fan cowl 24 is significant in promoting the
return of the pressure to the ambient freestream value
without promoting boundary layer separation. More
specifically, the aft portion 46 of the fan cowl 24
includes a chordal angle Y defined as that angle between
the chord 38 and a line connecting the outer surface 40
at the m~x; mllm thickness TmaX and the trailing edge 36.
The chordal angle Y according to the invention is within
the range of about 6 to about 11 and is preferably
about 9. The chordal angle Y is appro~imately twice as
large in magnitude as compared to that of the prior art
nacelle 52 shown in Figure 2. In addition, the aft
portion 46 of the outer surface 40 has a trailing edge
angle Z, defined between the chord 38 and a line tangent
to the outer surface 40 at the trailing edge 36. The
trailing edge angle Z according to the invention is
less than ~hat of the chordal angle Y and is preferably
about 8.
The profile of the outer surface 40, as
illustrated in the Figures and as above-described,
will provide a nacelle having reduced aerodyna~ic drag
compared to typical prior art nacelles. It is to be
appreciated that no one factor alone is effective for
providing extended l~mi n~r flow without boundary layer
separation. The combination of factors as above
described according to the present invention is preferred.
The above description of the profile of
the outer surface 40 is applicable for any longitudinal
section of the fan cowl 24. However, with respect to
any sections about the circumference of the fan cowl 24
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which are influenced by the wing 12, pylon 14, or
fuselage, the profile of the outer surface 40 as
illustrated in Figure 4 can include suitable variations
thereof to account for these influences and still be
within the scope of the present invention.
The nacelle 20, or fan cowl 24 in particular,
provided in accordance wi~h the present invention can
result in a reduction oE aerodynamic drag at cruise of
approximately 50% when compared to prior art nacelles.
However, the leading edge 34 as above-described in less
effective for off-cruise operation of the aircraft. To
improve the effects of the leading edge 34 during off-
cruise operation of the aircraft, a suitable conventional
leading edge device (not shown) can be provided. The
leading edge device is effective for modiying the flow
over the forward portion 42 of the fan cowl 24 for
maintaining a non-separated boundary layer during off-
cruise operation of the aircraft.
Although the invention has been described with
respect to a nacelle 20 comprising a fan cowl 24 of a
high bypass separate flow gas turbofan engine 10, it is
to be appreciated that a suitable l~mi n~r flow nacelle
can be provided for other engine applications.
For example, a laminar flow nacelle 82 according
to another embodi~ent of this invention can be provided
for a single exhaust turbojet or turbofan engine 84 as
is illustrated in Figure 8. The contour of the nacelle
82 is generally similar to the contour of the fan cowl
24 illustrated in Figure 2 and conforms to the normalized
laminar flow proile 72 of Figure 4. Also, the laminar
flow nacelle 82 is effective for producing the laminar
flow Cp distribution 62 as illustrated in Figure 3.
Furthermore, although an annular l~m;n~r flow
nacelle 20 has been disclosed, nacelles having shapes
other than annular can also be provided. For example,
a two-dimensional nacelle (not shown) defined by a
plurality of cowl members can be provided wherein each
cowl member has a profile effective for promoting the
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laminar flow Cp distribution 62 as illustrated in Figure 3.
Of course, it is to be understood that to obtain
and maintain laminar flow over any nacelle surface, the
surface should be designed substantially smooth for
avoiding any discontinuities or sites for propagation
of turbulent flow and boundary layer separation
While there have been described herein what
are considered to be preferred embodiments of the present
invention, other embodiments will become apparent from
the teachings herein.
