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Background
With the frantic development
of advanced aircraft in World War II, the speed
of sound became an operational barrier characterized
by severe aerodynamic problems, including substantial
increases in aerodynamic drag and buffet, rapid
increases in structural loadings, and potentially
catastrophic loss of controllability. During the
war and immediately thereafter, the Langley Research
Center conducted extensive wind-tunnel and flight
investigations to provide a fundamental understanding
of and solutions to the physical phenomena causing
these problems.
A breakthrough occurred in
1950, when Langley modified the operational Langley
8-Foot Transonic Pressure Tunnel with an innovative
slotted throat transonic test section to permit
valid aerodynamic testing in the complex transonic
regime. Using this unique facility, Richard T.
Whitcomb and others conducted experimental studies
of flow fields about aircraft at transonic conditions
to understand the problems that had first been
experienced during the war and to improve aerodynamic
efficiency by reducing or delaying the transonic
drag rise. A particularly informative source of
data was photographs of the extensive shock waves
observed about aircraft models during the tests
in the 8-Foot Transonic Pressure Tunnel. Instead
of individual shock waves for the wing and the
fuselage, as had been expected by many researchers,
a single strong shock wave was observed nearly
normal to the flow direction, crossing the flow
field near the tip of the wing. This observed shock
wave was very similar to that exhibited by a body
of revolution without wings. Given this important
clue, researchers turned their attention to defining
the equivalent body of revolution to minimize the
increased drag caused by the shock wave at transonic
conditions.
Cross-sectional area for
wing-body configuration and for
equivalent of revolution. Note bump in cross-sectional
area
of body of revolution caused by addition of wing
area.
Inspired by a presentation
on transonic flows made at Langley by Adolph Busemann,
a world famous German aerodynamicist who had come
to Langley following World War II, Whitcomb realized
that the transonic disturbances and shock waves
produced by aircraft were a function of the longitudinal
variation of cross-sectional area. As a result
of this phenomenon, the drag near the speed of
sound for a wing-body combination was the same
as that of a body of revolution with the same longitudinal
distribution of cross-sectional area. For most
airplane configurations, adding the cross-sectional
area of the wing to that of the fuselage results
in an abrupt increase, or bump, in the overall
longitudinal area distribution. Thus, to obtain
the minimum shock wave drag, the overall distribution
should be that of a smooth body with minimum drag.
Whitcomb theorized that the most obvious way to
achieve this distribution was to remove the equivalent
wing cross-sectional area from that of the fuselage
cross-sectional area in the region of the wing;
thereby the abrupt bump was avoided in area distribution.
This approach resulted in a pronounced “wasp-waist”
or “Coke-bottle” fuselage shape. The
cross-sectional areas of other aircraft components
(nacelles, etc.) are also included for analysis
of typical aircraft configurations, and the total
area distribution is examined for compliance with
the area rule.
Richard T. Whitcomb with
area-ruled F-106 aircraft (NASA 816) at the retirement
of NASA 816
(used for flight research at NASA Glenn and NASA
Langley) at Langley in 1991.
Whitcomb’s discovery
was initially highly classified, but the aircraft
industry was immediately notified and briefed on
the results of wind-tunnel tests that verified
his hypothesis. Whitcomb was subsequently awarded
the coveted Collier Trophy for his discovery and
the development of the area rule, and history has
recorded numerous applications to military aircraft
beginning with the U.S. Navy’s F11F Tiger,
which almost flew faster than speed of sound without
an afterburner in August 1954. Perhaps the most
dramatic application of the area rule was for the
U.S. Air Force’s delta-winged F-102 aircraft.
After a contract was awarded for the advanced interceptor,
wind-tunnel tests in the Langley 8-Foot Transonic
Pressure Tunnel in 1953 revealed that the transonic
drag was much higher than predicted, and that the
aircraft would not be able to penetrate the speed
of sound. Subsequent flight tests in August of
that same year verified the wind-tunnel predictions
when the YF-102 could not exceed the speed of sound
in level flight. On December 21, 1954, the F-102
with a modified, area-ruled fuselage (known as
the YF-102A) flew through the speed of sound while
still climbing. Whitcomb’s area rule had
saved a critical national military program and
had proven to be the major breakthrough for routine
supersonic flight. Following this famous application,
other famous military aircraft, such as the F-105,
F-106, F-4, B-58, and B-1, were designed with the
area rule as a guiding principle.
Today, the operational flight
envelopes of high-performance supersonic military
aircraft still require consideration of the principles
of the area rule. In the early 1970s, an interest
in higher cruise speeds for commercial transports
resulted in extensive NASA and industry studies
of near-sonic transports that incorporated the
area rule. Today, however, the more limited subsonic
flight speeds used by civil aircraft have not resulted
in any significant use of the area rule for fuselage
shaping of large transports. On the other hand,
as the speed and altitude capabilities of today’s
business jet aircraft continue to increase, the
area rule has entered the design process.
Langley Research and Development
Activities
At the time of Whitcomb’s
discovery of the area rule, the dominant theme
of the user community for both military and civil
aircraft was “higher, faster, and farther.”
Therefore, having successfully applied the area
rule to military aircraft in the 1950s and 1960s,
Whitcomb turned his efforts to potential applications
for subsonic civil transports. Unfortunately, the
relatively low cruise speeds at the time precluded
the application of the concept.
When the supercritical airfoil
permitted serious consideration of higher cruise
speeds, Whitcomb and his staff explored the advantages
of area ruling for advanced transport aircraft.
Several generic models were tested in the 8-Foot
Transonic Pressure Tunnel, and the results indicated
that the concept of area ruling, together with
supercritical wing technology, might permit near-sonic
cruise capability. The integrated principles of
area ruling resulted in configurations with geometries
that provided vivid visual evidence of the careful
tailoring of the cross-sectional area distribution
of the total aircraft. These exciting results and
data were quickly disseminated to the U.S. airframe
industry. Meanwhile, the growing national interest
in faster cruise speeds for commercial transports
maintained Langley’s interest in the area.
Near-sonic transport wind-tunnel
model with area ruling and advanced supercritical
wing.
Area distribution for near-sonic
transport design.
Note variations in fuselage area required to provide
relatively smooth area variation for total aircraft.
Boeing configuration in
ATT studies included area-ruled fuselage.
A series of industry studies
by Boeing, Lockheed, and General Dynamics in the
NASA-sponsored Advanced Technology Transport (ATT)
Program resulted in candidate near-sonic cruise
configurations that employed all of the geometric
principles dictated by the area rule. Each individual
industry design incorporated the graceful, curved
fuselage and shaping characteristics of area-ruled
aircraft.
Langley research on advanced
area-ruled subsonic transports continued until
the fuel crisis of the 1970s virtually eliminated
worldwide interest in near-sonic transport development.
Langley then turned its research emphasis to improving
aerodynamic efficiency at lower cruise speeds by
using the beneficial characteristics of the supercritical
wing. The principles of the area rule, however,
continued to be employed by designers for solutions
to configuration integration issues.
Applications to Civil Aircraft
None of the current U.S. large
commercial aircraft operate at cruise speeds high
enough to require the radical area-ruled fuselage
shapes necessary for a near-sonic transport. However,
designers of large commercial transports have used
the principles of area ruling to solve “local”
flow problems and interference effects—especially
nacelle integration issues for wing- or fuselage-mounted
engines. Following the early development of the
area rule, Whitcomb continued his remarkably intuitive
approach to transonic aerodynamics in efforts that
showed how the principles involved in the area
rule could be used to enhance the overall performance
of transport aircraft, without the radical reshaping
of the entire aircraft required for the near-sonic
transport configurations. For example, in 1958
he developed a special fuselage addition on the
forward part of the upper fuselage that significantly
reduced the shock-induced separation noted on the
inboard upper-wing surface for representative transport
configurations. The fuselage addition resembled
the upper forward fuselage fairing that was later
incorporated into the Boeing 747 transport. Whitcomb
also used area-rule principles in studies of the
beneficial impact of “antishock” wing-mounted
bodies on raising the drag-divergence Mach number
for representative fuselage-wing configurations.
In a series of wind-tunnel studies, he validated
the potential beneficial effects of semiconical
bodies located at several spanwise and chordwise
wing locations. The bodies reduced the local curvature
of the upper surface, a characteristic that favored
the potential for supercritical flow—a concept
that Whitcomb would later explore in the development
of the supercritical airfoil. Fundamentally, the
beneficial effects of these bodies included a deceleration
of the supersonic flow ahead of the shock wave
above the wing, and a decrease in the strength
of the shock and the associated flow separation.
Furthermore, the local pressure fields produced
by the bodies greatly reduced the adverse outward
flow of the separated boundary layer on swept-back
wings. Experiments in the Langley 8-Foot Transonic
Pressure Tunnel were conducted for Mach numbers
from 0.60 to 1.00 for several configurations. The
shapes of the auxiliary bodies were carefully designed
by Whitcomb in adherence to a special extension
of the area rule. In this application, he carefully
chose specific areas of the wing to be considered
in the development of cross-sectional area distributions.
For example, he omitted the cross-sectional areas
of the bodies downstream of the wing trailing edge
because the aerodynamic effects of those sections
were relatively complex and unknown; however, these
effects were probably secondary to those of the
sections of the bodies above the wing surface.
The semiconical forward and upper surfaces of the
bodies were accompanied by a flat lower surface
aft of the wing trailing edge.
Wind-tunnel research model
in Langley 8-Foot Transonic Pressure Tunnel
showing upper forward fuselage fairing and antishock
bodies on wing.
Oil flow visualization
of model wing with 35∞ sweepback at Mach
number of 0.90 and angle of attack of 4∞
(flow is left to right). A significant amount of
unacceptable flow separation is evident for basic
wing (left)
on rearward part of wing; the addition of antishock
bodies (right) greatly reduces separation.
The results of wind-tunnel
tests verified Whitcomb’s intuitive local
application of area-rule principles. For a representative
lift coefficient, the drag-rise Mach number was
increased by approximately 0.05 (from Mach 0.85
to 0.90). A very significant additional benefit
of the bodies was that they eliminated an unacceptable
pitch-up instability exhibited by the high-aspect-ratio
swept-wing models for Mach numbers of 0.80 and
greater. In fact, the configuration with added
bodies experienced significant pitch-down at several
of the test Mach numbers. The pitch-up of the basic
swept wing was expected and caused by severe separation
on the outboard region of the wing, which resulted
in a greater loss of lift on the outer sections.
This favorable effect of the bodies was attributed
to reducing the separation on the outboard region,
which resulted from the lessening of the strength
of the local shock and the retardation of the outflow
of the boundary layer into the outer region.
NASA’s Convair 990
aircraft in 1992. Note antishock bodies on wing.
View from beneath Convair
990 showing flattened lower surface of semiconical
antishock bodies.
Design trade-offs for applications
of the body concept include an assessment of the
additional parasite drag (including interference
effects) caused by the additional bodies. Data
on this novel antishock body concept were quickly
disseminated to the U.S. industry, and Whitcomb
was subsequently awarded a patent for the antishock
body concept.
One of the more significant
examples of the application of the area rule for
local flow problems involved the four-engine Convair
990 jet transport. The 990 was an attempt by the
Convair Corporation to compete with Boeing and
Douglas in the highly competitive jet transport
marketplace of the late 1950s. Unfortunately for
Convair, Boeing and Douglas had captured the early
market with sales of the 707 and DC-8, respectively,
whereas Convair’s initial attempt to enter
the rapidly growing industry was marred by massive
losses of over $425 million on its Convair 880
transport. When Boeing marketed their new 720 transport
it threatened to eliminate Convair from the competition;
Convair responded with a new design designated
the Convair 990, which would be marketed on speed
and luxury. The aircraft would differ from the
earlier 880 in having a stretched fuselage for
increased capacity, a larger wing, and the first
turbofans ever used by a civil transport. The new
turbofans were supplied by the General Electric
Company.
During briefings with Convair
engineers, Whitcomb advised them to incorporate
his concept of concial wing-mounted antishock bodies
for local area ruling of the wing and enhanced
high-speed performance. Impressed with the potential
of the antishock body concept, Convair designed
the 990 with the wing-mounted bodies. The first
flight of the new aircraft occurred on January
24, 1961, and even with the beneficial effects
of the bodies, high-speed drag problems were immediately
noted during the flight tests. The top speed was
limited to 580 mph (40 mph less than the guarantee)
and a serious range deficit was also noted that
would prevent coast-to-coast operations. An extensive
drag reduction program was initiated that led to
modifications that resulted in the achievement
of cruise performance in excess of the original
guarantees. The modifications included a sharper,
less-drooped wing leading edge; a nacelle afterbody
extension; a wing-fuselage fillet redesign; and
the addition of engine nacelle and pylon fairings.
During the drag reduction
program, General Electric representatives requested
the assistance of Whitcomb in minimizing an extremely
large nacelle-wing-pylon interference drag problem
that had been identified in flight tests. Pressure
measurements made around the nacelle afterbody,
pylon, and wing indicated the presence of a strong
shock wave with significant wave drag for aircraft
Mach numbers from 0.80 to 0.90. In addition, shock-induced
separation contributed to the drag problem. The
new turbofan engines had the fan located toward
the rear of the engine; this location resulted
in a sudden increase in area distribution near
the wing trailing edge. Essentially, flow encountering
the convergent-divergent channel between the nacelle,
pylon, and lower wing surface was being accelerated
to supersonic conditions, which resulted in a standing
shock. The level of drag rise for the entire aircraft
with increasing Mach number above 0.80 was approximately
equal to the nacelle afterbody pressure drag. Whitcomb
analyzed the problem using the principles of his
area rule on a local basis. In particular, the
area contained by the nacelle upper surface, pylon
side surface, and wing lower surface was analyzed
for each nacelle in terms of smoothness of the
area distribution and found to have abrupt changes
in area distribution (due to the pylon and fan
location) along the length ranging from the nacelle
intake to the trailing edge of the wing for both
the inboard and outboard nacelles. Any fixes for
the problem could not change the wing or the nacelle
basic lines, but auxiliary fairings could be added
to the pylons and nacelles. Following applications
of the local area rule, several pylon, nacelle,
and wing fairings were proposed to smooth out the
area distribution, and the most effective configurations,
consisting of forward and aft pylon fairings, were
adopted for production aircraft. This configuration
resulted in a significant increase in the drag-rise
Mach number for the aircraft, from about 0.80 for
the basic configuration to about 0.89 for the modified
aircraft. NASA later acquired a Convair 990 aircraft
for use in its research programs at the Dryden
Flight and Ames Research Centers for activities
ranging from evaluating new landing gear and brake
designs for the space shuttle to direct lift control
and medium-altitude research missions.
Sketch of 990 pylon and
nacelle fairings used for production aircraft.
Cross-sectional views at right are looking forward.
The Douglas DC-8 Super 62
aircraft with long nacelles.
Another successful example
of the use of the area rule for local interference
drag analysis occurred in the Douglas DC-8 transport
program. During a prototype flight investigation
of a new long duct nacelle for the DC-8, flight
results obtained with a proposed new nacelle afterbody
resulted in a much greater interference drag than
had been indicated by wind-tunnel tests. In fact,
the penalty measured in flight was double the wind-tunnel
value for representative cruise conditions. Examination
of pressure distributions on the nacelle in the
channel between the wing and nacelle indicated
that the shock in the channel was significantly
stronger and farther aft in flight than in the
wind tunnel; this caused very high levels of drag.
The difference between the tunnel and flight results
was attributed to the differences in boundary-layer
growth because of corresponding differences in
Reynolds number. Applications of Whitcomb’s
local area-rule methodology resulted in fairing
candidates that eliminated the problem. The successful
application of the area-rule process and the elimination
of what would have been a major performance penalty
for the long duct nacelle configuration provided
Douglas with the confidence and enabling technology
to proceed with the new versions of the DC-8, the
highly successful “Super Sixties” (DC-8-62
and DC-8-63).
Illustration of wing-pylon-nacelle
interference drag flow phenomenon for DC-8 indicating
difference in shock wave locations and relative
strengths for wind tunnel and flight.
Cessna Citation X with local
area-ruled features incorporated in its lower and
aft-fuselage shapes.
Another example of the use
of local area ruling was the successful design
of the centerline engine installation on the DC-10.
Application of the area-rule concept by McDonnell
Douglas provided the guidance needed to properly
locate the various components (i.e., the inlet
cowling relative to the support strut, strut shaping
versus horizontal tail location). This approach
significantly contributed to the aircraft meeting
its nominal (not just guarantee) performance levels.
Yet another, more recent, application of this principle
by McDonnell Douglas was for the engine pylon design
for the MD-90.
Aircraft designers quickly
absorbed the lessons learned through application
of the principles of the area rule for local flow
interference solutions, and the approach became
a general design technique that has been used for
the analysis and improvement of high-speed aerodynamics
for pylon-mounted engines on wings, pylon-mounted
engines on fuselages, and externally mounted stores.
In recent years, competing
elements within the business jet community have
pushed the cruise speed and altitude capabilities
of advanced business jet aircraft to near-sonic
conditions, requiring the incorporation of Whitcomb’s
principles for efficient cruise. A current application
of area ruling within the civil community is the
advanced Cessna Citation X business jet aircraft,
which nominally cruises at a Mach number of about
0.92 at altitudes of about 30,000 ft. Careful tailoring
of the fuselage, wing, engine pylon, and engine
nacelle geometries according to the general and
local principles of the area rule, along with the
implementation of other innovations such as the
supercritical wing, has resulted in the fastest
civil aircraft in the world (excluding the supersonic
Concorde). In recognition of the outstanding performance
and design of the aircraft, Cessna was awarded
the Collier Trophy for 1996 for the most significant
aeronautical achievement in the United States.
The SX-30, Raytheon Premier I and Hawker Horizon,
and Dassault Falcon 50, 900, and 2000 aircraft
all exhibit significant contouring in the aft-fuselage
area to minimize nacelle interference drag at transonic
speeds.
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