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A9946057
AIAA 99-0118 Applied Aerodynamics at the Douglas Aircraft Company - A Historical Perspective
Roger D. Schaufele Douglas Aircraft Company (Retired)
37th AIAA Aerospace Sciences Meeting and Exhibit
January ii-14,1999 / Reno, NV
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Applied Aerodynamics at
the Douglas Aircraft Company -
A Historical Perspective
Roger D. Schaufele ’
Douglas Aircraft Company (Retired)
Abstract The application of the fundamental principles of
aerodynamics in the design and development of the Douglas
Commercial ( DC ) series of aircraft and other related
designs is reviewed. The aerodynamic design methods and
procedures utilized in the early years are outlined. along with
some of the more notable aerodynamic design features of
the aircraft. Later developments in aircraft design with more
sophisticated aerodynamics led to more detailed methods.
which found application on later transports in the series. The
arrival of the jet transport era brought new challenges for the
aerodynamicists, who again had to come up with the new
methods to cope with the requirements of the new speed
regime. The use of basic aerodynamic concepts in the
solution of some interesting and unique problems that arose
in the design and development of the jet transport models is
also discussed.
Introduction Before getting into the history,it is important to
define “applied aerodynamics” as it will be used throughout
this paper. The term “applied aerodynamics” describes the
use of methods and procedures based on fundamental
aerodynamic theory or principles in the design of a specific
aircraftTheoretical aerodynamics serves as a marvelous
basis for calculating any number of important quantities
needed in aircraft design. However, the assumptions made
in order to develop the theory often result in differences
between the calculated values and those obtained from
experimental data. These differences were accounted for
insofar as possible with methods and procedures that were
based on the fundamental principles of aerodynamic
theory,corrected where needed by factors derived from a
correlation of measured data with available theory. These
correction factors, also known as “fudge” factors, embodied
a certain amount of risk if applied without some
understanding of the basis of the correction, and quite often
the need for the correction factors was eliminated as more
’ Fellow, AMA
detailed knowledge of the parameters was gained. As an
example, thrust horsepower (THP) is determlned ,from
(BHP) from the familiar equation
THP=BHPx 7
The value of T] was estimated from a propeller chart that
gave ?‘J as a function of advance ratio, J, and power
coefficient, C, . This was then multiplied by fq to account for
the fact that the propeller blade geometry was different than
the propellers used to experimentally determine the propeller
efficiency. Then there was another factor k ,., which
accounted for the compressibility effects on the propeller
efficiency. And finally there was another factor k, which was
called the engine manufacturers “honesty” factor, the
performance engineers audit of actual versus published BHP
output of the engine. For takeoff performance, THP was
further multiplied by a factor F, to match flight test. This factor
was eliminated when NACA published improved experimental
results on compressibility effects on propeller efficiency, and
the BHP factor went away when torquemeters became
available to measure engine BHP.
Applied aerodynamicsas practiced in industry, really
focused on several aspects of the aerodynamics discipline.
First there is the aerodynamic analysis associated with
aircraft performance. The basis for aircraft performance
analysis is found in the flight mechanics equations that
govern the various regimes of flight, namely takeoff, climb,
cruise, descent, and landing. In order to calculate aircraft
performance,detailed information on a number of
aerodynamic and propulsion parameters must be available.
Of prime importance is the drag determination for all
regimes of flight, followed by the determination of the
maximum lift coefficient in cruise, takeoff. and landing
configurations, In addition there are aerodynamic analyses
to determine the lift curve slope, zero lift angle.and flap
effectiveness. Also important in the performance area are
the aerodynamic aspects of the propulsion installation. In the
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early days, this had to do with the design of the engine
nacelle and cowling, the engine cooling drag, propeller
efficiency, exhaust recovery, and the drag associated with
carburetor and oil cooler air scoops, and in later years the
placement of the jet engine nacelle for minimum interference
drag, the design of the inlet, the assessment of the inlet and
exhaust nozzle losses, and the determination of engine
thrust reverser characteristics The aerodynamics associated
with performance prediction has always received a great
deal of attention, since the performance of a new design is
calculated two to three years prior to first flight and usually
guaranteed with very tight tolerances in the contractual
purchase agreements. Any undue optimism or pessimism in
the predicted performance can have dire consequences to
the program. In addition, there are the aerodynamic
analyses associated with the stability and controllability of
the aircraft. Here the emphasis is on determining the static
longitudinal, directional, and lateral stability characteristics of
the aircraft, and the related control surface effectiveness and
hinge moments needed to meet the specific design
requirements for the flight control system Following the
control system design, analyses must be made of the critical
control conditions in pitch, yaw and roll And finally, there are
the aerodynamic analyses associated with aerodynamic
loads, needed to design the structure Another very
important aspect of applied aerodynamics is in the
“troubleshooting” and fixing of aerodynamic problems which
show up in detailed wind tunnel and flight testing. Here a
good understanding of the fundamentals plus a great deal of
imagination are required to resolve some very difiicult
problems well beyond available theory
The DC-1 and DC-2 The history of applied aerodynamics at
the Douglas Aircraft Company really began with the design
of the DC-1 transport in 1932. Prior to this time,
aerodynamics was not really a major part of the design
effort, as shown by some of the earlier Douglas designs, for
example the Douglas C-l, the U.S. Army’s first cargo
airplane, Fig. 1, The DC-1 had a number of aerodynamic
design features new to Douglas airplanes, including NACA
4-digit series airfoil sections (Ref.1) fully cowled radial
Fig. 1 The Douglas C-l
aircooled engines (Ref.2) constant speed propellers,
retractable, but not fully enclosed, main landing gear. a large
span trailing edge split flap, and a very generous wing-to-
fuselage fillet. The DC-l was also remarkably free of
external protuberances such as flap and control surface
mechanism fairings. While many of these features were new
to commercial transports, they had been developed and
documented by wind tunnel testing by both the NACA in their
Langley wind tunnels and by Douglas in the GALCIT’ IO -
foot wind tunnel at Caltech Not much is known about the
method used to predict the aerodynamic drag characteristics
of the DC-I, but the method certainly involved estimating the
airplane parasite drag coefficient by summing up the
contributions of each of the various elements of the aircraft
using turbulent skin friction coefficients at the appropriate
Aeynold’s number, adjusted for the surface condition
(brazier head rivets, skin lap joints, etc.), form factors to
account for the zero lift pressure drag of the major elements,
as well as estimates of the engine cooling drag, oil cooler
drag, carburetor intake drag, and miscellaneous drag items
such as radio antennas, non-retracting tail wheel, and
~~ engine exhaust pipes. The induced drag was certainly
estimaled using the classical Prandtl wing theory, corrected
by the empirical “e” ior non-elliptic span loading and the
increase in parasite drag with lift. The maximum lift
coefiicient flaps up was probably estimated from the
available NACA section data, adjusted for three- dimensional
effects and downward horizontal tail loads for trim. Wind
tunnel data was undoubtedly useful in checking the
‘Guggonhoim Aoronautlcal laboratory, California institda of Technology
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estimates, especially for the flaps down configuration. These
relatively simple procedures, now explained in detail in
nearly every text on aerodynamic performance, were
adequate to allow the aircraft to meet all of the performance
requirements with some margin to spare.The aerodynamic
stability and control methods must have been quite minimal,
although legend has it that the distinctive wing planform of
the DC-l, with the straight trailing edge and swept leading
edge on the outer panel came about in order to maintain the
center of gravity forward of the aerodynamic center under all
loading conditions, thus achieving the condition for static
longitudinal stability. Adequate directional control for takeoff
with one engine inoperative was also achieved, undoubtedly
with the help of the wind tunnel tests, and a modest
aerodynamic horn balance was used on the rudder to reduce
pilot forces. Some believe that Jack Northrop had a hand in
the aerodynamic design of the DC-l, since he and Donald
Douglas had joined forces in early 1932 to establish the
Northrop Corporation with Douglas owning 51% of the stock.
Indeed, many of the advanced features of the DC-1 were
also found on Northrop’s earlier “Alpha”, “Beta”, and
“Gamma” shown in Fig 2 The application of these advanced
Fig. 2 The Northrop Gamma
(for that time) features to the prototype DC-l, Fig 3. and the
nearly identical production DC-2, Fig. 4. resulted in an
aircraft with greatly improved performance and economics
over the other commercial transports of that era.
In the later stages of the DC-l development, Donald
Douglas hired William Bailey Oswald as the first full time
aerodynamicist for the Douglas Aircraft Company. “Ozzie”.
as he was known throughout the company, had received his
Fig. 4 The DC-2
PhD. from Caltech in 1932. and had conducted some of the
wind tunnel tests on the DC-l. In his doctoral dissertation on
a systematic approach to the calculation of aircraft
performance, he made a significant contribution to the
application of the Prandtl three-dimensional wing theory. His
concept of the aircraft efficiency factor,“e”. was used to
empirically correct the theoretical induced drag to account
for increases in the parasite drag with lift coefficient,
The DC-3 The next of the DC series, the DC-3, Fig. 5, did
not really incorporate any really new aerodynamic
applications, although there were some minor changes in the
aerodynamic balance on the rudder and elevator to reduce
pilot forces, and the addition of a small dorsal to the vertical
fin to avoid rudder lock at high sideslip angles. However, the
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well developed overall design, combined with an increase in
passenger capacity of 50% brought a new level of efficiency
and passenger acceptance. Features carried over from the
DC-2. such as the cabin interior with true stand up height
and a floor uninterrupted by steps to climb over the wing
spars, and the use of stressed skin construction where the
wing skin carried much of the wing airloads to reduce the
wing weight. and the partially retracted landing gear to
reduced cruise drag, combined with the increased
passenger capacity resulted in outstanding direct operating
costs and made the DC-3 a great step forward in air
transportation.
Fig. 5 The DC-3
Under Dr. Oswald, the Aerodynamics group at the
Douglas plant in Santa Monica participated in the
development of a number of new aircraft in the 1930’s. The
B-18 was a medium bomber based on the use of DC-2
wings and engines, and the DC-3~empennage fitted to a new
bomber fuselage. The B-19, a one-of-a-kind experimental
long range bomber, the B-23, a 300 mph medium bomber,
and the first attempt at the DC-4. later called DC-4E for
experimental, Fig. 6, another one-of-a-kind aircraft. a four
engine commercial transport whose development was
financially supported by American, Eastern. Pan American,
TWA, and United. The aerodynamic design features of all of
these aircraft were very similar to the DC-3. However, much
progress was being made in developing aerodynamic design
methods based on fundamental thoery. but modified as
Fig. 6 The DC-4E
required based on correlation of both wind tunnel and flight
test data with the theory.
It should be noted that in 1938. Jack Northrop
severed his business relationship with Donald Douglas and
formed a new Northrop Corporation. The assets of the
former Northrop corporation became the El Segundo division
of the Douglas Aircraft Company, The El Segundo division
was a complete design and production organization and as
such had it’s own Aerodynamics group. This organization
designed and produced some notable military aircraft in the
following 25 years that they were a separate division of the
Douglas Company. including the SBD scout bomber, the
DB-7 I A20 , A26 , AD, A3D, and A4D attack aircraft, the
F3D and F4D navy fighters. but the aerodynamic design
features and the design methods were similar to those used
at Santa Monica.
The DC-4 In-1939 the. Santa Monica division undertook the ~- development of the production DC-4 four-engine commercial
transport Fig. 7. It’s aerodynamic design features included
the use of newer NACA 5-digit series airfoils (Ref. 3) and
large chord single-slotted flaps(Ref 4) for increased
maximum lift capability. a fully enclosed retractable landing
gear, and an aerodynamic boost, linked tab flight control
system. While not exactly an aerodynamic design feature,
the DC-4 also incorporated a constant cross-section
fuselage in the passenger cabin, which allowed easy
incorporation of additional fuselage “stretch” on later models
with increased passenger capacity. The wing design was
3
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Fig. 7 The DC-4
straightforward. defined by two airfoil sections, an NACA
23018 at the root, and an NACA 23010 at the tip, arranged
with 3 degrees of twist The design of the single-slotted flap
was based entirely on available NACA wind tunnel data,
since a siuitable theoretical multi-element airfoiranalysis was
not yet available For the design of the aerodynamic boost
linked tab flight control system, there were several
documents which laid the foundations for the design and
analysis (Refs 5,6,7.6) However, correlation of the theory
with experiment was not very far advanced, and practical
design methods were just being developed. As an example,
the approach utilized to determine the static longitudinal
stability characteristics of the DC-4 involved the calculation
Of contribution of the various elements of the configuration to
the aerodynamic center location for the complete aircraft
from available data,such as (Ref. 9, 10, 11, 12). Then wrnd
lunnel tests were conducted in the GALCIT 10-R wind lunnel r to measure the individual contributions, thus providing a
direct check on the calculation procedure. Fig 8 shows the
wind tunnel data for the aerodynamic center buildup, and the
effect of each element on the overall configuration
aerodynamic center. The testing also included the effect of
operating propellers at takeoff, climb, and cruise power,
which provided data on Ihe loss in longitudinal stability,
especially al high power settings. Additional data was
obtained on elevator effectiveness and hinge moments at
various elevator deflectionsin the cruise, takeoff, and
landing configurations for the calculation of control tab
deflection and control forces.Similar approaches were used
to develop suitable directional stability and control
characteristics That is. calculation of the contributions of the
various elements of the configuration to stabilily,
documentation of the calculations by wind tunnel lest data.
and verification of the calculated critical control requirements
by the wind tunnel test data All production aircraft, over
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Fig 8 DC-4 Aerodynamic Center Buildup
1300 in total, were diverted to the USAAF as C-54 military
transports. The flight test program again provided significant
flight test data which allowed further correlation and
refinement of the aerodynamic design and performance
methods then being used.
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The DC-5 Also In 1939, the Douglas El Segundo division
undertook the design of a small, short range commercial
transport, the DC-5 Fig.9. Aerodynamically, the design of the
DC-5 was roughly similar to the DB-7 / A20 light bomber.
The aerodynamic design and performance methods utilized
were the same as those used at thesanta Monica division,
and the projected performance was quite attractive.
However, the wind tunnel testing failed to reveal rather
severe tail buffet in certain configurations, which required the
addition of dihedral to the horizontal tail during the flight test
program. That change, plus additional structural
modifications and the start of WW II limited the production to
just 12 aircraft.
Fig. 10 The X8-42
bomber, the XBi43, Fig Il. with a maximum speed of well
over 500 miles per hour or nearly 0.7 Mach number, which
began to create additional challenges for practicing
aerodynamicists. Another notable design initiated at Santa
Fig. 9 The DC-5 Fig. 11 The XB-43
During the war years, the Douglas Santa Monica
division deslgned, built, and flight tested a noteworthy new
aircraft for the military, the XB-42 bomber, Fig. 1 O.The XB-42
was unique aerodynamically in that it had a clean straight
wing unaffected by engine nacelles or propeller slipstream.
since the two engines were housed inside the fuselage and
drove two counter-rotating propellers located at the aft end
of the fuselage. The wing design featured NACA 6 series
“laminar flow” airfoils and single-slotted flaps. Although only
three of these aircraft were built, the combination of high
installed power and very clean lines resulted in a maximum
speed capability of over 400 miles per hour or nearly 0.6
Mach number. With the development of jet engines during
the war, the XB-42 design was convened to a twin engine jet
Monica was the C-74 military transport, Fig.12. The C-74
was a very large, long range aircraft with very conventional
aerodynamic features; NACA 6 series airfoils, single slotted
flaps, aerodynamically boosted linked tab flight controls, and
neatly cowled air cooled radial engines. One notable design
concept was the use of full span trailing edge flaps, with the
outer sections of flap operated differentially as ailerons.
During flight tests, it was concluded that the gain in
maximum lift capability due to this feature was nol worth the
additional mechanical complexity, and the idea was dropped
from production aircraft. It should be noted that the detailed
design of the C-74 was done at the Douglas Long Beach
Division, which, like the El Segundo Division, became a
complete engineering design and manufacturing facility,
-
-
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Fig. 12 The C-74
designing and producing the C-124, C-133, and B-66 aircraft
for the U.S. Air Force. The design of these aircraft, each with
some unique aerodynamic features, further stimulated the
development of aerodynamic prediction methods for
performance and stability and control. The flight lest data
from these aircraft further enhanced the ability to correlate
predictions with actual results, allowing additional
refinements to be made to the methods.
The,DC-6 The next in the series, the DC-6, Fig. 13. was an
extension of the DC-4, featuring more powerful engines.
cabin pressurization, NACA double-slotted flaps, and an
improved aerodynamic boost linked tab flight control SyStem.
The aerodynamic methods were refinements of those used
on the DC-4, and except for the maximum lift coefficients in
the flaps down configurations, the aerodynamic
characteristics were not much affected. Minor cleanup of the
engine nacelles, including neater exhaust stack treatment
and redesigned carburetor air intakes, plus a reduction in the
number of radio masts and antennas resulted in a lower
overall drag level compared to the DC-4. The new double
slotted flaps resulted in a modest increase in the maximum
lift capability which helped keep takeoff and landing
distances in line at the higher weights of the E-6. The
aerodynamic boost linked tab flight control system was
simple and reliable and provided comfortable control forces
without the complication of hydraulic operation. The
development of the aerodynamic boost concept and its
application to the DC-6 is outlined in Rei, 9 The
aerodynamic wind tunnel and flight test programs on the
DC-6 went very smoothly, and the airplane met all of its
performance objectives easily The DC-6. but more notably
the slightly stretched DC-GB, set the standard for transport
performance and operating cost in the early 1950’s.
The DC-7 The next model, the DC-7, Fig. 14, was
conceived as a re-engined version of the DC-GB, designed
to fly LAX- to- JFK non-stop in just under eight hours. The
aerodynamic changes were confined to the engine
installation, with different nacelle lines to accomodate the
more powerful turbo-compound piston engines. One new
feature was the use of the main gear as a speed brake to
slow down to the rough air speed or to descend more rapidly
from cruise altitude. The last of the piston engine Douglas
airliners was the DC-7C. Fig. 15, designed in the mid 1950’s
Fig. 13 The DC-6 Fig. 14 The DC-7
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Fig. 15 The DC-7c
as a long range version of the DC-7. In order to meet the
longer range requirements, the fuel capacity, the maximum
takeoff gross weight and the wing area had to be increased.
This was accomplished by increasing the wing span through
the addition of a ten foot section to the wing at the centerline.
This change moved the engines further away from the
fuselage, resulting in less noise in the passenger cabin. The
landing gear also moved further from the centerline, giving
the airplane improved ground handling characteristics,
especially in crosswinds. The added section increased the
wing aspect ratio for improved cruise L/D, and the additional
wing area with trailing edge flaps installed increased the
maximum lift capability significantly. The DC-7C was the
ultimate piston engined transport, with a full passengers and
bags range of over 5000 statute miles, and a high speed
cruise of 350 miles per hour. The aerodynamic performance
was better than predicted in all areas, lending further
credence to the methods developed at Douglas to calculate
all of the elements of piston engine aircraft performance.
In retrospect, the applied aerodynamics of the piston
engine Douglas commercial transports focused on these
main elements. First, the use of the best available
aerodynamic technology, usually the result of NACA
research, well documented by experimental data. Secondly,
the development of detailed methods to predict the
aerodynamic performance, stability and control
characteristics, based on the best available theory,
correlated with appropriate wind tunnel and flight test data.
Thirdly, the use of systematic wind tunnel testing of a new
model to confirm insofar as possible the predicted
characteristics prior to flight test And finally, the progressive
improvement of a basic design, (DC-l, DC-2, DC-3). and
(DC-4, DC-6. DC-7, and DC-7C) through the application of
new piston engine technology and refinements in the
aerodynamic design features.
The Cominq of the Jet Transports In the late 1940’s and
early 1950’s the commercial transport industry was
transitioning from the pre-World War II piston engine designs
to the post war turbine engine configurations. Jet transport
studies were conducted by several manufacturers in the U.S.
and Great Britain By mid 1949, the DeHavilland Comet I
had made its first flight, followed almost immediately by the
first flight of Avro of Canada’s Jetliner However, because of
the limited thrust output and poor specific fuel consumption
of the available centrifugal flow jet engines, these aircraft
had payload-range performance and operating economics
that were inferior to the best of the piston engine designs, so
that they had a very limited market But jet engine
technology was advancing rapidly and by the early 1950’s
larger, more fuel efficient axial flow jet engines that made
possible larger capacity jet transports with competitive
payload-range performance and operating
economicsBoeing was sufficiently interested in the potential
jet transport market that the company invested a large
amount of it’s own funds to design, build. and flight test a jet
transport prototype, the Model 367-80 to gain experience
with this new type. Douglas continued to do design studies
and conducted some preliminary wind tunnel tests. Finally,
by early 1955, the major world airlines were convinced, on
the basis of the demonstrated performance of the Boeing
367-80 prototype and the design studies by Douglas, thal a
fast, efficient, economically competitive jet transport could be
built, and they urged the manufacturers to offer specific
designs for their consideration,
The DC-6 The DC-E. Fig. 16, incorporated a number of new
aerodynamic features associated with the expanded speed-
altitude envelope. A new swept wing, using Douglas
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,_ - . ” . “~ I . . _ _” . . . - _ . ~ . - “ . . “ . ~ ,111 . . ”
Fig. 16 The DC-6
designed airfoils was needed to meet the high Mach number
cruise requirements. An improved double slotted flap was
added for good low speed pet-formance.The jet engine
nacelles were suspended from pylons below the wing.
Aerodynamic boost control was retained for the elevator in
conjunction with the hydraulically trimmable stabilizer,but
increased control surface deflection and low control force
levels required the ailerons and rudder to be hydraulically
operated through fully powered irreversible actuators with a
unique feature that reverted control to an aerodynamic boost
system of reduced capability in the event of loss of hydraulic
power. Small upper surface spoilers on the wing were used
on the ground to increase the load on the landing gear for
landing and rejected takeoff (RTO) braking. Anti-skid braking
was also used for the first time to further improve stopping
pet-formance.ln-flight thrust reversing was used for slowdown
and rapid descent, as well as for additional stopping
capability, especially with adverse runway conditions. The
design and development of these features of the DC-8
proved to be a formidable challenge.
Interest in compressible flow phenomena by
practicing aerodynamicists began to grow in the early years
of World War II, and the fundamental aerodynamic
relationships were set forth in Ref. 10. By the late 1940’s
the concepts of an additional element in the drag equation,
the compressibility drag, AC,, and drag divergence Mach
number, Morv, were well established, but there was no
available theory which could be used lo calculate transonic
flow with local supersonic zones and shock waves, the
situation at cruise conditions. In addition to the cruise
requirement, there was the challenge of achieving good
maximum lift capability in the takeoff and landing
configurations, as well as good stall characteristics. The
wing design was a major aerodynamics group task for which
there were only very approximate methods for establishing the
geometry needed to meet the program performance
objectives. Nevertheless, there was an empirical method,
Ref.1 1, developed at Douglas in the late 1940’s. based
almost entirely on small scale wind tunnel model tests, that
provided an estimate of the wing Morv as a function of airfoil
thickness ratio, sweep, and lift coefficient for wings using
different types of NACA airfoils. The first swept wings
designed by Douglas were used on the D 558-11 research
aircraft and the A3D attack bomber. These wing designs
were done at the El Segundo Division in the late 1940’s but
produced little flight test data to verify their behavior at
transonic cruise conditions. Further study of the Morv
characteristics of airfoils, Ref. 12. led to the conclusion that
certain airfoils had higher values of M orV than others, for a
given thickness ratio and lift coefficient. An explanation for
this behavior was found in Ref. 13, which was the basis for
the “crest line” concept used by Douglas in transonic wing
design for the DC-8~ Briefly stated, the “crest line” concept
relates the Morv for any airfoil section of the wing to the
condition where the pressure coefficient at the airfoil “crest”,
the point on the airfoil that is tangent to the free stream
velocity, first indicates sonic velocity normal to the sweep
angle. Airfoils that carried a lot of negative pressure (lift) on
the upper surface forward of the crest had the the highest
Morv for a given thickness ratio and lift coefficient. With this
experimental evidence in hand, a method was devised to
design the wing with the highest possible Morv while taking
into account the loss in aerodynamic sweep in the root and
tip areas. The procedure involved the construction of curves,
based on wind tunnel model chordwise pressure data for
swept wings that related the measured pressures at
conditions just prior to Morv to those calculated by Ref. 14 for
the 2- dimensional airfoil section in incompressible flow
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01 0
I I I 20 40 60
PERCEM CHORD
Fig. 17 Wlng Airfoil Pressure Growth Curve
(M=O). These “growth curves”, Fig.17. were constructed for
the root, midspan, and tip areas of the wing and were used
to tailor the 2-dimensional airfoil incompressible pressure
distributions to achieve the desired 3-dimensional pressure
distributions near MI-J”, Using this design approach, it was
reasoned that the desired Molv of 0.82 could be achieved
with a wing sweep angle of 30 degrees and an average
thickness ratio of just under 11%. The wing sweep was less
than the 35 degrees used on nearly all of the production
military swept wing aircraft of that time, and the competing
Boeing 707. but the lesser sweep resulted in higher
maximum lift capability, more conventional stall
characteristics, and more favorable lateral-directional “Dutch
Roll” characteristics in cruise.
The wing design approach to achieve high maximum
lift capability, CL,,,~~, and good stall characteristics was
based on a concept that was outlined in Ref. 15. The key
ideas are that the CL,-,,~ for the basic wing can be estimated
by determining the wing CL where the span loading, Fig.1 8,
expressed in terms of local lift coefficient, becomes tangent
to the curve of airfoil section maximum lift coefficients across
the span, and that the initial stall would be expected to occur
on the portion of the wing where the span loading becomes
tangent to the curve of airfoil section maximum lift
coefficients across the span. The initial stall point should be
located such that there is some margin between the outer
panel airfoil clmax values and the span loading of c,at the
initial stall point. The maximum airplane lift coefficient is
Fig. 18 Wing Span Load and Maxlmum Llft
achieved by the use of airfoils with high section maximum lift
coefficients.
The wing spanwise lift distribution was estimated
using the method of Ref.16. The airfoil section maximum lift
coefficient for the three defining airfoil sections was
estimated using the method outlined in Ref. 17, which is
based on a correlation of measured airfoil section maximum
lift coefficients with the theoretical pressure difference
between the peak pressure near the airfoil nose, and the
pressure at 90% of the airfoil chord. The theoretical
pressures were calculated by the method of Ref. 14.ln fact,
the method of Ref. 17 was further developed at Douglas to
relate the peak pressure near the airfoil nose to the
geometry of the airfoil nose shape between 0.15 % chord
and 6.0% chord points. The change in height of the upper
surface coordinate between these two stations,A y is
correlated with measured airfoil c~,,,,~ data, Fig. 19, and
shows surprisingly good agreement.This so-called A y
method of estimating airfoil section c Imax has become a
standard method described in numerous texts on
aerodynamics_ Fortunately, the airfoil section shapes
required for the cruise conditions and the nose shapes
required to achieve high values of section cL max were
(c)l999 American Institute of Aeronautics & Astronautics
2.0
16
-: 12 u
09
04
Tern dam from NACA TR 924. RN = 9.OW.MlO AirfoIl thickness 12% or less
0.15% C from L.E
0 4
Fig. 19 Alrfoil Section Maximum Lift Correlatlon
compatible, so the wing design was based on three specially
designed airfoil sections, located at 25%, 55%. and 95%
semispan, Fig.20. Incidentally, over the years, there were
well over 1000 special airfoils designed by Douglas
-. .-
- x ( VP ---
Flg. 20 DC-8 Wing Planform and Alrfolls Fig. 21 DC-8 Wing Leading Edge Slots
aerodynamicists for a variety of research, development, and
production applications.
During the wind tunnel model test phase of the DC-8
program, it was apparent that the wing design method had
failed to account for the detrimental effect of the engine
nacelle pylons on both the low speed and high speed
aerodynamic characteristics of the wing. Concurrent wind
tunnel tests had indicated that the nacelle position for
minimum interference drag at cruise was below and forward
of the wing, resulting in a pylon that fitted over the wing
leading edge and extended forward of the wing. At low
speeds, the pylons interfered with the outward spanwise flow
near the leading edge stagnation point, causing exlremely
high suction peaks on the wing upper surface just inboard of
the pylons at high angles of attack, This situation resulted in
premature stalling of the wing just inboard of the pylons, and
consequently lower than predicted &,ax values. A solution
for this unanticipated problem was developed in the wind
tunnel and consisted of a short span leading edge slot,
-located just inboard of each pylon, which opened as the
flaps were extended, Fig. 21. These slots relieved the
interference caused by the pylons, and allowed the wing to
achieve it’s design Cu.,,ax capability, At high speed, the wind
tunnel gave mixed indications regarding the behavior of the
compressibility drag rise in the presence of the nacelle
pylons, depending on model scale, transonic tunnel facility,
test Reynolds number, and type of boundary layer transition
fixing used. After many wind tunnel tests and a number of
flight tests, it was concluded that although the design value
Slots Closed - Flnps UP i
(c)l999 American Institute of Aeronautics & Astronautics
of Motv was achieved, the interaction between the pylon and
the wing airfoil nose shape produced a significant
compressibility drag increment prior to Motv. The wing
airfoils forward of the front spar were extended by 4% of the
wing chord, and the nose shapes sharpened, Fig.ZZ.which
eliminated about half of the undesireable compressibility
drag rise prior to Mow Finally the pylon was redesigned to
Original Mldspan Airfoil
4% Extended Mtdspan Airfoil
Fig. 22 DC-8 Origlnai and Modified Airfoils
intersect the wing lower surface aft of the cruise stagnation
point, rather than fitting around the wing leading edge, Fig.
23.This “cutback” pylon essentially eliminating the
Original Pylon
Cutback Pylon
Fig. 23 Origlnal and Cutback Pylon
.0060
Original
.0040
G,
.0020
4% Leading Edgo
4% Leading Edge
.5 .6 .7
Mach number
.I3 ‘9
Fig. 24 Compressibility Drag Rise Comparison
interference between the pylon and the airfoil nose in cruise,
and allowing the wing to achieve it’s design compressibility
drag, Fig. 24.
Another new aerodynamic design issue that came
with jet transports was the definition of the buffet boundary,
described by a single curve of lift coefficient for the start of
buffet, or buffet onset. versus Mach number. Inside the
buffet boundary, the airplane can operate smoothly over a
range of speeds and altitudes. Outside the buffet boundary,
the airplane is subjected to significant separated unsteady
airflow over the wing, which results in noticeable shaking or
“buffetting” of the structure and flight controls. This buffetting
can be severe enough to cause minor structural damage to
control surfaces, and can be associated with longitudinal
pitch-up, or lateral wing drop. For the DC-B, early estimates
of the buffet boundary were made using a method based on
the physical phenomenon involved in the flow separation. At
low Mach numbers the lift coefficient for buffet onset was
related to the approach to the airplane maximum lift
coefficient. At higher Mach numbers, the lift coefficient
for buffet onset was related to a margin beyond the Motv for
that lift coefficient. This method gave a reasonable
preliminary definition of the buffet boundary, until wind tunnel
model wing pressure data could be obtained to define the lift
coefficient and Mach number where trailing edge separation
occurred. The wind tunnel model data was also used to
assess the longitudinal and lateral stability beyond buffet
onset with respect to pitch-up and roll off.
(c)l999 American Institute of Aeronautics & Astronautics
The aerodynamic boost control system for the
elevator worked out well, with control surface and and tab
showing well behaved characteristics, Ref 18. up to a Mach
number of 0.96. the design dive Mach number. In fact, the
aerodynamic boost elevator system functioned well during a
supersonic demonstration dive to a Mach number of 1 ,012.
The hydraulically powered ailerons and rudder also worked
out well, although the sideslip angles developed in the
landing configuration with full rudder deflection produced
higher than expected rolling moments, which required the
use of the ground spoilers for lateral control in the landing
configuration.
The effect of in-flight thrust reverser operation
on the aerodynamic characteristics of the DC-8 was
investigated in a low speed wind tunnel test of a large scale
model. Since model turbine thrust simulators had not yet
been developed, properly scaled exhaust mass flow ratios
were obtained using a unique non-metric ejector system
which fed high pressure air from and outside air source
directly into the inlet of the model flow through nacelles. At
maximum values of reverse thrust, the aerodynamic
characteristics showed some change, but none that
seriously affected the aircraft stability or controllability.
However, asymmetric reverse thrust in the landing
configuration. representing a failure in the thrust reverser
system, resulted in uncontrollable rolling moments with the
flaps down, causing the in-flight thrust-reverser system to
be locked out In this configuration. The flight test program
essentially confirmed the wind tunnel results, and in-flight
thrust reverse became a standard operational procedure.
The DC-6 program initially offered four models, Series
IO and 20 suited for U.S. transcontinental routes, and
Series 30 and 40 for international overwater operations. Low
bypass turbofan engine development led to the higher gross
weight, longer range Series 50. while fuselage stretches,
aerodynamic refinements, and still higher gross weights led
to the Series 60. The final version the Series 70, came about
through the re-engineing of nearly all of the Series 60
models with higher thrust, high bypass ratio turbofans.
Further discussion of the DC-8 aerodynamic development
can be found in Refs.19, 20, and 21. In retrospect, the DC-8
~when introduced was the right concept in terms of payload,
range, speed, and economics It easily adapted to improved
engines, was stretched for greater payload and improved
economics, and was in production for 13 years. These
virtues, plus the outstanding structure, have made the later
models quite popular, with about 300 of the total production
of 556 still in operation.
The DC-9 In the early 1960’s the separate aircraft divisions
of the Douglas Aircraft Company were consolidated at the
Long Beach facility, and design studies were focused on a
short range jet transport to supplement the long range, high
capacity DC-B By mid 1963. the DC-9 configuration had
been defined, Fig. 25The aerodynamic features included a
moderately swept clean wing, with a large chord, long span
double slotted flaps deflected about a simple external hinge
point, rear mounted turbofan engines, and a T-tail. -- -
Fig. 25 The DC-9
Aerodynamic boost flight controls were used for the elevator
and ailerons, although hydraulically operated wing upper
surface spoilers were used to augment lateral control. The
rudder was hydraulically operated with a reversion to
aerodynamic boost in the event of hydraulic system failure.
In the absence of the wing mounted nacelle pylons,
the design and performance prediction methods used on the
DC-8 were carried over with high confidence to the DC-g.
When questioned about the lack of experience with the clean
wing, aft engine, T-tail configuration, the response was.“If
they could do it at Boeing with the 727, and do it at BAC with
(c)l999 American Institute of Aeronautics & Astronautics
the 111, we can do it here at Douglas”. This proved to be the
case, but not without solving some really difficult
configuration-related problems.
The wing sweep of 24 degrees was selected , with an
average thickness ratio of just over 1 I%, providing an Morv
of about 0.80 at high speed cruise CL,The wing was defined
by three Douglas airfoil sections similar to the DC-8,with
slight modifications to the inner panel airfoils to promote an
inboard stall prior to Cu,,ax The trailing edge flap design
was a straightforward application of the experience gained
from previous programs The aft engine nacelle and pylon
design was studied extensively in the wind tunnel, Ref. 22. A
relatively simple short inlet configuration that was developed
showed good presure recovery and only a few counts of
drag rise at cruise conditions. The T-tail arrangement was
the subject of both analytical and wind tunnel studies. No
anomalies surfaced and the design proceeded without
difficulty.
The initial high Reynolds number tests of maximum lift
and stall characteristics showed that the model Chax values
exceeded the predicted full scale numbers at all flap
settings. However, the stall characteristics, judged by an
abrupt drop in CL immediately after the stall, combined wtih
fairly large rolling moment excursions and a slight pitchup
right after Chav were deemed unsatisfactory, and additional
tests of several modifications were made. These
modifications included variations in airfoil nose shapes
across the span, fences on upper and lower wing surfaces at
several spanwise locations, leading edge stall strips, and
several vortex generator configurations.
Part way through the wind tunnel tests to improve the
normal stall characteristics, a serious situation associated
with conditions at very high angles of attack well beyond the
normal stall, called the “deep stall”, was highlighted by the
unfortunate crash of the BAG1 11, a short range jet transport
with aft mounted engines and a T-tail similar to the DC-9 The
“deep stall”, discussed in detail in Ref.22, is defined as a
stalled condition at angles of attack ranging from 25 degrees
to 50 degreesfig. 26. For aircraft designs that employ aft
engines and a T-tail, the nature of the pitching moments
ANU .I
0
-.4
-.b
AND ..6
0 5 10 15 20 25 30 35 40 45 50 A,,c,,o .I Allack (DEG)
Fig. 26 Deep Stall Pitching Moments
beyond the normal stall are such that these typescan enter
this deep stall region. Furthermore,as was the case of the
BAG1 11, there may be a point at 40 degrees to 50 degrees
angle of attack where there is insufficient elevator control
power to pitch the aircraft down to an angle of attack below - the normal stall point. This latter situation is known as a
“locked in” deep stall. Considerable effort was expended in
the wind tunnel to understand the nature and cause of the
deep stall and to develop a configuration that would not be
subject to a “locked in” deep stall situation. In order to meet
the requirement for normal recovery capability at any angle
of attack. the DC-9 horizontal tail was modified to have a
larger span, approximately 20% larger than the original, and
a hydraulic power augmentation system was developed to
provide full down elevator control only under the most
adverse high angle of attack conditions, where the
aerodynamic boost elevator control loses effectiveness.
A comparison of the original and modified horizontal tails is
shown in Fig. 27.
For the normal stall, the wind tunnel program focused
on acheiving a strong nose down pitching moment just
beyond the stall, while maintaining as high a CLmax as
possible. During this testing, a special type of underwing
fencecalled a vortilon for vortex generating pylon, was
developed. This underwing device, Fig 28, does not create
a vortex except at angles of attack very near stall. As the
(c)l999 American Institute of Aeronautics & Astronautics
Fig. 27 Original and Modlfied Horizontal Tail
stall angle is approached, the wing lower surface stagnation
point moves aft of the intersection of the leading edge of the
vortilon and the wing, and the interference of the vortilon with
the leading edge crossflow creates a strong vortex that goes
over the top surface of the wing. This vortex appeared to
have two beneficial effects without impacting C~ex. First, it
energizes the wing spanwise boundary layer flow, reducing
the detrimental effect of this boundary layer on the outer
panel section maximum lift coefficients Second, the vortilon
creates an upwash field inboard of the vortilon which acts on
the horizontal tail to produce an increment of nose down
pitching moment just prior to and immediately after the stall.
Fig. 29 DC-9 -10 stall Control Configuration
Early flight test data on the spoiler-aileron lateral
control system indicated that pilots had a tendency to
overcontrol and had difficulty in maintainlng bank angle in
gusty landing approaches. Intensive effort to redesign the
spoiler-aileron mixer, lateral control spring forces, and
reduced friction levels resulted in an improved system with
light forces, low friction, and near optimum response.
Before the initial model of the DC-g, the Series 10.
had been certified, a higher capacity, higher gross weight
version, the Series 30 was in development. In order to
maintain the extremely short takeoff and landing distances of
the Series 10 at the higher takeoff and landing weights, full
span leading edge slats, Fig.30, were incorporated on the
Series 30 and all subsequent models.Fortunately, Douglas
Fig. 28 DC-9 Vortilon
In addition to the vortilon. wing leading edge stall strips were
also tested in the wind tunnel, and as expected, improved
the nose down pitching moments at the stall, but with a
significant loss in CLmaK ,During the flight development
program, a shot-l chord leading edge fence at 43%
semispan, combined with the vortilon and a very small, short
span leading edge stall strip, Fig:29, produced good stall
characteristics with only a small adverse effect on Chax.
Flg. 30 DC-980 L.E. Slats
had earlier conducted wind tunnel model tests of full span
leading edge slats and had an solid data base for the Series
30 slat design. In order to provide adequate slat chord ,
airfoil chords were extended by 6% of the original chord, all
forward of the front spar,Ref.23, and new airfoil nose shapes
(c)l999 American Institute of Aeronautics & Astronautics
with improved Motv characteristics were designed. Wind
tunnel and flight test data on the Series 30 showed no
aerodynamic problems, and FAA certification was received
quickly.
Two higher capacity, higher gross weight models, the
Series 40 and Series 50 were developed in the late 1960%
using essentially the same design methods and
incorporating the same features as the earlier models. One
exception was the addition of fuselage nose strakes to the
Series 50. As the DC-9 fuselage was lengthened, the
vortices shed by the forward fuselage at high angles of
attack had an increasingly detrimental effect on the vertical
tail, reducing the directional stability at moderate sideslip
angles. A wind tunnel investigation provided an
understanding of the problem, and a solution was developed
in the form of small, low aspect ratio strakes mounted on the
lower quadrant of the fuselage nose, Fig. 31. The strakes
generate their own vortices which alter the flow at the
vertical tail and eliminate the loss in directional stability at
high anoles ofatta& ;
DC-9 Series 50 Nose Strakes
The DC-9 proved to be a very successful design, with
over 900 aircraft of various models being delivered over a
span of 17 years. A detailed summary of all models is found
in Ref.24
The,DC-10 In late 1965, following an unsuccessful proposal since elegant transonic CFD methods were not yet available.
to the USAF for the C-5 military cargo program, Douglas The predicted performance and customer performance
began design studies on a number of large wide body guarantees were reviewed by Mr Mac himself in December
commercial transports. This activity was focused in early 1967, two months before the program was oficially launched
1966 by the appearance of a general specification for a by an order from American Airlines.
medium range, twin engine widebody jet transport capable of
carrying 250 passengers from Chicago to Los Angeles, but
also capable of operating from New York LaGuardia
tochicago with full passengers. By mid 1967, Douglas had
merged with McDonnell Aircraft, and the design studies had
progressed to a well defined three engine aircraft capable of
the the LaGuardia-Chicago mission, but with
transcontinental US. range
The DC-lo. Fig. 32, used aerodynamic design
features carried over from the DC-8 and DC-9 with a few
exceptions. The engine nacelles for the high bypass ratio
turbofan engines were much larger with respect to the other
components of the airplane, Fig. 33, than their predecesors,
placing a great deal of attention on potential wing-nacelle-
pylon interference for both cruise and maximum lift
conditions. Also the aft center engine installation was
different from anything that Douglas had done previously.
Fig. 32 The DC-10
The aerodynamic boost flight controls, used so successfully
on the DC-6 and DC-g. finally gave way to fully powered,
triple redundant, hydraulically actuated flight controls.
The aerodynamic design procedures and -
performance prediction methods used for the DC-1 0 were
further refinements of those used on the DC-8 and DC-g,
(c)l999 American Institute of Aeronautics & Astronautics
Low Bypass Engine Nacelle
High Bypass Engine Nacelle
Flg. 33 Wing Nacelle Comparison
The wing design was configured for a high speed
cruise Mach number of 0.85 and a long range cruise Mach
number of 0.82. Wing sweep was chosen as 35 degrees,
with an average thickness ratio of about 1 l%, with five
specially designed airfoil sections defining the wing.
Numerous modifications to the airfoils were required to
obtain the desired compressibility drag rise characteristics,
and the amount of aerodynamic twist was increased to
obtain satisfactory longitudinal pitching moment
characteristics beyond buffet onset. The wing nacelle pylon
installation, using cutback pylons, worked out well in the
wind tunnel, and the aft engine installation required very little
development in the wind tunnel to eliminate any premature
drag rise. The leading edge slat configurationwas similar to
that used on the DC-9 series 30, except that there were two
extended slat positions; full extension for the best possible
landing maximum lift coefficient, and an intermediate slat
extension for takeoff, which was selected for the best
possible takeoff climb (UD) over the range of takeoff flap
maximum lift valuesDuring the low speed wind tunnel model
testing, it was apparent that the close proximity of the large
englne nacelle to the wlng leading edge, and the very ugly
geometry of the nacelle, pylon, wing leading edge, and
extended slat juncture was generating a flow disturbance
that caused premature stalling of the wing directly behind the
nacelle. All sorts of fixes were tried in the wind tunnel, but
the only devices that eliminated the premature stalling were
a pair of strakes mounted on the upper portion of the wing
engine nacelles, Fig. 34. These strakes were subjected to
quite high local angles of attack due to the extreme
crossflow around the nacelles at airplane angles of attack
approaching maximum lift. The low aspect ratio planform of
the strake promoted a strong leading edge vortex from each
strake that went directly over the wing just behind the
nacelle.These twin vortices energized the wing airflow
sufficiently to eliminate the premature stalling and allow the
wing high lift system to achieve it’s full maximum lift
capability. The configuration developed in the wind tunnel
served as the starting point for the flight development
program to achieve the lowest possible stall speeds with
acceptable stall characteristics. Several flights were made
with various sizes of strakes. and with slight variations of
strake orientation, very quickly leading to the production
configuration. In later years, we at Douglas were pleased to
note that nacelle strakes became a standard feature of
Boeing and Airbus transports.
4. 34 DC-10 Nacelle Strakes
The only other aerodynamic item that required
attention during the flight development program was the
excessive amount of stall warning In the clean (flaps up.
(c)l999 American Institute of Aeronautics & Astronautics
slats retracted) configuration. The stall itself was acceptable
with a good pitch down and minimum roll off, but the initial
Stall warning occurred at around 12% above the stall speed,
The relatively simple fix involved automatic deployment of
the outboard leading edge slats at a fixed angle of attack,
which eliminated the separated flow causing the early stall
warning. Stall speeds were not affected, since the stall was
determined by the inboard portion of the wing as before.
The remainder of the flight test program proceeded
smoothly, and the now designated DC-1 O-1 0 met or
exceeded all of the predicted performance.
Early in the DC-10 program, higher gross weight,
longer range versions, the DC-10-20, DC-10-30. and DC-IO-
40 were developed. These models had the same
aerodynamic features as the DC-l O-l 0 except for the wing,
which was redefined to increase the overall span by 10 feet.
This change was made primarily to reduce the induced drag
in the critical one engine inoperative takeoff climb
configuration, although it did improve the criuse (UD) as
well. Modifications to the wing and aft engine inlet lines were
made during the detailed design of the Series 20, (later
called Series 40) to accommodate the engine selected for
these models.
Of the 386 commercial DC-l O’s built, nearly 300 are
still in active service, with many passenger models currently
being converted to cargo versions.
THE MD-80 and MD-90 In the mid 1970’s the Douglas
Aircraft division of McDonnell Douglas undertook a series of
studies aimed at extending the DC-9 product line beyond the
Series 50. The focus was on an aircraft with increased
passenger capacity, reduced seat mile operating costs,
maximum commonality with previous DC-9 models, while
maintaining or increasing range performance, and reducing
airport noise signatures. In order to meet these objectives.
the DC-9 fuselage was lengthened (again), the wing area
was increased by about 20%, Fig. 35. to accommodate the
higher gross weights required. and a higher bypass version
of the standard JTBD engine was employed, providing higher
thrust, lower airport noise levels, and improved cruise
24dN. (61 cm,
Fig. 35 MD-NJ Root Plug
specific fuel consumption. The design configuration was
finally set in mid 1977 and was designated as K-9-80, an
extension of the DC-9 product line. Eventually. in order to
acknowledge it’s corporate identity, it was redesignated as
MD-80.
The basic aerodynamics of the MD-80 are those of
the DC-g, with a few exceptions, the most obvious being the
wing modification. The additional area was obtained by the
use of a wing root “plug” which extended the DC-9 wing
lines into the side of the fuselage, and incorporated an
inboard trailing edge extension fitted with a constant chord
flap segment. A new inboard defining airfoil was designed,
using refinements of the methods used on the original DC-9
wing development, to maintain the aerodynamic sweep
(isobars) inboard. The aerodynamic design and performance
prediction methods were those used successfully on the
derivative models of the DC-8. DC-g. and DC-lo. The wind
tunnel test program for the MD-80 showed no unexpected
items, although the deep stall recovery margin was impacted
by the larger engine nacelles of the MD-80. The addition of
strakes to the lower outboard quadrant of each nacelle,
produced a reduction in the blanketed area of the
horizontal tail at extreme angles of attack, and restored the
deep stall recovery margin to an appropriate level. Another
new aerodynamic design feature for the MD-80 was
incorporation of the DC-1 0 slat system concept of having two
extended flap positions, full deflection for landing to achieve
(c)l999 American Institute of Aeronautics & Astronautics
minimum stall speeds, and partial deflection for takeoff to
optimize takeoff climb (UD)s. During the flight development
program, the stall characteristics with the takeoff slat
configuration were somewhat marginal, so an “autoslat”
extension mode was added. In the takeoff slat configuration,
when a predetermined aircraft angle of attack is reached, the
slat is automatically repositioned to the landing position,
providing the desired stall characteristics while maintaining
the desired (UD)s at the takeoff climb conditions. The only
other noticeable aerodynamic design difference came later
in the MD-60 production program, when a CFD developed
shape for the aft fuselage tailcone closure was substituted
for the original DC-9 design as pan of a drag cleanup
program.
The MD-90 (Ref. 25), first ordered in 1989. is
essentially a slightly stretched MD-60 with a modern high
bypass engine with greater thrust installed. Aside from the
obvious nacelle changes, the aerodynamic design-features
are the same as those of the MD-60, with two minor
exceptions. The aerodynamically boosted elevator was
replaced with a hydraulically powered elevator to maintain
pitch control responsiveness in the landing configuration.
Increased airframe and engine weight in the rear of the
aircraft and the slightly longer forward fuselage increased
the pitch inertia to a point where the rapid response of the
powered system was required. Deep stall recovery was
again impacted by the larger engine nacelles, but this time
the solution lay in the incorporation of a hydraulically
operated flap on the pylon trailing edge. The flap is operated
when the control column is moved to a predetermined
position near full forward travel.
The MD-80, and to a lesser extent the MD-90. has
enjoyed considerable success in the short to medium range
market segment, with over 1200 in operation at this time.
The MD-1 1 In the late 1980’s the Douglas Aircraft division
studies of a longer range, higher capacity, derivative of the
DC-10 led to the definition of the MD-l I. The major
aerodynamic deslgn features of the MD-1 1, Fig. 36. were
carried over from the DC-lo. New features included wlnglets
to improve cruise (L/D). a smaller horizontal tail made
Fig. 36 The MD-11
possible by the Incorporation of more sophisticated flight
control system with relaxed static stability in pitch, and the
modification of the outboard wing airfoil sections In the trailing
edge area to approximate “Divergent Traillng Edge’ alrfoil
performance.While not an aerodynamic design feature per
se, the introduction of a fuel tank in the horizontal tail to
maintain the aircraft center of gravity near the most aft llmit
for cruise also contributed to improved cruise efficiency.
Although seemingly mlnor modifications. these changes
resulted in an improvement in cruise (UD) of nearly 9%. In
spite of this improvement, the predicted cruise drag level was
not achieved, and this situation, combined with a higher than
expected engine sfc in cruise, resulted in a reduction in
range. Later increases in fuel capacity and takeoff gross
welght combined with further drag cleanup, essentially
brought the range capability back to the original level. As part
of a drag reduction program, a short splitter was added to the
wing trailing edge in the area of the modified airfoils.
In retrospect, the applied aerodynamics effort at
Douglas during the jet transport era was characterized by a
continuing drive to improve the design and prediction
methods associated with operation at high subsonic Mach
numbers, and at low speed , high lift conditions. In the mid
1950’s, transonic wind tunnel data provided much of the
information (often times wrong) that aerodynamic theory
could not provide.In the 60’s and 70’s. more empirical
methods, combined with more sophisticated wind tunnel test
techniques, sufficed. Finally In the late 70’S the rapid growth
of digital electronic computers and the development of a
number of more accurate CFD codes made life a blt more
routine for the aerodynamics design engineers. Ironically, at
Douglas, modern computational methods were never
(c)l999 American Institute of Aeronautics & Astronautics
utlllzed In the design of an entirely new commercial jet transport that reached production status, although serious preliminary design studies using modern methods were conducted and documented by wind tunnel tests in a cooperative effort with NASA, Refs. 26 and 27.
Concludina Comments The long history of applied aerodynamics at Douglas Aircraft , spanning some 66 years, was highlighted by the application of the best aeodynamic theory avallable,supplemented by specific wind tunnel and
flight test data, to develop reasonable aerodynamic design and performance prediction methods. This approach was a key element in the design of a long line of transport aircraft that made a significant contribution to the progress of commercial aviation. Mention should be made of the individuals that inspired the work of the Douglas transport aero group; Bailey Oswald, George Worley, Richard Shevell, Orville Dunn, Harold Luskin, Harold Kleckner were all leaders in the search for ways to “do it better” than the last time by understanding advances in the state of the art and applying them to the aerodynamic design and prediction methods. They were also the stimulus for galning understanding of aerodynamic problems which showed up in wind tunnel model testing, and for developing the configuration changes necessary to eliminate the problems. It should also be noted that in addition to the applied aerodynamics work on specific designs, significant contributions to the literature of aerodynamic theory were made by members of the Douglas Aerodynamics Research Group such as A.M.O. Smith, Tuncer Cebecl, John Hess, Zig Bleviss. Ellis Lapin, Ed Rutowski, Ernie Graham, Martha Graham, Beverly Beane,
Mllton Van Dyke. Joe Gieslng, and Preston Henne. just to mention a few. For those of us who lived through it, it was a marvelous experience, a great sense of accomplishment, and we were happy to have been a part of It.
References
I. Jacobs, Eastman N., Ward, Kenneth E., and
Pinkerton, Robert M., “The Characteristics of
78 Related Airfoil Sections from Tests in
the Variable Density Wind Tunnel”, NACA
Rept. 460, 1932.
2. Weick, Fred E., ” Drag and Cooling with Various
Forms of Cowling for a “Whirlwind” Radial Air-
Cooled Engine”, Part I and Part II, NACA
Rept. 313 and 314, 1929.
3. Jacobs, Eastman N., and Pinkerton, Robert M.,
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