applied aerodynamics at the douglas aircraft company

(c)l999 American Institute of Aeronautics & Astronautics

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

FOP permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 1801 Alexander Bell Drive, Suite 500, Reston, VA 20191


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


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


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


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


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.


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,

-

-


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


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


,_ - . ” . “~ 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


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


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


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.


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


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


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


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,


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.


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


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


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.,

“Tests in the Variable Density Wind Tunnel of

Related Airfoils Having the Maximum Camber

Unusually Far Forward”, NACA Rept. 537,

1935.

4. Wenzinger, Carl J, and Harris, Thomas A., “Wind

Tunnel Investigation of an NACA 23012 Airfoil

With Various Arrangements of Slotted Flaps”

NACA Rept. 664, 1939.

5. Harris, Thomas A.. “Reduction of Hinge Moments

of Airplane Control Surfaces by Tabs”, NACA

Rept. 528, 1935.

6. Goett, Harry, and Reeder, John P., “Effect of

Elevator Nose Shape, Gap, Balance, and Tabs

on the Aerodynamic Characteristics of a

Horizontal Tail Surface”, NACA Rept. 675.

1939.

7. Ames, Milton 6.. and Sears, Richard I.,

“Determination of Control Surface

Characteristics from NACA Plain Flap and Tab

Data”. NACA Rept. 721, 1941.

8. Gilruth. Robert R., and White, Maurice D..

“Analysis and Prediction of Longitudinal

Stability of Airplanes”, NACA Rept. 711, 1941.

9. Dunn, Orville R. “Aerodynamically Boosted

Surface Controls and their Application to the

DC-6 Transport”,lAS-RAeS International

Conference on Aerodynamics, 1949

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applied aerodynamics at the douglas aircraft company

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