northruptheory

35th WILBUR WRIGHT MEMORIAL LECTURE

The Development of All-Wing Aircraft

John K. Northrop

The thirty-fifth Wilbur Wright Memorial Lecture was delivered before the Society by Mr. John K. Northrop
on Thursday, May 29, 1947 at 6 p.m. in the Lecture Hall of the Institution of Civil Engineers, Great George
Street, London. The chair was taken by Sir Frederick Handley Page, C.B.E., President of the Society.

The President: They had now reached the highlight of the 1946-47 session, the Wilbur Wright Memorial
Lecture. It seemed incredible to think that these memorial lectures had been going on for 35 years, no sooner
was one over than another one came round. The Lecture was usually given in alternate years by an Englishman
and an American, and this year they were fortunate in having as their lecturer that distinguished American, Mr.
John K. Northrop.

Before introducing Mr. Northrop he had another duty to perform. As was customary on the occasion of the
Wilbur Wright Memorial Lecture, as President, he had cabled Orville Wright as follows:--

"On May 29th 1947 your distinguished countryman John K. Northrop will read the 35th Wilbur
Wright Memorial Lecture on The Development of All-Wing Aircraft. The reading of this annual lecture
is a constant reminder of those early years of this century when you and your brother laid down so
clearly, yet so simply, the firm foundations of the art of flying upon which must ultimately be built that
structure for world peace which can never be destroyed." He had received the following reply:--

Heartiest greetings to the Society and to all assembled to hear my esteemed fellow countryman
Northrop deliver the 35th Wilbur Wright Lecture. I believe Mr. Northrop will bring to you a good
message on matters useful in peace.
ORVILLE WRIGHT.

He had also sent the following cablegram to the Institute of the Aeronautical Sciences:

On behalf of the Council and Members of the Royal Aeronautical Society I send you our greetings on
the occasion of the reading of the 35th Wilbur Wright Memorial Lecture by your distinguished member
John K. Northrop on The Development of All Wing Aircraft.

They had cabled in reply:--

On the occasion of the reading of the Wilbur Wright Memorial Lecture by our distinguished American
contemporary John Northrop the Officers and Council of the Institute of the Aeronautical Sciences
extend heartiest greetings to the Royal Aeronautical Society. We are all looking forward to a period of
greater collaboration between the Societies on either side of the Atlantic.
R R. BASSETT, President

He now had great pleasure in introducing Mr. John K. Northrop. Mr. Northrop was well-known on both
sides of the Atlantic for his work on all wing aircraft. Indeed, he was one of the great pioneers in that field and
probably knew more about this particular type of airplane than anyone. He was chief designer, President and
everything else--in fact, he was the Northrop Aviation Company. He had been designing and developing the
all-wing type since about 1923. This evening they were to have the pleasure and privilege of hearing from Mr.
Northrop something of the difficulties and successes connected with that development.

He had pleasure in calling on Mr. Northrop to deliver his lecture.

One cannot undertake the presentation of one of the long series of Wilbur Wright Memorial Lectures without
a deep sense of appreciation of the tremendous contributions made by the illustrious group of scientists and
engineers who have given such great distinction to this event. The happy precedent of inviting individuals from
without the United Kingdom to make this presentation in alternate years has gone far in the past toward
improving the understanding, cooperative effort and fellowship of the English-speaking peoples, and I am
deeply honored to have been among those chosen to further this very worthy cause.

INTRODUCTION
In choosing the title, "The Development of All-Wing Aircraft," as the subject of my lecture I run some risk of
being accused of writing a company history rather than a paper of the broad scope ordinarily presented before
this time-honored institution. This is far from my intent, but being sincerely convinced that the all wing airplane
is a valuable step in the development of aeronautics, and desiring to contribute a maximum amount to the
available data in the limited time at my disposal, my paper must be confined largely to experience gained by
our company in its work on this subject.

Outside the efforts of the Horten Brothers in Germany there has, until a comparatively recent time, been little
physical accomplishment in the development of the all-wing airplane except by our company. The
contemporary Horten development has been fully described in technical reports emanating from Germany
since the close of the European war. In many instances the Horten conclusions were surprisingly similar to our
own. Their work was not carried so far, however, and I doubt that they had the sympathetic and responsible
governmental backing and the resultant opportunities for development accorded us.

In considering the development of all-wing aircraft I would like first to distinguish between all-wing and tailless
airplanes. Most tailless airplanes are not all-wing by our definition. There is a tremendous background of
development in tailless types, which has been fully reported by Mr. A. R. Weyl in Aircraft Engineering. These
articles outlined a surprising number of reasons for building tailless aircraft which have motivated the various
designers and constructors over the years. Only one of the many advantages to be gained through such
development has inspired our work, namely improved efficiency of the airplane.

More recently, through the rapid development of turbojet power plants, a second advantage has arisen,
which is the elimination of design difficulties attendant upon the impinging of high speed high-temperature jets
on tail surfaces. Still more recently a third possible advantage has appeared, this being the (as yet unproved)
probability that problems of stability in the transonic and supersonic ranges may be somewhat more simple of
solution in the tailless type than in the older and more conventional arrangements.

Only the first of these basic advantages, namely that of improved efficiency, has been readily apparent over a
number of years and, as a result, virtually all our efforts have been directed toward the reduction of parasite
drag and the improvement of the ratio of the maximum trimmed lift coefficient (Clmax) to the minimum drag
coefficient (CDmin). It is natural, then that we were not interested particularly in tailless airplanes as such; if we
could not eliminate vertical tail surfaces, fuselages, and a substantial portion of interference drag, the gains to
be made seemed not worth the effort necessary for their accomplishment.

Our work, therefore, through the years has been directed solely to all-wing aircraft, by which I mean a type
of airplane in which all of the functions of a satisfactory flying machine are disposed and accommodated within
the outline of the aerofoil itself. Of course, we have not as yet built any pure all-wing aircraft. All have had
some excrescences, such as propellers, propeller drive shaft housings, jet nozzles, gun turrets and the like. We
have, however, built a number of airplanes in which the minimum parasite drag coefficient has been reduced to
approximately half that ordinarily attained in the best conventional aircraft of like size and purpose, and in
some of the designs completed and tested the excrescences and variations from the aerofoil contour have
been responsible for less than 20 percent of the minimum airplane drag.

BASIC ASSUMPTIONS
A surprisingly large number of people, both within and without the aircraft industry, still appear to question the
economic reasons for going to all the trouble to build an all-wing airplane. "Sure," they say, "after a lot of
practice people can learn to walk on their hands, but it's most uncomfortable and unnatural, so why do it when
nothing is gained thereby?" Actually, there are startling gains to be made in the aerodynamic and structural
efficiency of an all-wing type, provided that certain basic requirements can be fulfilled by the type under
question. These requirements can be simply stated as follows:

First, the airplane must be large enough so that the all-wing principle can be fully utilized. This is a matter
closely related to the density of the elements comprising the weight empty and the useful load to be carried
within the wing.

The dimensions of the average human body may also at times be the limiting factor but, ordinarily, in the larger
types of transport or bombardment aircraft in which we are most interested, it will be found that excessive
sizes are not necessary in order to secure, within a wing of reasonable thickness ratio, adequate volume for a
commercial cargo or bomb load plus the necessary fuel.

The extremes explored and satisfactorily flown to date in our experience range from a "buzz" bomb having a
span of 29 feet, in which the warhead was cast as a portion of the aerofoil to the 172-foot XB-35 long-range
bomber airplane. The buzz bomb was practical because of the comparatively high specific gravity of the
warhead, plus the fact that the configuration permitted almost all of the wing to be used as a fuel tank.

The XB-35, on the other hand, is considerably larger than would be necessary to provide ample space for
passenger and crew comfort and ample volume for payload, be it cargo or bombs. It was designed larger than
necessary because we desired to keep the wing loading comparatively low in this first large experimental
venture. It has a normal gross weight of 165,000 lb., an overload gross weight of 221,300 lb., and sufficient
volume within the wing envelope so that the maximum gross weight at takeoff might well be increased to over
300,000 lb., somewhat over half of which could be devoted to bombs, fuel and miscellaneous payload. It may
be seen, therefore, that there is a practical range of size within which the all-wing airplane can be used. If the
requirements of space and volume do not permit the full use of the all-wing principle, a rudimentary nacelle
may be added without losing its economic advantages.

The second basic requirement is that the all-wing airplane be designed to have sufficient stability and
controllability for practical operation as a military or commercial airplane. We believe this requirement has
been fully met by hundreds of flights completed with this type, and we are fully convinced of its practicability
after having built a dozen different airplanes embodying scores of different configurations incorporating the
all-wing principle.

In comparing all-wing and conventional types, we may fairly assume that spans of comparative aircraft having
the same gross weight are equal, and as a further simplification we may for the moment neglect compressibility
effects in our comparison to the advantages of all-wing and conventional types of large bombardment or
transport aircraft having maximum velocities up to approximately 500 m.p.h.

COMPARISON OF MINIMUM DRAG AND MAXIMUM TRIMMED LIFT
Based on these assumptions and on the following proved data on the all-wing type, a comparatively simple
analysis of advantages may be made.

The ratio of the minimum parasite drag coefficient (CDmin) for all-wing airplanes to that for conventional
types is approximately 1:2. Minimum drag coefficients for a number of large bomber and transport aircraft
such as the B-29, B-24, C~4 and others average approximately .023. The minimum drag coefficients for
several all-wing types have been measured both in model and full-scale configurations and vary from less than
.010 to about .0113, which is the figure for the XB-35 including armament protuberances, drive shaft
housings, rudimentary nacelle for gun emplacements, and so on.

The ratio of maximum trimmed lift coefficient (Clmax) for all-wing to conventional types is approximately
1.5:2.3. The latter figure is typical for a number of the large airplanes of conventional arrangement previously
mentioned. The former is readily attainable in a configuration such as that of the XB-35 and may be subject to
considerable improvement through the use of several types of high lift devices yet to be proved.

For comparative airplanes of the same span and gross weight the selection of the required wing area will
depend either on flight conditions, including takeoff without flaps, or landing conditions. If the flight conditions
govern, the ratio of required wing areas of all-wing to conventional aircraft will be 1:1 because the two wings
are equally effective except under conditions of maximum lift. If landing conditions govern, the ratio will be 21
3:1, assuming the same landing speed in each case. If takeoff with partial flap deflection governs, the ratio will
be somewhere between the above two figures.

In large all-wing bombers and transports, and a growing extent in conventional long-range transports as well,
the ratio of gross weight at takeoff to landing weight will approach 2:1. Therefore flight conditions are likely to
govern the selection of wing area more than landing conditions. In the following calculations both extremes are
used as indicative of the range of advantage to be gained by the use of the all-wing configuration. Referring to
Fig. 1, it may be seen from equation (1) that the total minimum parasite drag of the all-wing airplane in terms
of the conventional airplane will vary from 50 percent if flight conditions govern, to 77 percent if landing
conditions govern. In this equation (Dp)a and (Dp)c represent the parasite drags of all-wing and conventional
airplanes while Sa and Sc represent the respective wing areas.

It is a well-known fact, based on the Breguet range formula, that with conventional reciprocating engines and
propellers the speed for maximum range is approximately that at which parasite drag and induced drag are
equal. Therefore, at the same cruising speed as the conventional airplane the all-wing type will require from 25
percent to 11 percent less power, as shown in equation (2), and with the same amount of fuel will fly from 33
percent to 13 percent farther, as indicated by equation (3). In these equations P represents power required,
and D total drag. V is airplane velocity and R range, with the suffices a and c again denoting the all-wing and
conventional configurations. If the all-wing airplane is operated at its most economical speed, instead of the
most economical speed of the conventional airplane, it will fly 19 percent to 7 percent faster and the range will
be from 41 percent to 14 percent greater with the same amount of fuel as indicated in equation (4) of Figure
2.

ADVANTAGES OF LOW PARASITE DRAG
Under high-speed conditions with any type of power plant the parasite drag becomes a much larger
percentage of the total drag than for cruising conditions with reciprocating engines. At high speed the parasite
drag may account for 80 percent or more of the total, while the induced drag drops to 20 percent or less.
Using an assumed figure of 80 percent parasite drag, which is probably correct to + 10 percent for most
aircraft, the power required to drive the all-wing airplane at the same speed as the conventional airplane will
be from 40 percent to 18-1/2 percent less, as shown in equation (5), and the range, at the high speed of the
conventional airplane, will be from 66 percent to 22 percent greater, as indicated in equation (6). As turbojet
and turboprop power plants both operate at relatively high speed for best fuel economy, the advantages of the
all-wing configuration, when used in combination with these power plants, will closely approach the above
figures for maximum range as well as high speed.

These advantages are all based on the simple aerodynamic values obtained with all-wing airplanes; namely,
that; CDmin equals 50 percent of conventional CLmax equals 65 percent of conventional. The probabilities
are that the minimum parasite drag can, within a comparatively short time, be reduced, at least for commercial
types, to about 40 percent of the conventional figure and that the maximum trimmed lift coefficient (CLmax)
may, within a similar short time, be increased to at least 75 percent of conventional values.

METHODS FOR INCREASING MAXIMUM TRIMMED LIFT
One of the most interesting devices for increasing maximum lift is, of course, the judicious use of boundary
layer control in conjunction with turbojets or gas turbines. Another involves the development of a better
combination of low pitching moment flaps and trimming devices which will permit of "lifting ourselves by our
boot straps" in a more successful manner than we have achieved to date. Model configurations tested up to
this time, employing such methods, have shown improvements of .1 or .2 CL over the figure now used of 1.5.

A third possibility of rather unconventional nature remains to be proved in the all-wing airplane. This consists
of placing the C. G. behind the aerodynamic center of the wing, eliminating inherent longitudinal stability by so
doing and replacing this characteristic, which heretofore we have always considered as an essential to
satisfactory aircraft design by highly reliable (and perhaps duplicate) automatic pilots which take over the
function of stability from the airframe and may perhaps do a better job of maintaining the proper attitude than
the present classical method. While unconventional and possibly a bit horrifying to those unaccustomed to the
idea, it may have practical application to very large aircraft where the pilot's skill and strength are largely
supplanted by mechanical means of one sort or another, and wherein the pilot controls the mechanism which in
turn places the airplane at the desired attitude. If the C.G. is located aft of the aerodynamic center the airplane
will trim at a high angle of attack with the flaps or elevator surfaces deflected downward rather than upward
from their normal position, thereby increasing the camber and rendering the whole aerofoil surface a high-lift
device. It is possible that trimmed lift coefficients in the order of 2.0 may be achieved by this method, and
experiments completed to date with such a device on conventional aircraft show that the C.G. may be
displaced at least 10 percent of the mean aerodynamic chord aft of a normal position without any
uncomfortable results in the flying characteristics of the airplane.

When these improvements in CLmax and CDmin can be realized, further startling gains in performance will
accrue, as will be outlined later. It would seem, however, that the present accomplishments offer sufficient
incentive to warrant all they have cost in time, effort and money, and that the question, "Why bother with an
all-wing airplane?" is already well-answered.

OTHER MAJOR ADVANTAGES
There are other major advantages of the all-wing type which cannot be so definitely evaluated but which can
and do contribute appreciably to improvement in efficiency and range. Two of these, namely the elimination of
jet-tail surface interference, and the possible elimination of wing-tail surface shock wave interference, have
already been mentioned. The third, and the most immediately applicable to designs of the near future, is the
improved adaptability of all-wing types to the distribution of major items of weight empty and useful load over
the span of the wing. While such distribution can be made to a limited extent in conventional airplanes, it can
be much more fully accomplished in the all-wing type. Such weight distribution results in substantial savings in
structural weight which have important effects on the ratio of gross weight at takeoff to landing weight. An
analysis of the range formula indicates that this ratio is one of the most important range parameters. Competent
authority has shown that distribution of fuel in the wings instead of the fuselage of a large conventional modern
transport would allow an increase in gross weight of 16 percent without increase to weight empty, with a
corresponding increase in range up to 30 percent.

It is fairly obvious that the all-wing airplane provides comparative structural simplicity, plus the possibility of
structural material distribution in a most effective way at maximum distances from the neutral axis, plus an
opportunity to stow power plant, fuel and payload at desirable intervals along the span of the wing, which
cannot be equaled in conventional types. These matters are rather intangible and difficult to illustrate by
numerical relationships. They depend to a large extent on the type and size of the airplane, what it is designed
to carry, and what the desired high speed may be.

PROBLEMS INVOLVED IN ALL-WING DESIGN
Having demonstrated, perhaps, that the advantages of the all-wing type are fully worth striving for, let us
consider the problems involved and their solution. Based on our present experience these difficulties do not
appear now of surpassing magnitude, but in 1939 several of them seemed so serious as to discourage the
most hardy optimist.

To one testing a swept-back aerofoil having a desirable root thickness, taper ratio and symmetrical section,
together with reasonable washout at the tips such as might be designed from the then available data, the first
results were a bit terrifying. The elevator effect was erratic, changed in sign with varying deflections, and was
entirely unsuitable for the control of an airplane. It was also seen that the degree of static longitudinal stability
indicated by the average slope of the pitching moment curves was less than that considered desirable in a
conventional airplane. Experiments involving visual observation of tufts on the model indicated a separation
along the training edge of the aerofoil which was apparently due to the planform configuration, and which was
responsible for the erratic curves. In early experiments a simple addition of 10 percent to the chord length with
a straight line contour from approximately the 70 percent chord point to the new 110 percent chord point,
almost completely eliminated the difficulty.

FIRST FULL-SCALE AIRPLANE
It was soon determined that date applicable to conventional wings with little or no sweep were completely
unreliable for the degree of sweepback required in practical all-wing designs, and that a whole new technique
had to be developed to determine the limits within which taper ratio, sweepback and thickness ratio could be
combined for satisfactory results. All these variables were explored in a series of wind tunnel models, and
when a reasonably satisfactory group of configurations had been determined it was decided to build our first
piloted flying wing, the N-1M (Northrop Model 1 Mockup).

Because of the many erratic answers and unpredictable flow patterns which seemed to be associated with the
use of sweepback, it was decided to try to explore most of these variables full scale, and the N-1M provided
for changes in planform, sweepback, dihedral, tip configuration, C.G. location, and control surface
arrangement. Most of these adjustments were made on the ground between flights; some, such as C.G.
location, were undertaken by the shift of ballast during flight. The variations to which this first airplane was
subjected involved two extremes of arrangement in which the airplane was found to be quite satisfactory in
flight.

It is an interesting commentary on the comparative ease with which the basic problems of controlled flight
were solved to note that no serious difficulties were experienced in any flight attempt, or with any of the
various configurations used. Some "felt" better to the pilot than others, but at no time was the airplane
uncontrollable or unduly difficult to fly. The principal early troubles were related to the cooling of the small
""pancake" - type air-cooled engines which were buried completely within the wing, and because of the
pusher arrangement did not have the benefit of slipstream cooling in taxiing, takeoff and climb. Engine-cooling
problems seriously handicapped the early flights but later, somewhat larger engines were installed and the
design of the cooling baffles was sufficiently improved so that repetitive sustained flights were accomplished
easily.

The first flight was more or less an accident in that, while taxiing at comparatively high speed over the normally
smooth surface of the dry desert lake bed used as a testing field, the pilot struck an uneven spot. He was
bounced into the air and made a good controlled flight of several hundred yards before returning to earth.
Altogether, this first airplane was used in over 200 flights of substantial duration, during which numerous
configurations were tested and a great deal of work was done in the determination of the best types of control
surface and surface control mechanism.

ELEVONS AND RUDDERS
From the inception of the work, longitudinal and lateral controls were combined in the "eleven," which word
was coined to designate the trailing edge control surface members which operate together for pitch control
and differentially for roll control. At no time during early tests did control about the pitch or roll axes give any
appreciable difficulty. The control which was least expected to cause difficulty gave the most, namely the
rudder.

Early in the test program it was found that the airplane had quite satisfactory two-control characteristics that
is, a normal turn resulted from a normal bank without the use of rudder controls and as a result, throughout the
program we have often considered the elimination of rudder controls entirely. It was indeed fortunate that the
first airplane developed such docile characteristics, for many of the rudder configurations tried proved to be
ineffective -- or worse, affected the flight characteristics of the airplane adversely.

From the start it was determined to eliminate, to the greatest extent possible, vertical fin and rudder surfaces;
first, because they violated the all-wing principle and added drag to the basic airfoil; second, because with the
moderate sweepback employed in our early designs the moment arm of a conventional rudder about the C.G.
was small, and an excessively large vertical surface would have resulted had we tried to achieve conventional
yaw control moments. The rudder development was therefore concentrated on finding a type of
drag-producing device at the wing tips which would give adequate yawing forces without affecting pitch or
roll. To this end we tried 25 or 30 different configurations in flight which were first tested in the wind tunnel.
As a result of this experience it was concluded that dynamic reactions were likely to be very different from
static reactions; some of the configurations which looked best in the wind tunnel proved to be quite
unsatisfactory in flight.

The best and most practical rudder found was one of the simplest in concept and one of the first to be flown,
namely a plain split flap at the wing tip which could be opened to produce the desired drag. This flap was later
combined with the trimming surface needed to counteract the diving moment of the landing flaps, forming the
movable control surfaces at the wing tip of the XB-35.

Among the many flights accomplished with the first experimental airplane were several in tow of other aircraft
where the distance to be covered, or the altitude to be gained, made it impractical to depend solely on the
airplane's own engines. After a few minutes of acquaintanceship with the slight differences brought about by
the presence of the tow cable, the airplane behaved well in tow and several comparatively high altitude flights
were made to investigate the spin characteristics. These appeared to be quite normal, based on preliminary
tests of this airplane. Later experience, however, indicated that the spin characteristics of tailless types vary
from one design to another, in the same fashion as may be expected in conventional types, and that no broad
generalization as to spin behavior can be made with safety.

N-9M FLYING MOCKUP FOR BOMBER
The N-1M was first flown in July 1940 and for about a year was consumed in a combination of aerodynamic
tests and attempts to solve engine cooling problems. As soon as good sustained flight demonstrations could be
made on schedule the Army Air Forces took active interest in the program and top-flight officers, including
General H. H. Arnold and Major General Oliver P. Echols, encouraged us to investigate the application of the
all-wing principle to large bomber aircraft. To this end it was decided to construct four scale models of a
larger airplane. These were designated N-9M (Northrop Model 9 Mockup) and they duplicated, except for
the power plant and propeller arrangement, the aerodynamic configuration of the proposed XB-35 airplane.

The first of these aircraft was completed and test flown on December 27, 1942, and had completed about 30
hours of test flying with pilot (and sometimes an observer) when it crashed, killing the pilot. The machine had
been on a routine test flight across the desert away from its base, and was out of sight of technically qualified
observers at the time of the accident. However, all evidence pointed to a spin, and the attitude of the airplane
on the ground indisputably indicated autorotation at the time of impact.

This loss was a serious setback and work was started immediately to recheck the spin characteristics of the
airplane in a spin tunnel. It was later determined, both in the tunnel and in flight, that recovery was good,
although a bit unconventional (requiring aileron rather than elevator action), but that the spin parachutes which
had been attached to the airplane for the low-speed stalling and stability tests then in progress were ineffective
as to size and improperly located.

SPINNING AND TUMBLING CHARACTERISTICS
Subsequent models, over hundreds of flights, gave no trouble. The low-speed stall and spin tests with rear
C.G. positions were accomplished without further difficulty and the N-9M proved an invaluable test bed in
which various control configurations could be proved in detail. A large number of additional rudder
configurations were developed and tested on the N-9Ms; likewise different types of mechanical and
aerodynamic boost for the control surfaces were investigated, as well as the general behavior of the airplane in
all types of air, and with different C.G. positions.

In connection with the model spin tests of this airplane, an investigation of the tumbling characteristics of the
type was made in the spin tunnel. These tests showed that if the model was catapulted into the airstream with
an imposed high velocity about the pitch axis in either direction, it would continue to tumble or come out of the
maneuver, depending on comparatively minor differences in eleven and C.G. position. In other words, under
circumstances of induced rotation about the pitch axis the recovery was marginal. However, it would never
tumble from any normal flight condition, such as a stall, spin, or any other to-be-expected maneuver. In some
configurations, if dropped vertically trailing edge down into the wind stream, a tumbling action would be
induced which might or might not damp out. This was not judged a serious matter in view of the fact that a
vertical tail slide is hardly a maneuver to be courted, even by a fighter airplane, let alone a 100-ton bomber.

The three remaining N-9Ms have been flown almost continuously since their completion dates to the present.
Only recently have all desirable test programs been completed and the airplanes relegated to a semi-retired
status from which they are withdrawn only for the benefit of curious pilots.

XP-79, ROCKET-POWERED AIRPLANE
In September 1942 we conceived the idea of combining the newly developed liquid-rocker motors with a
flying wing in a high speed and highly maneuverable fighter. The physical dimensions of the human frame
immediately became a limiting size factor and for this reason, as well as because much higher accelerations can
be withstood for longer periods in the prone position, it was decided to place the pilot prone in this design.
Three experimental, full-size glider versions of this little airplane were rapidly completed and a long series of
glider tests undertaken. In order to achieve the utmost in low drag and light weight, the original airplanes were
mounted on skids and the first glider tests were attempted with an automobile tow. Because of the rugged
construction of the gliders they had a fairly heavy wing loading and the equipment provided for towing proved
to be incapable of achieving enough speed for takeoff.

As a second expedient, detachable dollies were built from which the airplane was expected to take off at
flight speeds. Minor crack-ups occurred with this configuration and it was finally decided to compromise the
aerodynamic cleanness of these first test airplanes in order to provide a rugged permanent and dependable
landing gear for experimental purposes. The unusually large fin used here was required to stabilize the fixed
landing gear, a substantial portion of which extended ahead of the C.G. After this gear was installed, and with
another airplane as the towing medium, the takeoff difficulties were eliminated and a number of successful
glider flights were made.

These airplanes were flown both with and without wingtip slots and slats which were tested for the purpose of
eliminating tip-stall difficulties, as will be described later. They were also flown with a wide variation in vertical
fin area, to determine the amount necessary or desirable for various flight conditions.

In one memorable test during which the airplane was equipped with a fixed slat, a rather peculiar accident
occurred. The pilot, as mentioned before, lay prone within the wing contour. Two escape hatches were
located approximately opposite the center of his body, one on the upper surface, the other on the lower
surface. The handle which released the escape hatches was located close to the handle which released the
towing cable from the tug airplane. At the start of this particular flight, after a successful climb to 10,000 ft.,
the pilot inadvertently released the escape hatches at the time of his release from tow, and as a result partially
fell out of the airplane. The instinctive grasp on the control mechanism resulted in an indescribable wing-over
maneuver. When things calmed down the pilot found himself in a steady, uniform glide with the airplane upside
down. Minor movement of the controls seemed to produce little effect and the much shaken individual
crawled out of the airplane, sat on the leading edge of the center section while he checked his parachute
harness, and then slid off to make a perfectly normal parachute descent. The airplane, undisturbed by the
change in C. G., continued a long circling flight of the test area and finally landed in a normal continuation of its
upside down glide, a short distance from the takeoff point. It was rather seriously damaged but not so much
so as to prevent repair. A later check in the wind tunnel indicated that there was a very stable region in
inverted flight with this particular slat combination. Later the slats were abandoned as unnecessary and
perhaps undesirable.

The airframe was considered suitable for the purpose intended long before the rocket motors had been
developed to a degree of reliability considered safe for use, but finally a small motor having about five minutes'
duration, was installed and a number of rocket-powered flights were accomplished. The first powered flight
occurred in July 1944.

Although the first concept of the XP-79 as this fighter was designated, was as a rocket-powered vehicle
(similar in basic idea to the Messerschmitt ME-163), it soon became apparent that the completion of the
rocket motors would be far behind schedule and that serious difficulties were attendant to this development.
One of the basic concepts for the full-size motor was that the fuel pumps would be driven by rotation of the
combustion chambers, which were set at a slight angle to the thrust axis in order to develop torque. It was not
foreseen that the rotation of the combustion chambers would have a serious effect on the combustion therein,
and this difficulty, never completely solved, caused the abandonment of the particular engine which was being
developed for the project.

XP-79B TURBOJET AIRPLANE
As no alternative rocket engine was available, it became necessary to modify the design to incorporate
turbojet power plants, and the second of the XP-79 series, called the XP-79B, was completed with two
Westinghouse B-19 turbojets and first airborne on September 12, 1945. The takeoff for this flight was
normal, and for 15 minutes the airplane was flown in a beautiful demonstration. The pilot indicated mounting
confidence by executing more and more maneuvers of a type that would not be expected unless he were
thoroughly satisfied with the behavior of the airplane.

After about 15 minutes of flying, the airplane entered what appeared to be a normal slow roll, from which it
did not recover. As the rotation about the longitudinal axis continued the nose gradually dropped, and at the
time of impact the airplane appeared to be in a steep vertical spin. The pilot endeavored to leave the aircraft
but the speed was so high that he was unable to clear it successfully. Unfortunately, there was insufficient
evidence to fully determine the cause of the disaster. However, in view of his prone position, a powerful,
electrically controlled trim tab had been installed in the lateral controls to relieve the pilot of excessive loads. It
is believed that a deliberate slow roll may have been attempted (as the pilot had previously slow rolled and
looped other flying-wing aircraft developed by the company) and that during this maneuver something failed in
the lateral controls in such a way that the pilot was overpowered by the electrical trim mechanism.

ALL-WING BUZZ BOMBS
Several other all-wing aircraft and variations of them were built and tested during the same period. Shortly
after the advent of the V-1 an all-wing "buzz" bomb was designed and built. This airplane housed the German
V-1 resonator in a duct in the center of the wing and carried twice the German warhead in cast wing sections
on each side of the power plant with fuel in the outer wings. Several were built and flown successfully.

The first of these buzz bombs was tested as a pilot-controlled glider with good success. It was very small and
incorporated a number of extra bumps which were originally conceived to be the best way to carry standard
2,000 lb. demolition bombs. In spite of its peculiar configuration, which departed appreciably from the
all-wing ideal, it had quite good flight characteristics, was flown on a number of occasions (the airplane was
successfully slow-rolled) and demonstrated the suitability of the type for the purpose intended.

The one difficulty experienced in this series of tests is worthy of note. The piloted version of the buzz bomb
naturally required some type of landing gear for takeoff and landing, and in this case we employed tiny,
low-pressure air wheels, rigidly mounted in the airframe structure and extending only a few inches below the
contour of the aerofoil or, more specifically, the bomb-shaped bumps thereon. Landing on this gear involved
bringing the airplane in at an altitude of approximately 15 percent to 20 percent of the mean aerodynamic
chord just prior to contact, and no amount of practice on the part of the pilot produced a technique
satisfactory for this purpose. In every case a change in airflow appeared to develop as the airplane
approached within a quarter-chord length of the ground. The drag was apparently reduced, the lift increased
and the airplane rose, in spite of anything the pilot could do, to a height of 8 or 10 ft. above the ground, at
which point it stalled and flopped down out of control. This maneuver resulted in a number of rough landings
but no damage to either the pilot or the airplane. It was later found that the only way to make any sort of
smooth landing was to bring the airplane in at comparatively high speed and actually fly it onto the ground.
This difficulty was not experienced in airplanes having normal landing height above the ground, such as the
N-9M and XB-35.

XB-35, LONG-RANGE BOMBER
During all this development and testing of other types and scale versions of the XB-35, the design and
construction of the big ship had been under way. N-9M airplanes had proved the practicability of the design.
They closely approached the XB-35 configuration with the exception that they mounted only two pusher
engines, located at positions corresponding to points midway between engines 1 and 2, and engines 3 and 4.

The problem of control-surface actuation on the big bomber involved the development and testing of a
complete hydraulic control system, as none of the aerodynamic boosts or balances developed and tested in
the N-9M models had proved satisfactory. The system used in the XB-35 employs small valves which are
sensitive to comparatively minute movements of the control cable and which, when displaced, permit large
quantities of oil to flow into the actuating cylinders. This arrangement eliminates any pilot "feel" of the load on
the control surfaces unless a deliberate arrangement for force feedback is made. Rather than undertake this
later step, a comparatively simple force mechanism, which is sensitive to accelerations and airspeed, was
developed. This device gives the pilot a synthetic feel of the airplane which can be adjusted in intensity to
anything he likes, and which has proved satisfactory in flight. For reasons to be outlined shortly, a synthetic
feel was much more satisfactory than the feedback of actual control surface loads, particularly at high angles of
attack.

The XB-35 was first flown from Northrop Field to the Muroc Army Test Base in June 1946. The first several
flights indicated no difficulties whatsoever with the airframe configuration. Indications of trouble with propeller
governing mechanisms were discerned at an early date and it was shortly discovered that flights of any
substantial duration could not be accomplished because of oil leakage in the hydraulic propeller governing
system. On the last flight difficulty with both propellers on one side caused a landing with asymmetrical power,
which was accomplished without trouble.

The next six months, from August to March, were spent in a vain attempt to eliminate these difficulties, plus
those caused by a series of engine reduction gear failures. To date the XB-35 has not had sufficient time in the
air to fully demonstrate its ability to meet its design performance guarantees. However, large-scale model tests
in numerous tunnels have indicated the low-drag figures presented earlier in this paper, and preliminary speed
versus power tests completed early this month have given gratifying confirmation of our original expectations.
Flights accomplished to date have included all maneuvers necessary for large bombardment airplanes. So far,
however, violent maneuvers have not been attempted and no exact evaluation of stability and control
parameters has been possible.

Two turbojet powered all-wing airplanes, having the same basic shape and size as the XB-35 are virtually
complete at this time and will be flying late this summer. They are powered by eight jets having a sea level
static thrust of 4,000 lb. apiece. They incorporate small vertical fins to provide the same aerodynamic effect as
the propeller shaft housings and propellers of the XB-35.

Let us now turn to considerations of stability and control of the all-wing airplane. They are quite different from
those of conventional types and, unless reasonably well understood, may lead to discouragement at an early
date concerning projects well worth further evaluation.

STATIC LONGITUDINAL STABILITY
In any airplane the primary parameter determining the static longitudinal stability is the position of the center of
gravity with respect to the center of lift or the neutral point. Obviously, the neutral point may be shifted aft by
adding a tail or by sweeping the wing, or the C.G. may be shifted forward by proper weight distribution, so
that from the standpoint of static stability no particular configuration has any special advantage except as it
affects the possibilities of proper balance. In an all-wing airplane the elimination of the tail makes the problem
of balance somewhat more critical but not excessively so. Unfortunately, for any given airplane the neutral
point does not ordinarily remain fixed with variations of power, flap-setting or even lift coefficient, so that the
aft C.G. limit for stability is often prescribed by some single flight condition has always occurred for power-off
flight at angles of attack approaching the stall.

CHARACTERISTICS AT HIGH LIFT
The pitching instability of a swept wing at high lift coefficients is by now a somewhat familiar phenomenon.
The complete mechanisms involved, however, are still somewhat obscure. There are apparently two opposing
effects which are of prime importance. They are the tendency for sweepback to increase the relative tip
loading and also (by creating a span-wise pressure gradient) to promote boundary layer flow toward the tip.
On a plain swept-back wing the latter effect apparently nullifies the former, so that there occurs in the tip
portion of the wing a gradual decrease in effective section lift-curve slope with a resulting progressive decrease
in stability. The tip, under these circumstances, never completely stalls, as evidenced by the stable pitching
moments occurring at the maximum lift coefficient. On the other hand the addition of end plates will prevent to
a large extent the effects of span-wise flow, thereby straightening the pitching moment curve but producing the
normally expected tip stall, as evidenced by the strongly unstable moments in the vicinity of the maximum lift
coefficient. Thus, any modification to the basic wing which affects the span-wise flow will have a noticeable
effect on the pitching behavior at high lift coefficients.

In the case of the XB-35 the propeller shaft housings act to inhibit span-wise flow and straighten out the
moment curve below the stall as in the case of the end plate; but in order to obtain stability at the stall, a
tip-slot is provided to increase the stalling angle of the tip sections. By raising the trim flap in the outer 25
percent span and lowering the main flap in the inner 35 percent span, the stability characteristics are noticeably
affected, presumably because of a decrease in spanwise pressure gradient and therefore in boundary layer
flow.

Recent investigations have indicated that the problem of static longitudinal instability near the stall for plain
swept-back wings depends not only on sweep but also on aspect ratio and it now appears that for a given
sweepback the magnitude of the unstable break in the moment curve decreases with decreasing aspect ratio,
eventually vanishing.

The possibility of controlling the stalled portions of the wing, as outlined, means that trailing edge flap controls
can be laid out to maintain their effectiveness at very high angles of attack. Since a certain portion of this flap
must be used to provide high lift and roll control, the amount available for longitudinal trim is limited, so that for
the XB-35, for example, the total available nose-up pitching moment coefficient is .15 as compared to .30 for
a conventional airplane. This limited control plus the fact that the main wing flaps apparently cannot be made
self-trimming and impose a diving moment in the landing condition reduces the available C.G. range in percent
of the m.a.c. as compared with conventional airplanes. The XB-35 has a C.G. range of only 5 percent or 6
percent as compared with conventional values in the order of 10 percent or 12 percent. This comparison is
somewhat misleading, however, because the all-wing airplane may have a greater comparative m.a.c. in view
of its somewhat lighter wing loading. It is also much easier to arrange weight empty and useful load items
spanwise within close m.a.c. limits than in conventional types.

Where manual control of the elevator is employed the stick-*free stability and control of all-wing aircraft are
impaired by separation of the flow from the upper surface of the wing near the trailing edge, causing
up-floating tendencies at higher lift coefficients. If not corrected these up-floating tendencies lead to stick-free
instability and, in some cases, to serious control-force reversal at high lift coefficient. Aerodynamic design
refinements devised and tested by us to date have not provided a satisfactory solution to the up-floating
tendency. For small airplanes these undesirable forces can sometimes be tolerated, but for large aircraft the
only solution found so far has been the employment of irreversible full power driven control surfaces.

LATERAL STABILITY DERIVATIVES
It is when considering the lateral stability and control factors that the difference between the all-wing and
conventional airplanes becomes most apparent. It is reassuring to state that despite the large differences
apparent between the XB-35 and conventional aircraft, the dynamic lateral behavior of the XB-35 type is
quite satisfactory, as will be discussed later.

Definite requirements for the weathercock stability CAB, depend to a large extent on the airplane's purpose,
but positive weathercock stability is always required. The swept-back wing has inherent directional stability
which increases with increasing lift coefficient; but this is not considered sufficient for satisfactory flight
characteristics under all circumstances and must be supplemented by some additional device. The wingtip fin
has been favored by some since it gives the largest yawing lever arm and provides a suitable rudder location.
However, as previously pointed out, wingtip fins may be unsatisfactory at the stall. For the XB-35
configuration, effective fin area is provided in large measure by the side force derivative of the pusher
propellers.

RUDDER DEVELOPMENT
Rudders for all-wing aircraft are perhaps the chief control difficulty. Unless large fins are used a conventional
rudder cannot be employed. If large fins and rudders are used, an objectionable adverse side force due to
rudder is inherent, since the rudder moment arm is small and the side force comparatively great.

The use of pure drag rudders is feasible on the all-wing type because it is not necessary from a performance
standpoint to fly at zero yaw. Thus in the case of an engine failure equilibrium conditions involving a yaw angle
and the resultant corrective yawing moment do not involve appreciable side forces and associated bank
angles, nor noticeable drag increases. Thus, the rudder is used only rarely for trim and its drag is therefore
unimportant.

Of the many types of drag rudder investigated, a simple double-split trailing edge flap at the wing tip has been
found to have the most satisfactory all-round characteristics. This arrangement permits the simplest
construction and allows combination of trim flap and rudder in the same portion of the trailing edge. One
disadvantage of this type is its comparatively low effectiveness at low angles of rudder deflection, which may
be remedied by the employment of a nonlinear pedal-to-rudder linkage in the case of power-operated
rudders.

EFFECTIVE DIHEDRAL
Considering now the effective dihedral CID, it is apparent that sweepback is the essential difference between
the all-wing and conventional airplanes -- a difference that will disappear as flight speeds increase and it
becomes necessary to employ the desirable high-speed characteristics of swept wings in conventional tailed
configurations. For swept-back wings C1,8 increases quite rapidly with lift coefficient which gives difficulty
only when its value becomes too large. It is unimportant for either flight ease or for dynamic stability and
control characteristics when it is near zero. Flight ease may indicate that a slightly positive effective dihedral is
desirable while dynamic considerations point toward a slightly negative dihedral. Our practice has been to
retain positive effective dihedral over the complete flight range.

ROLL CONTROL
The rolling control for all-wing airplanes is essentially normal. When elevons are used rather than separated
aileron and elevator control, certain variations from conventional craft appear, in that, with the upward
elevator deflection required for longitudinal trim, the adverse yaw ordinarily due to aileron deflection
disappears. On the other hand, if large up-deflections are required for longitudinal trim, the up-going eleven
used as aileron loses effectiveness rapidly, thus reducing the available roll control at high lift coefficients.. This
is particularly undesirable when considering the increased dihedral effects of swept wings at high lift coefficient.

SIDE FORCE EFFECTS
All-wing airplanes, particularly those without fins, have a very low crosswind derivative; thus a low side force
results from sideslipping motion. Some crosswind force is probably important for precision flight, such as tight
formation flying, bombing runs, gun training maneuvers, or pursuit. This importance arises because with low
side force it becomes difficult to judge when sideslip is taking place, as the angle of bank necessary to sustain
a steady sideslipping motion is small. This lack of side forces has been one of the first objections of pilots and
others when viewing the XB-35. After flying in the N-9M or XB-35 the objection is removed, except for
some of the specific cases mentioned above. For the correction of the lack of sideslip sense, a sideslip meter
may be provided for the pilot or automatic pilot, and for very long-range aircraft there is a valuable
compensating advantage in being able to fly under conditions of asymmetrical power without appreciable
increase in drag.

DYNAMIC LONGITUDINAL STABILITY
The free longitudinal motions of any airplane fall into two modes. The first of these is a short-period
oscillation. It is highly damped for conventional airplanes and also for all-wing airplanes in spite of the relatively
low pitch-damping, Cmq. This somewhat surprising result is due to a coupled motion such that the vertical
damping, Z.,,,, comes into play absorbing the energy from the oscillation. Also, low moment of inertia in pitch
makes the small existing Cmq more effective than a similar value would be in conventional types. In tests on
the N-9M airplane this short-period oscillation was too rapidly damped to obtain a quantitative check. The
combination of low static stability in pitch, as previously described, and low moment of inertia in pitch results
in periods of oscillation for all-wing airplanes that are comparable to those of conventional types.

The second mode of longitudinal motion is a long-period oscillation commonly called the phugoid. This is a
lightly damped motion even for conventional airplanes, and seems slightly less damped for all-wing airplanes,
because of the fact that they have relatively low drag, and drag is the chief means of energy absorption in this
mode. N-9M tests indicate that calculation is slightly optimistic in this matter, but still this phugoid motion is
sufficiently damped so as to give no serious difficulties.. Being a slow motion, it is easily controlled.

To date the criteria for the description of airplane dynamic stabilities are vague. In the past it has been thought
that consideration of damping rates and periods of oscillatory motion were adequate, but it has become
evident that some further criteria are necessary. Consideration of the angular response of airplanes to various
unit disturbances may supply this need.

DYNAMIC LONGITUDINAL RESPONSE
The criterion of response is probably the only category in which the flying wing is importantly different from
the conventional airplane for longitudinal motion. The action of the two types in an abrupt vertical gust is
especially interesting, two factors combining to reduce the accelerations experienced by all-wing airplanes.
These factors are the relatively larger wing chord and shorter effective tail length of the all-wing type. The first
characteristic increases the time for the transient lift to build up and is the more important in reducing
accelerations. The second decreases the time interval between the disturbing impulse at the lift surface and the
correcting impulse at the effective tail, so that the airplane tends to pitch into the gust. This latter characteristic
is a matter of concern to pilots, since a disturbance in the air is likely to leave them farther from trim attitude,
consequently requiring more active pilot control in rough air. It is believed, however, that automatic control will
effectively eliminate this difficulty.

The response of the all-wing airplane to elevator deflection seems entirely adequate. It errs, if at all, on the
side of over-sensitivity because of low Cmq and low moment of inertia in pitch. An abrupt control movement
giving the same final change in trim speed for a conventional and a comparable all-wing airplane results in a
larger initial swing in pitch for the all-wing.

DYNAMIC LATERAL STABILITY
As with longitudinal motion, there are two characteristic modes that are of interest laterally. the first of these is
the spiral motion which is usually divergent on modern airplanes, thus uncontrolled flight results in a tightening
spiral. This slight instability seems favored by pilots. All-wing airplanes have readily acceptable characteristics
in this mode requiring from 15 to 20 seconds to double amplitude. In general, any time greater than five
seconds to double amplitude is considered acceptable.

The second mode, the "Dutch Roll" oscillation, is more critical for all-wing airplanes, particularly at low speed,
high weight and high altitude. All-wing airplanes seem comparatively bad in this respect because of the
combination of relatively large effective dihedral and low weathercock stability and, for the conditions noted
above as critical, are likely to approach neutral damping in the Dutch Roll mode. However, analytical
determinations of this motion, using calculated damping derivatives, indicated less satisfactory characteristics
than were obtained in actual flight tests. Because of a relatively low weathercock stability, the Dutch Roll is of
a rather long period, in the order of ten seconds for the XB-35. It is usually assumed that for periods of such
length, it is not important to have a high rate of damping since control would seem easily "inside" the motion.
However, there may be particular instances where this is not true. For instance, in an all-wing airplane in which
the rudder is particularly weak, the time of response to rudder control may be of the same order as the period
of Dutch Roll motion. This would make directional control extremely difficult in a condition, such as landing,
where the roll controls are not usable for changing heading. It is notable that for the very low weathercock
stability commonly encountered in all wing airplanes, the conventional solution of increasing weathercock
stability to offset increased dihedral does not hold. Increasing Cur leaves the damping essentially untouched,
but reduces the period and increases the number of cycles required to damp.

Another factor contributing to the relative lack of damping of all-wing airplanes in Dutch Roll motion is the low
value of the damping coefficient in yaw, Cur This appears to be inherent in all-wing designs, particularly if the
use of fins is abandoned. For special occasions, when particular airplane steadiness is required (such as a
bombing run), it is probable that the equivalence of such damping in yaw may be supplied by an automatic
pilot, or by temporarily increasing the drag at the wing tips. This latter effect can be accomplished on the
XB-35 by simultaneously opening both rudders and gives deadbeat damping in yaw.

DYNAMIC LATERAL RESPONSE
As in the longitudinal motions, the amplitudes of response of an airplane in lateral motion are probably as
important as the damping rates in determining free-flight characteristics. All-wing airplanes seem slightly
rougher in turbulent air than conventional aircraft of similar weight. This is due chiefly to the reduced wing
loading, but high effective dihedral and low weathercock stability may have an added effect. This is a matter of
interest in fixing upon analytical criteria for the description of free-flight qualities. As mentioned above,
increasing the weathercock stability for all-wing airplanes has a slight effect on the damping rates; however, it
affects the amplitudes of response to gusts materially.

Some data from the free-flight tunnel of the National Advisory Committee for Aeronautics indicate that
increasing weathercock stability, even for all-wing airplanes, materially helps the "flyability" of the airplane.
Another bit of evidence that is of interest in this connection has to do with the magnitude of the side force
derivative, Cy.B. Increase of this parameter improves Dutch Roll damping very materially but has virtually no
effect on amplitude of response to gusts, according to calculations. Free-flight wind tunnel data again give
tentative support to the investigations of response as a criterion by showing little improvement of flight qualities
of models with increase of Cy'B.

Flight tests of the all-wing glider in which the vertical fin, located aft on the ship's center line, was varied in size
from approximately 2 to 7 percent of the wing area, left the pilot somewhat undecided as to fin requirements
except that the larger fin seemed somewhat easier to fly. Presumably, this was, in the light of the foregoing
discussion, primarily because of the increased CnD, the coincidental increase in Cy'0 not being effective.

AUTOMATIC PILOT CONTROL
The application of automatic pilot control to an all-wing airplane has certain difficulties which are associated
primarily with the low value of C ,B. In conventional applications the fact that the airplane is side slipping is
detected by either a lateral acceleration or an angle of bank. In an all-wing airplane neither of these indications
exists except in an almost undetectable amount. Accordingly, it is necessary, in order to fly the airplane at zero
sideslip, and therefore in the direction of its center line, to provide a yaw-vane signal to which the pilot or
automatic pilot will respond. This introduces some difficulty in automatic pilot design because for small
disturbances the sideslip angle with respect to the wind, and the yaw angle with respect to a set of fixed axes,
are nearly equal and opposite for a flying wing. The customary automatic pilot control on azimuth angle
therefore tends to oppose the necessary control on sideslip. To avoid this difficulty it is necessary only to
reduce the rate of control on sideslip to approximately one-third that on azimuth. This modification to a
conventional automatic pilot was flown on the N-9M with complete success.

PROBLEMS OF CONFIGURATION--SWEPT vs. NON-SWEPT WINGS
Let us now turn to a consideration of the practical limitations in arrangement of the tailless airplane. They may
be summarized briefly as sweepforward, sweepback, and a non-swept wing configuration. The sweepforward
arrangement requires the use of a large fixed load forward of the leading edge at the center section for proper
balancing of the airplane. Therefore, a fuselage with some substantial part of the weight empty of the airplane
disposed therein is required. The swept-forward wing itself is unstable directionally and requires some type of
fin for weathercock stability. To this must be added more fin area to stabilize the fuselage. In addition, it may
be noted that the moment arm of the fin about the C.G. of the airplane is necessarily comparatively small, still
further increasing the size of the required fin. If we add to the aerofoil a protruding fuselage and an unusually
large vertical tail surface, we have departed from our basic all-wing concept. We have incorporated virtually
all the elements of drag found in the conventional aircraft and have not accomplished our intent of improving
efficiency.. For the above reasons, which could be argued pro and con for hours, our company has done no
active design and development work on airplanes with swept-forward wings.

An all-wing configuration embodying a straight, or non-swept wing, has been proposed and flown
successfully in model sizes. It offers the serious disadvantage that suitable distribution of weight empty and
useful load items is difficult and, if proper balance is to be accomplished, most of the structural weight and
useful load must be included in the forward 30 percent or 40 percent of the wing, leaving a large volume of
space within the wing unusable. Such a configuration results in an unnecessarily large airplane to accomplish a
given job and for this reason has not been considered seriously.

The swept-back arrangement exemplified by the various airplanes previously illustrated and described seems
to offer the best configuration for a materialization of our all-wing ideal. It can be balanced satisfactorily within
quite wide ranges of sweepback, utilizing almost all available volume within the wing for storage of useful load
items. It seems to fly satisfactorily in many different configurations and the arrangement is such that large
payloads can be carried virtually over the C.G., with the weight empty items so distributed as to cause little
variation in C.G. position between the fully loaded and empty conditions.

WEIGHT DISTRIBUTION
As has been pointed out previously, the permissible range of C.G. location is not overly critical in this type of
airplane. It is, nevertheless, of great advantage to be able to load the airplane almost at will, without concern
as to how the useful load is disposed and the swept-back configuration lends itself most suitably to such
loading.

In the case of the XB-35, the useful load, consisting largely of bombs and fuel, can be readily disposed in
suitable position about the C.G. While some fuel is located well forward and other fuel well aft of the desired
C.G. location, under normal operating conditions the proper balance is readily maintained. In case of failure of
one or more engines, it is necessary to pump the fuel from unused tanks to those supplying the remaining
engines, but a simple manifolding system provides this facility.

Based on a great many studies of various types and applications of the all-wing principle, some practical
limitations may be approximately defined. Where very dense (high specific gravity) payloads are
contemplated, such as warheads or similar munitions, quite small units are practical, as demonstrated by the
all-wing buzz bombs to which reference has been made. Medium-sized units having a span of perhaps 100 ft.
and a gross weight of 50,000 to 60,000 lb., appear entirely practical for medium bombers and freighters.
Here again the density of the useful load, both in payload and fuel, is comparatively high.

Airplanes designed to carry people need the largest volume of all. Even individual reclining chair
accommodations require a minimum space of perhaps 40 cubic ft. per passenger, which is a density of only
about 5 lb. per cubic ft. This is one-half to one-quarter the density of typical air cargo, and only 4 percent or 5
percent of the density of a warhead.

IMMEDIATE APPLICATIONS--ALL-WING AIRCRAFT
It may be concluded, then, that the all-wing design is immediately applicable and practical for a number of
military and cargo-carrying versions, and that the passenger-carrying aircraft are likely to be of rather large
size and, in the immediate future at least, will provide only comfortable seating instead of the more luxurious
appurtenances associated with long-range ocean travel.

An airplane of the XB-35 configuration and size can carry 50 passengers in comfort in the existing aerofoil
envelope with adequate headroom for all, and with vision forward through the floor, and upward if desired.
Passenger vision in a flying wing may be more satisfactory than in conventional types if we get used to the idea
of forward vision rather than that provided by side windows. The really interesting views are likely to be
forward and downward rather than to the side. An airplane like the XB-35 will have cargo space for 40,000
to 50,000 lb. of air freight at a density of 10 to 15 lb. per cubic ft., in addition to the necessary crew and
space for 50 passengers.

FUTURE POSSIBILITIES
Turning now to future possibilities, it seems that considerable further aerodynamic refinement can be made
over that already accomplished in all-wing types. Particularly if turbojets are used as the motive power, the
minimum parasite drag may be reduced to .008 or less. This value is obtained by subtracting the drag of
propeller shaft housings, gun turrets and other military protuberances from the XB-35 configuration and
assuming an improved degree of aerodynamic smoothness of the aerofoil section. Boundary layer removal and
the use of somewhat thinner wing sections may further appreciably reduce this figure.

A maximum trimmed lift coefficient 1.9 for the all-wing configuration seems attainable by methods already
suggested and possibly may be further increased by judicious use of boundary layer control in combination
with turbojet power plants. It is our opinion that the ratio of C1max to Cd min may be increased to a value of
235 within the not-too-distant future from our present actual achievement of about 130. In contrast, the years
of intensive development of the conventional types already passed promise an improvement of less magnitude
within a comparable time. In our judgment a trimmed maximum lift of 2.8 vs. a minimum drag of .020 seems
reasonable to expect for large, long-range transport and bombardment aircraft of conventional type.

These estimates are, of course, completely arbitrary and controversial. However, if one cares to assume their
validity, the following conclusions may be reached, based on methods and calculations used in the early part of
this paper. The total minimum profile drag of the all-wing airplane in terms of the conventional will be from 40
percent to 59 percent. The power required by the all wing to maintain the same cruising speed as the
conventional will be from 70 percent to 80 percent and, conversely, the maximum range of the all-wing, at the
cruising speed of the conventional airplane, will be 143 percent to 125 percent. The maximum range of the
all-wing airplane at its best cruising speed will be 158 percent to 130 percent of the conventional, and the
most economic speed will be from 125 percent to 115 percent faster.

Under high speed conditions corresponding to full power of reciprocating, turboprop or turbojet engines,
where the induced drag is assumed to be 20 percent and the parasite drag 80 percent of the total, the power
required to drive the all-wing airplane at the speed of the conventional airplane will be 52 percent to 67
percent and, conversely, the range will be 192 percent to 149 percent of the conventional airplane. The
maximum speed of the all-wing airplane at comparable powers will be 124 percent to 114 percent of its
conventional counterpart.

Different assumptions of comparative maximum lift and minimum drag values can be made to suit individual
opinion, but it is believed that any reasonable assumptions will always result in an advantage to the all-wing
configuration of such magnitude as to fully warrant whatever trials and tribulations may be associated with its
development.

POSSIBLE SUPERSONIC APPLICATIONS

So far in this discussion we have purposely avoided transonic and supersonic considerations. The neglect is
possibly a reasonable one when discussing commercial ventures, in view of the cost of higher and higher
speeds. A reasonable degree of sweepback, such as is required in the type of aircraft under consideration, will
permit speeds up to about 500 m.p.h. without involving great compressibility drag increases. For military
aircraft, however, we cannot ignore the sonic "barrier" and its implications, and it is a reasonable assumption
that sooner or later improved fuels will permit higher and higher operational speeds, even in commercial
aircraft.

Based on present knowledge of supersonic flight, it will always be more difficult to carry a given payload for a
given range at supersonic speed because of the additional wave drag encountered at these speeds. At
transonic or comparatively low supersonic speeds, a plain swept-back wing appears to be one of the best
possible configurations, provided that sufficient is available within the wing. Since the flow normal to the
leading edge is subsonic over almost the entire wing surface, subsonic airfoils with reasonably good subsonic
flight characteristics can be used at these speeds. The all-wing design eliminates wing-fuselage interference as
well as adverse interference between the tail surfaces and wing or body.

At higher supersonic speeds the problem or providing adequate volume is more difficult because of the fact
that more and more fuel is required for a given range and the percentage of thickness of airfoils suitable for
such use is much less than that satisfactory for subsonic flight. Save for one compensating factor, this problem
of volume and size might well rule out the all-wing airplane for supersonic use, and certainly does limit its
usefulness for low altitude flight. However, an attractive field of operation exists at very high altitude where air
densities are low and therefore wing areas must be comparably great if suitable lift coefficients are to be
maintained. If we design a frankly supersonic airplane to fly at, say, a Mach number of 1.6, with supersonic
diamond-section airfoils, the maximum cruising lift coefficient will probably be no greater than .15, and the
corresponding loading must be held to 40 lb. per sq. ft.

The above figures are based on assumed operation at 60,000 ft. and an air density ratio of .094. Such an
airplane might likewise be suitable for landing and takeoff at low altitude, in view of its comparatively light wing
loading, which would eliminate the necessity of high-life devices. The practicability of the design depends on
the relative density of the air at the altitude selected for cruising operation. If a sufficiently high altitude is
chosen it seems quite possible that adequate volume can be secured in the wing, in spite of its small thickness
ratio, by using low aspect ratio planforms approaching the triangular.

We can compare data on two wings having the same physical depth at the root, and identical wing areas. The
conventional wing is of a type already proved practical for all-wing airplanes. The delta wing has thickness
ratios suitable for supersonic flight, identical thickness and only slightly reduced volume. It should be quite
suitable for all-wing aircraft of reasonable size. From the aerodynamic point of view it appears that with the
delta wing it is possible to eliminate a substantial portion of the wave resistance and thus realize fairly favorable
lift-drag ratios at supersonic speeds.

It is gratifying to those of us who have been working on all-wing projects for years to recognize the increased
interest in the type evidenced in Germany toward the end of the war, and more particularly in England and
Canada in recent years. For many years we received scant encouragement and often seriously questioned our
own judgment, as well as our ability to achieve a successful solution to the many problems involved in the
development of this type. The goals and rewards have always seemed well worth attainment, however, and I
believe accomplishments to date have justified the effort required.

I hope this discussion may provide encouragement and incentive to those in Great Britain who have pioneered
all-wing airplanes and that these projects, both here and in the United States, may profit by each other's
mistakes and successes, thus bringing the two countries to the forefront in this important phase in the
development of air transport.