Aircraft structure to improve directional stability

An aircraft structure has an arrangement of aircraft components that provide inherent directional stability for a flight vehicle throughout an angle-of-attack range, even at very high angles-of-attack where conventional means of stabilization are ineffective. Components attached to an aircraft fuselage include a wing, horizontal stabilizers and vertical stabilizers. The wing is mounted forward of the horizontal stabilizers and is carried high on the fuselage. The horizontal stabilizer is mounted toward the rear of the aircraft and is attached near the bottom of the fuselage. The wing and horizontal stabilizers are joined on either side of the aircraft by forwardly sweeping aerodynamically shaped surfaces serving as the vertical stabilizers. The inclination of the vertical stabilizers preferably ranges from 45 degrees (top edge canted outboard) to 90 degrees (panels vertical). Preferably, the surface area of the vertical stabilizers is concentrated aft such that the aerodynamic center of the vertical stabilizers is located behind the center-of-gravity of the aircraft.

TECHNICAL FIELD 
This invention relates generally to aircraft and more particularly to a 
particular arrangement of components in aircraft design for providing 
increased directional stability at high angles of attack, a means to 
reduce the overall drag of the aircraft, and the capability to build an 
aircraft structure having a greater stiffness for a given aircraft weight. 
BACKGROUND ART 
Many types of aircraft are required to be highly maneuverable to perform 
their functions. Examples include aerobatic light aircraft, trainers, and 
fighter or attack aircraft. Such aircraft are occasionally required to 
operate in flight regions where the angle-of-attack is large. 
Angle-of-attack refers to the incidence of an aircraft with respect to its 
velocity vector. For aircraft to be able to fly safely at high 
angle-of-attack and perform maneuvers, the aircraft must be stable and 
controllable. In some modern aircraft, stability is provided by artificial 
means. Artificial stabilization may be achieved with control effectors, 
which are used to generate forces and moments to oppose unwanted aircraft 
motions. While artificial stabilization may be effective, it has some 
unfavorable side effects. For example, as the level of instability 
increases, the amount of control moment required to stabilize the vehicle 
increases. Beyond a certain level of instability, sufficient control 
moments may not be available, and the aircraft can experience a departure 
from controlled flight. Even if sufficient control authority is available 
to prevent unwanted motions, artificial stabilization results in less 
control power available to maneuver the aircraft. A loss of available 
control power results in a loss in mission effectiveness. Artificial 
stabilization is not an option with most low cost aircraft since such 
aircraft do not have computers and control effector designs required to 
add artificial stability. It is noted that a loss in inherent vehicle 
stability occurs with virtually all aircraft at high angles-of-attack. 
While stability in pitch axis can be altered with careful design and 
control of aircraft center-of-gravity, directional stability must be 
provided by stabilizing surfaces or by artificial means. Directional 
stability is defined as the tendency of an aircraft to weathercock into 
the wind when disturbed. When used without reference to an axis system, 
directional stability is assumed to be in the flight path or stability 
axis system. Although other axis systems may be used for convenience, the 
flight path or stability axis reflects the motion of the aircraft. Flight 
path directional stability is composed of two parameters, i.e. body axis 
directional stability and body axis lateral stability or dihedral effect. 
The body axes of a vehicle are mutually perpendicular and are normally 
aligned along the fuselage axis, the approximate plane of the wing, and 
normal to the other two. Body axis directional stability refers to the 
tendency of the fuselage to point back into the wind when disturbed from 
equilibrium. Body axis lateral stability refers to a tendency for the 
aircraft to roll in a direction to eliminate any side component of the 
relative wind. If an aircraft is at an angle-of-attack other than zero, 
flight path axis directional stability is a combination of both body axis 
directional and lateral stability and is calculated from equation 1 as 
follows: 
##EQU1## 
From equation 1 it is apparent that both body axis directional stability 
and body axis lateral stability contribute significantly to flight path 
directional stability when angle-of-attack is large. This implies that 
aircraft stability can be improved by any device or component that 
increases either or both values. It is highly desirable to provide 
inherent directional stability to maximize the amount of control power 
available to perform out of plane maneuvers, such as rolling, and to 
prevent rapid aircraft departures from controlled flight, which often lead 
to spins. Current and prior design practice has relied on large fixed 
aerodynamic surfaces mounted on the aircraft fuselage to stabilize the 
aircraft. However, such surfaces lose their effectiveness at large 
incidence angles. 
The current invention relies on a novel arrangement of wing and tail 
surfaces to provide inherent directional stability even at high 
angle-of-attack. The invention achieves these results by joining the wing 
and horizontal tail panels with canted or vertical stabilizing surfaces 
that operate effectively over a wide angle-of-attack and sideslip range 
and provide stabilizing moments about both directional and lateral axes. 
These stabilizing moments will insure that the aircraft remains 
directionally stable, or nearly so, to much higher angles-of-attack than 
is current practice. 
Directional stability has typically been provided by a single vertical tail 
located on the centerline of an aircraft at or near the rear of an 
aircraft fuselage. The vertical tail provides directional stability by 
acting as a lifting surface. When an aircraft is perturbed in a manner 
such that sideslip occurs, a local angle-of-attack occurs in the plane of 
the vertical tail. The local angle-of-attack generates a lifting force on 
the tail panel and creates a moment about the center-of-gravity of the 
aircraft that opposes the sideslip and returns the aircraft to a zero 
sideslip condition. The larger the tail surface area, the larger the 
moment that is generated, and the greater the directional stability. 
Although the use of a vertical tail to provide directional stabilization 
has proven effective over the years, a vertical tail presents serious 
disadvantages at high angles-of-attack. As aircraft angle-of-attack 
increases, the fuselage of the aircraft tends to block airflow to the 
vertical tail. The blockage of airflow reduces the effectiveness of the 
vertical tail. If angle-of-attack exceeds a certain value, this value 
being dependent on the vehicle in question, the flow over the fuselage 
separates, or detaches. Under a detached flow condition, not only is the 
tail partially blocked by the fuselage, but flow separation results in a 
region of low energy air in the vicinity of the vertical tail, which 
further reduces the effectiveness of the tail. 
An additional factor resulting in reduced effectiveness of the tail is the 
aft sweep of a typically configured tail. High performance aircraft 
typically employ swept wing and tail surfaces to reduce drag at transonic 
and supersonic Mach numbers. Due to structural considerations, these 
surfaces, including the vertical tail panel, are usually swept aft. For a 
vertical tail, aft sweep results in much of the airflow at high 
angles-of-attack being directed along the span of the surface of the tail 
rather than along the chord line of the tail, which further reduces 
effectiveness of the tail. 
A single, centerline vertical tail has additional disadvantages for modern 
high performance aircraft. Highly swept surfaces at the front of typical 
modern high performance aircraft are intended to generate vortices, or 
regions of high energy rotational flow, at high angles-of-attack. These 
vortices have been shown to interact with downstream aircraft components, 
sometimes in an unfavorable manner. Using the F-16 aircraft as an example, 
under certain conditions, the vertical tail actually contributes a 
destabilizing directional moment. 
Several modern high performance fighter aircraft employ twin vertical tails 
to partially overcome the disadvantages of a single tail surface. For a 
given aircraft configuration, the total area of the twin panels is usually 
more than a single panel configured to provide equivalent directional 
stability at low angles-of-attack. In all known current applications, the 
twin vertical tails are mounted on the fuselage. Therefore, at high 
angle-of-attack, the fuselage still tends to block a portion of airflow to 
the tail surfaces. Additionally, energy of the flow in the vicinity of the 
tail surfaces is reduced. On aircraft configurations set up to generate 
strong vortex flow fields, twin tail surfaces can reduce but not eliminate 
the effect of unfavorable interference due to vortex impingement and 
interaction. An additional problem encountered by fuselage mounted twin 
vertical tail configurations is the loss of effectiveness due to sweep 
back. 
Other arrangements have been proposed or used to provide inherent 
directional stability, but most of these are intended for low performance 
aircraft operating at low angle-of-attack. Examples of other arrangements 
include twin horizontal stabilizer mounted vertical surfaces (e.g., the 
B-24 bomber of World War II), boom mounted tails (e.g., the P-38 fighter 
aircraft), and a wing tip mounted vertical tail (e.g., the Beechcraft 
Starship and numerous light plane designs). All of these concepts attempt 
to improve upon conventional means of stabilization. None are entirely 
successful. A new approach is required to provide directional stability at 
high angles-of-attack. 
Historical results have shown that aerodynamic advantages exist for the 
biplane configuration. A biplane consists of two lifting surfaces or 
panels separated in height and sometimes longitudinal location. Test 
results have shown that the effective aspect ratio of a biplane is higher 
than that of a monoplane of the same span. The change in effective aspect 
ratio has been shown to be a function of main surface wing span, vertical 
distance between the lifting surface panels, ratio of the span of one 
panel to that of the other, and relative lengthwise positioning of the two 
panels, i.e., stagger. 
Attempts have been made in modern designs to make use of the benefits of 
the biplane. Such attempts have largely consisted of designs that join 
together tips of forward and aft lifting surfaces. These attempts have 
generally not been successful for a number of reasons. First, the direct 
joining of the tips of two lifting surfaces results in no vertical 
separation between the panels at the tips. Test results indicate that a 
condition of no vertical separation between lifting surfaces eliminates 
the desired biplane effect. Second, joining of the lifting surfaces limits 
design options. For example, if the tips of the lifting surfaces are 
joined, it is usually necessary for both lifting surfaces to have the same 
span, which results in a tandem wing configuration that is not ideal for 
all applications. 
An additional attempt to make use of the benefits of a biplane is a 
boxplane configuration. A boxplane configuration joins upper and lower 
wing panels in a rigid structure. However, a box plane arrangement uses a 
conventional horizontal stabilizer and does not attempt to use the joining 
members as vertical stabilizers. Other modern designs have attempted to 
use a biplane arrangement for supersonic vehicles. These designs may take 
advantage of favorable shock wave interaction at supersonic Mach numbers. 
Due to the nature of shock waves and their inclination with respect to an 
aircraft's direction of motion, a supersonic biplane arrangement is not 
suitable for a conventional wing and tail arrangement where the separation 
of wing and tail surfaces is set by considerations of aircraft balance and 
control capability rather than simply being determined by shock wave 
inclination at a design Mach number. 
SUMMARY OF THE INVENTION 
Consequently, one object of this invention is to improve on the prior art 
of providing directional stability at high angles-of-attack for flight 
vehicles. This goal is achieved by a unique arrangement of aircraft 
components including the wing, horizontal stabilizers, and vertical 
stabilizers. In addition to greatly improving high angle-of-attack 
directional stability, application of this invention results in a flight 
vehicle configured as and behaving in a manner similar to that of a 
biplane, thereby increasing the effective aspect ratio of the wing and 
reducing the induced drag of the aircraft. 
A further advantage of this invention is that by joining all major lifting 
surfaces into one unit, an aircraft structure can be made stiffer for a 
given aircraft weight. Conversely, the weight of an aircraft can be 
reduced for a given stiffness. 
A further object of this invention is to provide for a component 
arrangement that enables the vehicle to behave aerodynamically in the 
manner of a biplane, with an increase in effective aspect ratio of the 
wing and a corresponding reduction of the drag due to lift of the 
configuration. 
In a preferred embodiment, the primary lifting surface, or wing, is located 
in a conventional position ahead of a horizontal tail and is mounted in a 
high position on the fuselage, that is, near the top of the fuselage. The 
horizontal tail surfaces, or horizontal stabilizers, are mounted behind 
the wing near the rear of the aircraft. The horizontal stabilizers are 
mounted in a low position on the fuselage, i.e., near the fuselage bottom. 
The directionally stabilizing surfaces consist of two aerodynamically 
configured panels joining the tips of the wing and horizontal stabilizer 
on each side of the aircraft. For simplicity and for historical reasons, 
the panels that join the tips of the wings and horizontal stabilizers will 
be referred to throughout this disclosure as vertical stabilizers although 
they may be also considered directional stabilizers. It is noted, however, 
that the orientation of the vertical stabilizers is not necessarily 
vertical, but may be canted out at the top. The arrangement of wing, 
horizontal stabilizers, and vertical stabilizers therefore forms a closed 
box structure. The required vertical separation between the wing and 
horizontal stabilizer causes the vehicle to behave in a manner of a 
biplane, with benefits as described above. The closed box configuration 
additionally produces a stiffer structure than a typical arrangement of 
aircraft components where the wing, horizontal stabilizer, and vertical 
stabilizer(s) are cantilevered out in space. 
The preferred arrangement of this invention is a wing having a span equal 
to or greater than the span of a horizontal stabilizer. Such an 
arrangement forces the vertical joiner panels or vertical stabilizers to 
be either perpendicular to the approximate plane of the wing or canted out 
at the top. Since the wing is ahead of the horizontal tail, the vertical 
stabilizers joining the wing and tail are inclined to produce a forward 
aerodynamic sweep. Note that such an arrangement is in contrast to a 
typical arrangement for high speed vehicles where the sweep of lifting 
surfaces is traditionally in the aft direction. 
In practice, the unique configuration of applicant's invention results in 
desirable flight characteristics. When an aircraft utilizing applicant's 
invention encounters a disturbance such as side slip, or a sideways 
velocity component at low angle-of-attack, incidence to the airflow 
develops on each panel of the vertical stabilizers. This incidence can be 
transformed to an angle-of-attack on a vertical stabilizer approximately 
equal to the sideslip angle times the sine of the dihedral angle of the 
vertical stabilizer. Dihedral angle is defined as the angle between the 
reference plane of the aircraft and a plane along the span of the vertical 
stabilizer. The sideslip perturbation results in an incremental force 
being generated normal to the plane of each vertical stabilizer. The 
incremental force is in addition to any side-to-side symmetric force which 
would occur on the panel during normal operation without sideslip. Since 
the center of pressure of the incremental force due to sideslip lies 
behind the center of rotation or center-of-gravity of the aircraft, the 
net effect is to produce a moment on the aircraft which tries to reduce 
the sideslip and roll the vehicle away from the sideslip. The first effect 
of reduction of sideslip is known as "body axis directional stability". 
The second effect is known as "body axis lateral stability". The net 
effect of the vertical stabilizers is to stabilize the aircraft in 
sideslip as desired. 
If angle-of-attack is not zero, then both body axis directional stability 
and body axis lateral stability contribute to flight path directional 
stability. For the present invention, the moment components generated by 
the vertical stabilizers provide stabilizing moments about both axes. If 
the vertical stabilizer panels are indeed vertical (i.e., approximately 
perpendicular to the plane of the wing), then a directionally stabilizing 
moment is generated but not a laterally stabilizing moment. 
The improvement provided by this invention is its effect at high 
angles-of-attack where a conventional vertical tail loses its 
effectiveness. In an embodiment of the invention where the vertical 
surfaces are canted outboard, the total forces on the panel will typically 
act upwards in a symmetric condition, i.e. when no sideslip is present. 
The incremental force introduced on each panel tends to stabilize the 
aircraft both directionally and laterally. 
Several differences are apparent between the two wing mounted vertical 
surfaces and the single centerline vertical tail found on conventionally 
configured aircraft. First, in the two wing mounted vertical surfaces 
embodiment, the fuselage tends to block air flow to the downwind vertical 
stabilizer panel. While the incremental force on the downwind panel is 
still stabilizing, the total force acting normal to the surface of the 
downwind panel attempts to rotate the aircraft in the wrong or 
destabilizing direction if the downwind panel has geometric dihedral and 
carries an aerodynamic upload. However, the upwind panel provides a 
stabilizing moment, both from the basic force carried on the panel at high 
angles-of-attack as well as the incremental force due to sideslip. 
Therefore, any blockage of the downwind panel due to the presence of the 
fuselage actually results in an improvement in directional stability, 
since the effect of the destabilizing moment is reduced. 
The upwind vertical stabilizer panel operates in undisturbed air away from 
the influence of the fuselage. The capability of operating in undisturbed 
air contrasts with conventional single or multiple vertical stabilizer 
panels mounted on the fuselage. For conventional arrangements, blockage 
renders vertical stabilizers ineffective at high angles-of-attack 
resulting in a loss of directional stability. In the current invention, 
the aerodynamic surface that provides directionally stabilizing moments 
(i.e., the upwind stabilizer panel) is well away is from the fuselage in a 
region of relatively undisturbed air and can continue to be effective at 
virtually any angle-of-attack. 
Second, the arrangement of components of this invention requires that the 
wing be located in a position higher than and forward of the horizontal 
stabilizer. Consequently, the vertical stabilizers must be swept forward 
aerodynamically. As angle-of-attack increases, a component of flow 
velocity normal to the leading edge increases for the vertical stabilizers 
of this invention. This increases the aerodynamic effectiveness of the 
surfaces. The increase in aerodynamic effectiveness due to the forward 
sweep of the vertical surfaces contrasts with the result of the more 
typical unswept or aft swept vertical stabilizing surfaces employed on 
other aircraft where an increased angle-of-attack results in more of a 
local velocity component being directed along the span of the vertical 
surface, resulting in a loss in aerodynamic efficiency. 
Preferably, the vertical stabilizers have a straight leading edge extending 
from a leading edge of the wing to a leading edge of one of the horizontal 
stabilizers. Additionally, each of the vertical stabilizers have a 
straight trailing edge extending from a trailing edge of the wing to the 
trailing edge of the horizontal stabilizers. The vertical stabilizers also 
preferably have a chord that is substantially constant from a lower end to 
an upper end. 
It is envisioned that the implementation of this invention in the design of 
a particular flight vehicle would be tailored to the overall design 
requirements, the operating angle-of-attack range desired, and other 
factors. This invention can be applied to any flight vehicle, aircraft or 
missile. It is appropriate for everything from a high performance, 
supersonic fighter aircraft to low speed vehicles powered by piston or 
turboprop engines driven by a propeller. 
Other and further features and objects of the invention will be more 
apparent to those skilled in the art upon a consideration of the appended 
drawings and the following description wherein several typical 
installations are shown and one constructional form of apparatus for 
carrying out the invention are disclosed. This invention consists of the 
construction, combination, and arrangement of parts all as in hereinafter 
more fully described, claimed, and illustrated in the accompanying 
drawings.

DETAILED DESCRIPTION OF THE INVENTION 
Referring now to FIG. 1, shown is an aircraft designated generally 10. The 
components of aircraft 10 include fuselage 12 having a top 14 and a bottom 
16, and right wing 18 and left wing 20, which are affixed to fuselage 12. 
Right wing 18 has right wing tip 22. Left wing 20 has left wing tip 24. 
Proximate fuselage bottom 16, are right horizontal stabilizer 26 and left 
horizontal stabilizer 28, which are affixed to fuselage 12. Right 
horizontal stabilizer 26 has right horizontal stabilizer tip 30 and left 
horizontal stabilizer 28 has left horizontal stabilizer tip 32. Spanning 
between right wing tip 22 and right horizontal stabilizer tip 30 is right 
vertical stabilizer 34. Spanning between left wing tip 24 and left 
horizontal stabilizer tip 32 is left vertical stabilizer 36. 
In this embodiment, inlet 38 feeds two rear mounted engines which exhaust 
through nozzles 40. Vertical stacking of the engines results in vertical 
separation 42 of wings 18 and 20 with respect to horizontal stabilizers 26 
and 28, which can be seen most clearly in FIG. 2. 
Referring now to FIG. 2, horizontal stabilizer span 44 is approximately 
seventy percent of wing span 46. Therefore, vertical stabilizers 34 and 36 
are canted outward toward the top as can be more clearly seen in FIG. 2. 
Referring back to FIG. 1, wings 18 and 22 use typical trailing edge devices 
48 such as outboard ailerons and inboard trailing edge flaps. The 
embodiment of the invention shown in FIG. 1 is an example of a subsonic 
trainer aircraft. Therefore, elevators 50 are used on the trailing edges 
of horizontal stabilizers 26 and 28. Each trailing edge of horizontal 
stabilizers 26, 28 has an inboard junction at fuselage 16 and an outboard 
junction at the tip of each horizontal stabilizer 26, 28. As shown in the 
drawings, the inboard junction is no further aft than the outboard 
junction. In the drawings, the outboard junction of the trailing edge of 
each horizontal stabilizer 26, 28 is farther aft than the inboard 
junction. For supersonic application, all moving horizontal stabilizers 
may be desirable. Movable rudders 52 are located on the trailing edges of 
vertical stabilizers 34 and 36 to provide control about the yaw axis. 
Referring back to FIG. 2, a front view of aircraft 10 is shown. The 
direction of relative wind is designated by relative wind component vector 
54. For a condition of no side slip velocity, relative wind component 
vector 54 is zero. In a condition of no side slip, right vertical 
stabilizer 34 and left vertical stabilizer 36 may carry a small load as 
indicated by right vertical stabilizer load vector 56 and left vertical 
stabilizer load vector 58. In the presence of a side velocity that 
produces a side slip angle, an incidence angle is produced on right 
vertical stabilizer 34 and left vertical stabilizer 36 that generates 
right vertical stabilizer incremental force 60 and left vertical 
stabilizer incremental force 62. Incremental forces 60 and 62 produce a 
moment or twisting reaction about directional axis 64 and lateral axis 66, 
which may be more easily seen in FIG. 3. Center of gravity 68 is the 
reference point for the moments. Forces acting on right vertical 
stabilizer 34 and left vertical stabilizer 36 act behind center of gravity 
68 since vertical stabilizers 34 and 36 are located towards the rear of 
the vehicle. Therefore, the resulting moment about directional axis 64 is 
in a direction that will try to point aircraft 10 into the side velocity, 
i.e. reduce the side slip angle. This is known as directional stability, 
thus vertical stabilizers 34, 36 may be referred to as directional 
stabilizers. The resulting moment about lateral axis 66 tends to rotate 
wings 18 and 22 such that the downwind wing moves downward. This is 
referred to as lateral stability. 
Referring back to FIG. 2, right net normal force 70 is the total force 
acting on right vertical stabilizer 34. Left net normal force 72 is the 
total force acting on left vertical stabilizer 36. Net normal forces 70 
and 72 are the total forces acting on vertical stabilizers 34 and 36 due 
to the incidence of the aircraft and the side slip angle. Since a downwind 
panel may be partially blocked by fuselage 12, any loss in efficiency of 
the downwind surface is actually beneficial to the directional stability 
of aircraft 10. 
To provide some of the induced benefits of a biplane configuration and to 
insure adequate surface area of vertical stabilizers 34 and 36, the ratio 
of vertical separation distance 42 to wing span 46 cannot be too small. A 
minimum value of 0.15 is recommended. In the preferred embodiment, the 
ratio of vertical separation distance to wing span ranges from values of 
0.2 to 0.25. Although there is no upper limit on the height to span ratio, 
design considerations will typically limit the height to span ratio to no 
more than 0.5 to 0.6. Additionally, although both wings 18 and 22 and 
horizontal stabilizers 26 and 28 are shown flat, i.e. without geometric 
dihedral, such dihedral (gull wing up or down) in either surface is 
permissible. If wings 18 and 22 or horizontal stabilizers 26 and 28 
possess geometric dihedral, then vertical separation 42 should be defined 
by the average vertical distance between wings 18 and 22 and horizontal 
stabilizers 26 and 28. Note that changes in the values of the height span 
parameter also impact the inclination of vertical stabilizers 34 and 36. 
In FIG. 2, geometric dihedral angle 74 is approximately 52 degrees. 
However, geometric dihedral angle 74 may be 90 degrees wherein vertical 
stabilizers 34 and 36 are vertical. Preferably, geometric dihedral angle 
74 should range from 45 to 90 degrees. Preferably, vertical stabilizers 34 
and 36 should be vertical or canted outboard as shown. An inboard cant is 
not preferred since an inboard cant of vertical stabilizers 34 and 36 
results in a generation of a destabilizing lateral moment with side slip. 
When right vertical stabilizer 34 and left vertical stabilizer 36 are 
positioned vertically, i.e. where geometric dihedral angle 74 is 90 
degrees, little or no lateral moment will be generated. When geometric 
dihedral angle 74 is 45 degrees, horizontal stabilizer span 44 is one half 
of wing span 46. 
Referring now to FIG. 3, shown is a side view of aircraft 10. Left wing 20 
is shown positioned proximate fuselage top 14. Left horizontal stabilizer 
28 is shown positioned proximate fuselage bottom 16 and rearward of left 
wing 20. Aerodynamic center or center of lift 76 of left vertical 
stabilizer panel 36 is shown positioned behind center of gravity 68 of 
aircraft 10. By positioning center of lift 76 behind center of gravity 68, 
incremental forces on left vertical stabilizer 36 due to side slip will 
create a stabilizing directional moment. 
The ratio of horizontal stabilizer tip cord 78 to wing tip cord 80 should 
be equal to or greater than one. By providing a horizontal stabilizer tip 
cord to wing tip cord ratio of greater than one, forces acting on vertical 
stabilizer 36 will stabilize rather than destabilize aircraft 10. However, 
in an extreme case where both wing 20 and horizontal stabilizer 36 are 
mounted well aft such that aerodynamic center 81 of wing 20 is near or 
behind center of gravity 68 of the aircraft, this geometric requirement 
can be waived. It has been found that it is important to keep forces 
generated on vertical stabilizers 34 and 36 behind center of gravity 68 of 
aircraft 10. 
Referring now to FIG. 4, a graphical representation of the effectiveness of 
a conventional center line vertical tail for providing directional 
stability is compared to the effectiveness of providing directional 
stability by means of twin vertical stabilizer panels of the present 
invention. Flight path axis directional stability is plotted on the y-axis 
versus angle-of-attack, which is plotted on the x-axis. It can be seen 
from the graph that the effectiveness of the conventional surface drops 
rapidly at angles-of-attack greater than 25 degrees. Above 35 degrees, the 
vertical stabilizer actually destabilizes the vehicle. The data for the 
proposed invention show that wing mounted vertical stabilizers 34 and 36 
of the invention become more effective as angle-of-attack increases and 
remain effective to the highest angle-of-attack shown in the data. The 
results shown are normalized to the value of stabilizing moment produced 
by each configuration at zero degrees angle-of-attack. The absolute value 
is different between the two configurations because of differences in 
surface size and geometry. 
The vertical stabilizers act to provide directional stability along the 
flight path and are much more effective at high angles-of-attack than 
conventional single or multiple vertical stabilizers. This arrangement of 
the invention achieves these results by providing vertical stabilizers 
that are less strongly influenced by fuselage blockage at high 
angles-of-attack and providing a component geometry in which the effects 
of fuselage blockage actually improve the overall effectiveness of the 
vertical stabilizers instead of decreasing the effectiveness of the 
vertical stabilizers. Additionally, stabilizing moments are provided in 
both the body axis directional and lateral axes, both of which contribute 
to flight path axis directional stability at high angles-of-attack. 
Finally, by providing vertical stabilizers having a forward sweep, wherein 
the forward sweep increases a component of flow velocity normal to the 
leading edge of the vertical stabilizer, the aerodynamic efficiency of the 
vertical stabilizers improves as angle-of-attack increases. 
The invention provides directional stability by passive means without 
control augmentation and is suitable for application to new designs of any 
performance level ranging from slow speed, highly maneuverable aerobatic 
aircraft to supersonic fighters. 
The vertical separation between the wing and horizontal tail and the 
joining of the same by the vertical stabilizers produces some effects of a 
biplane with reductions in induced drag. Additionally, the closed 
structure may also be made structurally more efficient than a conventional 
structure having cantilevered lifting surfaces. Although only the 
preferred embodiment of arrangements for carrying out this invention have 
been described above, it is not to be construed that the invention is 
limited to such embodiments. Other modifications may be made by those 
skilled in the art without departing from the spirit and scope of the 
invention defined below. No attempt has been made to incorporate any other 
such modifications or forms in this disclosure in the interests of 
clarity.