U.S. patent number RE35,387 [Application Number 07/828,219] was granted by the patent office on 1996-12-03 for superfragile tactical fighter aircraft and method of flying it in supernormal flight.
This patent grant is currently assigned to Dynamic Engineering, Inc.. Invention is credited to Thomas H. Strom.
United States Patent |
RE35,387 |
Strom |
December 3, 1996 |
**Please see images for:
( Certificate of Correction ) ** |
Superfragile tactical fighter aircraft and method of flying it in
supernormal flight
Abstract
A superagile tactical fighter aircraft and a method of flying it
are disclosed. The superagile aircraft is characterized by
articulatable air inlets, articulatable exhaust nozzles, highly
deflectable canard surfaces, and control thruster jets located
around the nose of the fuselage, on the top and bottom surfaces of
the propulsion system near the exhaust nozzles, and on both sides
of at least one vertical tail. The method of operating the
superagile aircraft comprises the step of articulating the air
inlets and exhaust nozzles, deflecting the canard surfaces, and
vectoring the thruster jets so that supernormal flight is attained.
Supernormal flight may be defined as flight at which the superagile
aircraft operates at an angle of attack much greater than the angle
of attack which produces maximum lift. In supernormal flight, the
superagile aircraft is capable of almost vertical ascents, sharp
turns, and very steep descents without losing control. Enhanced
survivability of the pilot and the aircraft is achieved by the
invention.
Inventors: |
Strom; Thomas H. (Hampton,
VA) |
Assignee: |
Dynamic Engineering, Inc.
(Newport News, VA)
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Family
ID: |
46248724 |
Appl.
No.: |
07/828,219 |
Filed: |
January 30, 1992 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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721664 |
Apr 9, 1985 |
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Reissue of: |
194734 |
Sep 15, 1987 |
04896846 |
Jan 30, 1990 |
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Current U.S.
Class: |
244/75.1;
244/12.5; 244/45A; 244/52; 244/53B |
Current CPC
Class: |
B64C
5/04 (20130101); B64C 5/16 (20130101); B64C
13/00 (20130101); B64C 15/14 (20130101); B64C
39/12 (20130101); F02K 1/002 (20130101) |
Current International
Class: |
B64C
5/00 (20060101); B64C 13/00 (20060101); B64C
15/00 (20060101); B64C 39/12 (20060101); B64C
15/14 (20060101); B64C 5/04 (20060101); B64C
5/16 (20060101); B64C 39/00 (20060101); F02K
1/00 (20060101); B64C 009/00 (); B64C 015/00 () |
Field of
Search: |
;244/45R,45A,12.5,23D,214,53B,52,7C,7B |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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2421524 |
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Nov 1974 |
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DE |
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2327612 |
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Dec 1974 |
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DE |
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2856033 |
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Jun 1980 |
|
DE |
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2833771 |
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Sep 1982 |
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DE |
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3009340 |
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Sep 1983 |
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DE |
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522296 |
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Jun 1940 |
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GB |
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959405 |
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Jun 1964 |
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GB |
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128344 |
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Jul 1972 |
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GB |
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2097863 |
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Nov 1982 |
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GB |
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Other References
SAE Technical Paper Series, No. 821469, "Flight at Supernormal
Attitudes" by Thomas H. Strom and William J. Alford, Jr., for
Aerospace Congress & Exposition, Oct. 25-28, 1982 Anaheim,
California. .
J. A. Laughrey, D. J. Drape and P. E. Hiley; Performance Evaluation
of an Air Vehicle Utilizing Nonaxisymmetric Nozzles; Feb. 1981; pp.
89-95 vol. 18, Journal of Aircraft. .
Herbst, "10 Jahre TKF/J--90 Vorentwicklung", 17-19 Oct. 1983, pp.
15-23, New Projects & Planning in the Air Force,
Munich/Germany. .
Scott, "NASA Researches Aircraft Control During Deep Stall", 31
Oct. 1983, pp. 53 & 54, Aviation Week & Space Technology.
.
Thomas A. Horne, "Breaking the Stall Barrier", May 1984, AOPA
Pilot. .
European Patent Office Technical Board of Appeal, "In re European
Pat. Publ. No. 45987 of Boeing Co.", 10 Oct. 1985, pp. 121-128,
Journal of EPO. .
Kraus et al., "High Angle of Attack Characteristics of Different
Fighter Configurations", 4 Oct. 1978, AGARD Conference,
Neuilly-sur-Seine/France. .
W. B. Herbst, "Future Fighter Technologies", Aug. 1980, pp.
561-566, vol. 17, Journal of Aircraft. .
Herbst, "Design for Air Combat", Oct. 1981, AGARD FMP Conference,
Florence/Italy. .
Kraus et al., "Stability and Control for High Angle of Attack
Maneuvering", 19 Apr. 1982, AGARD Conference,
Neuilly-sur-Seine/France. .
Strom et al., "Controllable supernormal flight: a future
possibility", Dec. 1982, pp. 48-53, vol. 90, for Automotive
Engineering. .
Moore et al., "X-29 Forward Swept Wing Aerodynamic Overview", 13-15
Jul. 1983, AIAA Applied Aerodynamics Conference,
Danvers/Massachusetts pp. 1-8. .
Herbst, "Supermaneuverability", 17-19 Oct. 1983, Jahrbuch 1983 I,
Munich/Germany..
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Primary Examiner: Barefoot; Galen L.
Attorney, Agent or Firm: Scafetta, Jr.; Joseph
Parent Case Text
CROSS-REFERENCE TO RELATED APPLICATION
This application is a continuation-in-part of U.S. patent
application Ser. No. 721,664, filed Apr. 9, 1985, abandoned
concurrently with the filing of this application.
Claims
What I claim as my invention is:
1. A superagile tactical fighter aircraft comprising:
a fuselage having a nose, a midsection, an aft section and at least
one vertical tail, said fuselage adapted to house a human pilot, a
payload, fuel, automatic flight control systems, a navigation
system, and a life support system to assist and sustain the
pilot;
fixed wings mounted on the aft section of the fuselage, behind a
center-of-gravity of the aircraft;
a high thrust-to-weight propulsion system being mounted to the
wings;
fully articulating air inlets at a front end of the propulsion
system, said air inlets being deflectable so as to face into a
local air stream so as to provide minimally distorted air flow to
the propulsion system throughout a complete flight regime but
particularly at very high angles of attack and low air speeds;
fully articulating exhaust nozzles at a rear end of the propulsion
system, said exhaust nozzles being rapidly deflectable so as to
allow the pilot the capability to vector and direct gross thrust
produced by the propulsion system; and
rotatable canard surfacce means, mounted on the midsection on the
fuselage in front of the center-of-gravity and separate from the
wings, for fully and rapidly deflecting air flow thereacross in an
angle-of-attack range with trailing edge about 90.degree. up to
trailing edge about 45.degree. down, said canard surface means
being located so as to provide pitch control forces and moments and
to cause the aircraft to be longitudinally unstable and capable of
trimming at very high angles of attack approaching the range from
70.degree. to 90.degree..
2. The aircraft according to claim 1, further comprising:
thruster jet means, or any other acceptable force means, arranged
around the nose of the fuselage, on the propulsion system near the
exhaust nozzle, and on at least one vertical tail for vectoring
thrust.
3. A method of flying a human-piloted longitudinally unstable
fixed-wing tactical fighter aircraft comprising the steps of:
initiating an aerial maneuver in a vertical or pitch plane to very
high angles of attack and pitch approaching a range from 70.degree.
to 90.degree. by:
highly deflecting rotatable canard surfaces on the aircraft to
provide an angle-of-attack greater than an angle-of-attack which
produces maximum lift;
increasing thrust of a propulsion system to a maximum
capability;
vectoring gross thrust of the propulsion system by deflecting
articulating exhaust nozzles to angles which augment turning
capability and assure control of the aircraft at low
velocities;
deflecting articulating air inlets at the appropriate angle to face
into a local air stream so as to provide minimally distorted air
flow to the propulsion system; banking the aircraft to angles
approaching 90.degree.; decelerating the aircraft to a velocity
less than stalling velocity;
turning the heading angle of the aircraft; and pointing the
aircraft in any direction necessary to be aimed at an opposing or
threatening aircraft; whereby the aircraft has, by developing very
high turn rates and decelerations, effectively become superagile in
supernormal flight.
4. The method according to claim 3, further comprising the step
of:
redeflecting the rotatable canard surfaces, articulating air
inlets, and articulating exhaust nozzles to positions so as to
rapidly return the aircraft to an angle-of-attack below the
angle-of-attack which produces maximum lift; whereby the aircraft
returns in a controlled manner to an unaccelerated, trimmed normal
flight condition.
5. The method according to claim 3, further comprising the steps
of:
vectoring thrust from control thruster jets, or any other
acceptable force means, located around a nose, near exhaust
nozzles, and on at least one vertical tail of the aircraft, in
directions to provide acceptable and necessary control and trim
about roll, yaw, and pitch axes of the aircraft at low
velocities.
6. The method according to claim 3, further comprising the steps
of:
deflecting further the rotatable canard surfaces rapidly to attain
an angle-of-attack sufficiently higher than the angle-of-attack
which produces maximum lift so that a controllable stable and
trimmable nose-high aircraft pitch attitude is attained;
whereby favorable ejection attitude and enhanced survivability for
a pilot is achieved in the event it appears that a hard contact of
the superagile aircraft with ground is unavoidable.
7. The method according to claim 3, wherein:
said angle at which the gross thrust is vectored by the
articulating exhaust nozzles is appropriate to provide necessary
longitudinal and trim forces and moments at very low aircraft
velocities.
8. The method according to claim 5, further comprising the steps
of:
rapidly actuating the highly deflectable canard surfaces; providing
roll, yaw, and pitch control by aerodynamic control surfaces;
rapidly actuating the thrust vectoring exhaust nozzles; and rapidly
actuating the thruster jets;
whereby the superagile aircraft is prevented from entering into
flight conditions under which it will begin spinning motions.
9. The method according to claim 5, further comprising the steps
of:
rapidly actuating the highly deflectable canard surfaces;
providing roll, yaw, and pitch control by aerodynamic control
surfaces;
rapidly actuating the thrust vectoring exhaust nozzles; and rapidly
actuating the thruster jets;
whereby, in the event of inadvertent spinning motions, any
equilibrium between aerodynamic and centrifugal forces that exist
are eliminated in said spinning motions, thereby allowing rapid
recovery from the spinning motions. .Iadd.
10. A superagile tactical fighter aircraft comprising:
a fuselage having a nose, a midsection, an aft section and at least
one vertical tail, said fuselage adapted to house a human pilot, a
payload, fuel, automatic flight control systems, a navigation
system, and a life support system to assist and sustain the
pilot;
fixed wings mounted on the aft section of the fuselage, behind a
center-of-gravity of the aircraft;
a high thrust-to-weight propulsion system being mounted to the
aircraft;
fully articulating exhaust nozzles at a rear end of the propulsion
system, said exhaust nozzles being rapidly deflectable so as to
allow the pilot the capability to vector and direct gross thrust
produced by the propulsion system; and
rotatable canard surface means, mounted on the midsection on the
fuselage in front of the center-of-gravity and separate from the
wings, for fully and rapidly deflecting air flow thereacross in an
angle-of-attack range with trailing edge about 90.degree. up to
trailing edge about 45.degree. down, said canard surface means
being located so as to provide pitch control forces and moments and
to cause the aircraft to be longitudinally unstable and capable of
trimming at very high angles of attack approaching the range from
70.degree. to 90.degree.. .Iaddend..Iadd.11. A method of flying a
human-piloted longitudinally unstable fixed-wing tactical fighter
aircraft comprising the steps of:
initiating an aerial maneuver in a vertical or pitch plane to very
high angles of attack and pitch approaching a range from 70.degree.
to 90.degree. by:
highly deflecting rotatable canard surfaces on the aircraft to
provide an angle-of-attack greater than an angle-of-attack which
produces maximum lift;
increasing thrust of a propulsion system to a maximum
capability;
vectoring gross thrust of the propulsion system by deflecting
articulating exhaust nozzles to angles which augment turning
capability and assure control of the aircraft at low
velocities;
banking the aircraft to angles approaching 90.degree.; decelerating
the aircraft to a velocity less than stalling velocity;
turning the heading angle of the aircraft; and pointing the
aircraft in any direction necessary to be aimed at an opposing or
threatening aircraft; whereby the aircraft has, by developing very
high turn rates and decelerations, effectively become superagile in
supernormal flight. .Iaddend.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
The invention relates to the field of aeronautics, in particular to
aircraft manuevering and control devices of a fixed wing aircraft
with a human pilot assisted by an automatic flight control
system.
2. Brief Description of the Related Art
Until about 1978, the region beyond stall was considered an
unacceptable flight regime frequently characterized by
uncontrollable flight in spins and by undesirable deep stalls. Any
deep stall condition is characterized by a stable trimmed flight
but at a high angle of attack from which return to normal flight
may be difficult or impossible. A deep stall may be defined as an
out-of-control condition at an angle of attack greater than the
angle of attack for maximum lift with no significant motion other
than a high rate of descent. Conventional airplanes usually stall
and lose control effectiveness at angles of attack in the range of
18.degree. to 20.degree..
However, according to U.S. Pat. Nos. 4,261,533 and 4,099,687, it is
now possible, through the use of a rotatable horizontal tail on
aft-tail configurations or through the use of tiltable engines on
the wings, to provide stable and controllable flight at extremely
high airplane angles of attack.
Because movement other than a high rate of descent can be
controlled by varying thrust levels and all moveable control
surfaces with large deflections, the safety and usefulness of
flight at extremely high angles of attack are being re-examined and
redefined.
The essence of the longitudinal control concept, as set forth in
U.S. Pat. Nos. 4,261,533 and 4,099,687, is to rotate the tail or to
deflect large chord elevons to magnitudes of approximately the same
order, but of opposite direction, as the airplane angle of attack,
so that the effective tail aerodynamic angle of attack is below the
tail stall angle and is thus capable of providing both stability
and control for the entire aircraft.
Although rotatable canard arrangements are known from U.S. Pat. No.
4,569,493, U.S. Pat. No. 4,281,810, U.S. Pat. No. 4,010,920, and
West German Offenlegungsschrift 2421524, such arrangements deal
strictly with the stability and control of aircraft and models in
level and unstalled low angle of attack regions of flight and do
not address the problems of stability and control of aircraft in
the high angle of attack regions of flight.
In most cases, the upper limit of normal flight is associated with
conditions for maximum lift, beyond which the wing is completely
stalled. For some aircraft configurations, however, for example,
those employing wings with high leading edge sweep angles or
incorporating strakes, i.e. a continuous band of plates on the
fuselage, partial flow separation of the wing or control surfaces
may induce stability problems below the attack angles for maximum
lift and impose lower limits on the normal flight regions.
Solutions of these problems will allow flight above these lower
limits. Flight above the normal limits is considered of a
supernormal nature.
SUMMARY OF THE INVENTION
The present invention is directed to a fixed wing aircraft with
canard control surfaces and a method of human-piloted operation
thereof permitting supernormal flight which is concerned with
flight at extraordinary angles of attack, resulting in substantial
changes in the pitch and the flight path angles and also resulting
in the attainment of flight paths and vertical velocities which are
not normally attainable.
In piloted supernormal flight of the aircraft of the present
invention, the wing of an aircraft, such as a superagile tactical
fighter, is either partially or completely stalled, while the
longitudinal control surfaces, such as in a rotatable canard
arrangement, are deflected to approximately the same magnitude, but
of opposite sign, as the angle of attack of the aircraft, so that
the canard arrangement remains effective to control the aircraft
through large ranges of angles of attack, pitch, and flight path.
Such angles may vary from descending flight to deep stall, i.e.
-45.degree., to ascending flight in vertical climb, i.e.
+90.degree..
Also, fully articulating air inlets at a front end of the
propulsion system and fully articulating exhaust nozzles at a rear
end of the propulsion system are articulated to appropriate
directions relative to the air flow so that the inlet operates
effectively throughout the large angle of attack range and also so
that thrust is vectored in the desired direction by the exhaust
nozzle deflections.
Thrust may likewise be vectored by small thruster jets or other
similar devices arranged around the nose of the fuselage, on the
propulsion system near the exhaust nozzles, and on at least one
vertical tail to provide control forces and moments at low speeds
where the aerodynamic control surfaces tend to lose
effectiveness.
The advantages of such supernormal flight include: improved safety
through prevention of spins; steep descents and approaches to
landings; precise, steep surivable recoveries of remotely piloted
vehicles; and enhanced high angle of attack control manueverability
and agility.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1A shows a front elevational view of the superagile tactical
fighter aircraft of the present invention in level flight.
FIG. 1B shows a top plan view of the superagile tactical fighter in
level flight.
FIG. 1C shows a side elevational view of the superagile tactical
fighter in descending flight.
FIG. 2A shows a diagonal top plan view of the superagile tactical
fighter at the center of a plane having horizontal (X) and lateral
(Y) coordinates.
FIG. 2B shows a diagonal front elevational view of the superagile
tactical fighter at the center of a plane having lateral (Y) and
vertical (Z) coordinates.
FIG. 2C shows a diagonal side elevational view of the superagile
tactical fighter at the center of a plane having horizontal (X) and
vertical (Z) coordinates.
FIG. 3 is a graph showing the horizontal turn characteristics of
the superagile tactical fighter in the X-Y plane.
FIG. 4 is a graph showing in more detail the contributions of lift
and thrust to the turning characteristics of the superagile
tactical fighter in the X-Y plane.
FIG. 5 is a schematic perspective view illustrating the turning
performance of the superagile tactical fighter in aerial combat
against a conventional fighter jet aircraft.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
As illustrated in FIGS. 1A through 1C, a superagile fixed wing
tactical fighter aircraft, suitable for flying according to the
supernormal method of the present invention, comprises a fuselage
11 to enclose a human pilot, a payload, and fuel, as well as
automatic flight control, navigation, and life support systems to
assist the pilot. The fuselage 11 has a nose, a midsection, and an
aft section. A jet engine propulsion system 12 has fully
articulating air inlets 13 at a front end for supplying a
continuous flow of minimally distorted air flow, thereby
maintaining efficient engine operation of the high thrust-to-weight
propulsion system 12 over the entire flight regime, but
particularly at high angles of attack and low speeds. Also, fully
articulating exhaust nozzles 14 at a rear end are provided for
modulating and directing thrust forces produced by the propulsion
system 12. Although a twin engine propulsion system 12 is shown, a
single engine or another multi-engine system may be used instead. A
swept or modified delta fixed wing 15 is mounted on the aft section
of the fuselage 11 and is equipped with flaperons 16, elevons 17,
and spoilers 18. The wing 15 has the propulsion system 12 mounted
thereunder and this fixed wing 15 is the primary surface for
producing the aerodynamic resultant force characteristics of the
aircraft.
Mutually perpendicular reference planes are defined by the three
intersecting axes: longitudinal or horizontal X'--X', lateral
Y'--Y', and vertical Z'--Z'.
As shown in all FIGS. 1A--1C, twin secondary control
force-generating canard surfaces 19, mounted on the midsection of
the fuselage 11, are separated from and located forward of the wing
15 to provide destabilizing and controlling forces about the
lateral Y'--Y' axis. These canard surfaces 19 are symmetrically
located about the longitudinal axis X'--X' and are pivotable or
rotatable about a lateral axis Y.sub.c '--Y.sub.c ', shown only in
FIG. 1B, to any canard control deflection angle .delta..sub.c, as
shown in FIG. IC, in a range from about +45.degree. (trailing edge
down but not shown) to about -90.degree. (trailing edge up shown in
FIG. 1C) relative to the longitudinal axis X'--X'. This large
canard control deflection range, i.e.
+45.degree..gtoreq..delta..sub.c .gtoreq.-90.degree., allows the
canard surfaces 19 to be nearly aligned with the direction of local
air flow over the aircraft so that the canard surfaces 19 remain
unstalled. In order for an aircraft that employs canard control
surfaces to be capable of trimmable and controllable flight at very
high angles of attack, the aerodynamic surfaces on the aircraft
(i.e. wing canards) and the distribution of the mass of its various
components and systems must be arranged relative to the center of
gravity so that the aircraft is longitudinally unstable, i.e. the
slope of its pitching moment as a function of angle of attack must
be positive.
Vertical tails 20 provide directional stability and damping.
Rudders 21 on each tail 20 provide yaw control and augment roll
control obtained by the flaperons 16, elevons 17, and spoilers 18
on the wing 15.
As best shown in FIG. 1B, control thruster jets 22 are located near
the nose of the fuselage 11, near the exhaust nozzles 14 of the
propulsion system 12, and on the vertical tails 20. These control
thruster jets 22 may be powered by an auxiliary power system (not
shown) or by bleed from the propulsion system 12. The control
thruster jets 22 provide control of the aircraft at velocities
approaching zero.
Although not shown, an automatic flight control system including
conventional programmable, pilot interactable, automatic avionic
sensors, computers, effectors, and actuators inside the fuselage 11
help to provide rapid control and stability of the aircraft about
the mutually perpendicular axes X'--X', Y'--Y' and Z'--Z' of FIGS.
1A-1C.
Utilization in several combinations of the propulsion system 12,
fixed wing 15, canard surfaces 19, vertical tails 20, rudders 21,
and the various stability and control augmentation systems by the
human pilot allows the aircraft to operate in flight at angles of
attack significantly greater than those associated with maximum
lift. For this reason, the aforementioned elements and systems of
the present invention may be defined as supernormal flight
controls.
The nomenclature, symbols, and equations of motion applicable to
the superagile tactical fighter aircraft, when flying according to
the supernormal method of the present invention, are represented
immediately hereinbelow: ##EQU1## Where: ##EQU2##
D, Drag=C.sub.D qS
g, Acceleration due to gravity
L, Lift
S, wing area
T, Thrust
t, Time
V, Flight path velocity
V, Flight path acceleration ##EQU3##
X, Y, Z, Flat earth axes
X, Y, Z, Velocity components along flat earth axes
W, Weight
q=1/2 .rho.V.sub.2, dynamic pressure
.rho., density of air
.alpha., angle of attack
.gamma., flight path angle
.gamma., flight path angle velocity
.theta., pitch attitute angle
.PHI., bank angle
.psi.heading angle
.psi., heading angle velocity (turn rate)
For purposes of these calculations, the superagile tactical fighter
aircraft is represented as a point mass, without sideslip,
operating over a flat, nonrotating earth.
The thrust T, lift L, drag D, velocity V, and various angular
attitudes associated with and acting upon the aircraft in piloted
supernormal flight are best illustrated in FIG. 2C. In such
supernormal flight, the aircraft operates at an attack angle
.alpha. much greater than the angle of attack for maximum lift so
that the fixed wing 15 is either completely or partially stalled
while the canard surfaces 19 are deflected in a negative sense
through the deflection angle-.delta..sub.c. The absolute deflection
magnitude of the canard surfaces 19 is approximately the same as
the attack angle .alpha. for the entire aircraft so that such
canard surfaces 19 are nearly aligned with the local air flow and
are, therefore, unstalled. Thus, the canard surfaces 19 remain
effective as lift surfaces in providing the required forces and
moments for controlling the entire aircraft.
The velocity components of the aircraft along the X, Y and Z axes
are related to the aircraft flight path velocity V, flight path
angle .gamma., and heading angle .psi., according to equations 1, 2
and 3 given above.
The flight path acceleration V of the aircraft is related to the
thrust T, drag D, weight W, attack angle .alpha., and flight path
angle .gamma., as stated by equation 4.
The flight path angle velocity .gamma. of the aircraft is related
to the thrust T, lift L, flight path velocity V, weight W, attack
angle .alpha., bank angle 101 , and flight path angle .gamma., as
stated by equation 5.
The heading angle velocity .psi. or turn rate of the aircraft is
related to the thrust T, lift L, flight path velocity V, weight W,
attack angle .alpha., bank angle .PHI., and flight path angle
.gamma., according to equation 6.
As can be seen from equation 7, the pitch attitude angle .theta. of
the aircraft is the sum of the attack angle .alpha. and the flight
path angle .gamma..
The attributes characterizing the superagility of the tactical
fighter aircraft are quantified by the magnitudes of the velocity
V, acceleration V, flight path angle velocity .gamma., and heading
angle velocity .psi., in particular, and by the rapidity with which
these attributes can be changed. A graphic illustration of the
superagility, particularly the turning characteristics, of the
tactical fighter aircraft in the horizontal-lateral plane X-Y is
presented in FIG. 3. The heading angle velocity .psi. or turn rate
of the aircraft is a function of the Mach number M of the aircraft
at an altitude of 20,000 feet and with a wing loading W/S of 68
pounds per square foot. The aircraft thrust-to-weight ratio T/W
varies from a value of 0.5 at M=0 to 0.9 at M=0.9, as shown in the
inset table in the upper right-hand corner of FIG. 3.
The region in which normal flight occurs is bounded by curves
associated with the maximum lift coefficient C.sub.Lmax for a
conventional aircraft and with the maximum structural load factor
.eta. for the aircraft. The Mach number M and the turn rate .psi.
existing at the point where these curves intersect are referred to
as "corner" conditions. For a conventional fighter aircraft, not
equipped with the elements of the present invention, the turn rate
.psi. at the corner condition is the maximum instantaneous turn
rate. For the specific example shown in FIG. 3, the maximum
instantaneous turn rate for the conventional aircraft in normal
flight is about 23.degree. per second and occurs when the Mach
number M is approximately 0.7. The turning radius R associated with
these corner conditions for the conventional aircraft is about 1900
feet.
The region inn which supernormal flight occurs in FIG. 3 is bounded
by the curve associated with the maximum structural load factor
.eta. as an upper limit and by the line associated with the turn
rate .psi. at the corner condition as a lower limit. The change in
the heading angle .psi. with the change in time is described by
equation 6 above and is defined as the heading angle velocity or
turn rate .psi.. For example, if the angle-of-attack .alpha. is
70.degree. and the bank angle .PHI. is 90.degree., the change in
the turn rate .psi. or the instantaneous turn rate variation is
illustrated by the dashed curve. The superagile tactical fighter of
the present invention, when flying at this angle-of-attack .alpha.
of 70.degree., would exceed the normally maximum corner turn rate
.psi. at a Math number M of less than about 0.18 as the flight path
angle .gamma. reaches about 65.degree.. At a flight path angle
.gamma. of 78.degree., the turn rate .psi. approaches 50.degree.
per second and the turning radius R is less than 125 feet. For
reference purposes, there is also shown the sustained turn rate
characteristics for the aircraft when the thrust-minus-drag is
zero, i.e. T--D=0. In other words, the specific excess power is
zero, i.e. P.sub.s =0.
To better describe the physical aspects of supernormal flight,
additional characteristics of a turn by a superagile tactical
fighter aircraft are presented in FIG. 4. Basically, the data in
FIG. 4 are the same data presented in FIG. 3 with the addition of
the turn rates .psi. associated with the thrust T and the lift L.
Equation 6 above shows the relationship of the heading angle
velocity or turn rate .psi. to both the thrust T and the lift L.
For the particular flight conditions illustrated in FIG. 4, the
thrust-related term in equation 6 predominates when the Mach number
M is below 0.28 but the lift-related term predominates when the
Mach number M is above 0.28.
A pictorial turning performance illustration of the characteristics
shown in FIGS. 3 and 4 is presented in FIG. 5 where an elementary
one-on-one combat maneuver between the superagile tactical fighter
of the present invention and a conventional fighter jet which does
not have features for supernormal flight is shown. Upon positive
identification, the opponents have the options of firing a long or
medium range missile and or beginning to maneuver at their maximum
capability to evade the incoming missile while simultaneously
trying to position themselves for close-in combat. The conventional
fighter jet is limited to flying at its corner Mach number or
velocity because it does not have the capability of supernormal
flight at an angle of attack beyond that permitted by the maximum
lift coefficient. On the other hand, the superagile tactical
fighter aircraft of the present invention begins its maneuvering by
increasing thrust, preferably to a maximum. Next the pilot will
climb by increasing the angle of attack of the aircraft to a very
high level, say 70.degree., and then progressively increasing its
flight path angle .gamma. and bank angle .PHI. to values
approaching 90.degree..
Thus, as shown pictorially in FIG. 5 and graphically in FIGS. 3 and
4, the superagile tactical fighter aircraft in supernormal flight
accomplishes a decelerating, steep climb with a turn rate .psi.
over twice that of the conventional fighter jet attempting to
operate at its corner Mach number M or velocity V. The turning
maneuver in supernormal flight is analogous to a skidding turn of a
powerful decelerating wheeled vehicle. This turning maneuver allows
the superagile tactical fighter aircraft of the present invention
to turn tightly inside the flight path of the opposing conventional
fighter jet and to launch a "supershot" short range missile before
the opponent can turn and get into a firing position, thereby
scoring an aerial victory.
Although not illustrated herein, it is certain that innovative
aerial combat tacticians will develop other multi-aircraft
encounter techniques which will use the features of the present
invention in order to achieve greater advantages for the superagile
tactical fighter aircraft.
The method of the present invention relates to flying and
controlling a superagile tactical fighter aircraft employing highly
deflectable canard surfaces 19 and having control thruster jets 22
on the nose of the fuselage 11, near the exhaust nozzles 14 of the
propulsion system 12, and on both sides of vertical tails 20 so
that the aircraft may engage in supernormal flight in order to
provide extraordinarily agile maneuverability characteristics or
"superagility", relative to the maneuverability characteristics of
a conventional fighter jet not capable of the present inventive
method of flying because it lacks the aerodynamic structures of the
present invention. Thus, such a conventional jet is restricted to
flying, at best, at its corner Mach number M or velocity V, i.e. at
the maximum permissible limit of normal flight at the highest angle
of attack associated with maximum lift.
For example, the method of flying the superagile tactical fighter
aircraft of the present invention in supernormal flight may
comprise several steps. First, the superagile aircraft initiates an
aerial maneuver or responds to the initiation of an aerial maneuver
by an opposing fighter jet of essentially equal technological
development except that the latter does not have the supernormal
control system of the present invention and therefore is restricted
to maneuvers associated with angles of attack less than or equal to
the angle of attack at which maximum lift occurs. The opposing
fighter jet is also limited to a maximum turn rate .psi. at the
corner Mach number M or velocity V where the turn rate .psi.
provided by maximum lift and the turn rate .psi. allowed by the
ultimate structural load factor .eta. the conventional fighter jet
are equal.
According to FIG. 4, a representative value of a maximum turn rate
.psi..sub.max for a conventional fighter jet not equipped with the
present invention but flying in normal flight at 20,000 feet
altitude is about 22.degree. to 25.degree. per second and the
corresponding corner Mach number M is about 0.7, i.e. a velocity V
of about 725 feet per second at such altitude.
The second step of the inventive method of flying the superagile
tactical fighter aircraft is that, upon sighting the opposing
conventional fighter jet either visually or by electronic systems,
the superagile aircraft increases its thrust to a maximum level and
uses coordinated deflections of its canard surfaces 19, vectoring
of its control thruster jets 22, and articulating of its air inlets
13 and of its exhaust nozzles 14 in order to increase the
angle-of-attack .alpha. of the superagile aircraft to about
35.degree. so as to effect a large increase in the flight path
angle .gamma. and in the rate of climb.
As the superagile aircraft begins to decelerate because of the
large increase in induced drag associated with the high
angle-of-attack .alpha., the angle-of-attack .alpha. is further
increased to values in the range of 60.degree. to 70.degree. and
the superagile aircraft is banked at a high angle .PHI. approaching
90.degree.. This high banking maneuver allows the superagile
aircraft to decelerate further to a very low velocity so that it
can turn and redirect itself rapidly. Since the heading angle
velocity .psi. or turn rate of the superagile aircraft is
determined primarily at low speed below M=0.28 by the
thrust-dependent term T/WV sin .alpha. in equation 6 because the
velocity V is in the denominator, the ability of the superagile
aircraft to turn rapidly and redirect itself allows such superagile
aircraft to gain a favorable position for weapon firing
opportunities against the opposing conventional fighter jet.
One advantage of the present invention relates to the stresses
endured by the pilot when the superagile aircraft is decelerated
along the steep flight path shown in FIG. 5. Deceleration inertial
forces produced by thrust vectoring and aerodynamic drag associated
with high angles-of-attack are much larger than the deceleration
forces associated with the angle-of-attack which produces maximum
lift. The deceleration inertial forces cause the pilot of the
superagile aircraft to experience "eyeballs-down" stresses that are
more endurable than "hang-in-the-belt", "eyeball-out" stresses that
are nornally associated with the deceleration of a conventional
fighter jet which is constrained to fly at or below the
angle-of-attack which produces maximum lift.
The method of the present invention further involves the flying and
controlling of the superagile tactical fighter aircraft by
redeflecting the rotatable twin canard surfaces 19 to an
angle-of-attack below the angle-of-attack which produces maximum
lift in order to provide extraordinary unaccelerated, or return to
unaccelerated, trimmed flight conditions.
The inventive method of supernormal flight comprises the further
steps of applying, modulating, and vectoring thrust, controllably
orienting the superagile aircraft at a pitch attitude in the range
of 0.degree. to about 90.degree., and causing the aircraft to
descend steeply and rapidly in altitude, as shown in FIG. 5, while
stability and control are maintained. Eventually, the pilot levels
the aircraft out so that it returns to normal unaccelerated trimmed
flight conditions.
Further steps of the present inventive method involve controlling
roll, yaw, and pitch by deflecting the aerodynamic controls 16, 17,
18, 19 and 21 and by vectoring the thruster jets 22 in the
directions opposing any undesirable motions. For example, as best
shown in FIG. 1C, operation of the control thruster jets 22 on the
bottom surface of the nose of the fuselage 11 in combination with
the articulating air inlets 13 and the articulating exhaust nozzles
14 will force the nose of the aircraft up while, as best shown in
FIG. 1B, operation of the control thruster jets 22 on the top
surface of the rear of the propulsion system 12 near the
articulating exhaust nozzles 14 will force the tail of the aircraft
down so that an undesirable nose-down pitch of the aircraft is
counteracted and eliminated. Operation of the other thruster jets
22 will counteract and eliminate other undesirable movements of
roll, yaw, and pitch. These operations need not be detailed herein
because they should be discernible to persons of ordinary skill in
the field of aerodynamics from the example given immediately
hereinabove. By operating the thruster jets 22 in combination with
the articulating air inlets 13 and the articulating exhaust nozzles
14 to control roll, yaw, and pitch, the superagile aircraft is
prevented from attaining aerodynamic stall of the primary wing 15.
In the event of inadvertent stall, the pilot can rapidly actuate
the highly deflectable canard surfaces 19, the articulatable air
inlets 13, the articulatable exhaust nozzles 14, and the thruster
jets 22 to attain an angle-of-attack .alpha. sufficiently higher
than that associated with maximum lift.
Thus, a nose-high aircraft pitch attitude angle .theta. is
attained, thereby providing a favorable ejection attitude and
enhanced survivability for the pilot in the event that hard contact
of the superagile aircraft is anticipated to be unavoidably made
with the ground. A favorable ejection attitude is one in which the
pilot will be ejected from the aircraft in a direction away from
the ground.
The method of the present invention also involves the steps of
actuating the highly deflectable canard surfaces 19 and operating
the thruster jets 22 to control roll, yaw and pitch so that the
superagile aircraft is prevented from entering into flight
conditions whereby it will begin or sustain a spinning motion. In
the event of an inadvertent spinning motion, the pilot can rapidly
actuate the canard surfaces 19 and the thruster jets 22 to
counteract and eliminate the equilibrium between aerodynamic forces
and centrifugal forces that exist in a sustained spinning
motion.
The highly deflectable canard surfaces 19 and the thruster jets 22
my also be operated to return the superagile aircraft from the high
angle-of-attack .alpha. of supernormal flight to a normal flight
condition existing below the angle-of-attack .alpha. for maximum
lift.
The foregoing preferred embodiments of the superagile aircraft and
of the methods of flying it are considered illustrative only.
Numerous other modifications and changes will readily occur to
those of ordinary skill in aerodynamic technology after reading the
foregoing disclosure. Consequently, the disclosed aircraft and
method of flying it are not limited to the exact constructions and
steps shown and described herein but are intended to embrace other
embodiments within the purview of the appended claims without
departing from the spirit and scope of the present invention.
* * * * *