Note: Descriptions are shown in the official language in which they were submitted.
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HYBRID AIRCRAFT
Field of the Invention
The invention relates to a hybrid aircraft and, in particular, a hybrid
aircraft with vertical
take off and landing (VTOL) and/or running vertical take off and landing (R-
VTOL) and/or super
short take off and landing S-STOL capabilities. The term "Hybrid" refers to
the fact that the four
known lift principles, dynamic, ground effect, thrust lift and static lift
have been incorporated into
the aircraft design.
Background of the Invention
The need to transport substantial cargo, and/ or a large number of passengers
over
considerable distances quickly, efficiently and in a cost effective manner has
led to a variety of
medium to large (30 -100 ton payload) aircraft designs restrictively
successful in medium - long
range applications, where large ground infrastructures are available to
support their operations.
No design is currently existent as to make available transport services with
an
appropriate balance so as to provide a medium (30-40 ton) payload capacity,
with substantial
cruise speeds of up to about 400 km/h, with wide body cabin comfort, heavy
lift VTOL and
Super-Short Take Off Landing (S-STOL) operational capabilities, with typically
short to medium
(150 - 1000 mile) range at good transport economics and with the ability to
operate from either
unprepared field sites, and/or underdeveloped and/or existing aviation ground
structures while
being functionally compatible with commercial aircraft traffic and operating
patterns.
Novel airships designs emerged in the seventies and eighties, which can be
seen as an
effort to arrive at an aircraft design having the specified features. These
airships, commonly
termed lighter than air (LTA) aircraft, are based on predominantly buoyant
lift principles. Given
the use of static lifting gas, such LTA aircraft must be of enormous size and
volume to obtain
substantial (30-50 ton) lifting capacity. These aircraft have a number of
severe deficiencies.
They have poor low speed control characteristics, are very difficult to handle
on the ground, and
ballasting procedures make loading and unloading impractical. In addition,
these aircraft cannot
be accommodated in existing aviation support structures and, are unable to
maintain speeds
higher than 160 km/h because of the enormous drag penalty introduced by their
large volumes
resulting in poor transport productivity.
Efforts have been made to overcome the speed deficiency of LTA aircraft
through the
creation of "hybrid airships" as described in U.S. Patent 4,591,112 by
Piasecki et al. in which
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propulsive means are added to the LTA such that the static lift provided
offsets the empty weight
of the structures and propulsive means. This design is still unable to sustain
higher speeds
because it retains a large cross-section and con-esponding substantial drag
penalty. This hybrid
airship also retains the con-esponding ground handling problems due to
excessive physical size.
Its slow speed capabilities (110- 130 km/h), are not particularly suitable for
commercial use in
the transport of passengers. It remains a typical special mission aircraft
concept suitable for
craning operation. A further critical engineering problem confronted by such
design is the ability
to address vibration forces introduced by the helicopter type propulsive means
acting on the not
well integrated largely space propulsive support and airframe structures.
Further, an effort to improve on hybrid airship resulted in the design of
partially buoyant
airships which derive some limited lift from their hull shape. An example of
such airship is shown
in U.S. Patent 4,052,025 by Clark et al. This airship is truly a long range
aircraft with extremely
large dimensions which is unable to use existing aviation infrastructure. This
airship is very
complex and costly to construct with each fuselage panel having a different
configuration. In
addition, the helical wound base structure fuselage lacks rigidity which
prevents the aircraft from
attaining medium-high cruise speeds of about 400 km/h. Engineering problems
are created as
a result of its enormous wing span and other dimensions and, like LTA
airships, this partially
buoyant airship creates considerable ground handling problems. Further, this
partially buoyant
airship has no VTOL nor R-VTOL capabilities. This type of airship can carry
very large payloads
over very long distances. However, when compared to other aircraft, such as
the large jumbo
jets, the partially buoyant airship is not competitive in terms of payload
capacity and speed and
overall productivity.
In yet another effort in the same category, a partial buoyant aircraft has
been designed
with emphasis on the application of a "jet flap" and in combination with a
lifting body , as referred
to in U.S. Patent 4,149,688 by Miller, Jr . While the aircraft seems suitable
for improved short
take off and landing, its claim for VTOL capability will lack efficiency as
the deltoid shape is not
well suited for fitting with thrusters, particularly large scale thrusters,
efficient in vertical thrust
production. The rearward positioned thrusters, impede on a good VTOL
performance, because
their slipstream impinges on the rear top end of the hull. Further, thrusters,
when in VTOL
function, as in proposed positions fore and at the stem will lead to downwash
ground effects
severe unbalancing pitch moments of the aircraft in hover. Further,
structurally, such deltoid
fuselage are inefficient and costly to build.
CA 02235307 1998-04-16
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Helicopters are convenient ti OL aircrari but Gre cc~.~plex and costly to
operate. Since
a i'IcIICDpter obtains all Of IiS 1l i ir0lTl ItS cnglneS, It iS linaIrIE i0
Carl'y SUbStantial payICcdS (~ 10
ton payload) over medium (1000 miles) ranges as most useful load is consumed
by fuel. It is
particularly expensive to carry people and bulk break low density cargo using
helicopter
because of the severe restriction in available cabin space. Helicopters are
nonetheless
advantageous for carrying priority payloads into remote areas lacking aviation
infrastructure or
ground access, where their VTOL capabilities justify the expense. Attempts
have been made
to improve helicopter transport by increasing the size and numbers of rotors.
However, these
aircraft have not greatly improved the commercial viability of helicopter
transport.
A further aircraft design which seeks to address some of the regional short
haul ( up to
500 miles) air transport problems is the tilt rotor aircraft, also known as
the Bell Boeing V-22.
The tilt rotor aircraft have large tiltable rotors which allow the aircraft to
have both VTOL
capability and horizontal thrust. This type of aircraft has fixed wings which
provide for some
dynamic lift. Like helicopters, the tilt rotor aircraft is costly to build and
operate, it cannot carry
heavy lift, large size cargo, and offers limited cabin space and comfort. It
does not offer cost
effective passenger transport capabilities. The primary advantage of the tilt
rotor aircraft over
the helicopter, is the ability to transport payloads at higher speeds.
However, because of the
substantially high cost of the tilt rotor aircraft, their only application has
been limited to military
delivery mission, where delivery speed is a consideration, and restricted to
some selected
reconnaissance special mission operations. As a result, to date there are
believed to be no
commercial applications of tilt rotor aircraft.
A tilt rotor aircraft with tiltable wings is disclosed in NASA Tech Brief
entitled "Tiltable-Wing,
Tiltable-Rotor Aircraft" May, 1986.
A hybrid aircraft is described in German application 3,508,101 by Bothe. The
fuselage forms
a lift gas container and is roughly egg-shaped. Four tiltable propellers are
fit onto outriggers
mounted on the aircraft.
U.S. Patent 2,462,201 by Kilgore et al teaches an electrical aircraft
propulsion system. U.S.
Patent 3,110,456 of Creasey teaches a vertical take off aircraft which is
driven by ducted fans
arranged with vertical axes embedded in the wing sections of the aircraft.
,~',~li=. .'~~'-J .:~ :L.C I
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Summary of the-lnvention
~~ n aifCrafi h4S be°n !n\/cnt~a 'N1 iIC;1 nrOVIdeS ~n . CprCDrla:~
~C2lanC° ~~T yr ~VICInG
payloaa capacities of between 5 and ~0 tons, useful cruise speeds over a range
or distances,
the ability to operate with both existing and underdeveloped aviation
infrastructure and the
ability to operate in unprepared field sites.
The aircraft is capable of VTOL or R-VTOL and S-STOL, using about 30 % runway
length of conventional aircraft and has the ability to attain a medium cruise
speed range of 280 -
370 km/h. The aircraft has a significantly enlarged cabin and freight hold
space over
conventional air planes and helicopters, and has reasonable production,
operation and
maintenance costs. The aircraft of the present invention can be made to be a
size comparable
with conventional air planes to allow accommodation in existing aviation
maintenance and
support structures.
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A.~lr...~ _
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Because of the versatility of the aircraft and its high cost effectiveness,
its applications
for use can vary widely. For example, the aircraft of the present invention is
useful for
combination transport of passengers and light priority cargo delivery or
bringing heavy loads with
combined R-VTOL and VTOL operational capabilities in regions lacking
conventional air transport
infrastructure. The aircraft is useful in stow or low flying operations, such
as surveying, patrolling,
or search and rescue. The aircraft is useful in highly developed industrial
areas for point delivery
of passengers and cargo on extreme short haul (150 - 300 km multiple stop
shuttle type service
routes.
An aircraft has been invented which has an airframe geometry - tandem wing
configuration which provides for increased crash worthiness, due to
substantially reduced take
off and landing speeds of typically 90 - 150 km, a lifting body hull shape
which substantially
cannot stall and provides an air safety cushion beneath the hull in VTOL and
R/VTOL.
In o ~~ aspect of the invention, the aircraft is comprised of a hull with a
plurality of the
wings shaped as airtoils about the hull. Each of the wings has a propelling
means, such as for
example, a propeller or rigid prop-rotor, mounted thereto. The wings are each
pivotal moveable
about their neutral aerodynamic pressure axis and the propelling means are
also each
independently pivotally mounted. The pivotal movement of the wings and
propelling means is
controllable and in combination provide for lift thrust forces, control thrust
forces and forward
thrust forces. In one embodiment, the pivotal movement of all elements is
computer controlled
so as to provide substantially instantaneous control forces, for example, in
the presence of side
gusts and the like.
The wings are pivotal about their neutral axis so that minimal force is
required to make
significant changes in the attitude of an individual wing. This allows
substantially instantaneous
application of forces vector generated as the prop slipstream acts on the wing
to achieve control
moments, for example, differential deflection of two wing sections, left and
right, to counteract
the rotational moment introduced by a side gust. Where greater forces are
required to maintain
the desired attitude of the aircraft, the rotors can be actuated to pivot to
produce additional and
substantial control forces.
In another aspect of the invention, the hull shape of the aircraft provides
for the
generation of a significant ground lift effect to assist in VTOL operations
and significantly
improving VTOL lift performance. The ground lift effect is provided by a
plurality of propellers
spaced about the hull and disposed to create substantially vertical thrust.
The propellers are
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positioned in spaced relation relative to the hull such that their slipstreams
substantially do not
impinge on the hull. With such a propeller configuration, at initial take off,
50 % of the air mass
being forced vertically down by the each of the propeller thrust columns is
deflected by the
ground surface to move inwardly beneath the hull. These air masses from each
of the propellers
impinge and are forced upwardly to cause a cushion of air beneath the hull
acting upwardly to
create a ground lift effect. Preferably, the propellers are selected to each
have a direction of
rotation toward the center point, in plan view, of the hull. For example, the
four propellers can
be counter rotating fore and aft and left and right.
In another aspect of the present invention, the hull is to carry the main
aerodynamic lift
in R-VTOL and S-STOL. After reaching cruise, the hull is unloaded and up to 50
% dynamic lift
need will be supported by the wing sections. This improves the lift-to-drag
ratio to values (8-11)
comparable to wings of conventional design, The hull is shaped to provide
varying degrees of
camber in the upper and lower surfaces of the hull to thereby provide more of
aerodynamic lift
in an air stream. The aerodynamic hull lift allows the aircraft, when used in
R-VTOL or S-STOL,
to increase its payload capacity by 100 - 120 % compared to design load
capacity in VTOL
operations. In a preferred embodiment, the dimensions of the aircraft hull are
selected to both
maximize the aircraft's cargo capacity and to minimize the drag picture
thereby to optimize it
transport productivity.
In yet another aspect of the present invention, the hull is constructed to
have a rigidity to
withstand flight speeds of about 400 km/hr. The hull of the aircraft is
constructed of a plurality
of transverse rings with modular torsion members disposed therebetween to form
a triangulated
geodesio-type, space shell frame. A hard outer composite shell is semi-rigidly
mounted around
the space shell frame. Added means of intemai pressurization make the airtrame
a pressure
tensioned tensile structure under normal cruise load conditions which
increases the stiffness by
about 50 % compared to a non-pressurized vessel. The space shell frame is
dimensioned to
provide structural integrity to the airframe in the event of pressure failure.
Safe flight operations
can be continued with cruise speeds reduced to about 200 - 220 km/h.
The aircraft of the present invention can be driven by any suitable system.
For example,
a conventional drive train can be employed. In another aspect of the
invention, a turbo-electric
drive system can be used. A turbo-electric drive system comprises a central
gas turbine engine
and means to convey power to the propelling means. Such a system has various
advantages
over a conventional system including a reduction in weight and the avoidance
of the requirement
for a cross coupled transmission. In addition, the capability of installing
the turbines internal of
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the fuselage provides for internal ducting of heated gases from the turbines.
The exhaust gases
in such an arrangement can be used for heat exchange purposes for heating the
cabin, for
channelling to prevent icing on critical surtaces or for use in the heating of
gases for use in a .
static lift system, if desired.
In yet another aspect of the invention, a design of a lifting body hull
geometry has been
developed, leading to innovative construction methods in aviation which can be
termed "large
component airframe approach". This method allows the production of an airtrame
from a
sign~cantly reduced amount of different air frame components. The various
parts of the airframe
can be broken down in such a way that a smaller number of medium sized beams
plate and
panel elements can be produced using composite forming techniques other than
expensive auto-
cleave curing. Jointing techniques and self aligning components facilitate the
ease of assembly
of such an aircraft. Due to lower speed and reduced aerodynamic loading, less
expensive
materials such as Kevlar~ , E-Glass and formable thermo-plastics can be used.
The propelling means of the present invention can be any suitable type, for
example
propeller rotors or high speed fans. Where high speed fans are used, they are
disposed about
the hull body and this aircraft has been termed the "advanced hybrid aircraft"
(ANA Ship).
Preferably, 8 to 12 fans of, for example 2.5 to 4 m diameter, are mounted
about the horizontal
center line of the hull and are fitted with thrust deflectors for a range of
directional thrust. Other
fans can be mounted to provide directional or forward cruise thrust, for
example at the stern of
the hull. This aircraft provides excellent directional control in all flight
modes.
Thus, in accordance with a broad aspect of the present invention there is
provided an
aircraft comprising: a hull, a plurality of wings shaped as airfoils mounted
about the hull in spaced
apart relation, each wing being mounted to be pivotal moveable about its
neutral aerodynamic
pressure axis and a propelling means mounted on each wing and being pivotally
moveable
independent of the wing.
In accordance with a further broad aspect of the present invention there is
provided an
aircraft comprising: a hull having an upper surtace and a tower surtace and a
geometrical center ,
point, a vertical axis passing through the center point, a plurality of
propelling means mounted
on the hull in spaced apart relation about the center point, the propelling
means each being ,
disposed to provide thrust substantially parallel with the vertical axis and
to create a slip stream
directed toward the lower surtace of the hull, the lower surtace being shaped
to trap the slip
stream which is deflected beneath the hull.
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In accordance with another broad aspect of the present invention there is
provided an
aircraft comprising: a hull having a longitudinal axis and shaped to provide
substantial
aerodynamic lift in an air stream, the hull having an aspect ratio of between
about 1 to 2.5 and
a hull chord thickness ratio of between about 3 to 4.5.
In accordance with a further broad aspect of the present invention there is
provided an
aircraft hull having a cross sectional shape comprising four arc segments
connected tangentially.
In accordance with another broad aspect of the present invention there is
provided an
aircraft comprising a hull having a longitudinal axis, the hull including a
plurality of frame sections
in series each positioned substantially orthogonal to the longitudinal axis
and a plurality of torsion
members disposed between adjacent frame sections in series, the rings and
torsion members
interconnecting to form a triangulated frame.
In accordance with another broad aspect of the present invention there is
provided an
aircraft comprising: a hull having a plurality of wing sections attached
thereto and a plurality of
propelling means, the propelling means being driven by a turbo-electric drive
system including
a gas turbine, an alternator and a power conditioning and transmission system
to deliver power
generated by the gas turbine to the propelling means.
In accordance with a further broad aspect of the present invention there is
provided an
aircraft comprising: a hull shaped to provide substantial aerodynamic lift in
an air stream, a
plurality of wings mounted about the hull, a plurality of high speed fans
disposed about the hull
and having a thrust deflecting means mounted in association with the fans to
provide a range of
directional thrust.
Brief Description of the Drawings
A further, detailed, description of the invention, briefly described above,
will follow by
reference to the following drawings of specific embodiments of the invention.
These drawings
depict only typical embodiments of the invention and are therefore not to be
considered limiting
of its scope. In the drawings:
Figure 1 is a perspective view of the aircraft showing the overall
external configuration, having the propellers in the horizontal
thrust position;
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Figure 2 is a side elevation view of an aircraft according to Figure
1, having its propellers in the vertical thrust position and showing
the cabin and cockpit an-angement;
Figure 3 is a perspective view into the lower fuselage with the top ,
portion of the lifting body hull surface removed;
Figure 4a shows the geometric principles in the construction of
the hull as seen in cross section;
Figure 4b is a schematic perspective view showing the geometric
principles in the construction of the hull components;
Figure 5a is a front elevational view illustrating the ground lift
effect principles;
Figure 5b is a plan view illustrating the vortex patterns created by
the quad position thrusters beneath the lifting body hull;
Figure 6a is a perspective view of the fuselage showing the
construction elements;
Figure 6b is a cross section along line 6b - 6b of Figure 6a;
Figure 6c is a perspective view of a clustered assembly of the box
plate hull surface;
Figure 6d is a perspective view of a typical stand alone bay and
shows the keel, the shell space frame and the transverse ribs;
Figure 7a is a perspective, sectional view through a frame node .
connector with a cable guide/clamping device and a plurality of
space frame members; ,
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Figure 7b is a cross-sectional view through a frame node
connector with a cable guide/clamping device and plurality of
space frame members;
Figure 8a is a perspective view of a shell frame nodal connection
showing a rib section intertacing to the outer cover;
Figure 8b is cross section view of a shell frame nodal connection
with a rib section interfacing to the outer cover and outer cover
panel connector element;
Figure 8c is a side view showing a rib section interfacing between
the space shell frame and the outer cover panels;
Figure 8d is sectional view of an outer cover panel connector;
Figure 8e shows a cross section of an alternate hull surface
panel;
Figure 8f shows cross section of an alternate hull surface panel
with an integrated air duct;
Figure 9a is a perspective view of a propeller and a wing section
according to the aircraft of Figure 1, showing pivotal ranges useful
in the present invention;
Figure 9b is a perspective view showing the independent wing section vertical
tilt
capability and differential vertical tilt range;
Figure 9c is a side elevational view showing an alternate propeller
arrangement
having a stabilizer attached to the engine nacelle;
Figure 10 is a block diagram showing the main elements of the flight control
system;
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Figure 11 is a block diagram showing the system circuit for Turbo Electric
Drive
System (TEDS);
Figure 12a is a cross section through a hull station with an integrated carry
through beam; ,
Figure 12b is a perspective view of the internal frame at the interface of a
cant'
through beam and an outrigger;
Figure 12c is a perspective view of an external outrigger and a pivotal wing
section;
Figure " ~a is a Qlan view of an alternate AHA ship;
Figure ''3b is a schematic front elevational view showing the thruster
an-angement in an alternate AHA ship;
Figure 13c is a schematic side elevational view showing the thruster
arrangement in an alternate AHA ship; and,
Figure 13d is a cross sectional view through a fan unit useful in the
alternate AHA ship of Figure 13a.
Detailed Description of the Preferred Embodiments
The detailed description of the invention will start with a global overview of
the main features of
the aircraft and the main building blocks of the airtrame. It will then
proceed to the aspects of
"simple geometry" being the base for ease of production, and then proceed to
specific inventive
mechanical aspects of the pressurized hull structure, describe then the
control aspects, followed
by the description of an advanced propulsion system particularly suited to the
invented aircraft.
Last, a second alternate embodiment of the hybrid aircraft of the present
invention will be
presented.
Referring to Figures 1, 2 and 9a, a preferred embodiment of the Hybrid
Aircraft (HA), also called
the "aircraft", according to the present invention is shown. The aircraft is
comprised of a lifting
body hull 1 and four of thrust generating prop-rotors 23a, 23b, 23c, 23d
mounted fore and aft
about the center line on both sides along the hull at the end of outriggers 74
(Figure 12a and
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12b). Four wing sections 20a, 20b, 20c, 20d configured in a tandem fashion,
best seen in Figure
1, are mounted to rotate pivotally about the outriggers at their neutral
aerodynamic pressure axis
65 (Figure 9a). The four thrust generating prop-rotors 23a, 23b, 23c, 23d are
driven by engines
housed by nacelles 21a, 21b, 21c, 21d and are installed to provide, in
horizontal position (shown
in phantom in Figure 9a), thrust to propel the aircraft forward, and, in the
vertical position, static
v°rtinal t1 ~n ~yt lift end .~'.o.~arnl t n pct cim4~lt~y°,. ch'
in \/T<11 hnvcr and R. ~/Tfl1 T4,e ' s
I':,...~a b"d."~r ,.. .. . v~, ........, v~: . ~~r. 'axrs o.
rotation 24a of the prop-rotors can be rotated individually and are
independently pivotal about
axis 22a passing through engine nacelle 21a through typically a range of from -
10° to 90°,
relative to vertical. Preferably, axis 65 of wings 20 is common with axis 22
prop-rotor assemblies
21, 23. As will be discussed in more detail hereinafter, each prop-rotor
assembly 21 a, 23a
consists of the prop-rotor and an engine, a gear box, a lubrication system, an
interface to cross
shafting housed with the nacelle. Cross shafting 19, 19' is installed in
internal of the hull in carry
through beams.
Referring to Figures 1, 2, 3, 6a and 6b the layout of the main subsystem
components and
the main load carrying and distributing stnrctural elements of the overall
"HA" airframe are shown.
In the lower portion of the lifting body hull 1, a large keel 25 is
integrated. Above keel 25, two
carry through beams 26, 26' run perpendicular to the keel 25, fore and aft
through the hull 1.
These beams 26, 26' are connected to the keel 25 and the space frame shell
structure 41.
Additional beam truss structures 75, 75' (Figure 12a) interface with the keel
25 and shell
structure 41 and run across the lower bottom hull, outwardly from the left and
right sides of keel
at the position of the rear landing gear 8b, 8c. Beam truss structures 75, 75'
absorb and
redistribute the loads from the landing gear 8a, 8b, 8c into the keel 25 and
the lower portions of
the hull 1. The landing gear is a conventional tricycle land gear arrangement
8a, 8b, 8c. A hard
point, for external cargo lift operations, is provided and includes a hook 10
external to the lower
25 mid section of hull and a truss reinforcement 10' within the hull.
Structurally it is integrated into
the keel 25.
Transverse ribs 50, 50' are also formed in the lower hull. These ribs 50, 50',
having truss
construction, are integrated to the shell 41 and the keel 25 and follow the
lower hull curvature.
Keel 25, beams 75, 75' and transverse ribs 50, 50' together provide the
strongest portion of the
airframe. The lower hull is defined by a shell 30 formed by box plate elements
54 which are
connected in such a way as to contribute additional stiffness to the hull.
This lower hull shell 30
of the aircraft functionally requires the highest mechanical surface
stiffness. Significant
aerodynamic and mechanical surface forces are acting on the lower shell 30 due
to impact loads
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occurring in ground handling, by ground effect, for example, impact of gravel
at VTOL, and when
landing on water.
Preferably the aircraft also includes a close out structure 48, 48' which
separates the upper half
of the hull interior from the lower half of the hull interior, a lower cabin
13 formed within keel 25, -
upper cabins 14 formed by decking 16 and floor beams supported over ribs 50,
50'. The double
deck concept ofFers significant advantages over conventional aircraft holds
since the light volume
loads, for example passengers or parcels, can be carried in the upper cabins
14, while lower
cabin space 13, is configured to be practically suited for denser cargo and
can tolerate heavy
point loads. Assuming a typical hull width of 6-7 m of the lower deck cabin
for a 30 ton
commercial payload "HA", substantial additional cabin surfaces, for example up
to 110% of the
size of lower cabin can be realized by adding decking 16 to form upper cabin
14. In this way, 100
of the design load of the aircraft by weight can be realized. Upper cabin can
accommodate
low density cargo of 7-10 Ib. per cu.feet or additional passengers. This is of
particular operational
advantage when the aircraft is operated in R-VTOL or S-STOL mode, since it can
lift 100% to
120% more payload compared to its operations in VTOL.
To facilitate transport of passengers, staircases 12a, 12b are provided for
access between lower
cabin 13 and upper cabin 14. In addition, a rear access door 5 is provided.
Loading and
unloading of freight can be performed through a large front ramp 3 which opens
into lower cabin
13. This can be done without special loading or unloading equipment and can
shorten reduce
the time for loading and unloading. Also, large windows 6 can be conveniently
installed in the
upper cabin 14, feasible due to much lower pressure differential between cabin
and ambient air
space than in conventional aircraft. This design feature of the main top
passenger cabin 14
having large windows 6, for example having dimensions of 4 x 6 ft, can allow
for a particularly
attractive viewing experience from the aircraft, which typically may travel at
altitudes between
8,000 and 12,000 feet. Viewing ports 7 can be integrated into the front ramp 3
to allow for
viewing from the lower cabin 13.
The shell 41 extends from about 18% to about 85% of the hull length. It is
formed by series of
transverse rings 43 and torsion elements 44 connecting these rings and
resembles a geodesic
type structure in appearance. The shell structure 41, due to its basic
inherent triangulation,
provides a very stiff, rigid internal body shell which is optimized to absorb
torsion forces about
100 times better than a Gassical Zeppelin-type ring and longitudinal beam
construction. For ease
of construction purposes, this shell structure part 41 of the hull 1 can be
broken down into
segments 43' (Figure 6d) containing 3-4 rings 43 with corresponding torsion
members 44 forming
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together a hull segment of typically 5-7.5 m in length. An assembly of 5 - 10
segments,
depending on hull size, typically forms the whole shell. About the shell 41,
cover panels 42 are
mounted in a semi flexible manner as to provide a pressure tight and
aerodynamic cladding of
the hull 1.
The space shell frame structure 41 is provided with stiffened rings (not
shown) at its ends which
allow the attachment of the tapering front hull cap 57 and tapering rear hull
cap 5T. Caps 57,
5T are self supporting and are produced from the same box plate elements 54 as
lower hull shell
30. Referring to Figure 6c, these elements 54 are formed of sandwich
composites including
Kevlar, graphite, honeycomb and Kevlar combinations. These elements facilitate
modular
construction of the end caps 57, 5T.
In the relative voluminous upper interior portion of the hull 1, significant
space is available for
accommodation of a static lifting gas, for example, helium or hot air, if
desired. The inclusion of
such lifting gas is particularly useful if the aircraft is to be used for
VTOL. Alternately, or in
addition, large equipment, such as for example, wide aperture radar equipment
useful in sensor
platform missions, can conveniently be installed in this space, if desired.
The hull, including the shells 30, 41, covering 42 and caps 57, 57" are
preferably maintained in
tension. Even while on the ground some residual amount of internal pressure,
for example 1 -
1.5 "WC is required, to keep the airframe and the hull in this preferred
tensioned state. Thus,
for all practical purposes, close-out structure 48, 48' is required to be
pressure tight to separate
the cabins 13, 14 and doors 3, 5 from the upper interior of the hull. The
material of the close-out
structure 48 is a flexible membrane structure, or sandwich composite, which is
thin having a
thickness of for example, 1 to 3 mm and formed of a fewer layers than
conventional fuselage
sandwich composites. The close out structure is supported away from the cabin
13, 14 by keel
and a tent like frame work of beams 27 and suspension cables 49.
25 Definition of a fuselage shape controlling geometry
Referring to Figures 1, 2, 4a and 4b, the hull shape 1 is based on a simple
and locally symmetric
geometry and is generally elliptical in cross section and tapered toward the
bow 1' and stem 1 ".
The overall hull 1 size, i.e. its equivalent wing surface, is determined by
the required R-VTOL
or S-STOL lift capability of the aircraft and the desire to take off at low
speeds of typically 90-135
km/h. The hull geometry further provides for good aerodynamic cruise drag
performance in
terms of a satisfactory lift-to-drag-ratio, for example 8-12 designed to
balance, with the propulsive
power required in VTOL, at a medium high cruise speed of typically 300 - 370
km/h. Further, the
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hull geometry facilitates the formation of large "air cushion" with symmetric
ground effect patterns
which impinge beneath the aircrafts center of gravity and the aerodynamic
center of the hull. The
hull 1 shape is rounded to make use of internal pressurization to stiffen the
lifting body hull and
provides a shape suitable for landing on water surfaces with a minimum draft
of typically, 25 cm.
The shape supports low production cost, the semi axis symmetric cross section
supports the use '
of repetitive plate elements in the production of the airframe needing only a
small number of
different molds, typically 40-60 units.
For convenience, the standard geometry convention is followed in the
description of the hull
geometry, as follows: the x-axis is the horizontal axis across the width of
the craft; the y-axis
being the vertical axis; and the z-axis is the axis along the length of the
craft. The hull cross
section contour consists of 2 smaller arc segments 30, 30' and 2 larger arc
segments 31, 31' with
tangential end conditions. Angles a and p quantify the rotational distance of
the ends of arcs 30,
30' from the x-axis.
It is important to notice that the cross section is close to being elliptical,
but mathematically not
congruent with the contour of an ellipse. The cross sectional dimensions of
the hull change with
a given station length "z" value along the z-axis of the hull. According to
the known laws of
geometry, a relationship for the maximum hull width X (x) and maximum height Y
(x) value of
each cross section can be determined, as follows:
The hull width value is: X = (0.5 * x, * Romp + r"(x)
The upper hull height value is: Y = (f) a, (x, * Romp, r~(x)
The lower hull height value is: Y'= (f) Vii, (x, * Romp, r~(x),
wherein x, can assume values of between 2 and 3.5; Rort,~ is the selected
maximum diameter
of the original body of revolution; and r~(x) is the discrete radius of each
smaller arc (30, 30' in
each hull section, n, along the z-axis.
The hull cross section is generally a body of revolution cut in half, with an
original maximum
radius r~ 30 with a constant mid section 32 inserted in between the centers of
the two arcs 30
and 30'. The selection of the base radius romp determines the basic maximum
height of the hull.
The determined ra"~ multiplied by a factor x determines the width of the
constant mid section 32,
and thus the aspect ratio of the fuselage. The discrete value x depends mainly
on the amount
of dynamic lift to be supported by a given hull at the desired take off speed.
To achieve overall
good structural low airframe weight and good aerodynamic performance, the
constant mid
section 32 has to assume values between 2 * ro",~ and 3.5 * romp varying the
aspect ratio
between about 0.75 and 2.5.
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Another variable of use in varying the hull cross section geometry is the
selection of the angles
a and Vii, of the arcs about the x-axis. In order to improve dynamic lift
conditions for the same
size lifting body hull, different arc angle values a and ~i can be used. For
the lower portion of the
hull, arc angles ~3 = 75-85° have been determined most suitable, while
for the upper portion of
the hull arc angles a = 60-70° have been found to be preferable. This
geometry allows the hull
1 tn h4ye m~rv rGmber ~n rtc t~~ 4r~'~4 ~yill ~reute ~r° dv n~min 1i#
Dc 4n~...., i.:..i. 4...
.r vl.r ., mv.v ~imcm..v m. rw.~ nW vv~W , W y1' ca'i Vcl
airfoils increase the lift coefficient significantly for the same wing
reference surface area. At the
same time the increased camber of the upper hull relative to the lower hull
facilitates the
production of a substantially fail safe fuselage by making the underlying
shell frame members
more arched, and thus less prone to what is known as "snap-through" failure.
This is particularly
significant in modes under bending stress induced by maximal vertical gust
loads and where the
pressurization system of the hull has failed.
Likewise, resulting from above method, the lower angle (3 at 75 - 85 °
lets the lower portion of
the hull to have a less pronounced curvature 31' than the upper arc 31. The
resulting,
substantially flat hull bottom surface facilitates ground lift effect by
trapping air to create a
pressure build up beneath the lifting body 1. The potentially weaker structure
of the flatter lower
hull 31' is offset by the presence of the massive stiffening keel 25 and ribs
50, 50' installed along
the middle lower portion of the hull 1.
As shown in Figures 2 and 4a, the succession of variations in cross section
with typical hull
segment thicknesses of 1.5 - 3 m, form in their overall assembly typically a
fully or a half
symmetrical air foil shape. This airfoil of short aspect ratio, can have chord-
length to chord-
thickness ratios of between 1:3.5 to 1:5, i.e. 20 - 28 % chord thickness in
side view. Such an
aircraft having a body of medium slenderness is more efficient to resist
structural deformation
due to bending moments acting on the hull in cruise flight.
Applying the above-described method, optimum surtace to volume ratio hull
shapes with good
aerodynamic performance capabilities can be obtained while having finesse
ratios of between
about 3 and 4.5.
Referring to Figure 4b, each arc length 30, 30', 31, 31' can be subdivided
into a selected number,
n, of arc segments, typically to yield optimized arc length of 1-1.5 m, and
the length of the hull
can be divided into segments of typically 2.5 m. By use of such segmentation,
the hull can be
covered with panels requiring only 2 or 3 variations of curvatures. This
segmentation can also
apply to the hull ring 43 and torsion members 44. Significant reduction in
component production
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and assembly costs result. Additionally, for certain applications, the hull
design 1 can be made
essentially symmetric fore and aft, as shown in phantom and indicated as 39 in
Figure 5b, thus
reducing the number of airframe components which are different in shape by 50
%.
Lifting body hull and quad-rotor thruster arrangement
Referring to Figures 5a and 5b, in VTOL or hover mode, with an aircraft, as
shown, having a
symmetric hull and four prop-rotors, termed a "quad-rotor" arrangement, or in
an aftemative
embodiment having more than four thrust generators, 50 % of the air moved
downward in all
thrust down wash columns 35a, 35b, 35c, 35d, after hitting the ground, is
being deflected inward
beneath the hull, as shown by the arrows in Figure 5a. All these combined
counter rotating
vortex disks of air impinge, as indicated at 36, beneath the hull and
intersect below the center
of gravity 38, which is substantially the same as the aerodynamic center. When
the air vortexes
impinge, they create a zone of increased air pressure which is thrust 90
° upward and form a
supportive cushion of air, as indicated at 36'. This causes a fountain effect.
The proper
placement of the prop-rotors 23a, 23b, 23c, 24d and resulting downwash columns
35a, 35b, 35c,
35d act as "curtains" to trap the air under to the relatively wide hull.
Counter rotation of
propellers fore 23a, 23d and aft 23b, 23c and left and right improves this
effect, and is preferred.
With such counter rotating props only two gaps 37, 37' are available for the
air to escape. The
preferced hull shape embodiment therefor should be a symmetrical contour 39,
or close to being
a symmetrical body, as shown by hull 1.
Experimentation with a 6 m model representing the "HA" huff-rotor geometry has
confirmed the
validity of this concept and provided detailed scientific data on the actual
lift improvement
achievable. It has been found that where the distance between the ground and
the lowest point
of hull curvature is equal to the prop diameter, a ground lift effect force
equal to 30% of the
original thrust force delivered by the propellers can be obtained. With closer
ground proximity,
for example distances between the ground and the lowest point of the hull of
between 0.35-.50x
the prop diameter, even stronger ground lift effects, for example of up to
additional 36°fo of
original thrust generated, have been measured. In particular, prop disk
loading of Z 80 kg/mz
support this effect.
Geometry to minimize hull interterence during prop-rotor vertical thrust
Referring to Figure 2, to avoid lift toss induced by air being drawn over the
hull to feed the
propeller, particularly when propellers are positioned for aircraft hover or
VTOL, the rotors are
placed outside of the outer perimeter of the hull in plan view. In the
preferred embodiment, the
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horizontal propeller plane of rotation is substantially at the same vertical
position as the height
of the outer hull curvature at the corresponding locations along the length of
the hull. The
resulting airflow is almost tangential to the propeller plane and thus
minimizes down draft effects,
where the air feeding into the propellers is drawn over the hull surface. The
clearance, indicated
as C in Figure 5b, between the outer limits of the propeller rotation and the
outer perimeter of the
hull, in plan view, should preferably be at least 15 - 20 % of the prop
diameter.
Differential hull wing section, propeller tilt capability.
The lifting body airframe 1 provides the basic structural frame work, to which
externally, the
propulsion system is attached. Referring to Figures 9a, 9b and 12a-12c, the
large diameter prop-
rotors 23a, 23d are mounted at the end points of cantilevered outriggers 74a,
74d. Tilt-able wing
sections 20a, 20d are mounted between the hull and the prop-rotors 23a, 23d.
The chord center
line 40 of these wing sections can be pivoted independently from the tilting
of the prop-rotors.
Preferably, the wing sections are mounted to rotate and be positioned within
the range of from
-10° to 130°, relative to a horizontal axis with the usual
positioning in cruise being 1 - 5°, 40 -
75° in R-VTOL and up to 130° in certain hover flight control
conditions. The rotor axis 24 is
typically tilted in a range of from 0° to 110°, relative to
horizontal, with the prop axis being generally
0° in cruise flight, 90° in hover and various other positions
for control and take off modes.
Typically, a differential angle of up to 23° between rotor axis
positioning and wing position are
usefully practical.
Numerous benefits result from this differential tilt capability. In
particular, there are three benefits
which represent a major improvement over the existing tilt rotor technology.
Referring to Figure
9a, showing a vertically oriented wing section 20a', and a horizontally
oriented wing section 20a,
shown in phantom. First; the present invention allows the reduction of pylon
lift loss in VTOL and
S-STOL. With wing section 20a' in this vertical position, blockage of the
propeller slipstream is
substantially avoided and down wash drag over the outrigger 74 of the wing is
reduced to about
1.5% VTOL lift loss compared the values of 8% VTOL lift loss commonly
resulting where a fixed,
non-tiltable wing is used as a support pylon for a tiltable prop.
The wing section angle of rotation can be kept positive above prop angle of
rotation at all times
at S-STOL and in transition to cruise. In contrast to conventional fixed wing-
tilt rotor aircraft, this
arrangement not only prevents thrust lift losses but increases the dynamic
lift generated, or,
inversely, can be used to decrease lift at S-STOL landing. Additionally, the
wing sections 20a,
20b, 20c, 20d in presence of the a propeller thrust slipstream becomes a blown
wing section
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20 providing lift and control at very low speeds, for example below 80 km/h.
This is particularly
beneficial in maintaining attitude control in the final moments of the landing
process. Dynamic
lift improvements in S-STOL and transition flight and better aircraft handling
characteristics
result.
Thirdly, the independent tilting of the four wing sections 20a, 20b, 20c, 20d
allows the distribution
of required dynamic lift between the four wings to sustain cruise flight.
Depending on the speed
and the flight modes, the dynamic lift required for any given point of the
flight envelope to keep
the aircraft airborne is being supplied by the sum of the lift forces produced
by the lifting body and
the four wing sections. As the lifting body 1 has a significantly smaller
aspect ratio than the wing
sections 20a, 20b, 20c, 20d, it therefore produces more induced drag for a
certain amount of
dynamic lift produced. It is, thus, advantageous to unload the lifting body at
cruise speed and
allow the 4 wing sections make up the portion of the dynamic lift shifted away
from the hull. The
four wing sections 20a, 20b, 20c, 20d will produce the same amount of dynamic
lift with
significantly less overall drag penalty then the lifting body 1 carrying 100%
of the dynamic load
by itself.
The pressurized rigid fuselage with flexibly suspended outer shell
Referring to Figures 6a and 6b, the hull of the present invention consists of
a rigid self supporting
space frame shell 41 around which semi-flexibly attached a semi-rigid covering
formed of panels
42, which is able to substantially maintain its overall surface geometry when
under no internal
pressure. A gap 53, of typically .5 - .75 m, is formed between the shell 41
and the panels 42.
The shell 41 and the panels 42 act together with internal pressurization to
form a lightweight, rigid
pressure vessel providing significant increase (about 50-55 %), over
unpressurized vessels with
the same mechanical structure, in resistance to bending moment and
streamlining of the frame
41 to achieve a suitable aerodynamic hull surface shape. This mechanical
approach has been
termed a "Pressure Tensioned Shell Frame" (PTSF) fuselage by the inventor.
While internal pressurization is a known method in mechanical engineering to
stiffen vessels and
has been employed in space and aviation technology prior to the present
invention, no suitable
construction has been developed which is useful for large sized aircraft
travelling at medium to
high speeds. Generally, the shell 41 is a series of rings 43 in positioned
orthogonally to the z-
axis. Each ring 43 is formed, in the preferred embodiment as shown in Figure
12a, of a plurality
of tubes in the form of a polygon with 12 to 18 comer vertices 78 (18 vertice
polygon rings are
shown). Every second ring is rotated by 30° or correspondingly
20°, thus creating a frame cross
section (in front view) having 24 - 36 rows of vertices 78. In alternate
embodiment as shown in
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Figures 6a and 6b, the ring 43 vertices 78 are regularly spaced about the top
and lateral portions
of the cross section of the inner contour of the hull, while the lower
sections of the rings 43' are
modified to inGude a greater number of members to accommodate connection of
structural ribs.
Each vertex 78 is aligned for each second ring along the z-axis of the hull
curvature. The rings
43 and the interconnecting torsion beams 44, form a web of interlocking
triangles connected by
elevated nodal points 45, at the vertices, providing some amount of depth
space to the shell
structure.
Referring to Figures 7a, 7b and 8a to 8d, each node 45 has substantially the
shape of a ring and
has 6 bore holes which allow the insertion of bolts 45' which are fitted from
the inside of the ring
node 45 and are engaged in tapped counter holes formed in the end cones 43',
44' of members
43, 44. Each node 45, being a pin-type joint, acts to connect six tube members
of the shell frame
41. Perpendicular to the plane of intersection of the ring tube members 43
with the node 45,
preferably a latch mechanism 63 is inserted into a slot 62 formed across the
top side of the node
45. The latch 63 is maintained in the slot by a pin bolt 46 and is part of a
clamping device 60
which allows local clamping of a cable 58 and the tip or "toothed" portion of
the rib 52 which is
semi-rigid and which bridges the gap 53 between the inwardly located rigid
space frame shell 41
and the semi-rigid outer covering of panels 42. The panels 42 take the shape
of gores running
from bow cap 57 to stem cap 5T. The cable guide/clamping device 60 is
typically 0.65 m in
length and .15 m in height and is formed of a clamp 60' which secures a pair
of planar members
60". Planar members 60" are bonded to rib 52 and are engaged by clamp 60'
during assembly.
To further understand the mechanics of the attachment of the outer covering
panels 42, the
construction and mechanical particularities of the outer cover panels 42,
their connection to the
longitudinal interface ribs 52 and their mechanical function will be
explained.
The outer cover panels 42 are formed as a sandwich composite of lightweight
layers. The
preferred sandwich provides the best strength to weight ratio with currently
available materials
and consists of, from outwardly facing surface to inwardly facing surface: a
very thin layer, for
example 0.005mm, of an abrasion resistant film (e.g. Tedlar~); a layer of a
tri-axial woven fabric
using pol~aramide fiber (e.g. Kevlar~); a honeycomb core sheet of about 10 to
15mm thickness;
another layer of the tri-axial woven fabric using poly-aramide fiber; and a
final inner layer of an
effective gas barrier film, such as metallized Mylar~, which is commercially
available. The layers
of the sandwich are bonded with an adhesive. The honeycomb sheet contributes
enough "in-
plane" stiffness to the sandwich to resist wobble and flutter stresses which
are introduced into
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the outer surface during cruise at speeds of about 360 km/h. Such a panel
construction has a
weight of typically about 0.45-0.65 kg/m2. The panel has an equivalent tear
resistance of a 1.5
mm thick aluminum sheet with a fraction of the aluminum sheet's weight. The
panel has several .
times the unit tension strengths of modern Blimp flexible structures.
In an alternate embodiment of the outer hull cover panels are formed of two
sheets of flexible
material with a sheet of low density honeycomb disposed therebetween. In
another alternate
embodiment, as shown in Figure 8e, two thin sheets of flexible material 42,
42' are used to form
an inner and an outer covering layer. Closed cell foam 28, having a tow
density of between
about 15-30 kg/m3, is placed between sheets 42, 42'. The space between the
flexible sheets is
inflated prior to injection of the foam and the sheets thereby act as a mold
until the injected foam
cures. Such an aerangement provides enough structural stiffness to maintain
the cross sectional
shape of the hull. In yet another embodiment, shown in Figure 8f, the covering
is formed by two
spaced sheets 42, 42' of flexible, high-tensile material and the space 28'
between the sheets is
pressurized separately from the overall hull volume by independent means of
pressurization. In
this way, the higher internal pressure in the space acts to stiffen the outer
cover locally without
effecting the whole hull volume.
The sandwich composite, as described above, is the preferred material for use
in the formation
of panels 42. To assemble the covering on the hull, the sandwich material it
is cut into repetitive
panels 42' which follow the geometry defined by the sequence of vertices 78
along the perimeter
of the shell 41. Referring to Figure 8b, a interface seam of two adjacent
panels 42 and a
perpendicularly extending "toothed" rib 52 is shown. The "toothed" ribs have
typically a height to
span ratio of about .20:1 to .25:1 and can be produced from the same material
as the panels 42
or other light weight sandwich materials. The construction method useful for
interconnecting a
rib 52 with the panels 42 makes use of integrated joining edge technology
which has pertected
the fabrication techniques of heat and ultrasonic welding and facilitates the
joining of these
separate parts. Referring to Figures 8b and 8d, the interconnection of two
panels 42 with a rib
52 is accomplished by use of a connector 64 which is integrated to rib 52 and
engages enlarged
edges 42a of panels 42. Edges 42a of panels are enlarged by means of
integrated edge
technology. In particular, connector 64 is formed as an upper section 64a and
a lower section
64b which are joined by a fastener 2, such as a screw. Lower section 64b is
securely integrated
to rib 52. Corresponding grooves 64a' and 64b' are formed in sections 64a,
64b, so that when .
joined, a channel is formed between sections 64a and 64b which is shaped to
receive and firmly
hold panel edge 42a. During assembly of the covering, edges 42a of panels 42
are placed in
grooves 64b' and section 64a is placed over this arrangement so that edges 42a
also fit within
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grooves 64a'. Fastener 2 is then inserted to firmly join the sections 64a and
64b. By use of
connector 64, installation of the covering can be accomplished from outside of
the hull.
Preferably, connector 64 is formed in extended lengths by extrusion of
polyaramids, such as
Kevlar~.
To connect the ribs 52 with the nodes 45 of the space shell 41, the tip
portion of a "tooth" is
bonded into clamping device 60. The latch 63 formed on device 60 fits the
geometry and
location of each of the node connector points 45 and slot 62. Cable 58, formed
of Kevlar~, is
incorporated along the edge of the rib 52. This cable 58 acts as the main
tension member to
transmit the outer hull surface tension forces, generated in part by the
internal hull pressurization,
into the cable guide/clamping device 60 which, in turn, conducts all
concentrated tensile forces
through the latch 63 into the node 45. The latch 63, after insertion into slot
62 of node 45, allows
for some lateral rotational motion around pin bolt 46. This provides some
lateral deflection
flexibility for the semi-rigid rib 52 between the outer panels 42 and the
rigid frame 41. The
geometry of the latch 63 is determined by the amount of stresses the internal
pressurization of
the hull generates per unit area of the covering. Further clamping devices 60
can be attached
along the ring members between nodes 45, as desired to further distribute the
load into the
frame. About 36 rib connecting points are installed along the upper portion of
each ring 43.
The outer covering is formed by a complete set of panels 42, each longitudinal
panel is formed
of 12 - 24 gores. These panels can be installed from the hull top centerline
down to the
intersection of the lower shell 30. After the whole hull surface is completely
enclosed by the
upper hull covering formed of panels 42, lower shell 30, bow hull cap 57 and
stem hull cap 5T,
the hull can be internally pressurized. Pressurization means are well known
from blimps and air
supported domes, and do not require further description. With a slight
pressure built up over
ambient pressure, for example of about 3-12 inches of WC (3000 Pascal), the
outer cover panels
42 will be stretched tight. The hull geometries have been developed such that,
under internal
pressure, the hull enclosed gases (air and/or a lifting gas) effect a uniform
perpendicular surface
pressure throughout the inner hull and act to force the panels 42 radially
outwardly. The in-plane
surface tension load acting on each of the panels 42 is transferred into the
ribs 52. This tension
is consequently transferred through the cable guide/clamping device 60 to the
latch 63 and into
the node 45. This stresses the underlying shell frame structure members 43,
44. The arched
shape of the ribs 52 between two adjacent nodes 45 allows distribution of the
accumulated
tension toad between two rings 43.
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Referring to Figures 6b and 7a, at least some of the space frame nodes 45 have
rods 51 with
cable eyes 51' formed thereon installed about their pin bolts 46. The rods 51
are free to assume
the geometry driven angles to align with the keel 25. Thus, from the center
point inside of the
node 45, to the cable eye 51' of the end rod 51, a tension wire 47 runs to the
upper right 25a and
left 25b corner of the keel 25 at a location corresponding to the ring 43
position on which the
node 45 is located. These arrays of cables 47, act to offset the hoop stresses
created by the
internal pressure and retained by the surface tension and also act to
distribute some of the force
created by a concentrated payload housed in lower cabin 13 and upper cabin 14
through the keel
structure 25 and into the shell 41. Additionally, the tension wires 47 act to
render the shell frame
structure 41 stable without internal pressure present.
Application of detailed engineering, including finite element techniques and
vertical gust
simulation of the forces the aircraft hull is subjected to in cruise flight,
have quantified that the
combination of above-described covering panels 42, ribs 52, and the shell
members 43, 44, 45
substantially all remain pre-stressed due to the presence of internal
pressurization over a wide
range of operational speeds, for example up to about 360km1h cruise speed.
Only when vertical
gusts, particularly gust having speeds beyond 60 feet/second, are encountered
at such speeds,
some of the members connected to the keel go into compression. The
characteristic tensioned
nature of the elements of the shell 41 during typical cruise speeds, shows the
superior weight
to strength ratio of hull. As known, components made from materials such as
Kevlar~ and
carbon graphite, carry many more times the load of a force applied in tension,
than that applied
in compression. Averaged surface assembly unit weights of typically 2.2-2.5 kg
per m sq. (at
defined load conditions) are in this way achievable with the described
construction method.
Which consequently will result in a lightweight airtrame structure having
favorable empty weight
fractions of typically 0.45 to 0.5 at a 40 ton maximum take off weight
aircraft size. In the event
the pressurization fails, the local buckling strength of the space frame
members 43, 44, which
have typically 8 - 12 cm beam diameter and 0.5 - 2 mm wall thickness when
aircraft aluminum
alloys are used, is defined such that the full structural and shape integrity
of the lifting body hull
can be maintained at a lower cruise speeds of 110 - 125 knots with a 60
feet/sec vertical gust
moment. This will allow the aircraft to return to base even in adverse weather
conditions.
Load bearing integrated outrigger with tilt-able wing section
Referring to Figures 12a to 12c, due to the fact that the vertical static lift
component in a smaller
"Hybrid Aircraft" plays no, or, only a minor role in the balance of lift
forces in VTOL, a very
significant amount of force (typically 10 -20 ton) and vibrations are acting
on the outriggers 74.
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The design of such outriggers 74 should be sufficiently stiff to withstand the
forces and yet be
relatively lightweight.
The hull intemai cant'-through beams 26 (only one shown) are preferably of
conventional
modular truss construction or a large diameter tube. Beam 26 is closely
integrated to the keel
25 and frame 41 by a series of tension cables 73 connect the beams 26, 26' to
the shell 41 while
rigid tube elements 72 connect the keel 25 to beams 26, 26' to provide a well
supported beam
26. The ends 71 of beam 26 directly intertace with shell structure 41 and
create a very stiff local
web of beams suitable to act as a "hard-point" 76 at which the outriggers 74b,
74c are supported.
Some of the beam members in web 76 allow tangential distribution of moments
into the space
shell hull structure members 43, 44 which are particularly suited to absorb
forces acting "in
plane". The hard point 76 provides enough stiffness against bending moment in
its vertical, as
well as, its horizontal planes to withstand the rotor thrust forces which will
act in both direction
at this point.
The external outriggers 74 can be a square truss, as shown, or a tube.
Outrigger 74 is limited
in maximum height and width to the inner maximum chord height geometry of the
wing sections
airfoil type. A typical outrigger 74 has a diameter of 0.8 m to 1.8 m. The
specific diameter
is selected with reference to size of wing section 20, thrust required and
propeller sizes.
Because each wing section 20 can be rotated differentially from the propeller
axis, but is often
tilted at about the same angle as the propeller axis itself, the thrust forces
acting on the tip of this
20 outrigger 74, and its resulting bending moments, are very much in line with
the position of the
wing section. They can deviate in VTOL hover and S-STOL flight mode usually to
a maximum
angle of about 18-22°. Thus, the force vector introduced into the
outrigger, is the same as the
thrust vector which is the largest load acting on this structure. The wing
section is preferably
made an integral part of this arrangement. The wing section, when
approximately aligned with
the thrust axis, provides a much higher moment of inertia in the plane of its
chord length, than
the outrigger beam itself.
The wing section 20 is designed as a stiff wing box with stiffened ribs 66,
panels 77 extending
between ribs 77 and composite sandwich surface covering 68. The wing section
has a high
cross sectional moment of inertia along its length. Both the moment of inertia
for the outrigger
74 and the moment of inertia for the wing section 20 are summed for that
position in whatever
angle the wing section is rotated to at a given moment. This significantly
increases stiffness
results for a given structural weight. The combination of wing 20 and
outrigger 74 these
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components add to each other their individual maximum bending moment strength
capability,
when compared to acting alone to support a corresponding load. The wing ribs
66 installed
bearings are surface type bearings 69 optimized to sustain torsion loading.
Bearings 69 can be
produced from composites, to reduce weight, or can be made from classical
roller bearing
materials, such as steel.
With this outrigger and wing geometry, in VTOL flight mode, when the wing
sections 20 are in
vertical position, to also act under certain conditions as control moment
generators, adequate
outrigger stiffness is assured. Likewise, in forward flight, when the wing
sections act as a
conventional airfoil to produce dynamic lift, sufficient outrigger stiffness
is available to
alternatively accommodate both thrust and dynamic lift forces acting on the
wing sections 20 and
thus the outrigger, depending on flight mode and dynamic lift load
distribution.
Integrated cruise, VTOL propulsion and hover flight thrust control system
The aircraft of the present invention is configured to deliver precision hover
and good station
keeping ability in 80 % of all prevailing wind speeds. Control system hardware
and software
means are provided to supply a propulsion and force vector controlling in
combination, which can
deliver thrust vector changes rapidly to achieve attitude control. Smaller
vector changes can be
delivered in a fraction of a second, while larger vector changes can be
delivered within about .5 -
1.5 seconds.
Commercially available, large diameter tiltable prop-rotors are usually
limited to tilt rates of
typically .7 ° - 1.5 ° per second, to avoid excessive stresses
induced through the inertia of
gyroscopic forces. When considering the effect of sharp edge gusts or wind
direction changes,
which can occur in about .5 - 1 seconds, it becomes Gear that the tilting of
props 24 alone cannot
provide the yaw moments required, to maintain the hull in the pre-turbulence
position in such
conditions. This is particularly, when it is desired to match closely a target
over the ground
during hover.
Here the pivotal installation of the wing sections have their supplementary
function; allowing rapid
rotation around their neutral axis with rotational speeds of up to
22°/second. Referring to Figures
1, 9a and 9b, at the onset of the rotation in "y-axis" of the aircraft, an
onboard computer based
sensing system measures the rate of acceleration and determines the forces
required to arrest,
or, to slow down this adverse rotation. Within about 0.2 second about 6-7% of
the thrust
produced by the propellers 23a, 23b, 23c, 23d is available as yaw moment, by
the combination
of positive and negative lifting forces acting almost perpendicular on the
tilted wing sections 20a,
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20b, 20c, 20d which are pivoted up to 22 ° forward, or respectively
backward, from a vertical
position, in opposite directions on the left and right side of the aircraft.
This immediate activation of yaw moment will either prevent the start of a
rotation in the y-axis
in perturbations up to about 20 knots, and/or significantly slow down the
rotational speed in this
axis, until the slower pivoting propellers have reached a pivot angle of 2-
3° from vertical in about
1.5 - 2 seconds, again tilting differentially forward and backward on each
side of the hull, after
the onset of the perturbing wave front, gust or wind shift has collided with
the aircraft. Application
of the standard cosine function of the rotor tilt angle, which determines the
horizontal thrust
vector available in x-axis in a vertical acting thruster for a given amount of
vertical thrust
available, shows the following picture: at 3 ° prop-rotor tilt,
additionally to the vectors already
created by the immediate wing sections tilt, 5.2 % of vertical thrust vector
is available. The
percentage of each of the two control vector producing sub-systems 20a, 20b,
20c, 20d and
23a, 23b, 23c, 23d added, is sufficient to create strong combined yaw moments.
It is known from
control of modem helicopters, that typically 10-12% of overall thrust has to
be reserved to assure
good controllability. Similar figures, for example, about 7 % from wing
sections plus 5 % from
rotors, are achievable in the "Hybrid Aircraft", as demonstrated. Also, it has
to be kept in mind
that, in the design case of a typically commercial "HA" having a wing span of
about 35 - 40 m,
the available thrust vector is applied at the end of a very long moment arm of
up to 20 m long.
This delivers very powerful control moments, to rotate the aircraft back into
alignment with the
main wind direction.
When the prop-rotors 23a, 23b, 23c, 23d have differential tilt angles fore and
aft of 2 3° from
vertical, up to 5 % of vertical thrust component can be made available to
produce moments in
'yav~' by the prop tilt alone. With larger prop tilt angles from vertical, the
wing sections 20a, 20b,
20c, 20d continue to produce the moments as determined before, as they can be
rotated parallel
with the increasing rotation of the prop-rotor axis. The relative downwash
vector over the wing
sections remains unchanged. For a typical commercial sized "HA" having about
40,000 kp of
thrust, in combined wing and thrust axis tilt, more than 225,000 meter kg
control moment in 'y-
axis° is available, which corresponds to the perturbation moments of
winds gusting up to about
54 kts (28 mlsec, or, 93 ft per second) acting on the hull. This is close to
the highest gust speed
in which modem aircraft are designed to safely fly. This is also equal to the
best precision hover
performance of modem helicopters, and sufficient to assure a 85% "on station
time" typically
desirable for commercial operations.
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Flight control modes and related control elements
Table I shows the systematic symmetric "half' and "differential" combinations
and variations of
magnitudes of propeller thrust vectors, rotor axis vectors and the wing
section negative and
positive dynamic lift vectors for the two main flight modes to be controlled:
Hover and Cruise
flight. A description of components involved in the relevant attitude control
requirements follows
Table I.
Table I presents the elements involved in flight attitude control in a table
form.
Table I
H OVER CRLI1SB LItIHT
- Fl.It~3T F -
-
Ytou~ iw rru xau PitdiYtw vd. A1t$.
PZO(l. P~ICb X X ~ f Xp X .X X
C~lfgG ~
_.-__ X -
- Z X I X Z Z ~ X
X _
-- _.
X ~ __ ._.-X_. _,X
X . .. _
. X _.....
_.X
I
...
x
VTOL Flight mode and hover
The attitude control, pitch, roll moment generation in z, and x axis is
accomplished through
differential collective thrust changes of rotors side-to-side and fore-to-aft.
Main yaw moment for
rotation around the 'y-axis" is generated through tilt of the right side props
backward to maximum
of 10° (from vertical) and the tilt of left side propellers by the same
degree forward. Yaw moment
in hover control involves additionally the wing sections as described in
detail above. Forward
translational slow speed mode is achieved through collective tilt, of all four
props forward,
typically 2-5°. Translational motion backwards, likewise, is achieved
through collective tilt of all
four props or aU four wing section backwards. In case of backwards motion the
propeller tilt is
preferably limited to a maximum angle of 10° backward.
Translational movement sideways is achieved by introducing first a roll moment
with differential
thrust between right and left rotors and then keeping collectively thrust in
the same proportion.
Other solutions would be added cyclic in lateral for one axis only or
installation of other means,
such as fan thrusters (not shown) in the bow and stem caps perpendicular to
the center line of
the hull, in this way providing translational sideways movement without roll.
In another
embodiment, vertical stabilizers 88, 88' (Figure 9c) are mounted onto the
engine nacelles 21.
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When prop axis 24 is in vertical position, the rudder integrated tabs 89,89'
on the stabilizers 88,
88' collectively lift or right, to cause deflection of the propeller
slipstream and thus translational
movement in "x" axis without roll angle.
R-VTOL or S-STOL flight modes
In S-STOL the four propellers are tilted to a position of about 70-75°
from horizontal to maintain
an advantageous ground air cushion effect and to optimize lift and forward
acceleration vectors.
The four wing sections are tilted collectively to about a 15° higher
tilt angle than their
corresponding propeller axis. This provides very effective acting "blown"
control surtaces, over
which the air stream of the propeller is accelerated and thus the wing
sections can provide pitch
and roll control moment in the S-STOL situation, even at very low flight
speeds between 80 and
110 km/h (45-60 kts), when aerodynamics control surtaces typically are lacking
effectiveness,
due to missing air pressures almost not acting on them in very low speeds.
Directional control,
(yaw-control) is provided by differential thrust between right and left
propellers.
In an optional embodiment, vertical stabilizers 88, 88' (Figure 9c) having
rudders 89, 89' can be
used to provide a blown rudder arrangement. This embodiment would provide
directional (yaw)
control as known in conventional aircraft designs. This stabilizer rudder
configuration is
considered feasible, but not the preferred embodiment for structural reasons
and reasons of
applying a preferred advanced computer controlled stability enhancement design
approach.
Transition flight mode (from VTOL and Hover only)
Starting from a VTOL, or hover flight mode with the prop axis at about
90° from horizontal, the
thrust vector can be rotated by slowly tilting the four propellers
collectively forward. The x-axis
forward component of this vector accelerates the aircraft forward. This
results in increasing
forward speed and in the generation of dynamic lift by the lifting body hull
and the four wing
sections. This starts to reduce the magnitude of thrust lift required and
permits further downward
tilt towards horizontal of the propeller axis. During the tilting process, the
effectiveness of axis
bound generation of control vector changes. These are, however, governed by
basic laws of
geometry. Full transition is reached typically at speeds 2 165 km/h (90 kts).
Cruise Flight mode:
In cruise flight pitch control is provided by differential tilt of one pair of
wing sections, for example,
the left and right wing sections closest to the bow, fore and the other pair
aft. Roll control and
coordinated turns can be effected by differential tilt of the wing sections on
each side of the hull
and by differential thrust of right and left side propellers. Trim in cruise
flight is achieved by
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differential fuel tank filling of fuel tanks fore and aft of the center of
gravity. (Not shown).
Directional stability is provided through differential thrust variations of
right and left propellers
pairs and, in an alternate embodiment, can be assisted by deflection of
stabilizer 88 mounted
rudder tabs 89.
Basic flight control and guidance concept
The aircraft control concept is built around a concept known as Active
Computer Augmented
Stability concept. Figure 10 shows a block diagram of the main flight control
system components
of the hardware layout described above. The basic control concept includes the
control
elements, as follows:
Control means:
4 x variation of prop blade pitch, thrust regulation
4 x variation of propeller tilt axis
4 x variation of tilt angle of wing sections
4 x variation of rudders mounted in prop slip stream (optional)
Control means available to pilot:
1 control stick for roll and pitch control and to facilitate control in VTOL
hover
for transitional side movement and local translation fore and aft
1 button mounted on stick for collective prop axis tilt
1 button for collective thrust change
1 pedal for directional control
Mixers:
the 5 input variables of the pilots) have to converted info control actuator
control
signal size, to be specified for each of the 12 or 16 means of control
available.
Actuators:
due to the relatively slow cycle times of the aircraft, actuators activated by
electrical motors are sufficient. This will allow the omission of a
hydraulics, a
major contributor to maintenance burden in any aircraft
Cockpit and flight control system layout
Referring to Figure 2, the cockpit 17 is installed in the front portion of the
bow hull cap 57 of the
hull 1. The cockpit layout and pilot interfaces may be any suitable
arrangement such as that
arrangement recently developed by Boeing-Bell Helicopter in the realization of
the V-22 Osprey.
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Referring to Figure 10, the 12 or 16 elements listed above are controlled by
electric actuators and
configured with dual redundancy. The interaction of all elements controlled is
coordinated by a
computer 80. The system approach is typically based on state-of-art Fly-by-
wire or Fly-by-Light
control concepts. The computer 80 to which all components of the Fly-By-Wire
system are
linked, is housed in the cockpit.
The central core of the system is a three-axis laser gyro 83 for measuring the
rotation angles in
the x, y, z axis, which is linked to the computer 80 which monitors
continuously the attitude
changes of the aircraft. Auto-pilot functions are installed with typical
automatic pilot capability
for engagement by the pilot, if desired. The auto-pilot functions are
preprogrammed to
automatically handle certain recurring flight modes which arise as part of the
various operational
routine requirements. Further, the control system, is essentially based on
"Rate Change Control"
(RCC) principles, whereby the laser gyro 83 supplies the data regarding the
rate of changes of
relative movement in the x, y, z spatial coordinate system. The pilot, via a
control stick and the
other interface means, can set the desired values for rates of change in
regards to various
control parameters of the aircraft to be modified and/or maintained at any
given time. The "control
laws" computer program subroutines 85 deal with the preprogrammed equations
describing the
motion of the aircraft and provide data to the mixer 86, which will provide
discrete signals 87a,
87b, 87c, 87d to a combination of actuators to achieve the flight control
desired.
Onboard optical sensor 82 is preferably included in the flight control system
to provide data to
the computer concerning the translational movement of the aircraft in close
ground proximity.
Sensors 84 are conventional and provide data to the computer 80 concerning the
prevailing
atmospheric conditions.
Modem "steady state" laser-based gyro technology also monitors the continuous
moving
reference point, for example the actual location of the aircraft at any given
point in time or the
latest progression of position of the aircraft in its flight path.
Alternately, an onboard Global
positioning system 81 can provide information on the real time position of the
aircraft. This
sophisticated control technology is recently available at moderate commercial
systems costs, and
is currently available in "steady-state" hardware versions, having high
reliability and requiring little
maintenance, compared to earlier models using a combination of rotating
mechanical and
electrical parts.
Figure 10 is shows a superior flight control system without application of
mechanical means of
linkage. The above mentioned Fly-by-light is the same system approach as the
fly-by-wire
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system. However, the fly-by light uses optical fibers to transmit the data
from the central control
processors to the local processors and actuators. This is an advantage for an
aircraft of
significant physical size and which, from time to time, may fly close to
strong magnetic fields, for
example in the inspection of high voltage power lines.
The control system is fully digital and has a layout of triple, or quadruple
redundancy. Such a
system could have also have an advanced "teaming mode", which would enable the
computer
system to "learn" and "save" certain responses to gust, ground effect,
turbulence, etc.
To support ease of flight operations, particularly in minimum ground
infrastructure support
environments, load cells are in installed in the landing gears. These load
cells provide automatic
updates of the center of gravity shifts due to the rapid changing loading and
unloading situations
of the aircraft. These data will assure safe and flexible changing of payloads
with a minimum of
supervision by the crew or a "load master".
The drive systems
Two drive systems are preferred for use in the present aircraft. The first
drive system includes
conventional main drive train components. They include a conventional gas
shaft turbine and
auxiliaries, a gear box, a clutch and components for required "simplex" cross
shafting for each
pair of propellers front and aft. The shafting 19, 19' (Figure 3) allows
transmission of about 50%
of the propulsive power from any of the two paired fore and aft engines, over
to the opposite
located propeller in case of engine failure at that unit. The shaft 19, 19'
can be supported by the
carry through structures 26, 26'. These installations are quite conventional
and require no further
description.
The second preferred drive system is termed a Turbo Electric Drive System
(TEDS). Over the
past 10 years, very significant progress has been made in light weight
electric engine and new
drive technology. Permanent magnet brushless motors and high speed power
generator
technology combined with semi conductors (thyristors), used for manipulation
of high voltages
and currents has evolved significantly in capability and seen drastic
reduction in cost. In contrast
to conventional electric motors , these drive units are capable of running at
high rpm of between
about 10,000 and 40,000. The weight algorithms for the turbo-electric drive
outputs have
reached the 0.2 - 0.25 kg/kW range per shaft power for electric motors and a
level about 0.10 -
0.15 kg/kW for alternator output power in the Z 1000 kW range. While TEDS
systems are
known, the utilization of such system as primary drive systems for aircraft
has not previously
been realized. For a number of compelling engineering and operational reasons,
such an turbo-
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electric drive system (TEDS) approach is useful as an alternate system
approach for a "Hybrid
Aircraft" of the present invention.
The TEDS layout is shown in a basic block diagram in Figure 11. The system
includes
conventional (aircraft certified) gas turbine 90, 90' in redundant layout,
high speed, high density
alternators 93 in redundant layout which are each directly coupled through
gear boxes 92, 92',
which optimize alternator rpm, to electric power conditioning and control
units 91, 91'. A high
voltage power transmission system 96 in redundant layout transmits power to
four brushless
permanent magnet motors 99a, 99b, 99c, 99d and gear boxes 95a, 95b, 95c, 95d
to drive the
constant rpm variable pitch propellers 23a, 23b, 23c, 23d. This system
preferably has electronic
fuel injection governors controlled by the central flight computer 80. The gas
turbine engines 90,
90' and alternators 93, 93' are preferably housed in an engine room 104
suitably located in the
stem of the "HA". The engine room is preferably positioned within the hull to
be accessible from
within the hull during flight to allow for in-flight repairs and
modifications.
The shaft gas turbine 90 is the prime drive for the high speed alternator 93.
This constant speed
drive, configured as a direct shaft coupled alternator, can run at 2 10,000
rpm.
The electric power conditioning and control units 91, 91' include circuits
containing a set of high
performance thyristors and condition and manipulate the currents, wave forms
and outputs
instantaneously. It is controlled by a computer which receives, in turn,
control signals from the
flight guidance computer 80. The control signals have to be delivered to the
various actuators
to enact the power settings, etc., as a response to the pilot inputs.
At the power user level, the alternators 93, 93' supply energy, modulated by
units 91, 91', to the
four the brushless permanent magnet motors 99a, 99b, 99c, 99d. The drive
shafts of these
motors rotate at between about 10,000 - 12,000 rpm. The motor rpm is geared
down by
conventional two-stage gear boxes 95a, 95b, 95c, 95d to match the propeller
tip speed of an
optimized propellers size. High voltage power supply fines 96 run from the
engine room 104 to
the four motors. The drive motor windings can be arranged in two separate
segments and in
such a manner as to have a built in 50% power redundancy, in case of failure
of one of the
windings.
The propellers are of the constant speed (rpm) type. In order to maintain the
propeller speed
constant, when varying lift/controf require that the blade pitch be changed to
provide more or less
thnrst, the electronic fuel governor injects an amount of fuel into the
turbine which corresponds
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to the load condition signal produced by the flight computer. Systems having
constant rpm prop
with electronic fuel governor are very well established technology.
The hull internal power plant makes features of lifting gas heating or the air
within the hull
possible. The location of the gas turbines inside the hull and their position
in such a way that
some of the waste heat generated by them, can be extracted before the exhaust
gases leave the
stern of the aircraft, makes a combination with stem mounted heat exchangers
mechanically
practical enough, to have much easier installation of "super-heating"
techniques compared to the
conventional layout, in which waste heat has to be ducted from the exterior
location of the turbine
to heat exchangers located in the hull. The more practically feasible
installations of heat
exchange means and simpler ducting enhances the possibilities of heating the
hull internal gas,
thus to improve VTOL lift, if desirable.
Alternate HA embodiment having VTOL serial lift fans with integrated stern
propulsion
system
An alternate embodiment of the aircraft of the present invention is shown in
Figures 13a - 13d.
This aircraft is termed the Advanced Hybrid Aircraft ("AHA" Ship). This
embodiment has a
variation in the arrangement of propulsion and control elements from those
discussed earlier
herein.
The overall hull of the "ANA" is configured as a lifting body hull of small to
medium aspect ratio,
for example AR = 1-3, and finesse ratios of for example, between 1:4 - 1:6
with a symmetric
body for ground effect optimization. The hull can accommodate a lifting gas to
providing up to
15 % of its maximum take off weight by static lift, if desired. Its overall
design considerations
follow the main principles presented above for the preferred "quad-rotor"
embodiment.
The lifting body hull 105 has a plurality of pivotal stub wing sections 106a,
106b, 106c, 106d
mounted thereon in tandem arrangement, fore and aft of the center of gravity
124. Wings 106a-d
provide pitch control and coordinated turns in cruise and low speed. At the
end of the stub wings
106a, 106b, 106c 106d, vertical stabilizers 115a, 115b, 115c, 115d with
integrated rudder tabs '
116a, 116b, 116c, 116d are mounted to facilitate directional control in cruise
flight. The stub
wing sections are pivotally mounted at their neutral aerodynamic pressure
points and can be '
rotated from 0 ° (horiwntal) to 25 °.
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Along the maximum diameter middle portion of the hull in plan view, a
plurality of fans 108a -
108g and 108a' - 108g" are mounted. Typically 2-4 fan units are mounted fore
and aft and left
and right from the center of gravity 124. The fans are installed in horizontal
flanges 107 and 107'
which extend from the maximum perimeter of the hull 105. The fans are
preferably those
commonly used in the first fan stage of high-bypass-ratio gas turbines, for
example those
commercially available from General Electric Company. They are typically
deployed in current
large size passenger jet aircraft and providing together with the gas turbine
10,000 - 30,000 kg
of thrust each. These fans are very light weight and quiet. Generally a total
of 8-14 fan units,
having diameters of 1.5-4m would be required to produce the thrust required to
lift a 15 to 30 ton
aircraft.
Referring to Figure 13d, ducts 111 are formed in the mounting flanges 107,
107' for
accommodating the fans. The ducts are spaced so that there is a distance of
about 2 m between
fan locations to allow unhindered air supply to each fan. In the ducts, below
the fan disk 108 an
individual high speed brushless permanent high density magnet motor drive unit
99 can be
mounted and directly shafted to the fan. The motor drive units 99 preferably
have 800 - 2,500
kW-output capacity. Electric power to drive each of the fans 108a - 108g and
108a' - 108g", is
supplied by a central power stations 113 in the rear of the aircraft. A turbo
electric drive system
(TEDS), as previously described in detail with reference to Figure 11, can
provide the power
required.
Thrust deflectors 109 are installed in the ducts 11 below the fans. Thrust
deflectors 109 can
actively be controlled to deflect the thrust up to 25° to the left
right sides, seen from vertical
thrust axis. To allow the creation of thrust deflection creating yaw moment
independently from
creating translational movement fore and aft, preferably the thrust deflectors
for the four fans
closest to each of the bow and stem 108a, 108b, 108a', 108b', 108f, 108g,
108f, 108g' have their
pivot axis parallel to the z-axis of the hull 105 and the deflectors for the
remaining fans have their
pivot axis parallel to the x-axis.
Preferably, the thrust deflectors for the centrally located fans 108d, 108d'
are mounted beneath
the fans in circular frames 117 supported by roller bearings. The circular
frames 117 to which
thrust deflectors 109 are mounted, has the appropriate mechanical means to
allow quick rotation
around its axis thus to facilitate rapid change of direction of thrust about
at least 180° with
rotation rate of typically of 90° - 120° per second .
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The upper and lower openings of the duct 111 into which each fan installed,
can be closed by
a set of louvers 118, 118' to provide a fairing when the VTOL thrusters are
not in usage, typically
in cruise flight.
Power stations 113, based on high speed alternator 119 technology, produce
electric power .
required to run the electric fan motors 99. High speed brushless alternators
119 are directly
shafted to large scale gas turbines 120 which typically generate 5,000-10,000
kW output power.
An electronic current output conditioner is controlled by computers and is
coupled to the flight
control and guidance system of the aircraft to assure the supply of properly
modulated electric
power to each of the brushless motors 99.
At the center line rear section of the fuselage 105, further fans 108h, 108i,
108j (shown in
phantom), of similar type to those described above, can be mounted in ducts
112a, 112b, 112c
so the plane of rotation of the fan is in a vertical position. The fans are
located at about 95%
chord length of the lifting body hull 105. These fans 108h, 1081, 108j provide
a separate forward
propulsion system for the aircraft. The electric power required to drive these
fans is also
delivered from the same high speed alternator turbine combination housed in
power stations 113.
The air intake ducts 114a, 114b, 114c for these fan units 108h, 1081, 108j are
integrated into the
upper and lower rear hull surfaces at about 85-90% of the chord length of the
hull. The
slipstreams of the fans exit through duets 112a, 112b, 112c mounted at center
of the aircrafts
trailing edge. The location of air intake ducts 114a, 114b, 114c for the rear
fans 108h, 108i, 108j
facilitates the reduction of the boundary layer. The intake ducts could be
configured on the hull
surface as a boundary layer suction ring slot 118, assisting in efficient wake
propulsion and
boundary layer control, delivering all these features in a synergistic
fashion.
Directional control is accomplished very easily and precisely with the AHA
ship. Application of
differential and collective thrust vectors created by the thrust deflectors of
fans 108a - 108g and
108a' - 108g' acting in x and z-axis provides yaw moments to allow slow speed
translational
movement to the sides and fore and aft. Additional directional thrust in any
direction from 0 - 180
° can be provided by fans 108d, 108d' through rotation of its of
deflector. This allows fine tuning ,
of vectors to create an overall control vector picture which allows precision
control and any
desired flight direction in hover and VTOL. Additionally, the rear mounted
fans can also be
equipped with thrust deflectors and can be used to produce directional control
in forvvard flight
and hover. This is of particular use in the presence of head winds in hover or
VTOL.
CA 02235307 1998-04-16
WO 97/15492 PCT/CA96/00705
-35-
The high speed fans 108a - 108g and 108a' - 108g' operate at very high rpm of
about 6,000 -
10,000. The disc loading of these fans is increased by 50% over conventional
prop-rotors, from
about 80-90 kg/m2 to about 140-160 kg /m2, but fan downwash speeds and unit
lift performance
measured in kg lift per hp remain within economically feasible ranges. Thrust
column density
remains in a range which avoids soil erosion problems which should be
considered in any craft
capable of hover.
The criticality of the overall propulsion failure in hover triggered by any
individual engine failure
is decreased due to the higher number of overall propulsive elements available
and power
generation redundancy and the significantly smaller moment imbalance caused by
the failure of
one or two VTOL fan units. In the preferred embodiment, if one or two of the
electric motors
driving the fan units should fail, the remaining number of 12 fans would
assure that overall thrust
levels required for safe operations can be produced by the remaining units,
and overall thrust
available stays almost at 100%. Thrust performance can be maintained, in the
event of a failure
of one of the turbines, by increasing output of the remaining prime power
delivering turbines. In
the presented layout.
The alternate embodiment, includes cruise flight drag reduction aerodynamics
by means of
boundary layer control and wake propulsion. The geometry layout results in
major 20-30%
reduction of wake drag, particularly in an aircraft having a relatively large
lifting body hull. The
fact that the air breathing turbine engines are housed close to the stem and
preferably within 85-
95% of hull chord length, makes a relatively simple combination of a ring
suction slot and air
intake ducts, at about 85% of hull length, with the rear propulsion feasible.
It will be apparent that many changes may be made to the illustrative
embodiments, while falling
within the scope of the invention and it is intended that all such changes be
covered by the
claims appended hereto.
