NZ618737B2 - A helicopter - Google Patents
A helicopter Download PDFInfo
- Publication number
- NZ618737B2 NZ618737B2 NZ618737A NZ61873712A NZ618737B2 NZ 618737 B2 NZ618737 B2 NZ 618737B2 NZ 618737 A NZ618737 A NZ 618737A NZ 61873712 A NZ61873712 A NZ 61873712A NZ 618737 B2 NZ618737 B2 NZ 618737B2
- Authority
- NZ
- New Zealand
- Prior art keywords
- fuselage
- primary flight
- layers
- layer
- seat
- Prior art date
Links
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- B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
- B29K2077/00—Use of PA, i.e. polyamides, e.g. polyesteramides or derivatives thereof, as moulding material
- B29K2077/10—Aromatic polyamides [polyaramides] or derivatives thereof
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
- B29K2277/00—Use of PA, i.e. polyamides, e.g. polyesteramides or derivatives thereof, as reinforcement
- B29K2277/10—Aromatic polyamides [Polyaramides] or derivatives thereof
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
- B29K2307/00—Use of elements other than metals as reinforcement
- B29K2307/04—Carbon
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3082—Fuselages
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3088—Helicopters
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/712—Containers; Packaging elements or accessories, Packages
- B29L2031/7172—Fuel tanks, jerry cans
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/771—Seats
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B60—VEHICLES IN GENERAL
- B60N—SEATS SPECIALLY ADAPTED FOR VEHICLES; VEHICLE PASSENGER ACCOMMODATION NOT OTHERWISE PROVIDED FOR
- B60N2/00—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles
- B60N2/24—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles
- B60N2/42—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles the seat constructed to protect the occupant from the effect of abnormal g-forces, e.g. crash or safety seats
- B60N2/4207—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles the seat constructed to protect the occupant from the effect of abnormal g-forces, e.g. crash or safety seats characterised by the direction of the g-forces
- B60N2/4242—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles the seat constructed to protect the occupant from the effect of abnormal g-forces, e.g. crash or safety seats characterised by the direction of the g-forces vertical
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B60—VEHICLES IN GENERAL
- B60N—SEATS SPECIALLY ADAPTED FOR VEHICLES; VEHICLE PASSENGER ACCOMMODATION NOT OTHERWISE PROVIDED FOR
- B60N2/00—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles
- B60N2/24—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles
- B60N2/42—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles for particular purposes or particular vehicles the seat constructed to protect the occupant from the effect of abnormal g-forces, e.g. crash or safety seats
- B60N2/427—Seats or parts thereof displaced during a crash
- B60N2/42709—Seats or parts thereof displaced during a crash involving residual deformation or fracture of the structure
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B60—VEHICLES IN GENERAL
- B60N—SEATS SPECIALLY ADAPTED FOR VEHICLES; VEHICLE PASSENGER ACCOMMODATION NOT OTHERWISE PROVIDED FOR
- B60N2/00—Seats specially adapted for vehicles; Arrangement or mounting of seats in vehicles
- B60N2/68—Seat frames
- B60N2/686—Panel like structures
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/82—Rotorcraft; Rotors peculiar thereto characterised by the provision of an auxiliary rotor or fluid-jet device for counter-balancing lifting rotor torque or changing direction of rotorcraft
- B64C2027/8254—Shrouded tail rotors, e.g. "Fenestron" fans
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/82—Rotorcraft; Rotors peculiar thereto characterised by the provision of an auxiliary rotor or fluid-jet device for counter-balancing lifting rotor torque or changing direction of rotorcraft
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D11/00—Passenger or crew accommodation; Flight-deck installations not otherwise provided for
- B64D11/06—Arrangements of seats, or adaptations or details specially adapted for aircraft seats
- B64D11/0619—Arrangements of seats, or adaptations or details specially adapted for aircraft seats with energy absorbing means specially adapted for mitigating impact loads for passenger seats, e.g. at a crash
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D11/00—Passenger or crew accommodation; Flight-deck installations not otherwise provided for
- B64D11/06—Arrangements of seats, or adaptations or details specially adapted for aircraft seats
- B64D11/0689—Arrangements of seats, or adaptations or details specially adapted for aircraft seats specially adapted for pilots
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D37/00—Arrangements in connection with fuel supply for power plant
- B64D37/02—Tanks
- B64D37/04—Arrangement thereof in or on aircraft
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Abstract
composite structure forming a load bearing composite shell for a helicopter, with a shell which defines the exterior of a fuselage and includes a central fuselage section and tail boom, a fuselage which is adapted house an engine or drive train, a layered composite crashworthy seat and support structure and a fastening and a method of fastening by providing an adhesive between two layers which is allowed to flow through opposing holes in the layers to provide a chemical and mechanical attachment between the layers. The load bearing shell is formed as a continuous laminate structure. ucture and a fastening and a method of fastening by providing an adhesive between two layers which is allowed to flow through opposing holes in the layers to provide a chemical and mechanical attachment between the layers. The load bearing shell is formed as a continuous laminate structure.
Description
“A HELICOPTER”
FIELD OF THE INVENTION
The t invention relates to aircraft structure and in particular to composite aircraft structure.
BACKGROUND TO THE INVENTION
Prior art pter supporting ures and framework comprises materials such as wood,
aluminium, titanium, chrome molybdenum steel tubing and magnesium alloys. Fabrication and
manufacturing of helicopter structures are based around the extensive use of special jigs and frames
and certified holding fixtures where floors and jigs, frames and fixtures are frequently calibrated.
Such installations have inherent disadvantages including that they are fixed in a location and are not
The prior art ge fabrication and manufacturing process requires the fuselage to be
constructed by first assembling internal components and working outward. The prior art method
of constructing a helicopter begins by identifying a starting location or part such as a central floor
panel. The internal structure of the fuselage is then added systematically around that starting
on by adding sub frames and panels. The assembly is then strengthened by riveting or bolting
adjoining sub frames and panels to form a skeleton. Once all of the internal structure has been
completed the ge skeleton is enclosed with a skin that is either riveted or bolted into place,
usually by direct attachment to the skeleton. When the primary structure of the fuselage is
completed and the fuselage is structurally sound it would be removed from the fabrication or
cturing assembly jig.
Traditional helicopter fuselage manufacturing has numerous disadvantages. One disadvantage is
that construction is extremely labour intensive. The completed fuselage has a vast number of
individual parts, each requiring prior fabrication. To track and assemble these parts requires a
skilled work force. Further, the fabrication jigs have long set-up times and long breakdown times.
tion of helicopter ge in traditional manner is very expensive.
A further disadvantage to traditional helicopter manufacturing is the ed external surface of
the pter fuselage covered in a mass of domed rivet heads. This type of finish is both
unattractive and results in high drag penalties. Significant materials cost and time is associated with
the use of flush head rivets in the outside skin of the fuselage to remove the drag penalty.
A r disadvantage to traditional helicopter fuselage fabrication using sheet metal panels to
form the outer skin of the fuselage is the difficulty in achieving a smooth and thus aerodynamically
favourable shape.
A further disadvantage to ional pter ge fabrication is that door and window
openings are lly hand finished. Finishing by hand results in no two door or window openings
being the same. Each window or door therefore es individual shaping, usually by hand, to
ensure a fitment that allows closure without gaps.
A further disadvantage to traditional helicopter fuselage manufacturing using riveted structures and
thus lapped joints is the ingress of moisture. This moisture becomes trapped and corrosion will
ensue. Corrosion can lead to structural failure.
A prior art method of attaching empennage appendages to a fuselage is with a mechanical fastening
such as rivets, screws or bolts. ical ings such as riveted joints or bolted joints are
known to be labour intensive and require the use of special fixtures and tooling jigs.
A disadvantage associated with use of mechanical fastenings to secure such ages is that each
of the adjoining surfaces must have a plurality of holes formed for the fastening to pass though.
Such holes can cause weakening of the structure and may contribute to a point of structural failure.
To mitigate the risk of structural failure such methods of fastening often require regular
maintenance checks to ensure structural integrity is maintained, particularly for any cracks that may
be propagating between adjacent holes.
Mechanical fastenings have a further disadvantage in that substantial damage may be caused to an
aircraft by tearing the surrounding material should a secured appendage be struck by some external
object.
ical fastenings have a further disadvantage when attaching to curved surfaces together. The
ch in shapes between the curved surface and the lly flat fastener may create undue
stress on a location immediate to the fastener.
A further antage is that mechanically fastened surfaces or riveted surfaces are prone to
sealing issues where moisture can ingress or be trapped.
A r disadvantage is that mechanically fastened surfaces are prone to various types of
corrosion. Filiform, intergranular and surface corrosion can form between mechanically fastened
surfaces. Often, corrosion in these areas goes undetected even with periodic maintenance,
dismantling and inspections and can result in catastrophic failure of the fastening or region
proximate the ing.
Traditionally helicopter crew and passenger seats have been inbuilt structures with the helicopter
fuselage. Later years have seen crew and passenger seats evolve into stand-alone assemblies for the
forward seats and ay seats for the rear passengers. Certification standards require the
inclusion of a crashworthy seat for all occupants of the helicopter.
More recently since the introduction of new certification rules, seats in newly certified helicopters
are required to be worthy” meeting certain design parameters of maximum load factors,
inertial forces, and reactions between occupant, seat, and safety belt or s corresponding with
the applicable flight and ground load conditions, including the emergency landing conditions of the
category in which certification is sought.
As a result there have been several newly design crashworthy seats installed into helicopters as new
helicopter designs or retrofits to older helicopter designs. These new seat designs incorporate
s of shock absorber, collapsing lever mechanisms, brake, energy absorbing foams, and
collapsible metal structures.
One seat design in the prior art for meeting the crashworthy seat rd is known as a stroking
seat mechanism. Disadvantages of the stroking seat ism include the requirement for regular
inspection and ing, corrosion protection for metal surfaces, rtent jamming of the seat
action and injuries to limbs that occur during the stroke of seat.
Another design in the prior art for meeting the crashworthy seat standard is known as a braking
seat mechanism. Disadvantages of this seat mechanism include the friction pad loosing preload
over time and the regular requirement for inspection and readjustment, the metal frame requires
corrosion protection, , inadvertent jamming of the seat action and injuries to limbs that occur
during the stroke of seat.
r design in the prior art for meeting the orthy seat standard is an aluminium sheet
metal box . Disadvantages of this seat mechanism include allowing the occupant to fall
through the seat pan into the seat base. While the occupant may survive the crash, ce has
shown the occupant is subsequently trapped in the seat base and unable to escape the crashed
aircraft.
In this specification, where reference has been made to external sources of information, including
patent specifications and other documents, this is generally for the purpose of ing a context
for discussing the features of the present invention. Unless stated otherwise, reference to such
sources of information is not to be construed, in any jurisdiction, as an admission that such sources
of information are prior art or form part of the common general dge in the art.
It is an object of the t invention to provide a solution which overcomes or at least
ameliorates at least one of the abovementioned disadvantages or which at least provides the public
with a useful choice.
Other objects of the invention may become apparent from the following description which is given
by way of example only.
SUMMARY OF THE INVENTION
In one aspect the invention relates to a y flight structure consisting of a load g
composite shell defining at least the exterior of a fuselage, the fuselage defining at least a central
fuselage section and tail boom, wherein the central ge section is adapted enclose at least one
of an engine or drive train.
Preferably the tail boom has attached or is adapted to have ed at least one of an empennage,
fins, or tail rotor mechanism.
Preferably the empennage is adapted to support a ducted fan tail rotor ly, horizontal and
al stabilisers.
Preferably the tail boom is adapted to have attached an empennage to thereby form a monocoque
structure.
Preferably a forward section of the fuselage adapted to house one or more occupants and flight
controls.
Preferably the composite primary flight structure further comprises a central section and forward
section of the fuselage delineates at least one of a door, window or hatch opening.
Preferably the ure delineates an opening located in an upper region of the fuselage surface,
the opening d to allow at least one of an engine or main rotor gearbox or drive train module
to be at least partly inserted into the fuselage.
Preferably the composite primary flight structure further comprises at least two members
components extending between at least an upper and lower internal surface of the fuselage.
Preferably the members provide at least one attachment point for at least one of an engine or main
rotor gearbox or drive train module, or at least one attachment point for a frame ly to which
at least one of an engine or main rotor x or drive train module are adapted to be attached.
Preferably the composite primary flight structure r comprises at least two members ing
at least between side internal surfaces of the fuselage and transversely to the at least two beam
components.
ably at least one of the members is adapted to transfer load d by at least one of an
engine or main rotor gearbox or drive train module to the composite shell.
Preferably at least one of the members is arranged to at least partially create an enclosed space in
which one or more fuel cells can be located.
Preferably the members are adapted to create a structure attachable to the internal surface of the
composite shell such that the structure least partially absorbs aircraft impact energy and diverts
energy away from the enclosed space.
Preferably at least two members further comprise an integrally formed and forward ding
member d to extend from the central fuselage section to the forward section, the protruding
member having a lower region adapted to attach to a lower internal surface of the composite shell
and an upper region adapted to support a cabin floor panel.
Preferably the forward section of the fuselage and the cabin floor panel at least partly define a
cabin space such that seating for occupants and flight controls can be located or referenced.
Preferably the cabin floor is supported by the forward protruding member and a plurality of
structural members, the structural members adapted to extend between the lower side of the cabin
floor panel and the internal surface of the lower region of the composite shell.
Preferably the ural members are adapted to attach to the cabin floor panel and the internal
surface of the composite shell.
Preferably the structural members comprise a first set of members and second set or members and
the first set of members are adapted to extend substantially perpendicular to the second set of
members such that the structural members er, when combined, are adapted to form a lattice
structure extending between the cabin floor panel and the internal surface of the composite shell.
ably the cabin space is further adapted to allow a plurality of seats, flight ls and
occupant ints to be d.
Preferably the composite shell ses a laminate, the laminate comprising a plurality of fabric
layers.
Preferably the plurality of fabric layers se a first layer of continuous filament mat, a first
layer of carbon and Kevlar composite, a layer of unidirectional carbon fibre, a second layer of
carbon and Kevlar composite, a second layer of continuous filament mat and wherein the laminate
is, or is d to be infused with a cured or curable resin
Preferably the laminate further comprises additional laminates and/or core filler material including
Soric.
Preferably the laminate further comprises a layer of carbon unidirectional fabric between the first
and second layers.
ably the laminate further comprises a layer of carbon double bias fabric between the first and
second layers.
Preferably the laminate further comprises a filler layer between the first and second layers.
Preferably the layer of carbon double bias fabric comprises a plurality of layers of carbon double
bias fabric, including a plurality of layers that span a surface of an upper region of the central
fuselage section to a surface of an upper n of the tail boom, a plurality of layers that span the
upper region of the central fuselage surface section to a middle and a lower surface section of the
tail boom, and a ity of layers that span between a lower region of the central fuselage n
to at least some way into the lower surface of the tail boom .
Preferably the unidirectional carbon fibre fabric is approximately 200 g/mm2.
Preferably at least some of the support fabrics are approximately 200 mm to 300 mm wide.
Preferably the support s are integrated within the laminate layers of the fuselage.
ably the continuous filament mat layers are approximately 300 g/mm2.
Preferably the carbon and Kevlar composite layers are approximately 190 g/mm2.
Preferably the filler layer is approximately 2 mm thick.
Preferably the outer surface layer is formed as a substantially smooth and substantially continuous
surface that extends through at least the central fuselage n to the tail boom.
Herein described is an enclosure adapted to provide protection for one or more fuel cells
consisting in at least four upstanding panel s spanning between an upper and lower internal
surface of a substantially enclosed composite fuselage or outer shell of a flight structure, at least
two panel members having a surface extending laterally relative to the or of the fuselage or
outer shell of the flight structure interior and attaching to the al surface of the fuselage at
least in an upper and lower region thereof, and at least two panel members extending
longitudinally to the interior of the fuselage or outer shell of a flight structure interior and attaching
to the interior surface of the fuselage at least in an upper and lower region thereof, the panel
members defining an enclosed space.
Preferably each of the fuselage or outer shell and panel s comprise a composite laminate
ure.
ably the enclosure further comprises a floor member attached to a lower region of the
enclosed space.
Preferably the upper region of the panel members is adapted to support at least one of an engine or
main rotor gearbox or drive train module, or at least a frame assembly to which at least one of an
engine or main rotor gearbox or drive train module are attachable.
Preferably the ure is adapted to support at least one fuel cell.
Preferably at least one of the panel members has an aperture through which a fuel cell can be
inserted.
Preferably the fuselage or outer shell and panel members together form a structure such that, when
subjected to an impact load, divert impact energy away from the enclosed space.
Herein described is a method of forming a helicopter fuselage having at least a centre section
and tail boom n, the method consisting of providing at least two moulds, a first moulds
having an inner shape substantially corresponding to a port section of a desired outer fuselage shell
profile and a second mould having an inner shape substantially corresponding to a starboard
section of a desired outer fuselage shell profile, applying a plurality of fabric layers into each of the
first and second mould to form a port ge section and a starboard fuselage section, and
infusing a resin into the layers of fabric to create a composite ure.
Preferably the step of applying a plurality of fabric layers into the mould comprises applying a first
layer of CFM, applying a first layer of carbon and Kevlar composite fabric, applying a layer of CU,
applying a second layer of carbon and Kevlar composite fabric, applying a second layer of CFM.
ably the method r comprises applying additional laminates and/or a core material
including Soric.
Preferably the step of applying a plurality of fabric layers into the mould further comprises a layer
of CU between the first and second layers.
Preferably the step of applying a plurality of fabric layers into the mould the laminate structure
further comprises a layer of CDB between the first and second layers.
Preferably the method further comprising ng each of the first and second moulds together,
such that the port section and starboard section of the fuselage are adjacent.
Preferably each mould section has one or more of an alignment marker or datum, the method
further comprising aligning each of the ent markers or datums.
ably the method further comprising applying a plurality of fabric layers across the adjoining
port section and ard section of the fuselage.
Preferably the method further comprising providing a plurality of moulds corresponding to desired
al fuselage structure.
Preferably the method further comprises the steps of allowing the resin to cure and removing the
composite structure from the mould.
Herein described is a method of building a helicopter comprising the following steps:
providing a fuselage outer skin or shell, the shell comprising at least a fuselage centre section and
tail boom, the fuselage outer skin or shell delineating a plurality of openings and defining a
substantially enclosed space, inserting a plurality of members into the fuselage through at least one
of the plurality of gs.
Preferably the method further comprises assembling the plurality of members inside the fuselage
outer skin or shell to form an internal structure.
Preferably the method further comprises ing a plurality of floor panels into the fuselage
through at least one of the openings.
Preferably the method further comprises bonding the plurality of s to either the inner
e of the fuselage outer skin or shell, other members and floor panels, or a combination
thereof.
ably the step of g comprises applying fabric to overlap regions of the proximate
regions of the adjoining members and infusing the fabric with resin to form a composite structure.
Preferably the method further comprises arranging the plurality of members in the centre section
inside the fuselage to define an enclosed space.
Preferably the fuselage further forms a forward section of a ge adaptable to enclose a cabin
for housing a pilot, passengers and flight controls.
Preferably the plurality of gs include either a door, a window or a hatch opening.
Preferably the method further comprises attaching an empennage or any number of flight
stabilisation appendages to the tail boom.
ably the tail boom and empennage define a monocoque structure.
Preferably the empennage is adapted to support a ducted fan tail rotor assembly, horizontal and
vertical stabilisers.
Preferably the method further comprises inserting an engine or engine supporting frame into the
fuselage through one of the openings.
Preferably the method r comprises attaching the engine or engine supporting frame to at
least one of the members.
Preferably the method further ses inserting at least one fuel cell into the enclosed space
within the fuselage through one of the openings.
Herein described is a orthy seat ly for an aircraft comprising: a seat base
component, the seat base having a lower surface adapted to attach to the aircraft, a seat pan
component wherein the seat pan is adapted to attach to the top of the seat base, the seat
component formed of a composite laminate structure.
ably the seat base is of a form such that when subjected to an impact, the seat base crushes
to absorb impact energy.
Preferably the seat base ent having a base wall and four side walls, the side walls
comprising substantially continuous inner and outer layers that each extend around the perimeter
of the seat base, the side walls further comprising an intermediate layer, the intermediate layer
comprising a plurality of discrete wall sections disposed between the inner and outer layers such
that a crushing region is defined at or near the region where each wall transitions to the next.
Preferably the base wall and the side walls have a region of overlap, wherein the region of overlap
is provide by either the sidewalls extending into the base wall, or the base wall ing into the
side wall, or both.
Preferably the region of overlap of the side layers extend approximately 50 mm across the base
surface and the base layers extend approximately 30 mm up the side surfaces.
Preferably the seat base further comprises a upper layer ing around the upper region of the
four side walls.
Preferably the upper support layer is a continuous filament fabric of approximately 450 g/m2.
Preferably the seat component comprising a seat back and seat pan, the seat back and seat pan
formed from a plurality of substantially continuous .
Preferably the seat is an all composite upper seat comprising a seat pan and seat back structurally
bonded to a crushable composite seat base.
Preferably the base wall is d to be fastened to an aircraft cabin floor.
Preferably the base wall comprises one or more additional layers of carbon double bias fabric.
Preferably the base wall of carbon double bias fabric is approximately 400 g/mm2.
Preferably the base wall of carbon double bias fabric is at least 75 mm wide.
Preferably the base wall and four side walls define an enclosed internal space, the internal space
adapted to house an energy absorbing material, or luggage.
ably the energy absorbing foam is Confor CF45, Confor CF47 or similar.
Preferably at least one of the side walls has an opening providing access to the ed space, the
opening having a support layer disposed around the periphery thereof.
Preferably the support layer is a continuous filament mat fabric of imately 450 g/m2.
Preferably the rearward side wall extends upward at of the seat base such that additional support is
provided to the back section of the seat.
Preferably the inner and outer layers are as carbon and Kevlar fabric 0/90, bidirectional or l
type fabric.
Preferably the inner and outer layers are a Carbon and Kevlar fabric of approximately 180 g/m2.
Preferably the intermediate layers are a continuous filament fabric of approximately 450 g/m2.
Preferably the seat pan component has a substantially upstanding seat back portion al with a
substantially flat seat base portion, wherein the seat pan component is constructed from alternating
layers of fabric materials.
Preferably at least one of the fabrics is a flow .
Preferably at least one of the fabrics provides ical strength.
Preferably at least one of the fabrics prevents ive failure of the other fabrics when subjected
to an impact load.
Preferably the ating fabrics are Carbon and Kevlar fabric and continuous filament.
Preferably a core material spaces a portion of the layers of fabric apart.
Preferably the layers of fabric materials se: a first surface layer of carbon and Kevlar
composite fabric of imately 180 g/mm2, a second layer continuous filament mat of
approximately 300 g/mm2 , a third layer of laterally spaced portions of carbon unidirectional fabric,
each approximately 100 mm in width and approximately 300 g/mm2, a fifth and sixth layer of
carbon and Kevlar composite fabric of approximately 180 g/mm2, a sixth layer of continuous
filament mat of approximately 300 g/mm2, a seventh of continuous filament mat of approximately
300 g/mm2, an eighth layer of carbon and Kevlar composite fabric of imately 180 g/mm2, a
plurality of laterally spaced ninth layers of carbon ectional fabric of approximately 300
g/mm2, a tenth of continuous filament mat of approximately 300 g/mm2, and an eleventh layer of
carbon and Kevlar composite fabric of approximately 180 g/mm2.
Preferably the layers of fabric materials further comprise a layer of core material ed between
the sixth and seventh layers.
Preferably the core material comprises two laterally spaced members.
bly the core is a PVC foam material or similar.
ably the width of each ninth layer of fabric is approximately 50 mm.
Preferably there are at least three pairs of laterally spaced ninth layers.
Herein described is a fastening providing a mechanical and chemical attachment between at
least a first and second opposing surface, each of the surfaces having a plurality of holes extending
from an opposing side to a non opposing side, the fastening comprising an adhesive layer located
between the first and second surface and extending through at least some of the plurality of holes,
the ve layer having a head located on each non opposing side of the first and second surface.
Preferably the adhesive is a methyl methacrylates based epoxy adhesive.
Preferably the adhesive is at least one of ITW Plexus MA530, MA550, Clickbond, Permabond,
Parsons, ITW Ramset A7.
Preferably the adhesive has a viscosity of approximately of 0 cps.
Preferably the adhesive is mixed with a fibre additive.
Preferably additive is at least one of carbon, Kevlar or a composite glass fibre.
ably the gap between each opposing surface is not greater than 10 mm.
Preferably the gap between each opposing surface is approximately 3mm.
Preferably the fastening hole size is approximately 3 to 5 mm
Preferably the fastening the hole size is approximately 4 mm.
Preferably the fastening holes are spaced approximately 25 mm between hole centres.
Preferably the fastening holes are located first relative to an outside edge of a surface and spaced
inward.
Preferably the fastening holes located closest to an outside edge of a surface are approximately 10
mm from that outside side edge.
Preferably at least one of the fastening es is part of an aircraft structure or fuselage.
Preferably at least one of the fastening surfaces is an appendage d to be attached to an
aircraft structure or fuselage.
Herein described is a method of joining two es comprising providing at least two opposing
surfaces to be joined, providing a plurality of holes in each surface that extend from an opposing
side to a non opposing side, providing a layer of ve between the opposing es, wherein
the adhesive is of a type that chemically or mechanically bonds to the surfaces, the adhesive being
of a liquid form or at least paste, positioning the surfaces together such that the ve flows or
is ise forced through at least some of the plurality of holes from the ng side to a non
ng side, and the adhesive forming a head on the non opposing side, and curing the adhesive.
Preferably a portion the adhesive forms a ite rivet.
Preferably the method further ses mixing a fibre additive to the adhesive.
Preferably the step of oning the surfaces together comprises positioning the es not
greater than 10 mm apart.
Preferably the step of positioning the surfaces together comprises positioning the surfaces
approximately 3mm apart.
Preferably a method providing a plurality of holes in each surface comprises providing a size of
approximately 3 to 5 mm
Preferably a method providing a plurality of holes in each surface comprises providing a size of
approximately 4 mm.
Preferably a method wherein the holes are spaced approximately 25 mm between hole centres.
Preferably the holes are located first relative to an outside edge of a surface and spaced inward.
Preferably the holes located closest to an outside edge of a surface are approximately 10 mm from
that outside side edge.
Herein described is an aircraft e with appendages ed by a fastening mechanism
comprising providing a mechanical and chemical attachment between the aircraft surface and
appendage, the mechanism comprising a mechanical and chemical attachment between at least a
first and second opposing surface, each of the surfaces having a plurality of holes extending from
an opposing side of each of the aircraft surface and appendage to a non opposing side, sing
an adhesive layer located between the first and second surface and extending through at least some
of the plurality of holes, the adhesive layer, when cured, having a head located on each non
opposing side of the aircraft surface and appendage e.
Preferably the adhesive is a methyl methacrylates based epoxy adhesive.
Preferably the adhesive is at least one of ITW Plexus MA530, MA550, Clickbond, ond,
Parsons, ITW Ramset A7.
Preferably the adhesive has a viscosity of approximately of 150,000 cps.
ably the adhesive is mixed with a fibre additive.
Preferably an additive is at least one of carbon, Kevlar or a composite glass fibre.
Preferably the gap between each opposing surface is not greater than 10 mm.
Preferably the gap between each opposing surface is approximately 3mm.
Preferably the hole size is approximately 3 to 5 mm
Preferably the hole size is approximately 4 mm.
Preferably the holes are spaced approximately 25 mm between hole centres.
Preferably the holes are located first ve to an outside edge of a surface and spaced inward.
Preferably the holes located closest to an outside edge of a surface are approximately 10 mm from
that outside side edge.
The following embodiments may relate to any of the above s.
Other s of the invention may become apparent from the following description which is given
by way of example only and with reference to the anying drawings.
As used herein the term “and/or” means “and” or “or”, or both.
As used herein “(s)” following a noun means the plural and/or singular forms of the noun.
It is intended that reference to a range of numbers disclosed herein (for example, 1 to 10) also
incorporates reference to all rational numbers within that range (for example, 1, 1.1, 2, 3, 3.9, 4, 5,
6, 6.5, 7, 8, 9 and 10) and also any range of al numbers within that range (for example, 2 to 8,
1.5 to 5.5 and 3.1 to 4.7) and, therefore, all sub-ranges of all ranges expressly disclosed herein are
hereby expressly disclosed. These are only examples of what is specifically intended and all
le combinations of numerical values between the lowest value and the highest value
enumerated are to be considered to be expressly stated in this ation in a similar manner.
In this specification where reference has been made to patent specifications, other external
documents, or other sources of information, this is generally for the purpose of providing a context
for discussing the features of the ion. Unless specifically stated otherwise, reference to such
external documents is not to be construed as an admission that such nts, or such sources of
information, in any jurisdiction, are prior art, or form part of the common general dge in
the art.
The term “comprising” as used in this specification means “consisting at least in part of”. When
interpreting statements in this specification which include that term, the features, prefaced by that
term in each statement or claim, all need to be present but other features can also be present.
Related terms such as “comprise” and “comprised” are to be interpreted in the same manner.
The term “composite “as used in this specification means to a composition of two or more
materials including a matrix and reinforcement or at least one of the matrix or the reinforcement.
The term “matrix” means or includes a polymer material often called or interchangeably ed to
as a resin or resin solution. The term orcement” means or includes a fabric or fibrous
material including one or more constituents.
To those skilled in the art to which the invention relates, many changes in construction and widely
differing embodiments and applications of the invention will suggest themselves without departing
from the scope of the ion as defined in the appended . The disclosures and the
descriptions herein are purely illustrative and are not intended to be in any sense limiting.
BRIEF PTION OF THE DRAWINGS
The invention will now be described by way of example only and with reference to the drawings in
which:
Figure 1 shows a l empennage of a helicopter aircraft having one or more vertical stabiliser
fins.
Figure 2 shows two surfaces with an adhesive located between them.
Figure 3 shows a fin attached to a section of empennage using a coupling member.
Figure 4 shows an end view of a fin to be attached to an empennage having a plurality of holes.
Figure 5 shows the adhesive between upper opposing surface that may be, for e, a tail fin
and lower ng surface that may be, for example, an age.
Figure 6 shows the ve forms a head on the outward side of the surfaces.
Figures 7a and 7b respectively show perspective and front views of an energy absorbing seat
assembly for an aircraft, and particularly helicopters.
Figure 8 shows an exploded perspective view of the section of the seat.
Figure 9 shows a top cross nal view of the seat base.
Figure 10 shows a side cross sectional view of the seat base.
Figure 11 shows an exploded view of the seat base showing a most preferred arrangement of the
ural laminate layer.
Figure 12 shows a rear perspective view of the seat generally having a seat pan, seat back and a pair
of support ribs.
Figure 13 shows a rear view of the seat having cross sectional lines A-A and B-B.
Figure 14a shows in detail cross-section A-A.
Figure 14b shows in detail cross-section B-B.
Figure 15a shows two sections of an outer shell of a helicopter.
Figure 15b shows each section of the shell is preferably formed separately by a moulding process
then joined together to form a unitary fuselage structure.
Figure 16 shows a side view of the helicopter shell.
Figure 17 shows a member or bulkheads preferably installed in the rear of the central fuselage or
forward portion of the tail boom to provide al t.
Figure 18 shows a pair of members or keel beams having a forward protruding member.
Figure 19 shows a member or panel adapted to be inserted through the front window opening and
be located between the keel beams to thereby create a floor where ancillary items such as a fuel cell
may be placed or d.
Figure 20 shows a cabin floor panel which can be inserted through a window opening and
positioned above and joined to floor members.
Figure 21 shows a floor panel that can be inserted through the tion hatch delineated by the
top surface of the fuselage and optionally joined to the keel beams.
Figures 22 shows optional panel adapted to be d on the lower outer floor regions of the
central ge to facilitate a flat surface for cargo storage.
Figure 23 shows panel adapted to be upstanding on the rear region of the central fuselage to form a
ceiling to the space located below.
Figure 24 shows a floor panel adapted to be inserted through the opening in the top region of the
central fuselage section.
Figure 25 shows a panel adapted to be inserted through the opening in the top region of the central
fuselage section and close the internal region of the central fuselage from the tail boom region.
Figures 26 and 27 show front windows and cabin roof adapted to attach to openings ated by
the fuselage.
Figure 28 shows a plurality of covers adapted to close window openings and inspection hatches
delineated by the fuselage.
Figure 29 is a side view of one preferred embodiment of the helicopter fuselage.
Figure 30 is a bottom view of one preferred embodiment of the helicopter fuselage.
Figure 31 shows a side view of the preferred helicopter fuselage 300 and age 336 showing
an area where an additional layer of fabric 332 may be applied to the laminate structure.
Figure 32 shows a bottom view of the preferred helicopter and in particular illustrates a plurality of
openings delineated by the fuselage.
DETAILED PTION OF THE INVENTION
Figure 1 shows a typical empennage of a helicopter aircraft 100 having one or more vertical
stabiliser fins 101 and one or more horizontal stabiliser fins 102. Stabilising fins are primarily
required on air craft to provide aerodynamic stability but may provide secondary functions such as
housing devices such as sensors and lights or mechanisms used for ft flight control.
A prior art method of attaching age appendages such as fins 101, 102 to a fuselage or
aircraft such as the pter tail boom 103 and fan duct structure 104 is with the use of a
mechanical fastening such as rivets, screws or bolts. Mechanical fastenings such as riveted joints or
bolted joints are known to be labour intensive and require the use of l fixtures and tooling
jigs.
A disadvantage associated with use of ical fastenings to secure such appendages is that each
of the adjoining surfaces must have a plurality of holes formed for the fastening to pass though.
Such holes can cause weakening of the structure and may contribute to a point of structural failure.
To mitigate the risk of structural failure such methods of fastening often require regular
maintenance checks to ensure structural integrity is ined, particularly for any cracks that may
be propagating between adjacent holes.
Mechanical fastenings have a further disadvantage in that substantial damage may be caused to an
aircraft by tearing the surrounding material should a secured appendage be struck by some external
object.
Mechanical fastenings have a further disadvantage when attaching to curved surfaces together. The
mismatch in shapes between the curved surface and the generally flat fastener may create undue
stress on a location immediate to the fastener.
A further disadvantage is that mechanically fastened surfaces or riveted es are prone to
sealing issues where moisture can ingress or be trapped.
A further disadvantage is that mechanically fastened surfaces are prone to various types of
corrosion. rm, intergranular and surface corrosion can form between mechanically fastened
surfaces. Often, corrosion in these areas goes undetected even with ic maintenance,
dismantling and inspections and can result in catastrophic failure of the ing or region
proximate the fastening.
One way to fasten empennage appendages to a ge or aircraft structure and overcome the
problem of cracking which is associated with using traditional mechanical fasteners is by the use of
a al bond such as methyl methacrylates or a mechanical adhesive. Figure 2 shows two
surfaces 105 with an adhesive 106 located between them. Such adhesives are initially a liquid or
paste to allow them to flow between the surfaces 105 before chemically reacting with the surfaces
105 and curing to create a rigid bond. Each of the surfaces 105 should have a substantially nt
shape to ensure the best possible bond, but are not required to be a flush fit. When the adhesive
106 is d to the surfaces 105 the assembly is oned together in the direction of the
indicating arrows and the adhesive is allowed to cure. The adhesive typically requires a al
exothermic reaction for cure.
A disadvantage of using a chemical bond to join an appendage to an aircraft structure is that such
bonds typically e only shear strength. The chemical bond may still be prone to failure if
subjected to a tensile pressure or a peel pressure should a secured appendage be struck by some
external object.
One preferred embodiment of the invention relates to an attachment or method of attaching
appendages to a fuselage or aircraft structure which overcomes or at least rates at least one
of the abovementioned disadvantages or which at least provides the public with a useful . In
red embodiments the attachment or attachment method is used to attach empennage
appendages to an me, but may also include attaching various other items including brackets
and antennas.
Figure 3 shows a fin 101 attached to a section of empennage 104 using a coupling member 108.
The coupling member 108 is an integral combination of a mechanical fastening provided by a head
107 and a chemical bond provided by a cured adhesive 109.
To couple the es it is preferable, but not necessary that each of the surfaces are substantially
flush fitting with one another or at least have a substantially coherent shape.
Preferably the coupling member 108 is formed by at least some of the following steps:
1. The area the opposing surfaces to be joined is marked or otherwise noted.
2. The area the opposing es to be joined has a plurality of holes formed in each surface.
The holes are not required to be aligned. Figure 4 shows an end view of a fin to be
attached to an empennage having a ity of holes 111.
3. A portion of ve is applied between each of the opposing es. Figure 5 shows
the adhesive 109 between upper opposing surface that may be, for example, a tail fin 101
and lower opposing surface that may be, for example, an empennage 104. The adhesive
may be a glue or resin and may include fibres or be a fibre composite.
4. The opposing surfaces 101, 104 are then positioned relative to one another in the direction
indicated F. The positioning of the opposing surfaces 101, 104 causes adhesive 109 to be
forced through holes 111. As the adhesive flows through holes 111 surface attraction (via
the Coandă effect) causes the flow of the adhesive to follow the contour of the surface
h the hole and, having passed h the hole 111, flows radially outward.
Adhesive flowing beyond the extremities of the fin 101 may be coved or otherwise shaped to
e a smooth and namically favourable transition 110 between the fin 101 and
age 104.
The adhesive 109 forms a head 107 on the outward side of the surfaces as shown in Figure 6.
When the adhesive is cured the head 107 creates a coupling member 108. The coupling member
108 provides a mechanical fastening while the surfaces and resultant ly also enjoy a chemical
bond. The coupling member provides strength to resist both shear forces and tension forces by a
greater measure than mere bonding or strictly mechanical fasteners would otherwise provide.
Forming the coupling member to join two surfaces is particularly useful when the surfaces are
formed by a substrate and resin. The adhesive 109 may be the same or similar material to the resin
such that the assembly, when cured, becomes a structure of substantially uniform material.
A further age ed by the coupling member 108 is that improves any repair process that
may be required to take place, for example, if an appendage is required to be ed. Replacement
may be required due to fatigue weakening the appendage or by damage from an inadvertent strike.
The damaged appendage can be cut from the empennage using any known method and the
empennage ground or sanded to reproduce its original shape and r. Once the empennage
has been restored, a new appendage can be prepared and attached to the empennage by forming a
new coupling member 108 using the abovementioned process. Any material er from old
coupling members does not substantially affect the formation of new coupling members.
Preferably the adhesive is ITW Plexus MA530 or MA550. However, similar ves such as
Clickbond, Permabond, Parsons, ITW Ramset A7 or other methyl methacrylates based epoxy
adhesives.
ably the viscosity of the adhesive is approximately of 150,000 cps.
Optionally, the adhesive is mixed with a fibre additive. Preferably that additive is carbon, or
alternatively may be Kevlar or a composite glass fibre.
Preferably the gap between each positioned surface is not greater than 10 mm. Preferably the gap
between each positioned surface is approximately 3mm gap. Preferably the surfaces and can go up
to 10mm on fairing edges.
The hole size may be 3 to 5 mm and is preferably imately 4 mm. ably the holes are
located approximately 25 mm between hole centres. Preferably the holes are located first relative to
an outside edge of a surface and worked inward. Preferably the outermost holes are approximately
mm from an outer side edge.
Traditionally helicopter crew and passenger seats have been inbuilt ures with the helicopter
fuselage. Later years have seen crew and passenger seats evolve into stand-alone assemblies for the
forward seats and foldaway seats for the rear passengers. ication standards require the
inclusion of a orthy seat for all occupants of the helicopter.
More recently since the introduction of new certification rules, seats in newly ied helicopters
are required to be “Crashworthy” g n design parameters of maximum load factors,
inertial forces, and reactions between occupant, seat, and safety belt or harness corresponding with
the applicable flight and ground load conditions, including the ncy landing conditions of the
category in which certification is sought.
As a result there have been several newly design crashworthy seats led into helicopters as new
helicopter designs or its to older helicopter designs. These new seat designs incorporate
designs of shock absorber, collapsing lever mechanisms, brake, energy absorbing foams, and
collapsible metal structures.
One seat design in the prior art for meeting the crashworthy seat standard is known as a stroking
seat mechanism. Disadvantages of the stroking seat mechanism include the requirement for regular
inspection and servicing, corrosion protection for metal surfaces, inadvertent g of the seat
action and injuries to limbs that occur during the stroke of seat.
Another design in the prior art for meeting the crashworthy seat standard is known as a braking
seat mechanism. Disadvantages of this seat mechanism include the friction pad loosing preload
over time and the regular requirement for inspection and readjustment, the metal frame requires
corrosion protection, , rtent jamming of the seat action and injuries to limbs that occur
during the stroke of seat.
Another design in the prior art for meeting the crashworthy seat rd is an ium sheet
metal box design. Disadvantages of this seat mechanism include allowing the occupant to fall
through the seat pan into the seat base. While the occupant may survive the crash, evidence has
shown the occupant is subsequently trapped in the seat base and unable to escape the crashed
aircraft.
One preferred form of the invention is a crashworthy seat for nts of the helicopter that
overcomes or ameliorates at least one of the abovementioned disadvantages or at least provides the
public with a useful choice. Preferably the crashworthy seat also meets current certification
requirements for energy absorbing seat designs of 30 G downwards and 18.5 G forwards.
Another aspect of the invention is a seat residing on a mechanism that provides a vertical distance
to allow crushing. In an embodiment of the invention the seat is an all composite upper seat
comprising a seat pan and seat back urally bonded to a crushable composite seat base. The
seat is intended to be compatibly located in both d and aft cockpit positions.
Figures 7a and 7b respectively show ctive and front views of an energy absorbing seat
assembly 200 for an aircraft, and ularly helicopters. The seat 200 is generally constructed with
an upper composite laminate section 202 and a lower composite section 201. Each composite
section is ucted with a particular arrangement of fabric layers that are infused and bonded
together with a curable resin.
The lower composite section 201, or base, is crushable to aircraft impact energy to go at least some
way toward mitigating injuries nable by the seat occupant during a hard landing or crash.
Figure 8 shows an ed perspective view of the section of the seat 200. Preferably the upper
and lower seat sections are individually assembled and constructed then subsequently joined
together. An energy absorbing member 203, such as foam, may be installed into the open central
area of the seat base 201 to further improve the impact energy absorption properties of the seat.
Preferably the energy absorbing foam is Confor CF45, Confor CF47 or similar.
In the event energy absorbing characteristics are, for example in an experimental or r built
aircraft, and the foam 203 is not ed, the open central area of the seat base 201 a hatch
opening 209 and lid 204 may be provided to allow, for example, the open central area to be used
for luggage storage.
In use, of the seat is anchored through the lower composite section 201 to an aircraft cabin floor
and is ible in both forward and aft cabin positions A seat occupant is ined into the
seat in use by a seat belt assembly sing lap and shoulder belts which may be attached to the
seat.
The preferred arrangement of fabric layers for the lower composite section or seat base 201 will
now be described with nce to figures 9 to 11. Figure 9 is a top cross sectional view of the
seat base 201 and Figure 10 is a side cross sectional view of the seat base 201. Figure 11 is an
exploded view of the seat base 201 showing a most preferred arrangement of the structural
laminate layer. Preferably the seat base is constructed with a ity of fabric layers which are
infused and bonded together with resin to form a composite or .
Preferably the seat base 201 has a side structure with an outer layer of fabric 204 and an inner layer
of fabric 205 that extends around the outer periphery. An intermediate layer 206 is located between
the inner and outer periphery layers 204, 205 on each of the front, back and side surfaces. When a
hatch g 209 in the seat base 201 is desired, a layer of fabric 207 is required to be layered
around that opening and preferably on the internal surface.
Preferably the intermediate layer 206 is not continuous between each surface, that is, the
intermediate layer does not wrap around the ally orientated corner regions 208 of the seat
base. The corner regions 208 have a gap or thin section that provides a point of weakness in each
corner of the structure to thereby provide a local buckling mode when under pressure from a
substantially vertical direction. Figure 10 shows the arrangement of fabric base layers where an
upper layer 212 and lower layer 213 are arranged either side of an intermediate layer 214.
Preferably each of the outer edges of the base layers and lower edges of the side layers has a section
of fabric that provides a region of overlap for additional support. Preferably the overlapping
s of the side fabrics extend approximately 50 mm across the base surface and the base fabrics
extend approximately 30 mm up the side surfaces.
It should be appreciated that regions are shown to be pping, that region may instead be a
continuous portion of fabric that extends between base and side walls. Similarly, where regions of
fabric are shown to be continuous, those regions may instead be overlapping sections of .
Preferably any regions of overlap are approximately 30 to 50 mm in size.
ably the inner and outer periphery layers 204, 205 and upper and lower base layers 212, 213
are ucted with a resin infused bidirectional fabric such as carbon and/or Kevlar (CKC) 0/90,
bidirectional or coaxial type fabrics. The inner and outer periphery layers 204, 205 e
mechanical strength to the assembly. Preferably the CKC fabric is a Carbon and Kevlar fabric of
approximately 180 g/m2. Preferably the ediate layer 206 is a resin infused continuous
filament mat (CFM) fabric. The intermediate layers 206 help to provide additional rigidity to the
inner and outer periphery layers. Preferably the CFM fabric is approximately 450 g/m2.
A layer 210 is provided on the inner periphery of the upper edge of the seat base 201 for additional
support in the region where the seat pan 202 is to be joined to the seat base. Preferably the layer
210 also extends upward at the rear section of the seat base for providing onal support to the
back section of the seat pan 202. Preferably the additional layers 210 extend approximately 30 mm
beyond the flange where the seat base 201 meets the seat pan 202. ably the upper support
layer 210 and hatch support layer 207 are CFM fabric and approximately 450 g/m2.
Members 215 may be provided to facilitate a secure mount through which fasteners can be located
to attach the seat base to the floor of an aircraft cabin. s 215 are preferably constructed
with 75 mm wide carbon double bias fabric of approximately 400 g/mm2.
The preferred arrangement of fabric layers for the upper composite section or seat 202 will now be
described with reference to figures 12 to 15. Figure 12 shows a rear perspective view of the seat
202 generally having a seat pan 220, seat back 221 and a pair of support ribs 222. The t ribs
extend from the top of the seat back 221 to the front portion of the pan 220. The support ribs help
to maintain the ty of the seat structure in a survivable crash such that the seat back is
substantially restrained from folding toward the seat pan. Preferably the seat 202 is constructed
from a plurality of fabric layers which are infused and bonded together with resin to form a
composite or matrix. Preferably the seat is bonded to the seat base using a methyl methacrylate
structural adhesive such as ITW Plexus MA530 or MA550. However, ves such as Clickbond,
Permabond, Parsons, ITW Ramset A7 or other methyl methacrylates based epoxy adhesives are
suitable alternatives.
Figure 13 shows a rear view of the seat having cross sectional lines A-A and B-B. Cross-section A-
A is shown in detail in Figure 14a. Cross-section B-B is shown in detail in Figure 14b. Generally,
the seat is constructed from alternating layers of structural fabrics such as CKC and fabrics such as
CFM. A core material such as foam is used to space the layers of fabric apart to enable a composite
I-beam type structure in the central portion of the seat 202 structure, r, the foam maybe
d when using a spilt two mould method of infusion. Preferably each of the layers is arranged
to continuously extend from the upper portion of the seat back 221 to the forward portion of the
seat pan 220. However, it should be noted that the seat back and seat pan may be constructed
using non-continuous sections of fabric and those non-continuous section joined with pping
sections as described in on to the seat base above.
A layer of CFM is used as a flow medium. The CFM layer does not contribute significant th
in itself but is used as to separate two structural materials to create a sandwich construction. Resin
flows through and around carbon are problematic. The use of CFM as a flow medium addresses
and alleviates problems with the gap and flow.
As mentioned, CKC is a material including carbon and Kelvar materials. The carbon material is
used to provide strength and ess. The Kevlar material is used for its energy ing
characteristics or as a labyrinth around the carbon. When the carbon structure fails mechanically
the Kevlar holds the structure together and contains the crush. The Kevlar also prevents carbon
fragments from becoming spears or cutting edges which could seriously injure an occupant.
The seat back and pan has a first surface layer 223 and second surface layer 224. Preferably the first
surface layer 223 is of a al that provides mechanical strength such as CKC and approximately
180 g/mm2. Preferably the second layer 224 is of a material such as CFM and approximately 300
g/mm2.
A third layer 225 is pair of laterally spaced portions of fabric, each approximately 100 mm in width.
The third layer may consist of more than one layer, and preferably two layers. Preferably the third
layer is of a material such as Carbon uni (CU) and approximately 300 g/mm2.
Fifth and sixth layers 226 and 227 are ed behind the thirds layer or layers. Preferably the fifth
e layer 226 is of a material that provides mechanical strength such as CKC and approximately
180 g/mm2. Preferably the sixth layer 227 is of a material such as CFM and approximately 300
g/mm2.
Two laterally spaced core members 228 are provided behind the sixth layer as an optional. The
core layers provide a gap between the top and bottom layers to improve the mechanical strength of
the assembly. Preferably the core members 228 are a PVC foam material or similar. The core
member 228 may be d in a two piece split infusion mould system.
A seventh and eighth layer 229, 234 are provided behind the core members 228. ably the
seventh layer 229 is of a material such as CFM and approximately 300 g/mm2. Preferably the
eighth surface layer 234 is of a material that es mechanical strength such as CKC and
approximately 180 g/mm2.
A plurality of ninth layers 230 are laterally spaced and aligned with the core s 228. The nine
layers 230 are ed to give additional th to the seat back and pan. Preferably the ninth
layers 230 comprises three laterally spaced pairs of fabric. Preferably the ninth layer fabric is
Carbon Uni and approximately 300 g/mm2. Preferably the width of each section of fabric is
approximately 50 mm.
Tenth and eleventh layers 231, 232 close the rear side of the seat back and pan. Preferably the tenth
layer 231 is CFM of approximately 300 g/mm2. Preferably the eleventh layer is CKC and
approximately 180 g/mm2.
The seat may be d as a single action including all front, core and back laminate materials or
the seat may be infused as a two process two piece split mould system where the forward half of
the seat is infused as one part, the core material omitted and the aft half of the seat is infused as a
second part. The two parts are then ally bonded together using a methyl metacrylic
structural adhesive.
The seat may be upholstered for comfort. Preferably the seat is upholstered using energy absorbing
CF45 and CF47 foams as a cushion on the seat base to provide comfort and energy absorption
properties.
The assembled seat and seat base provides a collapsible or crushable structure to absorb energy
when an aircraft impacts the ground. The ng process occurs in a preferred sequence; firstly
the rear panel of the seat base will buckle; secondly the side panels will buckle; thirdly the front
panel will buckle. Each of the corner regions 208 allows each of the seat base side panels to buckle
substantially without causing buckling of the other panels.
If the loads on the seat and seat base during an aircraft impact are not substantial enough to buckle
any or all of the seat base side walls, the impact loads will be distributed to other areas such as
energy absorbing foam that may line the seat back and cushion.
ages offered by this embodiment include a construction that is light weight. The seat is less
than half the weight of conventional stroking or a braking seat design while meeting the FAA FAR
Part 27 crashworthy seat standard.
Preferably the energy absorbing seat assembly comprising an upper engineered composite laminate
section and a lower engineered composite crushable section, wherein the use of the seat is
anchored h the lower section to the cabin floor. Preferably the composite upper seat
comprises a seat pan and seat back structurally bonded to a crushable composite seat base.
Preferably the crushable seat base is compatible in both forward and aft positions. ably the
structurally engineered seat comprises a seat pan and seat back which is an engineered laminate to
produce calculated mechanical properties. Preferably the occupant is restrained into the seat in use
by a seat belt assembly comprising lap and shoulder belts which are attached to the seat. ably
the base n comprises structurally engineered seat base which is an engineered laminate to
produce calculated mechanical properties and a designed order of collapse under certain load
conditions. ably the purpose of the carbon laminate is to provide stiffness. Preferably the
purpose of the Kevlar is to contain the carbon laminate during crushing. Preferably the engineered
laminate comprises at least one carbon fibre laminate and at least one arimid fibre laminate.
ably the crashworthy seat is designed to meet the certification requirements of Federal
Aviation Authority (FAA). Federal Aviation Regulations (FAR), Title 14. Aeronautics and Space.
Part 27. Airworthiness Standards ; Normal Category Rotorcraft. Preferably the seat comprising
seat pan and seat back laminates are d using the RTM process. Preferably the seat base
laminates are infused using the RTM. Preferably the seat is bonded to the seat base using structural
adhesive HPR25A/B. Preferably energy absorbing foam Confor CF45 and Confor CF47 is
installed in the seat base to further assist energy tion. Preferably the seat is upholstered using
energy absorbing CF45 and CF47 foams in the seat base cushion. Failure is predicted in the
following modes and sequence - back panel buckle, side panel , and global buckle.
ably the local buckling modes (panels) do not result in ultimate failure if ing ure
can support buckled shape and redistribute load to other areas.
Prior art helicopter structures or ork ses materials such as wood, aluminium,
titanium, chrome molybdenum steel tubing and magnesium alloys. Fabrication and manufacturing
of helicopter structures are based around the extensive use of special jigs and frames and certified
holding fixtures where floors and jigs, frames and fixtures are frequently calibrated. Such
installations have inherent disadvantages including that they are fixed in a location and are not
mobile.
The prior art fuselage fabrication and cturing process requires the fuselage to be
ucted by first assembling internal ents and working outward. The prior art method
of constructing a helicopter begins by identifying a starting location or part such as a central floor
panel. The internal structure of the fuselage is then added systematically around that ng
on by adding sub frames and panels. The assembly is then strengthened by riveting or bolting
adjoining sub frames and panels to form a skeleton. Once all of the internal structure has been
completed the fuselage skeleton is enclosed with a skin that is either riveted or bolted into place,
usually by direct ment to the skeleton. When the primary structure of the fuselage is
completed and the fuselage is urally sound it would be removed from the fabrication or
manufacturing assembly jig.
Traditional helicopter fuselage manufacturing has us disadvantages. Of substantial note is
that construction is extremely labour intensive. The ted fuselage has a vast number of
dual parts, each requiring prior fabrication. To track and assemble these parts requires a
skilled work force. Further, the fabrication jigs have long set-up times and long breakdown times.
Production of helicopter ge in traditional manner is very expensive.
A further disadvantage to traditional helicopter manufacturing is the finished external surface of
the helicopter fuselage d in a mass of domed rivet heads. This type of finish is both
unattractive and s in high drag penalties. Significant materials cost and time is associated with
the use of flush head rivets in the outside skin of the fuselage to remove the drag penalty.
A further antage to traditional helicopter fuselage fabrication using sheet metal panels to
form the outer skin of the fuselage is the difficulty in achieving a smooth and thus aerodynamically
favourable shape.
A further disadvantage to traditional helicopter fuselage fabrication is that door and window
openings are typically hand finished. Finishing by hand results in no two door or window openings
being the same. Each window or door therefore requires individual shaping, usually by hand, to
ensure a fitment that allows closure without gaps.
A further disadvantage to traditional helicopter fuselage manufacturing using riveted structures and
thus lapped joints is the ingress of moisture. This moisture becomes trapped and corrosion will
ensue. Corrosion can lead to structural failure.
According to one embodiment the invention is a ge structure for an aircraft, and preferably a
helicopter. Figures 15 to 32 show the preferred helicopter fuselage, preferred components of that
fuselage and preferred process of assembly. The preferred fuselage forms the outer shell of the
helicopter and the outer shell is a load bearing structure. In this specification a load bearing
fuselage means the outer shell of the helicopter provides primary structural support for the
helicopter, including, but not limited to, supporting the engine and gearbox mechanisms, a cabin,
fuel cells, a tail boom and tail rotor assembly and an empennage. The preferred fuselage is
constructed from a ite, that is, a lamination of a plurality of fabric layers that are, or are to
be d with a r, including Polyester, Vinyl ester and Epoxy resins, that is cured or can be
cured to lock the layers together in a substantially rigid formation. Preferably the resins are UV
stable.
The ite fuselage ure may offer advantages such as reducing manufacturing effort
compared to traditional metal frame based aircraft. The composite fuselage is preferably formed, at
least in part, in a mould. A moulded airframe offers advantages such as repeatability of part size
and shape. Repeatability of parts goes at least some way toward ensuring accurate fitment of
adjoining parts or assemblies, thereby somewhat ameliorating the requirement for individual
attention to the t of adjoining parts which is often prevalent in traditional metal frame based
aircraft.
A further advantage that may be provided by the composite ge ure is durability. A
traditional metal frame based aircraft has a working life that is somewhat dictated by the es
imparted during flight and by engine and working surface d ion. Metal is known to
eventually crack in such environments, therefore dictating strict flight hour records to ensure the
aircraft materials are not used beyond a safe duration. The preferred composite fuselage has an
improved rate of oration ed with a metal fuselage.
A r advantage that may be provided by the composite fuselage structure is replacement of
some or all of the fuselage should mechanical damage occur, for example, by bird strike or
inadvertent strike from a ground based object. A damaged section of composite fuselage can be cut
and a ement composite section installed. The preferred replacement of a section of fuselage
includes cutting a damaged n from the airframe, preparing a section of the me that
substantially corresponds to the removed section and installing the section of airframe, preferably
with a region of fabric in the composite lamination structure to that overlaps with the existing
ure.
A further advantage that may be provided by the ite fuselage structure is the prepositioning
of key identifiers within the mould such as alignment stations, dimension marks and
one or more locating datums.
A r advantage that may be provided by the composite fuselage structure is the prepositioning
of window rebates, cabin and cargo door rebates and inspection panel rebates.
A further advantage that may be provided by the composite fuselage ure is the moulding
inclusion of opening sealing flanges for windows, cabin and cargo doors and inspection panels and
hatches.
Replacing a section of the fuselage also includes replacing large fuselage regions such as the tail
boom. To replace a tail boom, the ng tail boom is cut from the fuselage in the region
proximate the central fuselage section, or at least proximate a damaged section of the tail boom. A
replacement tail boom that substantially corresponds to the removed tail boom can be prepared
and abutted to the remaining fuselage structure. The grafting process includes ing section of
composite material and overlapping that composite material across the abutting region.
Figure 15(a) shows two sections of an outer shell of a helicopter 301, 302. Each n of the
shell is preferably formed tely by a moulding process then joined together to form a unitary
fuselage structure 300 as shown in Figure 15(b).
Preferably the fuselage sections 301, 302 are made of a composite material and by a moulding
process. Preferably the composite material has several layers of fabric including an outer layer of
CFM of approximately 300 g/mm2, a next layer CKC of approximately 190 g/mm2, a layer of any
additional supporting material, a foam or filler layer such as soric of approximately 2 mm, a next
layer CKC of imately 190 g/mm2 and a next layer of CFM of approximately 300 g/mm2.
The outer layer of CSM forms layer is to protect against impact and abrasion . Optionally a
layer of gel coat is d to the mould e before the laminate structure is laid up to promote
a smooth outer fuselage surface and ease of release from the mould.
Preferably each of the fabric layers is laid into a mould that substantially conforms to the shape of
the outer surface of the desired fuselage sections 301, 302. The layers of fabric are then infused
with a UV stable resin ition and optionally subjected to a vacuum s to lock the
structure together. Several methods of resin infusion can be used including a single vacuum bag
method, dual vacuum bag method, silicon bag method or Light RTM contra mould method. The
cured layers remain in their respective moulds. The moulds are then rotated and joined together
either by bolting or clamping. The infused sections 301, 302 are then laminated together. The
laminating joining process is performed either by infusion or hand lay. Once the whole of the
fuselage has cured the mould halves are withdrawn from the completed fuselage structure.
The preferred layup process for each ge half is as follows, and the opposing mould half
would be similar. The ge mould is preferably prepared by ng, polishing and the
application of a release agent. The mould is optionally sprayed with a sacrificial gel coat
approximately 0.018” to 0.022” thick and the gel coat allowed to cure. The fuselage outer layer or
shell laminate preferably comprises CFM 300, CKC 190, plus additional laminates, which are
location specific additional laminates throughout the length of the fuselage plus Soric 2mm, CKC
190 plus CFM 300. Additional laminates may be desired to strengthen or stiffen specific areas of
the structure dependant on loads and load paths.
Once all laminates are correctly laid and additional laminates correctly positioned the ge
mould is sealed with an optional single vacuum bag, a dual vacuum bag, a silicon vacuum bag or a
Light RTM conta mold. Preferably vacuum is applied at 100%. Preferably vacuum leak tests are
carried out. Preferably the entire ge is left under vacuum for a minimum of 12 hrs. The
fuselage mould is preferably heated during this time to +30oC. The inside of the ge mould is
also heated to +30oC. The preferred resin is a vinyl ester and epoxy blend resin Derakane 510C-
350FR is activated with 0.2% cobalt and retarded 0.7%. Preferably the resin is heated to +30deg c.
Immediately prior to on beginning a final leak test is carried out on the fuselage laminate
under vacuum. Leak test is to be <2-3mb / minute. Preferably the infusion resin as prepared is
catalysed at 2% immediately prior to infusion in sufficient quantities to maintain infusion of the
ge in a single smooth process.
Infusion is preferably performed in a specific order to ensure complete infusion and to avoid
lockout by beginning at the t part of the mould and progressing to the lowest vacuum port,
then progressively further resin valves are opened to influence the resin flow. The d and
joined fuselage outer shell or fuselage can be constructed, in a size suitable for a six person
helicopter, to less than 100kg.
Figure 16 shows a side view of the helicopter shell. Preferably the shell 300 has a central fuselage
region 303 and tail region 304 that are integrally formed. Preferably the central fuselage region 303
is adapted a support at least an engine and drive train, and may further support fuel tanks and
structural members such as beams and bulkheads. The drive train es the main rotor gear box
that is adapted to drive the main rotor and connect to a shaft that drives the tail rotor.
The fuselage outer shell is a hollow ess structure when d from the fuselage .
To complete a helicopter assembly various members including keel beams, floor panels and
bulkheads are to be installed in a red order such that the flight loads imparted to the fuselage
and those loads from the main and tail rotor systems and engine are distributed about the
helicopter airframe. Each member has a size and location within the ge to adequately
distribute that load. Each member is inserted in order into the hollow fuselage through an opening
delineated by the outer shell such as an opening for a cabin door. Each member is trimmed to size
and once inside the fuselage is permanently bonded in place.
The fuselage is preferably assembled in the following order: Tail boom forward bulkhead 311, Fuel
tank aft bulkhead 312, Keel beams 313, Keel beams 314, Transverse beams 315, fuel bay floor 317,
aft cabin bulkhead 318, cabin floor 319, upper bulkhead 321, central cargo floor 323, side cargo
hold floors 324, side cargo hold roof panels 325, aft hold floor 326, engine drip tray 327,
windscreens and windscreen mullion 328 and roof window panels 329.
Preferably the tail region 304 includes at least a tail boom. However, the tail region may also
include items such as a tail rotor mount and/or an empennage and/or flight isation
appendages.
The shell 300 may also e an integrally formed forward section 305 having a cabin for housing
a pilot, passengers and flight controls. Preferably the shell 300 includes an ally formed
forward section 305 and delineates a plurality of openings for windows 307, doors 306, inspection
hatches 309, 310, storage areas 308 and the like. Preferably the openings allow all other
ents of the pter to be installed. For example, items such as bulkheads, beams, walls
and supporting structure, together with ancillary components such as flight control mechanisms,
fuel cells, engines and drive train are to be led after the fuselage sections 301, 302 are joined.
Figures 17 to 28 show a number of components preferably installed internal to the fuselage 300.
The components are installed in a preferred order.
Figure 17 shows a member or bulkheads 311, 312 preferably installed in the rear of the central
fuselage or forward portion of the tail boom 304 to provide internal t. Preferably the
bulkheads 311, 312 are installed through the opening 313 defined at the top of the fuselage.
Bulkheads 311, 312 are preferably installed to e support or ng locations for an engine
and rotor gearbox.
Figure 18 shows a pair of members or keel beams 313 having a forward protruding member 316.
Preferably the keel beams are adapted to extend from the upper internal surface of the central
section of the fuselage to the lower internal surface. Preferably the members 316 are adapted to
extend d to the cabin area 305 to e support for a cabin floor. Preferably s 314
are adapted to compliment members 316 and provide further support for a cabin floor. Preferably
members 314 are adapted extend front the front region of the cabin to the rear region of the
central fuselage section. Preferably members 315 are adapted to extend transversely to members
316 and provide further t for a cabin floor and seats installation. Preferably each of the
members 314, 315, 316 are adapted to mesh or join with at least some overlap to form a lattice like
structure that s through the lower region of the cabin space 304. Preferably the members
314, 315, 316 are shaped such that they substantially conform to the contour of the lower inner
surface of the fuselage 300 such that the structure can be bonded together to form a substantially
rigid structure. Preferably each of the members 314 to 316 are adapted to be inserted into the
internal region of the fuselage through an g such as the front window opening 307 or a door
opening 306. The upper region of the beams 313 are preferably adapted to connect to a engine
and/or gearbox structure and distribute loads associated with the engine and gearbox to the
ge walls.
Figure 19 shows a member or panel 317 adapted to be inserted through the front window opening
307 and be located between the keel beams 313 to y create a floor where ancillary items such
as a fuel cell may be placed or mounted.
Figure 20 shows a cabin floor panel 319 which can be inserted through a window opening 307 and
positioned above and joined to floor members 314, 315, 316. Bulkhead panel 318 extends from a
region closer to top inner surface of the fuselage 300 to the bottom inner e, and preferably
also spans between the internal surfaces of the fuselage side walls. Panel 318 optionally includes an
inspection hatch 322 that and complimentary hatch cover 320. A further panel 321 is adapted to
attach between the upper edge of the panel 318 and the top inner e of the fuselage.
Preferably panel 321 is removable to facilitate inspection of the fuselage area immediately aft,
which may house an engine or drive mechanism. Preferably the beams 313, floor members and
bulkhead panel 318 form an enclosure for ing one or more fuel cells housed within the
enclosure from impact energy.
Figure 21 shows a floor panel 323 that can be ed through the inspection hatch delineated by
the top surface of the fuselage 300 and optionally joined to the keel beams 313.
Figures 22 shows optional panel 324 adapted to be located on the lower outer flor regions of the
central fuselage to facilitate a flat surface for cargo storage. Figure 23 shows panel 325 adapted to
be ding on the rear region of the central ge to form a ceiling to the space located
below.
Figure 24 shows a floor panel 326 adapted to be inserted through the opening in the top region of
the central ge section. Preferably the floor panel is adapted to join to the joined to the inner
surface of the fuselage and surrounding structure.
Figure 25 shows a panel 327 d to be inserted through the opening in the top region of the
central fuselage section and close the internal region of the central fuselage from the tail boom
Preferably the internal structure including panels, walls, member and beams 311 – 327 are made of
a composite material and by a moulding s. Preferably the composite material has several
layers of fabric including an outer layer of CFM of approximately 300 g/mm2, a next layer CKC of
approximately 190 g/mm2, a layer of any additional supporting al, a foam or filler layer such
as soric of approximately 2 mm, a next layer CKC of approximately 190 g/mm2 and a next layer of
CFM of approximately 300 g/mm2. The outer layer of CSM forms layer is to protect against impact
and abrasion damage. Optionally a layer of gel coat is applied to the mould surface before the
laminate structure is laid up to promote a smooth outer fuselage surface and ease of release from
the mould.
The members are y flat panel sections including primary structure, secondary structure and
tertiary structure and are an engineered laminate specific for each parts purpose. Primary ure
is direct load bearing structure design to carry flight engine and transmission loads through the
airframe. An example of primary structure is the keel beam 313 and fuselage structure 301 and 302.
Secondary structure carries indirect loads and shared loads from primary ure through the
me. An example of secondary structure is the floor 323. Tertiary structure is minor load
carrying structure the failure of which would not lead to subsequent failure of secondary or y
structure. An example of tertiary structure is the panel 327.
A typical layup process for each flat panel would be as follows. The laminating table is prepared by
cleaning, polishing and the application of a release agent. The table is then optionally sprayed with
a sacrificial gel coat approximately 0.018” to 0.022” thick. The gel coat is allowed to cure. Each flat
panel differs in its laminate structure. The basic flat panel laminate may have additional laminates,
which are location specific throughout the panel. Additional laminates may be desired to
strengthen or stiffen ic areas of the structure dependant on loads and load paths.
Once all laminates are correctly laid and onal laminates tly positioned the table mould is
sealed with an optional single vacuum bag, a dual vacuum bag, a silicon vacuum bag or a Light
RTM conta mold. Vacuum is then applied at 100%, vacuum leak tests are carried out and the entire
table is left under vacuum for a minimum of 4 hrs. The table mould is heated during this time to
+30oC. The vinylester / epoxy blend resin Derakane 510C-350FR is activated with 0.2% cobalt and
retarded 0.7%. The resin is heated to +30oC. Immediately prior to infusion beginning a final leak
test is carried out on the table te under vacuum. Leak test is to be <2-3mb / minute. The
infusion resin as ed is catalysed at 2% immediately prior to infusion in sufficient quantities to
maintain infusion of the fuselage in a single smooth process. Infusion must be done in a specific
order to ensure complete infusion and to avoid lockout. Infusion begins at the deepest part of the
laminate and is allowed to progress, then progressively further resin valves are opened to influence
the resin flow.
Figure 26 and 27 show front s 328 and cabin roof 329 d to attach to openings
delineated by the fuselage.
Figure 28 shows a plurality of covers adapted to close window gs and inspection hatches
delineated by the fuselage. Windows are vacuum formed to the moulded ge shape to match
the moulded window rebates. Windows are installed using a vulcanizing process. Cabin, Cargo and
hatch seals are a press fit to the preformed rebates created during the moulding process. Cabin,
Cargo and hatch hinges and latches are match drilled to mould stations and identifiers preformed
during the moulding process.
One disadvantage of the prior art is that recesses for windows and doors are typically finished by
hand. Hand finishing results in a structure that has a unique profile thereby requiring windows,
doors and the like that are to be fitted into the recess are to also be hand finished to ensure they
will fit, and fit without excessive gaps.
In this preferred embodiment each of the components and appendages internal to the fuselage are
formed by a composite moulding process. The process provides advantages such as accurate shape
reproduction when many such items are to be produced. The accurate reproduction of item shapes
provides the advantage of being able to e items on the ge without requiring any specific
attention to the fitment of that item. Accurate fitment allows manufacturing cost and time to be
y reduced.
One preferred embodiment of the helicopter fuselage is shown in Figure 29 as a side view and in
Figure 30 as a bottom view. Preferably the fuselage is a continuous laminate structure that includes
at least the l fuselage section 303 and the tail boom 304. An optional cabin region 305 is
included in the continuous fuselage section as shown in the gs. Preferably the continuous
central fuselage section 303 and the tail boom 304 include rcing members that span the
region of continuity. Preferably a first member 330 is located on the surface of an upper region of
the fuselage and spans the e of the upper region of the central ge section to the surface
of the upper section of the tail boom. ably a second member 331 is d on the surface of
the fuselage below the first member and spans the upper region of the l fuselage surface
section to the middle and lower surface sections of the tail boom. Preferably a third member 332 is
located on the surface of the fuselage and spans between a lower region of the central fuselage
section to at least some way into the surface of the tail boom region.
Preferably each of the first second and third members are made of a fabric such as unidirectional
carbon fibre fabric of approximately 200 g/mm2. ably each of the first second and third
members are approximately 200 mm to 300 mm wide. Preferably these members are integrated
into the laminate structure of the fuselage as an ‘additional laminate’. Preferably the s are
centrally located between inner and outer e layers. ably each of the first second and
third members span a transition zone define as the region between the central fuselage section of
the fuselage and the tail boom.
The members e additional strength and stability to the tail boom and any age
structure located on the tail boom relative to the central fuselage section.
The method of ucting the helicopter includes the following steps in the following preferred
order. Those skilled in the art will recognise the particular ly process may be reordered.
1. Preparing a load bearing fuselage outer skin or shell, including preparing and joining two
opposing sections of fuselage.
The joined fuselage includes a centre n, tail boom and optionally a forward section.
The tail boom optionally including an empennage. The fuselage preferably delineating door
and window openings. The optional forward section including a cabin for housing a pilot,
gers and flight controls. The centre section adapted a support an engine and drive
train. The tail boom and empennage being substantially of a que structure. The
empennage adapted to support a ducted fan tail rotor assembly, horizontal and vertical
stabilisers. The load bearing fuselage further delineating an opening located on an upper
surface adapted to allow the engine-main rotor gearbox and drive train module to be
inserted and internally attached.
Preferably the load bearing fuselage is ed by layering several fabric layers and infusing
those layers with a polymer or resin to form a ite structure. Preferably the ng
fuselage sections are joined while each still in their respective moulds.
2. Preparing at least two keel beam components.
Preferably the keel beam components are prepared by layering several fabric layers and
infusing those layers with a polymer or resin to form a composite structure.
3. Installing the at least two keel beam components by inserting them through a window or
door opening in the fuselage.
4. Connecting an upper region of the at least two keel beam components. The connection can
be made by applying a section of fabric to overlap regions of the proximate regions of the
fuselage and beams and infusing the fabric with a polymer or resin to form a ite
structure.
. Preparing at least two bulkhead components.
6. Installing the at least two bulkhead components by inserting them through a window or
door opening in the fuselage.
7. Connecting an upper region of the at least two bulkhead components. The connection can
be made by applying a section of fabric to overlap regions of the ate regions of the
bulkheads, fuselage and beams and infusing the fabric with a r or resin to form a
composite ure. Preferably the at least two bulkhead components include a rear
bulkhead and a forward bulkhead which define an enclosed space together with the at least
two keel beam components for housing fuel cell components
8. Preparing at least one floor panel
9. Installing the floor panel into the fuselage by inserting them through a window or door
g in the fuselage.
Preferably the fuselage or at least a section of the fuselage, including a centre section and tail boom
section, is constructed according to the following preferred order of steps. However, those d
in the art will recognise the assembly steps need not take place in strict order.
1. A mould is prepared and the inside e of that mould coated in a release agent such as
gel coat or liquid wax. Gel coat approximately 0.018” – 0.022” thick may optionally be
d.
2. A layer of CFM fabric is laid in the mould.
3. A layer of CKC fabric is laid in the mould.
4. An optional layer, or layers, of fabric is laid in the mould.
. A layer of core medium such as Soric is laid in the mould.
6. A layer of CKC fabric is laid in the mould.
7. A layer of CFM fabric is laid in the mould.
8. The layers of fabric are locked together, preferably by a resin infusion process.
Optional layers include additional structural stability enhancing fabrics embedded in the layers of
other fabrics. Optional layers may be d to then regions of the fuselage proximate
window and door openings, s and regions where other attachments or fastenings are d.
Optional layers may be a layer of CDB (carbon double bias), alternatively an optional layer may be
a layer of CU (carbon uni) to form a load bearing member.
Figure 31 shows a side view of the preferred helicopter fuselage 300 and empennage 336 g
an area where an additional layer of fabric 332 may be applied to the laminate structure. Preferably
the additional fabric 332 is located proximate the region where the central section of the fuselage
transitions to the tail boom. Preferably the additional layer 332 is CKC or CDB material.
Preferably the empennage is constructed using a laminate structure of CFM and CKC material. The
empennage 336 may be joined to the tail boom section of the fuselage 300 by at least one
overlapping layer of al 333. Preferably the region of overlap is at least 30 mm. Preferably the
overlapping layer of material is a tion structure of CKC and CKC material.
Figure 32 shows a bottom view of the red helicopter and in particular illustrates a plurality
of openings delineated by the fuselage. Preferably each of the openings has an additional layer of
fabric material 334, 335 applied to at least part of the edge of the opening. In such circumstances
where the opening has a lip or flange, the fabric layer preferably extends from the main surface of
the ge to the edge of the lip or flange. Preferably the fabric is a CFM material. Further, it may
be desirable for multiple layers of additional fabric material to provide additional strength to the
region of the fuselage proximate where the fabric is applied. For example, three layers may be
applied. Preferably the layers are at least 25 mm in width.
The preferred helicopter fuselage bears a load or force to which the fuselage is subjected to in
supporting the weight of the ge and associated components and in resisting externally applied
forces such as those created by the engine and drive train mechanisms.
Design engineers use the term load path to describe, in general terms, the way in which loads path
through a ure from the points of application to the points where they are reacted. In
contrast, stress trajectories are more clearly identified by the direction of the principal stress vectors
at a point.
Dynamic loads are imparted to the fuselage due to high forward speed. The red helicopter
fuselage withstands the sum of the loads imparted to the fuselage at 160 kts multiplied by a 3.5
times limit load plus a further 1.5 times ultimate load for 3 seconds. Load must be distributed into
the fuselage through the main rotor hub, mast, main rotor gear box, into the main rotor gear box
frame and into the keal beams 313 then into the fuselage via a multitude of load paths.
Principally primary structure of the aircraft butes the load. Secondary ure also carries
load but would not lead to a catastrophic failure of the aircraft if damaged.
The tail boom primarily supports the tail rotor. The purpose of the tail rotor is to oppose main
rotor torque in the hover and to provide directional control in the hover and at low speed. The tail
boom also has to support the c in-flight loads. Ie the upper vertical fin offsets the need for
tail rotor thrust in cruise. The lower vertical fin provides stability in an autorotation. The horizontal
stabilizers provide for longitudinal stability in forward flight.
The weight of the helicopter must be overcome by main rotor upwards thrust otherwise it will
remain grounded. This weight can be referred to as load. This load will vary dependant of the
vring of the helicopter. The main rotor loads are directed into a main rotor x frame.
The main rotor gear box frame is preferably attached at six locations to the left and right main keel
beams 313. All of the manoeuvring loads are transferred into the ge structure. The loads
developed on the tail boom are transmitted into the center section of the fuselage. The cabin
structure has to withstand the dynamic loads from d and sideways speed. The cabin must
also withstand the loads applied from the weight of the occupants on their seats or cargo loads on
the cabin floor. All the cabin loads are transferred into the keel beams 313, 314 and into the
fuselage.
Claims (32)
1. A primary flight structure comprising: a load bearing composite shell defining at least the exterior of a pter fuselage as a 5 continuous laminate structure, the fuselage defining at least a central fuselage section and tail boom, wherein the central fuselage section is adapted to enclose at least one of an engine or drive train. 10
2. A primary flight structure as claimed in claim 1, wherein the tail boom has attached or is adapted to have ed at least one of an empennage, fins, or tail rotor mechanism.
3. A primary flight ure as claimed in claim 2, wherein the empennage is adapted to support a ducted fan tail rotor assembly, horizontal and vertical stabilisers.
4. A primary flight structure as claimed in any one of claims 1 to 3, wherein the tail boom is adapted to have attached an empennage to thereby form a monocoque structure.
5. A primary flight structure as claimed in any one of claims 1 to 4, further comprising a 20 forward section of the fuselage adapted to house one or more occupants and flight controls.
6. A primary flight structure as claimed in any one of claims 1 to 5, wherein the central n and forward section of the fuselage ates at least one of a door, window or hatch opening. 25
7. A primary flight structure as claimed in any one of claims 1 to 6, wherein the structure delineates an opening d in an upper region of the fuselage surface, the g adapted to allow at least one of an engine or main rotor gearbox or drive train module to be at least partly inserted into the fuselage. 30
8. A primary flight ure as d in any one of claims 1 to 7, further comprising at least two beam components ing between at least an upper and lower internal surface of the
9. A primary flight structure as claimed in 8, wherein the beam components provide at least 35 one attachment point for at least one of an engine or main rotor gearbox or drive train module, or at least one attachment point for a frame assembly to which at least one of an engine or main rotor gearbox or drive train module are adapted to be attached.
10. A primary flight structure as claimed in any one of claims 8 or 9, further comprising at least 5 two members extending at least between side internal surfaces of the fuselage and transversely to the at least two beam components.
11. A primary flight structure as claimed in any one of claims 8 to 10, wherein at least one of the s is adapted to transfer load created by at least one of an engine or main rotor gearbox 10 or drive train module to the composite shell.
12. A primary flight structure as claimed in any one of claims 8 to 11, wherein at least one of the members is arranged to at least partially create an enclosed space in which one or more fuel cells can be located.
13. A primary flight structure as claimed in any one of claims 8 to 12, wherein the members are d to create a structure attachable to the internal surface of the composite shell such that the structure least partially absorbs aircraft impact energy and diverts energy away from the enclosed space.
14. A primary flight structure as claimed in any one of claims 5 to 13, wherein at least two members further comprise an integrally formed and d protruding member adapted to extend from the central fuselage n to the forward section, the protruding member having a lower region adapted to attach to a lower internal surface of the composite shell and an upper region 25 adapted to support a cabin floor panel.
15. A y flight structure as claimed in claim 14, wherein the forward section of the ge and the cabin floor panel at least partly define a cabin space such that seating for occupants and flight controls can be located or referenced.
16. A y flight structure as claimed in claim 14 or claim 15, wherein the cabin floor is supported by the d ding member and a plurality of structural s, the structural members adapted to extend between the lower side of the cabin floor panel and the internal surface of the lower region of the composite shell.
17. A primary flight structure as claimed in claim 16, wherein the structural members are adapted to attach to the cabin floor panel and the internal surface of the composite shell.
18. A primary flight structure as claimed in claim 16 or claim 17, wherein the ural 5 members comprise a first set of members and second set or members and the first set of members are adapted to extend substantially perpendicular to the second set of members such that the structural members together, when combined, are adapted to form a lattice structure extending between the cabin floor panel and the internal surface of the composite shell. 10
19. A primary flight structure as claimed in claim 5, wherein the cabin space is further adapted to allow a plurality of seats, flight controls and occupant restraints to be located.
20. A primary flight structure as claimed in any one of claims 1 to 19, wherein the composite shell ses a laminate, the laminate comprising a plurality of fabric layers.
21. A primary flight ure as claimed in claim 20, wherein the plurality of fabric layers comprise a first layer of continuous filament mat, a first layer of carbon and Kevlar ite, a layer of unidirectional carbon fibre, a second layer of carbon and Kevlar ite, a second layer of continuous filament mat and wherein the laminate is, or is adapted to be infused with a cured or 20 curable resin.
22. A primary flight structure as d in claim 21, wherein the laminate further comprises additional laminates and/or core filler material including Soric. 25
23. A primary flight structure as claimed in claim 21 or claim 22, wherein the laminate further comprises a layer of carbon unidirectional fabric between the first and second .
24. A primary flight structure as claimed in any one of claims 21 to 23, n the laminate further comprises a layer of carbon double bias fabric between the first and second layers.
25. A primary flight structure as claimed in any one of claims 21 to 24, wherein the laminate further comprises a filler layer n the first and second layers.
26. A primary flight structure as d in claim 24, wherein the layer of carbon double bias 35 fabric comprises a ity of layers of carbon double bias , including: a plurality of layers that span a surface of an upper region of the central fuselage section to a e of an upper section of the tail boom, a plurality of layers that span the upper region of the l fuselage surface n to a middle and a lower surface section of the tail boom, and 5 a plurality of layers that span between a lower region of the central fuselage section to at least some way into the lower surface of the tail boom region.
27. A primary flight structure as claimed in any one of claims 21 to 26, wherein the unidirectional carbon fibre fabric is approximately 200 g/mm2.
28. A primary flight structure as claimed in any one of claims 21 to 27, wherein at least some of the support s are approximately 200 mm to 300 mm wide. 29. A primary flight structure as claimed in any one of claims 21 to 28, wherein the support 15 fabrics are integrated within the laminate layers of the fuselage.
29. A primary flight structure as claimed in any one of claims 21 to 28, wherein the continuous filament mat layers are approximately 300 g/mm2. 20
30. A primary flight structure as claimed in any one of claims 21 to 29, wherein the carbon and Kevlar ite layers are approximately 190 g/mm2.
31. A primary flight structure as claimed in any one of claims 21 to 30, wherein the filler layer is approximately 2 mm thick.
32. A primary flight structure as claimed in any one of claims 21 to 31, wherein the outer e layer is formed as a substantially smooth and substantially continuous surface that extends through at least the central fuselage section to the tail boom.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
NZ618737A NZ618737B2 (en) | 2011-06-07 | 2012-06-07 | A helicopter |
Applications Claiming Priority (8)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
NZ59330011 | 2011-06-07 | ||
NZ593300 | 2011-06-07 | ||
NZ59330411 | 2011-06-07 | ||
NZ593304 | 2011-06-07 | ||
NZ59368311 | 2011-06-22 | ||
NZ593683 | 2011-06-22 | ||
NZ618737A NZ618737B2 (en) | 2011-06-07 | 2012-06-07 | A helicopter |
PCT/NZ2012/000087 WO2012169906A1 (en) | 2011-06-07 | 2012-06-07 | A helicopter |
Publications (2)
Publication Number | Publication Date |
---|---|
NZ618737A NZ618737A (en) | 2016-03-31 |
NZ618737B2 true NZ618737B2 (en) | 2016-07-01 |
Family
ID=
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