U.S. patent application number 11/799324 was filed with the patent office on 2008-08-28 for jig and method of manufacturing aircraft frames in a composite material.
This patent application is currently assigned to AIRBUS ESPANA, S.L.. Invention is credited to Ignacio Jose Marquez Lopez, Jesus Manuel Martin Martin.
Application Number | 20080203601 11/799324 |
Document ID | / |
Family ID | 39714966 |
Filed Date | 2008-08-28 |
United States Patent
Application |
20080203601 |
Kind Code |
A1 |
Martin Martin; Jesus Manuel ;
et al. |
August 28, 2008 |
Jig and method of manufacturing aircraft frames in a composite
material
Abstract
The present invention relates to a jig for the manufacture, by
means of injection and curing processes, of preforms of composite
material frames for aircraft fuselages by using the RTM (resin
transfer molding) technology. Two preforms are thus manufactured,
one with a C shaped section and another with a L shaped section,
together with the preforms of the stabilization ribs for
stabilizing the web of the frames and the preform of the roving or
staple fiber to cover the gap between the C shaped preform and the
L shaped preform. Theses preforms are previously manufactured by
any known process for manufacturing preforms. According to a second
aspect, the present invention relates to a method of manufacturing
composite material load frames for aircraft.
Inventors: |
Martin Martin; Jesus Manuel;
(Madrid, ES) ; Marquez Lopez; Ignacio Jose;
(Madrid, ES) |
Correspondence
Address: |
LADAS & PARRY LLP
26 WEST 61ST STREET
NEW YORK
NY
10023
US
|
Assignee: |
AIRBUS ESPANA, S.L.
|
Family ID: |
39714966 |
Appl. No.: |
11/799324 |
Filed: |
May 1, 2007 |
Current U.S.
Class: |
264/101 ;
425/543 |
Current CPC
Class: |
B29D 99/0003 20130101;
B29L 2031/3082 20130101; Y02T 50/40 20130101; B29C 70/48 20130101;
B29L 2031/008 20130101; B29C 70/085 20130101; Y02T 50/43
20130101 |
Class at
Publication: |
264/101 ;
425/543 |
International
Class: |
B29C 45/17 20060101
B29C045/17; B29C 45/03 20060101 B29C045/03 |
Foreign Application Data
Date |
Code |
Application Number |
Feb 28, 2007 |
ES |
PCT/ES07/70042 |
Claims
1. A jig for injecting and curing preforms of an aircraft fuselage
composite material frame (2), characterized in that the mentioned
jig shapes preforms comprising at least one C shaped preform (3),
at least one L shaped angular preform (4), at least one
stabilization rib (5) for stabilizing the web of the frame (2) and
at least one preform (6) of the roving or staple fiber in the joint
between the C shaped preform (3) and the L shaped angular preform
(4), and in that the mentioned jig comprises an injection and
curing jig (7) injecting and curing resin for the manufacture of
the preforms, a vacuum system (8) allowing to shape the previous
preforms by means of applying vacuum and a heating and closing
system such that the section of the structure of the frames (2) by
means of the previous preforms is integreated in one part.
2. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to claim 1, characterized in that the preforms
(3, 4, 5 and 6) are manufactured by means of RTM (resin transfer
molding) technology.
3. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to a claim 1, characterized in that the
preforms (3, 4, 5 and 6) comprise fabric and reinforcements with
unidirectional tape in the inner flange to increase their moment of
ineitia and, consequently, their rigidity.
4. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to claim 1, characterized in that the injection
and curing jig (7) comprises a tub (10) in which there are placed
the remaining elements shaping the jig (7) and the preforms (3, 4,
5 and 6) are placed, one male parts (11) assembly which is placed
under the C shaped preform (3), an assembly of upper male parts
(12) shaping the frame (2) on the side of the outer flange, an
assembly of upper male parts (13) shaping the frame (2) on the side
of the inner flange, an assembly of male parts (14) which are
arranged on both sides of the stabilization ribs (5) for
stabilizing the web of the frame (2) and a cover (15) sealing the
jig (7) against the tub (10).
5. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to claim 1, characterized in that the vacuum
system (8) comprises an assembly of sealing rubbers (16) arranged
at the upper part of the tub (10), one vacuum pump (17), a system
of tubes (20,22) joining the jig (7) with a resin injection machine
(19) and with the vacuum pump (17), and a vacuum circuit (21)
closing the jig (7) and from which the tubes (20, 22) joining said
jig (7) with the vacuum pump (17) and with the resin injection
machine (19) come out.
6. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to claim 1, characterized in that the heating
and closing system comprises a hot plate press (18).
7. A jig for injecting and curing preforms of an aircraft fuselage
frame (2) according to claim 1, characterized in that the heating
and closing system comprises an autoclave.
8. A method of manufacturing composite material aircraft fuselage
frames (2) comprising the following steps: a) Placing and closing
the injection and curing jig (7). b) Placing the jig (7) on the
heating and closing system. c) Connecting the vacuum system (8). d)
Applying pressure on the closing and heating system to close the
jig (7) and ensure tightness. e) Heating the jig (7) to the
injection temperature. f) Applying vacuum to the jig (7), through
the vacuum system (8). g) Injecting the resin. h) Constricting the
tubes (20) once the resin has overflowed through the injection
points. i) Applying compacting pressure to the injection machine
(19). j) Heating gradient up to the curing temperature. k)
Maintaining the curing temperature. l) Cooling. m) Demolding.
Description
FIELD OF THE INVENTION
[0001] The present invention relates to a jig for the manufacture
of composite material frames for aircraft, as well as to a method
of manufacturing said frames in a composite material.
BACKGROUND OF THE INVENTION
[0002] In addition to conferring shape and rigidity to the aircraft
fuselage, load frames are structural elements in charge of
withstanding and transferring the loads from other structural
elements of the aircraft, such as the wings or stabilizers.
[0003] In the state of the art, the manufacture of the frames is
carried out by using machined metal structures or shaped sheet
metal structures having in which the part that withstands most of
the load is reinforced with machined parts. In the case of shaped
sheet metal, the section that is normally used is obtained in two
pieces: on one hand the Z is manufactured, and on the other hand,
the brackets which are riveted to the former section are
manufactured.
[0004] This process has the drawback that long assembly times are
necessary and that the final weight is much greater than what would
be desired.
SUMMARY OF THE INVENTION
[0005] In a first aspect, the present invention relates to a jig
for the manufacture of aircraft load frames such that the section
of the structure of the obtained frames is done in an integrated
manner such that the resulting weight is optimized. At the same
time, the manufacturing method proposed by the present invention is
carried out by means of a repetitive process with a short times in
curing cycles, such that the necessary assembly times are
decreased.
[0006] Thus, the present invention develops a jig for the
manufacture, by means of injection and curing processes, of
composite material frame preforms for aircraft fuselages by using
RTM (resin transfer molding) technology. Therefore, two preforms
are manufactured, one with a C-shaped section and another one with
an L-shaped section, together with the preforms of the
stabilization ribs of the web of the frames and the preform of the
roving or staple fiber to cover the gap between the C shaped
preform and the L shaped preform. These preforms are previously
manufactured by any known preform manufacturing process.
[0007] According to a second aspect, the present invention develops
a method of manufacturing an aircraft load frame in a composite
material.
[0008] Thus, aircraft load frames made of a composite material are
obtained by means of the jig and the method of manufacture of the
present invention with the following advantages: [0009] Frames with
complex and integrated geometries are manufactured, meeting the
objective sought in any structure. [0010] The problem of the
surface finish on only one face is solved by adding a high
dimensional precision without the necking of radii, since it is a
closed framework. [0011] Control of thicknesses is improved
(reaching tolerances .ltoreq.0.2 mm), whatever these thicknesses
may be, therefore achieving a good fit between the outer flange of
the frames with the skin and the stringers making it rigid. [0012]
The process is repetitive with short curing cycle times, which
reduces the duration of the manufacturing process.
[0013] Other features and advantages of the present invention will
be understood from the following detailed description of the
illustrative embodiments of its object, together with the attached
drawings.
DESCRIPTION OF THE DRAWINGS
[0014] FIG. 1 shows the section of a known metallic aircraft load
frame.
[0015] FIG. 2 shows the section of an aircraft load frame made of a
composite material according to the present invention.
[0016] FIGS. 3a and 3b show perspective views of the aircraft load
frame made of a composite material that is finished, impregnated
and cured with the jig and the method of the present invention.
[0017] FIG. 4 shows a C-shaped preform of the aircraft load frame
made of a composite material according to the invention.
[0018] FIG. 5 shows an L-shaped preform of the aircraft load frame
made of a composite material according to the invention.
[0019] FIG. 6 shows one preform of the stabilization rib of the web
of the aircraft load frame made of a composite material according
to the invention.
[0020] FIG. 7 shows a preform of the roving or staple fiber of the
aircraft load frame made of a composite material according to the
invention.
[0021] FIG. 8 shows the integration of the preforms shaping the
aircraft load frame made of a composite material according to the
invention.
[0022] FIG. 9 shows a sectional view of the jig for the manufacture
of composite material frames for aircraft according to the
invention.
[0023] FIG. 10 shows a perspective view of the jig for the
manufacture of composite material frames for aircraft according to
the invention.
[0024] FIG. 11 shows a detail of the stabilization ribs of the web
of the composite material frames for aircraft according to the
invention.
[0025] FIG. 12 shows a general view of the vacuum system of the jig
for the manufacture of composite material frames for aircraft
according to the invention.
DETAILED DESCRIPTION OF THE INVENTION
[0026] In a first aspect, the present invention relates to a jig
for injecting and curing the preforms of an aeronautic fuselage
frame 2 made of a composite material.
[0027] The section to be manufactured is formed by a C-shaped
preform (FIG. 4) and by an L-shaped angular preform 4 (FIG. 5), in
addition to different stabilization ribs 5 of the web (FIG. 6) and
a preform 6 of the roving or staple fiber (FIG. 7). This
arrangement of preforms allows the manufacture of aircraft load
frames 2 with the parts for joining them to the following frame
integrated therewith, as can be seen in FIG. 8. The preforms are
made of fabric and reinforcements with unidirectional tape in the
inner flanges to increase their moment of inertia and,
consequently, their rigidity.
[0028] Likewise, the jig object of the invention comprises the
following elements: an injection and curing jig 7, a vacuum system
8 and a closing and heating system.
[0029] Injection and Curing Jig 7
[0030] The injection and curing jig comprises different members:
[0031] A tub 10. This is the base element, inside of which the
remaining elements shaping the jig 7 as well as the preforms 3, 4,
5 and 6 to be simultaneously injected and cured are placed. This
element integrates the resin injection and extraction boreholes,
the resin distribution channel, the sealing system for the
subsequent application of a vacuum and the thermocouples for
thermal control of the jig 7 during the heating cycle. [0032] An
assembly of male parts 11. This is the assembly of machined male
parts that goes under the C-shaped preform 3. They allow demolding
the frame 2 without difficulties. These male parts are
longitudinally cut to make demolding and handling easier. [0033] An
assembly of male parts 12. This is the assembly of upper male parts
shaping the frame 2 on the side of the outer flange, i.e. the one
in contact with the fuselage skin and the legs of the stringers. It
copies the entire geometry of the skin on which it rests, as well
as the shape of the legs of the stringers, while at the same time
it incorporates grooves on its upper surface for the injection of
resin. These grooves open into the inlet and outlet boreholes of
tub 10. The male parts 12 are longitudinally cut to make demolding
and handling easier. [0034] An assembly of male parts 13. This is
the assembly of upper male parts shaping the frame 2 on the side of
the inner flange, i.e., the one in the innermost part of the
fuselage and serving to stiffen the section of the mentioned frame
2. The reinforcements with unidirectional tape giving rigidity to
the frame 2 are placed on this flange. It incorporates grooves at
its inner surface allowing the extraction of the resin. These
grooves open into the resin outlet boreholes of the tub 10. [0035]
An assembly of male parts 14. They are the male parts in both sides
of the stabilization ribs 5 for stabilizing the web of the frame 2,
and therefore, such male parts are placed between the male parts
11. They are provided with a resin outlet channel to allow the
correct impregnation of the ribs 5. Their design must be such that
it allows demolding. [0036] Cover 15. It is the upper part of the
jig 7, sealing said jig 7 against the tub 10, to the sealing system
incorporating said tub 10. The cover 15 is planar to ensure in a
simple and efficient manner the vacuum level required inside the
jig 7. It incorporates thermocouples for the thermal control of the
jig 7 during the heating cycle thereof.
[0037] Vacuum System 8
[0038] The vacuum system 8 comprises the following elements: [0039]
An assembly of sealing rubbers 16, arranged in several grooves at
the upper part of the tub 10. [0040] A system of hollow silicone
tubes 20, 22 joining the jig 7 with the vacuum pump 17 and the
resin injection machine 19. [0041] A vacuum circuit 21. A
leak-tight (metallic or non-metallic) tube circuit 23 to which the
silicone tubes coming from the injection and curing jig 7 are led,
and from which comes another silicone 22 tube leading to the vacuum
pump 17. It is therefore a circuit arranged over the press 18 and
joining the different resin extraction points of the jig 7 to one
other. The connection between the vacuum circuit 21 and the
silicone tubes 20 leading to the jig 7 is made through leak-tight
connectors. To prevent the resin inform entering the vacuum system
21, the joining is done through expansion or draining vessels, and
the resin would fall on such vessels if it accidentally reached
this position. [0042] Vacuum pump 17. It is able to reach a vacuum
level of 0.5 mbar.
[0043] Closing and Heating System
[0044] According to the concept of the injection and curing jig 7
detailed in this invention, two processes for closing and heating
the jig 7 can be used for the resin injection and curing process:
[0045] A hot plate press 18. It consists of hydraulic or pneumatic
presses, with the geometry enveloping all of those frames 2 which
are to be manufactured, with the following basic operating concept.
[0046] i. Pushing cylinders at the upper part of the press,
reacting against columns connected to the floor. [0047] ii. A lower
carriage with horizontal movement, for inserting and extracting
tools in the press. [0048] iii. Upper heated plate. [0049] iv.
Lower heated plate. [0050] v. Pressure and temperature control
system with a programmable automaton. [0051] vi. Insulating hood to
prevent heat escape during the heating cycle. [0052] vii. A bushing
system to pass the resin injection and extraction tubes from the
injection system to the jig 7. [0053] viii. A connection system for
the thermocouples housed in the jig 7, such that the programmable
automaton controls the different heating areas of the plates of the
press, according to the local temperature of the jig 7. [0054]
Injection and curing autoclave. In this case, the autoclave exerts
the closing pressure, for which it is necessary to close the jig 7
with a vacuum lock. The resin injection and extraction tubes must
be able to withstand the pressure of the autoclave without
collapsing, for which the they will be connected to the bushing of
the autoclave to connect said autoclave with the resin injection
system.
[0055] According to a second aspect, the present invention develops
a method of manufacturing composite material load frames for
aircraft comprising the following steps: [0056] 1. Placing and
closing the injection and curing jig 7. [0057] 2. Placing the jig 7
on the injection and curing press 18. [0058] 3. Connecting the
vacuum system 8. [0059] 4. Applying pressure to the press 18, to
close the jig 7 and to ensure tightness. [0060] 5. Heating the jig
7 up to the injection temperature. [0061] 6. Applying vacuum to the
jig 7 though the vacuum system 8. [0062] 7. Injecting the resin.
[0063] 8. Restricting the silicone tubes 20 once the resin has
overflowed through the injection hoses. [0064] 9. Applying
compacting pressure to the injection machine 19, up to 3 bar, i.e.
the resin passing through the inlet tube of the jig enters with a
pressure of 3 bar. [0065] 10. A heating gradient up to the curing
temperature. [0066] 11. Maintaining the curing temperature. [0067]
12. Cooling. [0068] 13. Demolding.
[0069] Those modifications which are comprised in the scope of the
following claims can be introduced in the described preferred
embodiment.
* * * * *