Aircraft pressure containment assembly module

An aircraft pressure containment module is adapted to be mounted generally transversely within an aircraft fuselage for providing pressure containment to separate pressurized and non-pressurized sections of an aircraft. A plurality of web segments are joined together to form a web member. A ring member is arranged on a first pressurized side of the web member. A plurality of truss assemblies are also arranged on the first side of the web member. Each truss assembly is comprised of a first truss member adapted to extend outwardly from the ring member to a peripheral edge of the web member. The first truss member is joined to the web member and joins together adjacent web segments. A second truss member extends from the first truss member in a direction generally normal to the first side of the web member. The second truss member is adapted to be joined to an aircraft fuselage in a longitudinal sense or it may comprise a portion of longitudinally extending member of the fuselage such as a stringer. A third truss member is joined to the second truss member and the first truss member to form a truss assembly having at least in part a generally triangular shape. The truss assemblies are adapted to react pressure loading from the web member to the fuselage of an aircraft in a manner similar to a conventional dome type pressure containment module. An aircraft using the module has an elongated fuselage extending in a longitudinal direction and the pressure containment module mounted generally transversely within the fuselage.

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
The present invention relates to a novel structure for pressure containment 
to separate pressurized and non- pressurized sections of an aircraft. 
2. Prior Art 
U.S. Pat. No. 4,869,443 discloses fuselage sections. U.S. Pat. No. 
3,473,761 discloses a pneumatic tubular construction. U.S. Pat. No. 
2,090,038 discloses an aircraft structure. U.S. Pat. No. 2,121,670 
discloses a metal frame assemblage and a method of construction including 
convex metal plates. U.S. Pat. No. 2,006,468 discloses an airplane 
fuselage. U.S. Pat. No. 2,723,092 discloses a radome panel having 
structurally reinforcing elements. U.S. Pat. No. 4,531,695 discloses a 
helicopter fuselage with at least one mainframe member having two 
generally vertical side beams joined by generally horizontal top and 
bottom beams. U.S. Pat. No. 4,648,570 discloses a system for supporting 
interior passenger elements within an aircraft. Other U.S. Patents showing 
airplane fuselage section structures include U.S. Pat. Nos. 2,387,219, 
1,885,406, 1,866,534 and 1,622,242. The aforenoted prior art does not 
disclose a pressure containment structure corresponding to the present 
invention. 
Present day commercial and military transport category aircraft having 
pressurized cabins/cargo compartments typically utilize metallic concave 
dome type structures for pressure containment to separate pressurized and 
non-pressurized sections of the aircraft. These structures are typically 
located in the aft section of the fuselage just forward of the aircraft's 
empennage and are attached peripherally to the inside fuselage surface to 
form a structural boundary. The continuous reinforced shell like 
construction of the dome uniformly distributes and reacts pressure loads 
into the fuselage structure itself. This type of construction requires 
forming of double curvature segments which are assembled to form a concave 
reinforced shell. Added radial and circumferential straps are incorporated 
to enhance damage tolerance. Construction and assembly methods are 
complex. Design of efficient methods attaching the dome to the fuselage 
while minimizing radial offsets and longitudinal bending stiffness 
reduction effects add further complexity. In addition, this type of design 
does not lend itself to the reaction of planar loading such as that which 
would occur from a rear fuselage mounted engine attachment. 
SUMMARY OF THE INVENTION 
In accordance with this invention an innovative pressure containment module 
can provide the structural efficiency of the prior art dome approach for 
reacting pressure loading while at the same time simplifying manufacture 
and assembly. The module of this invention also offers the added benefit 
of providing a structural foundation for the reaction of planar loadings 
derived from external or internal attachment of other components including 
engines and support and/or mission equipment. The design of the module 
incorporates excellent durability and damage tolerance features including 
ease of inspection while minimizing secondary structural effects 
associated with the connection to the fuselage. It is readily repairable 
and by virtue of its modular construction can be easily retrofitted into 
aircraft being modified to meet new mission requirements. The module's 
structural concept is intended to comply with applicable Federal Aviation 
Administration (FAA) regulations for aircraft certification. 
In accordance with one embodiment of the invention the module comprises an 
aircraft pressure containment module which is adapted to be mounted 
generally transversely within an aircraft fuselage for providing pressure 
containment to separate pressurized and non-pressurized sections of the 
aircraft. The module comprises a plurality of web segments forming a web 
member having a first side adapted to be exposed to a first pressure and a 
second side adapted to be exposed to a second pressure less than the first 
pressure. The peripheral edge of the web member has a desired shape 
adapted to conform to the inside surface of an aircraft fuselage. A ring 
member is arranged on the first side of the web member. A plurality of 
truss assemblies are attached to the web member. 
Each truss assembly is comprised of first, second and third truss members. 
The first truss member is adapted to extend outwardly from the ring member 
to the peripheral edge of the web member. The first truss member is joined 
to the web member and joins together adjacent web segments. A first end 
portion of the first truss member is joined to the ring member. The second 
truss member is adapted to extend from a second end portion of the first 
truss member in a direction generally normal to the first side of the web 
member. The second truss member is joined to the second end portion of the 
first truss member and to a circumferential frame in the fuselage. The 
second truss member is adapted to be joined to an aircraft fuselage in a 
longitudinal sense or it may comprise a portion of a longitudinally 
extending member of the fuselage such as a stringer. The third truss 
member is adapted to extend from the second truss member to the first 
truss member to form a truss assembly having at least in part a generally 
triangular shape. The third truss member is joined to the first and second 
truss members. 
The greater the number of truss assemblies, the more the pressure 
differential can be uniformly reacted and the lower the reacting load in 
each truss assembly. In addition, the strength requirements of the web 
segments and first radial truss members are reduced. The truss assembly 
subtended angle with respect to the planar web member is also a 
consideration. The number of truss assemblies is established in 
conjunction with the center ring diameter through a structural sizing and 
tradeoff process involving these factors. 
Preferred embodiments of the invention include one or more of the following 
structural features: (a) the web member is generally planar to provide an 
improved structural foundation for attachment of equipment or engines; (b) 
the truss assemblies are arranged to extend radially outwardly from the 
ring member and are generally equally spaced from one another; (c) the 
third truss member is adjustable to compensate for fuselage tolerances and 
is adapted to be under tension whenever a fuselage in which it is 
installed is pressurized; (d) the diameter of the ring (preferably made up 
of ring segments 33) is selected to be proportional to the effective 
diameter of the fuselage in which it is installed and the magnitude of the 
pressure loading as well as to accommodate the structural loading of other 
components; (e) the number of truss assemblies is selected to provide 
substantial uniformity of load reaction; (f) the third truss member has a 
generally tubular shape with end fittings fastened to each end preferably 
adapted to be bolted to the respective first and second truss members; and 
(g) the web segments are generally sector shaped and each web segment has 
a truss assembly joined to it. The web segments provide for modular 
construction either inside or outside of an aircraft and provide the 
module with excellent durability and damage tolerance characteristics as 
well as crack arrestment capability. 
In accordance with another embodiment of the invention an aircraft is 
provided having an elongated fuselage extending in a longitudinal 
direction and a pressure containment module as described above mounted 
generally transversely within the fuselage for providing pressure 
containment to separate pressurized and non-pressurized sections of the 
aircraft.

DETAILED DESCRIPTION OF THE INVENTION 
Referring to FIG. 1 a typical present day commercial or military transport 
aircraft 10 is shown having a pressurized cabin or cargo compartment 11. A 
prior art metallic, concave dome type structure 12 is joined to the 
fuselage 13 for pressure containment. The dome structure 12 separates the 
pressurized cabin or cargo compartment 11 from a non-pressurized section 
14 of the aircraft 10. These structures 12 are typically located in the 
aft section of the fuselage 13 just forward of the aircraft's empennage 15 
and are attached peripherally to the inside fuselage surface 16 to form a 
structural boundary. The continuous reinforced shell like construction of 
the dome structure 12 uniformly distributes and reacts pressure loads into 
the fuselage 13 structure itself. This type of construction requires 
forming double curvature segments which are assembled to form a concave 
reinforced shell type dome structure. Added radial and circumferential 
straps (not shown) are incorporated to enhance damage tolerance. 
Construction and assembly methods for this type of dome structure 12 are 
complex. The methods for attaching the dome structure 12 to the fuselage 
13 while minimizing radial offsets and reduction in longitudinal bending 
stiffness add further complexity. In addition, this type of design does 
not lend itself to the reaction of planar loading such as would occur from 
a rear fuselage 13 mounted engine attachment (not shown). 
Referring now to FIGS. 1 and 2 in accordance with this invention an 
innovative pressure containment module 20 is provided which can provide 
the structural efficiency of the prior art dome structure 12 for reacting 
pressure loading while at the same time being simple to manufacture and 
assemble. The module 20 of this invention also offers the added benefit of 
providing a structural foundation for the reaction of planar loadings 
derived from external or internal attachment of other components including 
engines and support and/or mission equipment (not shown). The design of 
the module 20 incorporates excellent durability and damage tolerance 
features including ease of inspection while minimizing secondary 
structural effects associated with the connection of the module 20 to the 
fuselage 13. It is readily repairable and by virtue of its modular 
construction can be easily retrofitted into an aircraft 10 being modified 
to meet new mission requirements. The module 20 structural concept is 
intended to comply with applicable Federal Aviation Administration (FAA) 
regulations for aircraft certification. 
The innovative aircraft pressure containment module 20 of this invention 
generally consists of standard aluminum or other alloy components 
configured as a preferably flat web member 21, continuously supported by a 
series of three member radial truss assemblies 22 as shown in FIG. 2. The 
web member 21 comprises a number of segments 23, preferably in the shape 
of sectors (e.g. pie shaped), which are joined together by the first 
radial truss members 24. The web member 21 has a first side 25 adapted to 
be exposed to a first pressure (e.g. pressurized cabin), a second side 26 
adapted to be exposed to a second pressure less than the first pressure 
(e.g. unpressurized aircraft section). A peripheral edge 27 of the web 
member 21 has a desired shape adapted to conform to the fuselage inside 
surface 16. 
A ring member 28 is arranged on the first side 25 of the web member 21 at a 
central location and is joined thereto. In the case of a web member 21 
with a circular peripheral edge 27 the ring member 28 preferably is 
arranged substantially coaxially about the center of the web member 21. 
The ring member 28 can comprise a unitary member or most preferably it is 
made up of ring segments 33 joined to the web member 21 which when fully 
assembled provide a complete ring member 28 or it could comprise a 
combination of such approaches. The ring segments 33 are separate 
circumferential assemblies consisting of web and flange elements which 
when joined together constitute the 360-degree ring member 28. 
Referring now to FIGS. 2 and 3 the truss assemblies 22 will be described in 
greater detail. 
The truss assemblies are preferably comprised of: 
(a) a first truss member 24 which is adapted to extend outwardly 
(preferably radially) from the ring member 28 to the peripheral edge 27 of 
the web member 21. The first truss member 24 is joined to the web member 
21 and also joins together adjacent web segments 23. Joining in accordance 
with this invention can comprise any conventional means of joining, 
including without limitation, riveting, bolting, welding, brazing, or the 
use of structural adhesives. A first end portion 29 of the first truss 
member 24 is also joined to the ring member 28; 
(b) a second truss member 30 is adapted to extend from a second end portion 
31 of the first truss member 24 in a direction generally normal to the 
first side 25 of the web member 21. The second truss member 30 is joined 
to the second end portion 31 of the first truss member 24. The second 
truss member 30 is adapted to be joined to the aircraft 10 fuselage 13 in 
a longitudinal sense or it may comprise a portion of longitudinally 
extending member of the fuselage 13 such as a stringer 32; and 
(c) a third truss member 40 adapted to extend from the second truss member 
30 to the first truss member 24 to form a truss assembly 22 having at 
least in part a generally triangular shape, the third truss member 40 is 
joined to the first and second truss members 24 and 30. 
The preferably substantially flat or planar web member 21 of the module 20 
is stiffened by the first radial truss members 24 of the truss assemblies 
22. The truss assemblies 22 are adapted to react pressure loading from the 
web member 21 to the fuselage 13 of an aircraft 10 in a manner similar to 
a conventional dome type pressure containment module 12. The first truss 
members 24 react pressure loading from the plane of adjacent web segments 
23 to the fuselage 13 and third truss member 40 connections, respectively, 
thereby simulating the membrane effect that the traditional dome type 
structure 12 provides. The truss assemblies 22 are installed on the 
pressurized side 25 of web member 21 within the fuselage 13. Second truss 
members 30 may also act as stringers 32 depending on their number and 
location. The third truss members 40 are always under tension when the 
fuselage 13 is pressurized. The truss assemblies 22 are preferably 
arranged to extend radially outwardly from the ring member 28 and 
preferably are substantially equally spaced from one another. 
The center ring 28 provides the structural foundation for the truss 
assemblies 22. Its diameter is determined by the outer diameter of the 
fuselage 13; the magnitude of the pressure loading caused by pressurizing 
the cabin 11; and the need to accommodate structural attachment of other 
components (not shown). 
A larger center ring member 28 diameter allows for greater flexibility in 
its construction but results in less efficiency of the truss assemblies 
22. The ideal diameter is approached by consideration of a practical 
minimum size ultimately derived by analyzing structural requirements of 
the planar web member 21 and space limitations within the fuselage 13 
constraining the angle that the third truss member 40 subtends with 
respect to the web member 21. Since the preferred subtended angle is 45 
degrees, ring diameter and truss geometry considerations are tailored for 
each specific application. 
The number of truss assemblies 22 is also determined by these factors, 
wherein a practical minimum dictates the efficiency in achieving 
uniformity of load reaction. The greater the number of truss assemblies 
22, the more the pressure differential can be uniformly reacted and the 
lower the reacting load in each truss assembly 22. In addition, the 
strength requirements of the web segments 23 and first radial truss 
members 24 are reduced. The truss assembly 22 subtended angle with respect 
to the web member 21 is also a consideration as previously indicated. The 
number of truss assemblies 22 is established in conjunction with the 
center ring member 28 diameter through a structural sizing and tradeoff 
process involving the factors cited. 
One can perform in a conventional manner a structural sizing trade off 
analysis using the effective diameter of the fuselage 13 and the magnitude 
of the pressure loading to quantify the structural efficiency of various 
ring 28 diameters with respect to the enclosed web 21 portion formed by 
the ring 28 periphery and the number of truss assemblies 22 that can be 
practically installed to distribute pressure loading sufficiently so as to 
achieve a uniform reaction into the fuselage 13. A relationship based on 
these parameters may be determined by fixing one or more variables such 
as, for example: (a) the subtended angle of the third truss member 40, (b) 
the intersection of the third truss member 40 with the first truss member 
24, or (c) the desired out of plane stiffness for providing support for 
loading other than pressure loading. 
The diagonal or third truss members 40 preferably are of a simple tubular 
construction as shown in FIG. 4 with mechanically fastened conventional 
end fittings 41 adapted to be bolted to the first and second truss members 
24 and 30 through holes 42. The third truss members 40 are interchangeable 
with adjustment features to compensate for fuselage 13 tolerances. 
Adjustment is provided by means of discrete axial installation positions 
of the end fittings 41 mounted on each end of the tubular portion 43 of 
the third truss member 40. These installation positions are determined by 
quantification of assembly tolerances for the third truss member 40 
geometry selected. Each of the end fittings 41 and extremities of the 
tubular portion 43 of the third truss member 40 are match fitted to 
achieve these installation positions as defined by the boltholes 42. 
Selecting the axial length of the respective end fittings 41 results in 
the ability to vary the length of the third truss member 40 upon 
installation. 
Referring now to FIGS. 2 and 5, fuselage 13 stringers 32 are spliced at the 
web member 21 to simplify pressure sealing and installation of the 
assembly module 20. This is accomplished by utilizing back to back 
machined splice fittings 35 on either side 25 or 26 of the web member 21 
bolted together to mechanically join the stringer 32 to the assembly 
module 20 interface. The machined splice fittings 35 nest inside the 
stringers 32 and are mechanically fastened to the stringer web (not shown) 
and flanges 36. In this manner, the continuity of the web member 21 is 
retained and the potential for pressure leaks is minimized. Note that 
while FIG. 5 depicts a stringer having an inverted hat cross-section, any 
structural cross section suitable for a stringer 32 design could be used 
and may be spliced by conforming splice fittings 35. The assembly module 
20 is sealed to the fuselage 13 by means of an arrangement of separate 
peripheral edge members 37 aligned and fastened to the outer periphery of 
the assembly module 20 web member 21. The peripheral edge members 37 or 
"angles" in turn are fastened to the fuselage 13 inside surface 16. This 
arrangement provides a float feature for the entire module 20 which 
facilitates its installation and also provides two sealing surfaces to the 
fuselage 13, namely, each peripheral edge member 37 attaching flange 38 
and a positive pressure sealing surface 26 with respect to the web member 
21. Positive fuselage 13 cabin pressure forces the web member 21 against 
the upstanding joined legs 39 of the peripheral edge members 37 which form 
the attaching flange 38. 
Assembly of the module 20 is modular and can be accomplished either inside 
the aircraft 10 or outside the aircraft utilizing standard assembly 
fixtures. Each web segment 23 with its corresponding first truss member 24 
and ring segment 33 constitute a subassembly. Ring member 28 construction 
is based on the use of ring segments 33 preferably having a channel 
cross-section with one flange (not shown) of the channel attached to the 
web member 21. The channel portion of the ring segment 33 as shown in FIG. 
2 extends in a direction normal to the web member surface 25 and the 
opposite second channel flange of the web segment (not shown) is free. 
This arrangement permits the first truss member 24 to be connected to the 
ring member 28 in a simple splice arrangement wherein web and flanges of 
each respective ring segment 33 are mechanically fastened to each other 
with brackets or any other desired means (not shown). The ring member 28 
is made up of several segments 33 joined as separate assemblies or it may 
be fabricated as a single assembly depending on the diameter selected. As 
individual segments 33, it is possible to fabricate the channel portion of 
each ring segments 33 as an extrusion formed to the desired radius. The 
channel cross section (not shown) of the ring member 28 also adds out of 
plane stiffness to the web member 21 and supports the center section of 
the web 21. 
For assembly outside of the aircraft 101 the web segments 23 are joined 
together to provide a complete web member 22. Assembly of the module 
inside of an aircraft 10 involves setting up a locating tool of 
conventional design (not shown) within the fuselage. The locating tool is 
positioned on the floor structure of the fuselage 13 and optically 
aligned. It is assembled and built up from this position and contains 
fixturing with appropriate indexing features to permit the module 20 web 
segments 23 to be located inside the fuselage 13 for joining and 
subsequent attachment to the interior surface 16 of the fuselage 13. The 
tool itself is portable and can be dismantled and re-assembled for re-use. 
It is designed to position and hold the web segments 23 to enable the 
build-up of the module to be self contained inside the fuselage. The tool 
preferably utilizes a truss type construction with a square framework 
lined up adjacent to the web 21 plane. The framework preferably contains 
indexing mounts from which the web segments 23 are positioned and held in 
the fuselage 13. The assembly of the locating tool consists of hinged and 
bolted joints allowing the tool to be dismantled and moved in and out of 
the fuselage through existing openings (i.e. passenger doors, cargo 
doors). The locating tool concept is adaptable to a full range of aircraft 
sizes. The individual web sections 23 are positioned and then joined 
together to form the web member 21. In either approach, the second 30 and 
third 40 truss members are positioned inside the aircraft 10 for final 
installation. The locating tool may be of any desired design and does not 
form part of this invention. 
Locating, positioning and final installation incorporate float features to 
preclude built in stresses and preloading. The web member 21 location and 
separate attachment to the fuselage 13 by means of the peripheral edge 
member arrangement 37 and installation adjustment in the diagonal third 
truss members 40 provides a float capability to accommodate tolerance 
buildups and the prevention of preloading resulting in undesirable 
built-in stresses. The peripheral edge member 37 attachment as shown in 
FIG. 5 allows positioning flexibility, while the third truss member 40 
attachment to the second truss member 30 is made after the web member 21 
is secured in place. Thus the truss assemblies 22 are installed stress 
free and independent of the web member 21 installation. 
The design is sufficiently flexible to be applicable to a range of aircraft 
10 sizes and shapes wherein only dimensioning is affected. Assembly inside 
the fuselage 13 using a portable locating tool (not shown) provides the 
option of modifying existing aircraft 10 for new internal fuselage 13 
configurations involving separation of pressurized and non-pressurized 
sections of the fuselage 13. It allows replacement of existing damaged 
pressure dome structures which may be uneconomical to repair or replace in 
kind. Since all of the containment module 20 elements are modular and 
easily built up and assembled from a single locating tool on or off the 
aircraft, size and weight of the module can be readily accommodated and 
managed. 
By virtue of its modular design and use of discrete web segments 23 and 
truss assemblies 22, which preferably incorporate common features, the 
module 20 of this invention has excellent durability and damage tolerance 
characteristics. Redundancy exists in the number of radial truss 
assemblies 22 selected. Sectoring of the web member 21 provides crack 
arrestment capability. Uniformity of load distribution is provided through 
tooling and assembly techniques and eccentricities are minimized through 
simple geometry and conventional mechanical connections. Tooling and 
assembly, by building up the containment module 20 in a sequence that 
precludes preloads and built-in stresses, ensures that the installed 
assembly will react applied loads uniformly. In addition, the use of tight 
tolerance holes and attaching hardware for the truss assembly 22 enhances 
the ability of the structure to distribute loads uniformly. All joints are 
designed to be axially aligned using standard clevis arrangements so that 
no secondary bending effects due to offsets are produced. 
Truss members 24, 30 and 40 are readily inspectable because they are 
discrete and accessible. Overall fuselage 13 stiffness and structural 
continuity is maintained because the basic fuselage shell 13 is 
undisturbed. The design lends itself to both new production and retrofit 
because of the simplicity of its assembly. The web member 21 portion of 
the module 20 can easily accommodate add on structure (not shown) for 
equipment mounting, penetrations for systems and their installations as 
well as external mounting for rear fuselage engine configurations. This 
would not be possible with conventional pressure dome structures 12. The 
truss assembly 22 network provides an added stability feature, as well as 
the ability to react out of plane loads without inducing undesirable 
secondary effects into the plane of the web member 21, for these add on 
options. 
The pressure containment module 20 of this invention is comparable in 
weight to the prior art pressure dome configuration 12 while mitigating 
interface and assembly complexities and without introducing the complex 
geometry features associated with the dome configuration 12. 
The pressure containment module 20 is far superior both in weight and 
manufacturability to flat pressure bulkheads used where in-plane load 
reaction structure is required. This is because it combines the efficiency 
of a dome design 12 for reacting pressure with the in-plane structural 
members required to accommodate planar loads. Reparability of the module 
20 is enhanced because of its modular construction and the use of discrete 
truss assemblies 22, simple parts and subassemblies. No special tools are 
required and on-aircraft repair can be readily accomplished. 
In summary, the module 20 of this invention is lightweight, modular and 
retrofitable. It is compliant with FM aircraft certification requirements 
for structural integrity. The module 20 is readily inspectable and 
repairable. It minimizes structural complexities associated with geometry, 
eccentricities, assembly and installation in an aircraft 10. The module 20 
accommodates other structural applications such as mission equipment 
mounting, engine supports, system penetrations and installations, etc. It 
is a viable alternative to both traditional pressure dome 10 and flat 
pressure (not shown) bulkhead construction designs. 
While the invention has been described in reference to an airplane type of 
aircraft it is applicable to any desired aircraft having a need to 
separate higher-pressure portions from lower-pressure portions such as 
pressurized and unpressurized portions. Such aircraft include helicopters, 
gliders, missiles, rockets and spacecraft to name a few. While the module 
20 has been described with respect to metal or alloy components it could 
be constructed of any suitable material including fiber composites. While 
the web member 21 preferably has a circular peripheral edge 27 it can have 
any desired shape adapted to mate with the inside surface 16 of the 
fuselage 13. In the case of non-circular web members 21 the ring member 28 
is located at the effective center of the web member 21 and the truss 
assemblies 22 preferably extend radially outwardly therefrom. 
It should be understood that the foregoing description is only illustrative 
of the invention. Various alternatives and modifications can be devised by 
those skilled in the art without departing from the spirit of the 
invention. Accordingly the present invention is intended to embrace all 
such alternatives, modifications and variances which fall within the scope 
of the appended claims.