GB2224000A - Fuselage or pressure vessel structure - Google Patents
Fuselage or pressure vessel structure Download PDFInfo
- Publication number
- GB2224000A GB2224000A GB8824723A GB8824723A GB2224000A GB 2224000 A GB2224000 A GB 2224000A GB 8824723 A GB8824723 A GB 8824723A GB 8824723 A GB8824723 A GB 8824723A GB 2224000 A GB2224000 A GB 2224000A
- Authority
- GB
- United Kingdom
- Prior art keywords
- shell
- fuselage
- longitudinal
- outer portion
- structural members
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000005728 strengthening Methods 0.000 claims 2
- 238000000034 method Methods 0.000 claims 1
- 238000010276 construction Methods 0.000 description 4
- 239000000463 material Substances 0.000 description 3
- RTAQQCXQSZGOHL-UHFFFAOYSA-N Titanium Chemical compound [Ti] RTAQQCXQSZGOHL-UHFFFAOYSA-N 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 125000004122 cyclic group Chemical group 0.000 description 1
- 238000007689 inspection Methods 0.000 description 1
- 229910052751 metal Inorganic materials 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 229910052755 nonmetal Inorganic materials 0.000 description 1
- 230000000750 progressive effect Effects 0.000 description 1
- 230000002787 reinforcement Effects 0.000 description 1
- 239000010936 titanium Substances 0.000 description 1
- 229910052719 titanium Inorganic materials 0.000 description 1
Classifications
-
- 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
- B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
- B64C1/064—Stringers; Longerons
-
- 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
- B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
- B64C1/068—Fuselage sections
-
- 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
- B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
- B64C1/12—Construction or attachment of skin panels
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
- Tires In General (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Description
1 M&C FOLIO: 230P57504 2 ',2 22 4 0 0 0 FUSELAGE WANGDOC: 1352P This
invention relates primarily to an aircraft fuselage. The invention can find application in the construction of a pressure-sealed aircraft fuselage. The invention can also be applied to heavy duty boilers and other hermetic vessels capable of withstanding pressurisation.
There is known an aircraft fuselage comprising an outside skin or shell reinforced by longitudinal and transverse structural members (stringers and bulkheads). The transverse structural members are intended to arrest the development of a longitudinal crack in the shell and are backed-up by additional bands of titanium provided between the shell and bulkhead along the perimeter.
However, the construction of fuselage having no back-up arrangement is susceptible to lateral cracks. Such a crack can grow to an excessive size and result in a failure of the fuselage in flight.
Therefore. this fuselage construction ensures one side back-up of the shell only along the transverse structural members.
2 An aircraft fuselage is known comprising a shell strengthened by intersecting longitudinal and transverse structural members, and having joints in the longitudinal direction.
In order to make the shell more fail-safe. additional reinforcement straps are provided on the transverse structural members at either side of the joint connected by eyed rods or high strength bolts.
In case of a partial failure of the fuselage shell at the longitudinal joint the eyed rods or high strength bolts serve as a back-up strength line for transmitting loads acting in the failure zone. The load carrying capacity (strength) of the structure is therefore maintained when a crack develops in flight.
However. the absence of shell back-up in the longitudinal direction can cause a lateral crack to develop to an excessive size.
What is desired is an aircraft fuselage, in which a crack developing in the longitudinal or lateral direction could be effectively arrested.
The present invention provides an aircraft fuselage (or other vessel) comprising a shell strengthened by longitudinal and transverse structural members. in which the shell has the form of interconnected continuous outer portion and inner grid-shaped portion conforming to the shape of the outer portion and having grid ribs extending along the structural members.
K 11 3 Preferably, the thickness of each grid rib of the inner portion of the shell ranges from 0.2 to 1.0 times the thickness of the outer portion of the shell, whereas the cross-sectional area of the ribs ranges from 0.1 to 1.0 times and from 0.1 to 0.7 times the cross-sectional area of the longitudinal and transverse members, respectively.
Making the shell from interconnected outer and inner portions enables the development of a crack to be effectively arrested within allowable limits during failure of the outer portion of the shell inside the grid, as the crack runs to the border of double thickness of the outer and inner portions of the shell. The crack is thereby arrested as the thickness of the outer portion of the shell comes in'--o play ensuring reduced stresses at the mouth of the expanding crack. The proposed range of the cross-sections and thicknesses of the ribs of the inner portion of the shell is determined by the efficiency of crack arresting. If the crosssectional area of the ribs of the inner portion of the shell is less than 0.1 times the area of the longitudinal or transverse structural members, and the rib thickness of the inner portion of the shell is less than 0.2 times the thickness of the outer portion of the shell, the crack can be arrested as efficiently as in a smooth one-piece shell without the inner grid portion.
An increase in the cross-sectional area of the ribs of 4 the inner portion of the shell to over 0.1 and 0.7 times the area of the longitudinal and transverse structural members respectively. as well as an increase in the rib thickness in the inner portion of the shell to greater than the thickness of the outer portion of the shell in inadvisable because it entails a greater weight of the structure.
The invention will now be described in greater detail, by way of example, with reference to a specific embodiment thereof taken in conjunction with the accompanying drawings, in which:
Figure 1 is a lateral section of an aircraft fuselage; Figure 2 shows the outer portion of the shell of the fuselage; Figure 3 shows the inner portion of the shell of the fuselage; Figure 4 is a section taken along the line W-W in Figure 1; Figure 5 is a section taken along the line V-V in Figure 1.
An aircraft fuselage comprises a shell 1 (Figure 1) strengthened by longitudinal structural members 2 and transverse structural members 3. The shell 1 has the form of a continuous outer portion 4 (Figure 2) and an inner portion 5 (Figure 3) having the shape of the outer portion 4 (Figure 2) and fashioned as a network or grid the ribs of which are arranged along the structural members 2,3 (Figure 1). The outer portion 4 (Figure 2) and inner portion 5 (Figure 3) can be made of the same material or of different materials (metal, non-metal, or a combination thereof). Intersecting ribs 6 (Figures 4,5) of the inner portion 5 (Figure 3) of the shell 1 are positioned along longitudinal and lateral axes 7,8 (Figures 4.5) of the structural members 2,3 (Figure 1). The thickness of each rib 6 (Figures 4,5) of the inner portion 5 (Figure 3) of the shell 1 (Figure 1) ranges from 0.2 to 1.0 times the thickness of the outer portion 4 (Figure 2) of the shell 1 (Figure 1). The cross- sectional area of the ribs 6 (Figures 4,5) ranges from 0.1 to 1.0 times and from 0.1 to 0.7 times the cross-sectional area of the longitudinal and transverse structural members 2 and 3, respectively.
The shell 1 (Figure 1) of the aircraft fuselage in the zone of a pressurized cabin (not shown) has an interior pressurisation acting cyclically with a period of one cycle per flight. During the service life of the aircraft the shell 1 is subject to 20 - 30 thousand load cycles. Routine operation of the aircraft causes fatigue or accidental damage to the shell 1 resulting from cyclic loads or mechanical damage both longitudinally or transversely of the shell 1. In order to prevent catastrophic failures, the proposed fuselage construction has an inner portion 5 (Figure 3) service i 6 as a crack arrester. In the case of formation of a crack in the outer portion 4 (Figure 2) of the shell 1 in one grid cell, accompanied by progressive growth of such a crack, ends of the crack run to the edge of the rib 6 (Figures 4,5) and tend to stop. This phenomenon is associated with the stress level in the shell 1 (Figure 1) which determines the rate of crack growth both in the grid cell and at a portion of the rib 6 (Figures 4,5) where greater localised thickness acts to reduce the stress in the shell 1 (Figure 1) thereby ensuring effective reduction in the crack growth rate. Therewith, the preferred geometry of the ribs ensures a margin of strength at the crack length within one cell even if the crack develops during flight. After-flight inspection is supposed to detect such a crack.
The above-described structure enables an increase in the life of an aircraft fuselage and the more efficient arrest of crack development. Depending on the material of the grid the efficiency of crack arresting is increased by a factor of 5 to 27.
J1 t 7
Claims (4)
1. An aircraft fuselage or other vessel comprising a shell in the form of a continuous outer portion and a grid-shaped inner portion. conforming to the shape of the outer portion, connected therewith. and having grid ribs extending along longitudinal and transverse structural members strengthening the shell.
2. A fuselage or vessel as claimed in claim 1, in which the thickness of each rib of the inner portion of the shell ranges from 0.2 to 1.0 times the thickness of the outer portion of the shell, whereas the crosssectional area of the ribs ranges from 0.1 to 1.0 times and from 0.1 to 0. 7 times the cross-sectional area of the longitudinal and transverse structural members, respectively.
3. A fuselage substantially as described with reference to, and as shown in. the accompanying drawings.
4. An aircraft having a fuselage according to any preceding claim.
Published 1990 atThe Patent Office, State House, 6671 High Holborn. LondonWCIR4TP. Further copies maybe obtainedfrom. The Patent Office Sales Branch. St Mary Cray. Orpington. Kent BR5 3RD- Printed by Multiplex techniques ltd, St Mary Cray. Kent. Con. 1187
4. An aircraft having a fuselage according to any preceding claim.
Amendments to the claims have been filed as follows 1. An aircraft fuselage or other vessel comprising a shell in the form of a continuous outer portion and a grid-shaped inner portion, conforming to the shape of the outer portion. connected therewith, and having longitudinal and circumferential crosspieces extending along longitudinal and circumferential structural members strengthening the shell.
2. A fuselage or vessel as claimed in claim 1, in which the thickness of each crosspiece of the inner portion of the shell ranges from 0.2 to 1.0 times the thickness of the outer portion of the shell. whereas the crosssectional area of the longitudinal crosspieces ranges from 0.1 to 1.0 times the cross-sectional area of the longitudinal and structural member and the cross-sectional area of the circumferential crosspieces ranges from 0.1 to 0.7 times.the cross-sectional area of the circumferential structural members.
3. A fuselage substantially as described with reference to, and as shown in. the accompanying drawings.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR8814714A FR2638708B1 (en) | 1988-11-10 | 1988-11-10 | IMPROVEMENT TO A FUSELAGE |
Publications (3)
Publication Number | Publication Date |
---|---|
GB8824723D0 GB8824723D0 (en) | 1988-11-30 |
GB2224000A true GB2224000A (en) | 1990-04-25 |
GB2224000B GB2224000B (en) | 1992-05-27 |
Family
ID=9371782
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB8824723A Expired - Lifetime GB2224000B (en) | 1988-11-10 | 1988-10-21 | Aircraft fuselage or other pressure vessel. |
Country Status (3)
Country | Link |
---|---|
DE (1) | DE3900167A1 (en) |
FR (1) | FR2638708B1 (en) |
GB (1) | GB2224000B (en) |
Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5170967A (en) * | 1990-02-28 | 1992-12-15 | Fuji Jukogyo Kabushiki Kaisa | Aircraft fuselage structure |
US5223067A (en) * | 1990-02-28 | 1993-06-29 | Fuji Jukogyo Kabushiki Kaisha | Method of fabricating aircraft fuselage structure |
GB2279930A (en) * | 1993-06-18 | 1995-01-18 | Royal Ordnance Plc | Aircraft structures |
US8061035B2 (en) | 2004-09-23 | 2011-11-22 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
US8157212B2 (en) | 2004-04-06 | 2012-04-17 | The Boeing Company | Composite barrel sections for aircraft fuselages and other structures, and methods and systems for manufacturing such barrel sections |
US8168023B2 (en) | 2004-11-24 | 2012-05-01 | The Boeing Company | Composite sections for aircraft fuselages and other structures, and methods and systems for manufacturing such sections |
US8388795B2 (en) | 2007-05-17 | 2013-03-05 | The Boeing Company | Nanotube-enhanced interlayers for composite structures |
JP2019151321A (en) * | 2018-03-02 | 2019-09-12 | ザ・ボーイング・カンパニーThe Boeing Company | Stringer transition through common base support |
Families Citing this family (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE102008012282A1 (en) | 2008-03-03 | 2009-09-17 | Airbus Deutschland Gmbh | Hull structure for airplane |
DE102010018932B4 (en) * | 2010-04-30 | 2013-06-13 | Airbus Operations Gmbh | Perimeter stiffening for an aircraft fuselage |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB534810A (en) * | 1940-02-26 | 1941-03-19 | David Henderson Sandilands | Stressed skin panels |
GB555496A (en) * | 1942-02-20 | 1943-08-25 | Henry Kremer | Improvements in or relating to structures built up of plywood or other light-weight laminated material |
GB559954A (en) * | 1942-09-07 | 1944-03-13 | Vultee Aircraft Inc | Airplane structures |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR845332A (en) * | 1938-10-28 | 1939-08-18 | Henschel Flugzeugwerke Ag | Device for stiffening smooth sheets |
-
1988
- 1988-10-21 GB GB8824723A patent/GB2224000B/en not_active Expired - Lifetime
- 1988-11-10 FR FR8814714A patent/FR2638708B1/en not_active Expired - Fee Related
-
1989
- 1989-01-05 DE DE3900167A patent/DE3900167A1/en not_active Withdrawn
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB534810A (en) * | 1940-02-26 | 1941-03-19 | David Henderson Sandilands | Stressed skin panels |
GB555496A (en) * | 1942-02-20 | 1943-08-25 | Henry Kremer | Improvements in or relating to structures built up of plywood or other light-weight laminated material |
GB559954A (en) * | 1942-09-07 | 1944-03-13 | Vultee Aircraft Inc | Airplane structures |
Cited By (18)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5223067A (en) * | 1990-02-28 | 1993-06-29 | Fuji Jukogyo Kabushiki Kaisha | Method of fabricating aircraft fuselage structure |
US5170967A (en) * | 1990-02-28 | 1992-12-15 | Fuji Jukogyo Kabushiki Kaisa | Aircraft fuselage structure |
GB2279930A (en) * | 1993-06-18 | 1995-01-18 | Royal Ordnance Plc | Aircraft structures |
GB2279930B (en) * | 1993-06-18 | 1997-03-26 | Royal Ordnance Plc | The protection of aircraft structures |
US8157212B2 (en) | 2004-04-06 | 2012-04-17 | The Boeing Company | Composite barrel sections for aircraft fuselages and other structures, and methods and systems for manufacturing such barrel sections |
US8182628B2 (en) * | 2004-04-06 | 2012-05-22 | The Boeing Company | Composite barrel sections for aircraft fuselages and other structures, and methods for systems for manufacturing such barrel sections |
US8382037B2 (en) | 2004-04-06 | 2013-02-26 | The Boeing Company | Composite barrel sections for aircraft fuselages and other structures |
US8882040B2 (en) | 2004-09-23 | 2014-11-11 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
US8061035B2 (en) | 2004-09-23 | 2011-11-22 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
US10689086B2 (en) | 2004-09-23 | 2020-06-23 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
US9738371B2 (en) | 2004-09-23 | 2017-08-22 | The Boeing Company | Splice joints for composite aircraft fuselages and other structures |
US8303758B2 (en) | 2004-11-24 | 2012-11-06 | The Boeing Company | Methods for manufacturing composite sections for aircraft fuselages and other structures |
US8418740B2 (en) | 2004-11-24 | 2013-04-16 | The Boeing Company | Composite sections for aircraft fuselages and other structures, and methods and systems for manufacturing such sections |
US8168023B2 (en) | 2004-11-24 | 2012-05-01 | The Boeing Company | Composite sections for aircraft fuselages and other structures, and methods and systems for manufacturing such sections |
US8657990B2 (en) | 2007-05-17 | 2014-02-25 | The Boeing Company | Nanotube-enhanced interlayers for composite structures |
US8388795B2 (en) | 2007-05-17 | 2013-03-05 | The Boeing Company | Nanotube-enhanced interlayers for composite structures |
JP2019151321A (en) * | 2018-03-02 | 2019-09-12 | ザ・ボーイング・カンパニーThe Boeing Company | Stringer transition through common base support |
JP7202194B2 (en) | 2018-03-02 | 2023-01-11 | ザ・ボーイング・カンパニー | Stringer transition through common base support |
Also Published As
Publication number | Publication date |
---|---|
GB2224000B (en) | 1992-05-27 |
GB8824723D0 (en) | 1988-11-30 |
FR2638708A1 (en) | 1990-05-11 |
FR2638708B1 (en) | 1994-05-06 |
DE3900167A1 (en) | 1990-07-12 |
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Legal Events
Date | Code | Title | Description |
---|---|---|---|
PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 19931021 |