CA2031928A1 - Strain reduced airplane skin - Google Patents
Strain reduced airplane skinInfo
- Publication number
- CA2031928A1 CA2031928A1 CA002031928A CA2031928A CA2031928A1 CA 2031928 A1 CA2031928 A1 CA 2031928A1 CA 002031928 A CA002031928 A CA 002031928A CA 2031928 A CA2031928 A CA 2031928A CA 2031928 A1 CA2031928 A1 CA 2031928A1
- Authority
- CA
- Canada
- Prior art keywords
- foam
- skin
- bulkheads
- aircraft
- invention according
- 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.)
- Abandoned
Links
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
- B64C1/12—Construction or attachment of skin panels
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C44/00—Shaping by internal pressure generated in the material, e.g. swelling or foaming ; Producing porous or cellular expanded plastics articles
- B29C44/02—Shaping by internal pressure generated in the material, e.g. swelling or foaming ; Producing porous or cellular expanded plastics articles for articles of definite length, i.e. discrete articles
- B29C44/12—Incorporating or moulding on preformed parts, e.g. inserts or reinforcements
- B29C44/18—Filling preformed cavities
- B29C44/186—Filling multiple cavities
-
- 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
- B64C1/10—Bulkheads
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/10—Manufacturing or assembling aircraft, e.g. jigs therefor
-
- 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
- B64C2001/0054—Fuselage structures substantially made from particular materials
- B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
-
- 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
- B64C2001/0054—Fuselage structures substantially made from particular materials
- B64C2001/0081—Fuselage structures substantially made from particular materials from metallic materials
-
- 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
ABSTRACT
An aircraft (10) has the space between the bulkheads (12) on the interior of the aircraft (10) filled with a polyisocyanurate solid closed cell foam material (20). The foam (20) is applied so that it adheres to both the inside of the aluminum skin (16) of the fuselage and the facing sides (24) of the bulkheads (12). The foam (20) may be applied by either spraying (Fig. 6) or by pouring (Fig. 5) the resin with appropriate catalyst materials to cause the resin to form the foam (20). The foam (20) acts to significantly strengthen the aircraft (10) structure and thereby increase the time of usage of aging aircraft (10) before the catastrophic failure of the adherence of the skin (16) to the bulkheads (12).
An aircraft (10) has the space between the bulkheads (12) on the interior of the aircraft (10) filled with a polyisocyanurate solid closed cell foam material (20). The foam (20) is applied so that it adheres to both the inside of the aluminum skin (16) of the fuselage and the facing sides (24) of the bulkheads (12). The foam (20) may be applied by either spraying (Fig. 6) or by pouring (Fig. 5) the resin with appropriate catalyst materials to cause the resin to form the foam (20). The foam (20) acts to significantly strengthen the aircraft (10) structure and thereby increase the time of usage of aging aircraft (10) before the catastrophic failure of the adherence of the skin (16) to the bulkheads (12).
Description
~)3~9~g STR~IN REDUCED ~IRPL~NE SKI~I
This invention relates to covering the inside sur-face of the metal skin of an airplane with a polymer -foam material~ and more particularly~ t~ covering that sur-face with a palyisocyanurate foam entirely between adjacent S bulkheads ~or the purpose of reducing the chance of struc-tural failure due to metal fatigue caused by many cycles o-f pressurization and depressuriæatiQn of the aircraft~
~ 5 the fleet of commercial and military aircraft agest certain time delayed problems are becoming more and more current. ~mong these problemsl aircra-ft have been ex-periencing metal fatigue due to the cyclic cabin presur-ization when climbing to cruising altitude and cabin de-pressuri2ation when landing. The cyclic e-ffect of landing and climbing to cruising altitude already has caused some aircraft structural failure and i5 predicted to hecome a major cause for the retirement of otherwise good aircraft in the coming years. The pressurization to approximately 12 pounds per square inch di-fference at altitude has in effect created a cold work of the aluminum al,loy skin with microscopic crack propagation at certain points of the air-craft skin. The prop~gation of the cracks lead to metal fatigue and eventual catastrophic destruction of the skin~
~ ircraft operating in the commercial or military environment must, by design~ transport passengers or pay-load -from one location to another. Particularly modern com-mercial jet aircra-ft, that typically takeoff from an air-203~928 part and climb to a cruising altitude, require the cabin tobe pressuri2ed in order to maintain inside ca~in pressure cl~se to sea level pressure and thereby assure the comfort of the passengers and crew. Because of the necessity o-f pressuri2ation at higher altitudes, the pressure difference between inside the cabin and the outside ambient causes the metal skin of the aircraft to deflecta that is to bulge outward, in order to sustain the higher inside pressure.
The aluminum panels covering the fuselage section of the aircraft have sufficient strength to easily maintain structural integrity as long as the pressure is maintained constant. Howeverl the aircraft must of nece~sity land to discharge the passengers and this necessitates the pressure differential hetween the inside of the fuselage and the ambient being reduced to zern. With a zero pressure ~if-ference, no metal deflection, or hulging is occurring. ~s the cycle repeats, a certain amount of internal molecular rearrangement in the aluminum panels forming the aircraft skin, occurs ~similar to the cold working of metal) during each pressurization~depressurization cycle. This molecular rearrangement may displace some af the atomic crystals in the skins.
Normally, the minute amount of flexing occurring during cabin pressurization does not displace the crystal lattice of the metal significantly as long as the metal remains in an elastic deformation configuration with re-spect to metal stress. The pressurization/depressurizatian cycle continues over many tens of thousands of landings and ~331~2~
takeoffs with corresponding cycles of the flexing o-f the skin. ~s in cold work of metals, at some point the metal becomes harder due to the rearrangement af atoms in the crystal lattice. This of course increases the resistance of the metal to yield ta an applied force. When the cycles continue~ small fissures start to show up near the dis-placed layers of atoms. These -Fissures allow the remaining metallic crystals to carry the load of the separated atoms~
thereby increasing the stress. Stress is the product o~
force divided by area. The remaining atoms carrying the load now are under a greater stress. The atoms yield one by one, as would the constituents of a tug of war. The fis-sures are propagated into small cracks and su~sequently larger cracks. ~t some point the remaining atoms holding the load are overwhelmed and a catastrophic failure occurs.
One solution to this problem is that the manufac-turer could increase the thickness of the aluminum skin, which would give the outside wall additional strength and reduce the flexure, thereby preventing metal fatigue for close to an indefinite timespan. This solutinn would in-crease the cost and more importantly~ increase the weight of the aircraft. The increased aircraft weight would, in turn~ redu~e the payload of the aircraft, and thereby sig-nificantly increase the per passenger mile operating cost.
Moreover, this solution is not practical for the many thou-sands of aircraft already in service.
This invention relates to covering the inside sur-face of the metal skin of an airplane with a polymer -foam material~ and more particularly~ t~ covering that sur-face with a palyisocyanurate foam entirely between adjacent S bulkheads ~or the purpose of reducing the chance of struc-tural failure due to metal fatigue caused by many cycles o-f pressurization and depressuriæatiQn of the aircraft~
~ 5 the fleet of commercial and military aircraft agest certain time delayed problems are becoming more and more current. ~mong these problemsl aircra-ft have been ex-periencing metal fatigue due to the cyclic cabin presur-ization when climbing to cruising altitude and cabin de-pressuri2ation when landing. The cyclic e-ffect of landing and climbing to cruising altitude already has caused some aircraft structural failure and i5 predicted to hecome a major cause for the retirement of otherwise good aircraft in the coming years. The pressurization to approximately 12 pounds per square inch di-fference at altitude has in effect created a cold work of the aluminum al,loy skin with microscopic crack propagation at certain points of the air-craft skin. The prop~gation of the cracks lead to metal fatigue and eventual catastrophic destruction of the skin~
~ ircraft operating in the commercial or military environment must, by design~ transport passengers or pay-load -from one location to another. Particularly modern com-mercial jet aircra-ft, that typically takeoff from an air-203~928 part and climb to a cruising altitude, require the cabin tobe pressuri2ed in order to maintain inside ca~in pressure cl~se to sea level pressure and thereby assure the comfort of the passengers and crew. Because of the necessity o-f pressuri2ation at higher altitudes, the pressure difference between inside the cabin and the outside ambient causes the metal skin of the aircraft to deflecta that is to bulge outward, in order to sustain the higher inside pressure.
The aluminum panels covering the fuselage section of the aircraft have sufficient strength to easily maintain structural integrity as long as the pressure is maintained constant. Howeverl the aircraft must of nece~sity land to discharge the passengers and this necessitates the pressure differential hetween the inside of the fuselage and the ambient being reduced to zern. With a zero pressure ~if-ference, no metal deflection, or hulging is occurring. ~s the cycle repeats, a certain amount of internal molecular rearrangement in the aluminum panels forming the aircraft skin, occurs ~similar to the cold working of metal) during each pressurization~depressurization cycle. This molecular rearrangement may displace some af the atomic crystals in the skins.
Normally, the minute amount of flexing occurring during cabin pressurization does not displace the crystal lattice of the metal significantly as long as the metal remains in an elastic deformation configuration with re-spect to metal stress. The pressurization/depressurizatian cycle continues over many tens of thousands of landings and ~331~2~
takeoffs with corresponding cycles of the flexing o-f the skin. ~s in cold work of metals, at some point the metal becomes harder due to the rearrangement af atoms in the crystal lattice. This of course increases the resistance of the metal to yield ta an applied force. When the cycles continue~ small fissures start to show up near the dis-placed layers of atoms. These -Fissures allow the remaining metallic crystals to carry the load of the separated atoms~
thereby increasing the stress. Stress is the product o~
force divided by area. The remaining atoms carrying the load now are under a greater stress. The atoms yield one by one, as would the constituents of a tug of war. The fis-sures are propagated into small cracks and su~sequently larger cracks. ~t some point the remaining atoms holding the load are overwhelmed and a catastrophic failure occurs.
One solution to this problem is that the manufac-turer could increase the thickness of the aluminum skin, which would give the outside wall additional strength and reduce the flexure, thereby preventing metal fatigue for close to an indefinite timespan. This solutinn would in-crease the cost and more importantly~ increase the weight of the aircraft. The increased aircraft weight would, in turn~ redu~e the payload of the aircraft, and thereby sig-nificantly increase the per passenger mile operating cost.
Moreover, this solution is not practical for the many thou-sands of aircraft already in service.
2~3~28 C~
~ better sQlution would oe to find a technique and structure to strengthen the existiny the aluminum skin of aircr~ft. Any practic~1 solution to elimin~te, or extend ~ell into the future~ the problem of metal fatigue by strengthening the e~isting al~minum skin of an airLraft must meet several requirements~ First, the solution must be of reasnnaole cost and second, the solution must not significantly increase the weight of the aircraft~ Further, any solution must not increase the fire ha~ard in the air-1~ craft in the e~ent of a crashs and particularly must notemit any noxious gases when burning. In addition, the so-lution should be of a type which can be used to upgrade existing aircraft, as well as being included in the manu-facture of new aircraft.
1~ One known material which could be useful to provide a sol~ltion to the metal fatigue problem is a polyisocyanu-rate foam material. This mAterial has a low density, does not burn and adheres to the aluminum components of an air-craft. Furtherl polyisocyanurate foam material has a ten-~0 sile and compressive stréngth sufficient to pre~ent the aluminum skin From fracturing and is non-soluble in any solvent typically found on an aircra~t. In addition, ,ooly-i60cyanurate foam is an excellent insulating material and does not absorb moisture to any significant extent.
2S The prior art has suggested adding various materi~
als to the structural components and skin of an aircra-ft.
For example~ U.S. Patent 1,852t146 to E.~ ~arns et al, entitled~ "Resilient Coating", suggests coating the struc-2~31~
tural ~omponents of an aircraft with rubber to reduce noise, corrosion and reduce component failure resulting from the massive shock of a crash. ~ similar shock absorb-ing solution to crashes for vehicles, including aircraft, is suggested in United States Patent 3,687,401 in the name of ~.~. Moore~ entitleo, "Light-Weight~ Wreck-Resistant Caoin". This technique would not increase the slowly oc-curring metal fatigue failure due to the cyclic prQSsuri~a-tion~depressuri2ation of the aircraft, would burn and emit massive amounts of smoke and would add too much weight to the aircraft.
In United ~tates Patent 2,81q,03Z, in the name of Q.J. Detrie et al, entitled "~ircraft Fuselage Having Panel Damping Material", suggests adding sheets of a wave dampening material on the inner sur~ace of the aluminum skin for the purpose of preventing vi~rations from causing structural failure of an aircraft. There is no discussion in the Detrie et al Patent about metal fatigue due to ex-pansion and contraction during the life of the aircraft.
Further, the solution suggested by Detrie et al does not solve that problem oecause expansion and contraction cannot be dampened. ~ similar damping solution with other materi-als i5 also suggested in United States Patent 3,160,549 in the name of D.B. Caldwell et al~ entitled "Vibration Damp-2S ing Structure".
United States Patent 4,235~3~ in the name ofWilliam P. Johnson, entitled "rhermal lnsulation For ~ir~
craft Fuselage"~ shows an airplane fuselage having a poly-2l~31~28 ~srethane foam attached to the in~ide of the aluminum skinby a matrix of standoffs for the purpose of thermal insula-tion. The type of structure suggested by 30hnson will pro-vide no inhîbition against metal fatigue due to cyclic cab-in pressuri-~ation and depressuriz~tion. Further~ the poly-urethane material will burn and emit no~ious gases and should not be ~Ised in an aircraft. Other uses of polyure-thane used in aircraft i5 shown in United States Pat~nt 4~QQ1~581 in the name of H.E. Littell~ Jr., erltitled "Lami-nated Aircraft Windshield".
Polyisocyanurate resins are descrioed in United States Patent ~,211,703 in the name of L.G. Gilman et al~
entitled "Polyaliphatic Polyisocyanurate laminating Resin Prepared In The Presence Of Q Cocatalyst System".
Polymer foams have been used to strengthen aircraft wings, as indicated by United States Patent 3,519,228 in the name of L.J. Windecker, entitled "Airfoil Structure"
and United States Patent 3,645,481 in the name of L.R.
Purdy, entitled "~irfoil Structur-e". The problem being solved by these patents, however, is unrelated to the prob-lem of metal fatigue caused by the oyclic pressuriz~tion~depressurization of the fuselage of the aircraft.
In accordance with one aspect of this invention~
thQre is provided in an airplane body formed by a thin met-al ma~erial affixed to~ a plurality of bulkhe~ds, the im-provement of affixing a non-flammable polymer -foam material to the interior side of the metal material and to the oulk-heads.
~3i928 In accordance with a second aspect of this inven-tion, there i5 provided a method of strengthening the skin of an aircraft. such skin being affixed aver a plurality of spaced bùlkheads. The method comprises the step of ~ffix-ing a layer of a polymer foam material to the interior sideof the skin and the facing sides of the bulkheads.
One preferred embodiment of the subject invention i5 hereafter described with specific reference to the fol-lowing Figures, in which:
Figure 1 shows a perspective view of a conventional aircraft fuselage tructure of the prior art, Figure Z shows a cross-sectional view, taken across lines 2-2 of Figure 1, of the aluminum skin and bulkheads of the structure shown in Figure l;
Figure 3 shows a perspective view Df the aircraft fuselage s~ructure having a polyisocyanurate foam material affixed to the inner surface of the aluminum skin and to the facing sides of the bulkheads;
Figure 4 shows a cross-sectional view of the alumi-num skin and bulkheads of the structure shown in Figure 3;
Figure 5 show~ one technique of affixing the poly-isocyanurate foam;
Figure ~ shows another technique of affixing the polyisocyanurate foam;
S Figure 7 shows an axially loaded member having a force P from each end -thereof;
Figure 3 shows a beam member being deflected by a force P; and ~3192~
Figure 9A show a composite beam of fuam material affixed to a thin sheet of aluminum; and Figure 9~ shows the foam material equivalent of the composite oeam shown in Figure ~.
S Referring to Figures 1 and 2, a section of the fu-selage 10 o-F a conventional aircraft i5 shown. Fuselage 10 i5 framed by bulkheads lZ and stringers 14 in a known man-ner and enclosed by thin sheets of aluminum forming skin 16. The skin i5 typic~lly forty mils ~0.04 inches) thick surrounding the fuselage 10, although in certain high stress ar~as, such as around the cockpit at the front of the fuselage 10, the thicknes may be a5 much as eighty mils thick. The sheets of aluminum skin 16 are affixed to the oulkhe~ds 12 by conventional fastening means, such as rivets 18.
~ fter an aircraft takes off and climbs to altitude, the air pressure within fuselage 10 is maintained at ap-proximately one atmosphere in order to provide for the com-fort of the passengers. However, the air pressure outside20 of the aircraft decreases to Llose to nothing at cruising altitudes around 40,000 feet. This creates a pressure P
against skin 16 in a direction outward from fuselage 10.
The pressure P causes skin 16 to expand to a position 16', as seen in Figure 2. This expansion i5 limited by rivets 18 into bulkheads 12~ whereby the skin between each pair of adjacent bulkheads is expanded. This expansion creates particular stress at the rivets 18 holding the skin against the bulkheads. When the ~ircraft land6, the pressure P
2 ~ 8 difference ~etween the interior of the aircraft and the ambient is reduced to zero and the skin returns to the non-expanded shape 16 seen in Figure 1.
~ fter many cycles of expansion and contraction~
metal fatigue f~ilure can occur and the skin can become dislodged from the rivets 18 at the bulkheads 12 and fall off the ~ircraft. If this occurs when -the aircraft is aloft~ an almost instantaneous depressurization of the fu-selage lQ occurs? with the possibility of loss of life and property and almost assured damage.
Referring now to Figures ~ and 4, an improved con-struction of an aircraft fuselage 10 is shown. ln Figures 3 and 4~ identical components previously described with respect to Figures 1 and~ 2 are given identical reference numerals. The only thing that has been added to the struc-ture shown in Figures 3 and 4, relative to the structur~shown in Figures 1 and 2, is the layer of polyisocyanurate foam 20 which has been formed between the inner surface 22 of aluminum skin lb and the facing edge surfaces 24 of bulkheads 12. Polyisocyanurate foam 20 is a well known closed cell foam material and is generally used as an insu-lator in refrigeration systems. This material has many advantages over other foam materials, such as polyurethane, which would make it useful in an aircraft. However, it has not heretofore been used in any aircraft applications.
The ~ef~l pr~pertie ~ssociated with polyisocyanu-rate foam are that it will not hurn, it has a low density of not less than two pounds per cubic foot, and it adheres 2 ~ 2 ~
to aluminum very well. Fur-ther ? polyisocyanurate ~oam is an excellent thermal insulator and ab~orbs virtually no mQisture. In addition7 it i5 insoluole in any solvent typ-ically associated with an airplane, such as gaso1ine, kero-sene or cleaning agents typically used~ The one property of polyisocyanurate foam which make it particularly advan-tageous as shawn in Figures 3 and 4 is that the compressive and tensile strength is about 35 pounds per square inch, which is much greater than the approximately 12 pounds per square inch pressure dif~erential between the pressurized interior of fuselage 10 anà the ambient at cruising alti-tudes. While polyisocyanurate foam has particular advan-tages~ other closed cell polymer foams may be used, such as polyurethane foams~ if one i5 willing to accept the risk of flammability of such material in the event of a crash.
~ s seen in Figures 3 and 4~ foam 20 is provided in the channels 26 formed by the inner side Z2 of skin 16 and the facing 6ides 24 of bulkheads 12. Foam 20 may oe ap-plied by either pouring the polyisocyanurate r-esin in each channel 26, as will be explained hereafter with respect to Figure 5, or by spraying the polyisocyanurate resin in each channel 26, as will be explained herea~ter with respect to Figure 6. Regardless of how the polyisocyanurate resin is applied9 it expands by aoout thirty times and adheres to the aluminum of both the sur~ace 22 and surfaces 24 within the channel Z6~ thereby creating a unitary structure. The amount of stress required to de-laminate the foam Z0 from the aluminum surfaces ZZ and 24 is greater than the foam itself, and thus, the foam will be pulled apart before it de-laminates.
Referring now to Figures 5 and 6, the manner of applying polyisocyanurate foam 20 over aluminum skin 1 will now be des~ribed. There are two way6 in whirh the foam 20 may be applied and these two ways are by pouring7 as indicated by Figure 5, and by spraying, as indicated by Figure ~. Whether the foam 20 is to be applied by pouring or spraying, the polyisocyanurate resin initi~lly applied will increase in volume by about thirty times in forming the polyisocyanurate foam 20. In addition, the amount of strength required must be determined and this determination i5~ in part, based upon the distance between bulkheads. It i5 sugge~ted that where the bulkheads 12 are separated by three feet, two inches of foam 20 be applied; that where the bulkheads 12 are separated by six feet5 three inches of foam 20 be applied and where the bulkheads 12 are separated by up to nine feet, four inches of foam oe applied.
Referring first to Figure 5, the method of pouring the polyisocyanurate resin to form the polylsocyanurate foam 20 will be described. In order to pour any resin to form a foam, a form must be created by closing char,nel 2 in order to contain the resin and resulting foam 20. ~e-cause of the shape of bulkheads 12 in an aircraft are gen-erally closed loop non-circular shapes and further, since the bulkheads 12 are not uniform in size or shape throuyh-out an airplane~ apparatus must be developed adaptable for 2031~2~
use on many different size and shaped bulkheads~ hs seen in Figure 5, this apparatus includes a thin sheet 36, such as commercially available heavy duty aluminum Foil or fi-berglass sheeting. and a flexible backing 3~. ~acking 38 may be a series of laterally expandable narrow strips hingedly affixed together and si~ed to fit against adjacent bulkheads 12. ~acking 38 may be designed to position foil 36 either against the in~ard -fating edges of the bulkheads lZ, as shown in Figure 5~ or within channel 2~ if the de-sired thickness of foam 20 i5 less than the thickness of channel 26. By making backing apparatus 38 adJustable and flexible, it may be used to pour the resin for many differ-ent shaped bulkhead lZ shapes.
When utiliziny the pouring method of Figure 5, the aluminum foil 36 remains in place after the foam 20 has formed from the poured resin. Foil 36 then acts to provide additional strength and further as a vapor barrier to pre-vent any moisture absorption by foam 20. In pouring the polyisocyanurate resin, a calculation of the volume en-closed in channel 2~ by the backing apparatus 38 and foil 3L is first made and the proper amount of resin is added to fill that volume. ~pproximately two pounds of resin per cubic foot of enclosed channel 26 volume may be used. If too much resin i5 added7 the resulting excess foam 20 will ~5 have to be removed by mechanical means, such as knives, saws or other cutters7 since no solvent is av~ilable to remove the foam without damaging the aluminum components 12 and 20. Each pour should be for a portion of each closed ~3~92~
laop channel 26 in order to control the application of the material. However, a plurality of oacking appar~tus ~ and foil 3~ may be set in place fQr por-tions in different chan-nels 26 and each may be poured at generally the same time.
It should be noted that it only requires a few minutes for the poured resin to form the foam.
Referring now to Figure ~, the polyisocyanurate re~in may be applied by using existing spraying apparatus 40. Such apparatus is commonly used to spray urethane foam resins in other applic~tions, such as wall insulation.
Because converting the polyisocyanurate resin to a solid foam requires a two step cataly~ation~ extreme care ;n the spraying application is required. The amount of resin sprayed in channel 26 is to be commensurate with the de-sired thickness of the resulting foam 20 and this can belearned by the person operating the spraying equipment 40 with practice. It may be necessary to clean the alu~inum surfaces within channel Z6 with a solvent to remove grease and dirt prior to spraying the foam. This is particularly true for retrofitting existing aircraft and may also be necessary for new air~rafts as well. This cleaning may also be necessary when pouring the resin~ as shown in ~ig-ure 5. One solv~nt which may be used is methylene chlo-ride.
In the sprayin~ applica~i~n, th~ resin sh~uld be applied to a single closed loop channel 26 at the same ~im~. Thr~ ~im~ r~qui -~d ~r ~h~ 1 iquid rr~sin ~ rn in~
the solid foam i5 sufficient to permit an entire channel to ~3~28 be sprayed during a commnn application ~gain, i-F excess foam i5 formed, it may be removed oy mechanical cutter means~ such as knives and saws.
The applied Foam 2~ and aluminum skin 1~ act as a composite unit and the resulting deflectian between skin lb7 during no pressurization, anci skin 16~ a during pressur~
ization, is substantially reduced7 as seen in Figure 4.
This precludes additional slip dislocations -From Dccurring in the metal due to maintaining the stresse~ well into the elastic region of the stress/strain curve. The principal function of the palyisocyanurate foam layer 20 is to add mechanical strength. Not only does foam Z0 act as a com-posite7 but the aluminum skin 16 is strengthened by having a support at every point throughout its surface area~
Should a crack occur in the -Future for some reason7 it is held to the foam Z0 and not allowed to buLk1e into the slipstream. Further, foam 20 acts a5 a block to prevent the escape of air and subsequent depressurization of the Fuselage 10 as a result o-F a small crack.
In addition to adding strength and preventing de-pressurizationl the foam 20 acts as an excellent noise and thermal insulator~ as well as acting as a barrier to fire and heat in the event of a crash and provide those extra few minutes for the passengers, during which time firecrews may be able to extinguish the fire. Further, if a crash or ditch occurs over water, the low density of foam Z07 acts as additiDnal flotation and could prevent the aircra-Ft from sinking.
2~31~8 In order to calculate the theoretical stiffness of the aluminum/polyisocyanurate composite ma-terial, reference i5 made to Figures 77 3, 9~ and 9BA These calculations are based upon a sprayinq application in which no backing alu-minum foil sheet 3~. The calculations for this type of abeam would provide better results.
Referring now to Figure 7 an axial loaded member 2R
having a length 1 and a wirth ~ is shown. When a foroe P
i5 applied in opposite directions from the two ends of mem-ber 28, member 28 expands by an amount S. The amount S maybe determined according to the equation:
S - P~ E
where E is Young~s Modulu~ for the material of member 28.
The stiffness of member 2~ may then be expressed as:
5tiffness = P/~ E~13 Figure 8 shows a member 30, which may be either a single material, such as aluminum, beam or a composite ma-terial, such as foam covered aluminum, beam. For member 30, the deflection S may be expressed by the equation:
~ - P~13/48 E~I
and the stiffness may be expressed as:
.~
' 2~3~92~
Stiffness = Pt~ = 48 E~Itl3 To determine the increase in strength ~ a foam and aluminum composite member 30 over an aluminum member alone, ane must determine the ~tiffness ratio, which i~:
Stiffness R~tio - (Stiffness)~=f(5tiffness)~"
or = E,~-IC-13 ~lJE~ I~L~13~=
where I is the Moment of Inertia of the particular beam.
Referring to Figure ~, the aluminum skin 1~ and foam 20, as constructed according to Figures 3 and 4, i5 6chematically shown as the composite beam 32. This compos-ite beam 32 has a width W and the thickness of the foam 20is t~r~ the thickness of the aluminum 16 i5 t~ and the neutral axis 34, i5 y. Since the width of the alumin~lm skin lh and foam 20 is the same~ the stiffness ratio can be expressed as:
Stiffness Ratio = EC I~tE~ :I Cl]
One can determine the Young's Modulu6 E for bath ~luminum and polyisocyanurate foam from available handbooks and find that:
E rfE~ n = 1~40 ~ince n = 40, the composite beam of Figure ~ may be shown as in Figure 9B. Thus, the neutral axis 34 may be found by the equation:
(wt.~ t .~r J 2 ~ t~ I ) + ( nw t~ L ) ~ t~.~ .L / Z ) s y ~ ( n~w-t/~ :L . + W ' t1") or y = ~t.,v2fZ + n~t~ Z ~ t.t.~t~l~f~n tf.~, + t.,.) ~2]
From Figure ~B, the moment of Inertia I can be cal-culated as:
Ic a ( W ' t.f 3 )JlZ + (w-t.~)C(t ~rJ 2)+t ~t 3, - y } 2 ~ (n~w~t ft L3 ) J 1 Z
-~ (n-W~t~.tL)~y-(t~ 2)}2 C3]
The moment of inertia for the aluminum skin 12 is:
I ft:L = W~ t~.t L ~ 12 C4]
From equation Cl~, the stiffness ratio then b~-comes:
Sti~fness Qatio s ~lJn)~(I,.JIf.~ 5 2't and a calculation of the stiffness ratio of Equation C5 using equations ~3] and L4] yields the following results:
- ' , `
2~3~92~
1~
T~E~LE I _ __ Thickness ~inches)Stiffness Rati~
Ql ~ninum F~am 0.04 O.S 100 1 .01 172 1 . 5347 1 2.~ 7~61 0 . 05 0 . 5 99 1 0 1 . 0~42 1.5 1912 2.0 41ZZ
0.06 0.~; 61 1.0 39Z
1 .51 172 2 . 02534 0.07 0.5 40 1 . 0258 1 . 5773 2.0 1677 0.08 O.S 2a 1~0 180 1 . 5539 2.01 172 , ,' 2~
Thus, when one applies two inches of polyisocyanurate fo~m over a forty mil thick aluminum skin 1~ the composite material has a stiffness 7461 times as great as the alumi-num skin lb alone. Accordingly, the amount of deflection of the skin of fuselage 10 with foam 20 applied will be significantly less and the met~l fatigue problem will be correspondingly reduced A
~ better sQlution would oe to find a technique and structure to strengthen the existiny the aluminum skin of aircr~ft. Any practic~1 solution to elimin~te, or extend ~ell into the future~ the problem of metal fatigue by strengthening the e~isting al~minum skin of an airLraft must meet several requirements~ First, the solution must be of reasnnaole cost and second, the solution must not significantly increase the weight of the aircraft~ Further, any solution must not increase the fire ha~ard in the air-1~ craft in the e~ent of a crashs and particularly must notemit any noxious gases when burning. In addition, the so-lution should be of a type which can be used to upgrade existing aircraft, as well as being included in the manu-facture of new aircraft.
1~ One known material which could be useful to provide a sol~ltion to the metal fatigue problem is a polyisocyanu-rate foam material. This mAterial has a low density, does not burn and adheres to the aluminum components of an air-craft. Furtherl polyisocyanurate foam material has a ten-~0 sile and compressive stréngth sufficient to pre~ent the aluminum skin From fracturing and is non-soluble in any solvent typically found on an aircra~t. In addition, ,ooly-i60cyanurate foam is an excellent insulating material and does not absorb moisture to any significant extent.
2S The prior art has suggested adding various materi~
als to the structural components and skin of an aircra-ft.
For example~ U.S. Patent 1,852t146 to E.~ ~arns et al, entitled~ "Resilient Coating", suggests coating the struc-2~31~
tural ~omponents of an aircraft with rubber to reduce noise, corrosion and reduce component failure resulting from the massive shock of a crash. ~ similar shock absorb-ing solution to crashes for vehicles, including aircraft, is suggested in United States Patent 3,687,401 in the name of ~.~. Moore~ entitleo, "Light-Weight~ Wreck-Resistant Caoin". This technique would not increase the slowly oc-curring metal fatigue failure due to the cyclic prQSsuri~a-tion~depressuri2ation of the aircraft, would burn and emit massive amounts of smoke and would add too much weight to the aircraft.
In United ~tates Patent 2,81q,03Z, in the name of Q.J. Detrie et al, entitled "~ircraft Fuselage Having Panel Damping Material", suggests adding sheets of a wave dampening material on the inner sur~ace of the aluminum skin for the purpose of preventing vi~rations from causing structural failure of an aircraft. There is no discussion in the Detrie et al Patent about metal fatigue due to ex-pansion and contraction during the life of the aircraft.
Further, the solution suggested by Detrie et al does not solve that problem oecause expansion and contraction cannot be dampened. ~ similar damping solution with other materi-als i5 also suggested in United States Patent 3,160,549 in the name of D.B. Caldwell et al~ entitled "Vibration Damp-2S ing Structure".
United States Patent 4,235~3~ in the name ofWilliam P. Johnson, entitled "rhermal lnsulation For ~ir~
craft Fuselage"~ shows an airplane fuselage having a poly-2l~31~28 ~srethane foam attached to the in~ide of the aluminum skinby a matrix of standoffs for the purpose of thermal insula-tion. The type of structure suggested by 30hnson will pro-vide no inhîbition against metal fatigue due to cyclic cab-in pressuri-~ation and depressuriz~tion. Further~ the poly-urethane material will burn and emit no~ious gases and should not be ~Ised in an aircraft. Other uses of polyure-thane used in aircraft i5 shown in United States Pat~nt 4~QQ1~581 in the name of H.E. Littell~ Jr., erltitled "Lami-nated Aircraft Windshield".
Polyisocyanurate resins are descrioed in United States Patent ~,211,703 in the name of L.G. Gilman et al~
entitled "Polyaliphatic Polyisocyanurate laminating Resin Prepared In The Presence Of Q Cocatalyst System".
Polymer foams have been used to strengthen aircraft wings, as indicated by United States Patent 3,519,228 in the name of L.J. Windecker, entitled "Airfoil Structure"
and United States Patent 3,645,481 in the name of L.R.
Purdy, entitled "~irfoil Structur-e". The problem being solved by these patents, however, is unrelated to the prob-lem of metal fatigue caused by the oyclic pressuriz~tion~depressurization of the fuselage of the aircraft.
In accordance with one aspect of this invention~
thQre is provided in an airplane body formed by a thin met-al ma~erial affixed to~ a plurality of bulkhe~ds, the im-provement of affixing a non-flammable polymer -foam material to the interior side of the metal material and to the oulk-heads.
~3i928 In accordance with a second aspect of this inven-tion, there i5 provided a method of strengthening the skin of an aircraft. such skin being affixed aver a plurality of spaced bùlkheads. The method comprises the step of ~ffix-ing a layer of a polymer foam material to the interior sideof the skin and the facing sides of the bulkheads.
One preferred embodiment of the subject invention i5 hereafter described with specific reference to the fol-lowing Figures, in which:
Figure 1 shows a perspective view of a conventional aircraft fuselage tructure of the prior art, Figure Z shows a cross-sectional view, taken across lines 2-2 of Figure 1, of the aluminum skin and bulkheads of the structure shown in Figure l;
Figure 3 shows a perspective view Df the aircraft fuselage s~ructure having a polyisocyanurate foam material affixed to the inner surface of the aluminum skin and to the facing sides of the bulkheads;
Figure 4 shows a cross-sectional view of the alumi-num skin and bulkheads of the structure shown in Figure 3;
Figure 5 show~ one technique of affixing the poly-isocyanurate foam;
Figure ~ shows another technique of affixing the polyisocyanurate foam;
S Figure 7 shows an axially loaded member having a force P from each end -thereof;
Figure 3 shows a beam member being deflected by a force P; and ~3192~
Figure 9A show a composite beam of fuam material affixed to a thin sheet of aluminum; and Figure 9~ shows the foam material equivalent of the composite oeam shown in Figure ~.
S Referring to Figures 1 and 2, a section of the fu-selage 10 o-F a conventional aircraft i5 shown. Fuselage 10 i5 framed by bulkheads lZ and stringers 14 in a known man-ner and enclosed by thin sheets of aluminum forming skin 16. The skin i5 typic~lly forty mils ~0.04 inches) thick surrounding the fuselage 10, although in certain high stress ar~as, such as around the cockpit at the front of the fuselage 10, the thicknes may be a5 much as eighty mils thick. The sheets of aluminum skin 16 are affixed to the oulkhe~ds 12 by conventional fastening means, such as rivets 18.
~ fter an aircraft takes off and climbs to altitude, the air pressure within fuselage 10 is maintained at ap-proximately one atmosphere in order to provide for the com-fort of the passengers. However, the air pressure outside20 of the aircraft decreases to Llose to nothing at cruising altitudes around 40,000 feet. This creates a pressure P
against skin 16 in a direction outward from fuselage 10.
The pressure P causes skin 16 to expand to a position 16', as seen in Figure 2. This expansion i5 limited by rivets 18 into bulkheads 12~ whereby the skin between each pair of adjacent bulkheads is expanded. This expansion creates particular stress at the rivets 18 holding the skin against the bulkheads. When the ~ircraft land6, the pressure P
2 ~ 8 difference ~etween the interior of the aircraft and the ambient is reduced to zero and the skin returns to the non-expanded shape 16 seen in Figure 1.
~ fter many cycles of expansion and contraction~
metal fatigue f~ilure can occur and the skin can become dislodged from the rivets 18 at the bulkheads 12 and fall off the ~ircraft. If this occurs when -the aircraft is aloft~ an almost instantaneous depressurization of the fu-selage lQ occurs? with the possibility of loss of life and property and almost assured damage.
Referring now to Figures ~ and 4, an improved con-struction of an aircraft fuselage 10 is shown. ln Figures 3 and 4~ identical components previously described with respect to Figures 1 and~ 2 are given identical reference numerals. The only thing that has been added to the struc-ture shown in Figures 3 and 4, relative to the structur~shown in Figures 1 and 2, is the layer of polyisocyanurate foam 20 which has been formed between the inner surface 22 of aluminum skin lb and the facing edge surfaces 24 of bulkheads 12. Polyisocyanurate foam 20 is a well known closed cell foam material and is generally used as an insu-lator in refrigeration systems. This material has many advantages over other foam materials, such as polyurethane, which would make it useful in an aircraft. However, it has not heretofore been used in any aircraft applications.
The ~ef~l pr~pertie ~ssociated with polyisocyanu-rate foam are that it will not hurn, it has a low density of not less than two pounds per cubic foot, and it adheres 2 ~ 2 ~
to aluminum very well. Fur-ther ? polyisocyanurate ~oam is an excellent thermal insulator and ab~orbs virtually no mQisture. In addition7 it i5 insoluole in any solvent typ-ically associated with an airplane, such as gaso1ine, kero-sene or cleaning agents typically used~ The one property of polyisocyanurate foam which make it particularly advan-tageous as shawn in Figures 3 and 4 is that the compressive and tensile strength is about 35 pounds per square inch, which is much greater than the approximately 12 pounds per square inch pressure dif~erential between the pressurized interior of fuselage 10 anà the ambient at cruising alti-tudes. While polyisocyanurate foam has particular advan-tages~ other closed cell polymer foams may be used, such as polyurethane foams~ if one i5 willing to accept the risk of flammability of such material in the event of a crash.
~ s seen in Figures 3 and 4~ foam 20 is provided in the channels 26 formed by the inner side Z2 of skin 16 and the facing 6ides 24 of bulkheads 12. Foam 20 may oe ap-plied by either pouring the polyisocyanurate r-esin in each channel 26, as will be explained hereafter with respect to Figure 5, or by spraying the polyisocyanurate resin in each channel 26, as will be explained herea~ter with respect to Figure 6. Regardless of how the polyisocyanurate resin is applied9 it expands by aoout thirty times and adheres to the aluminum of both the sur~ace 22 and surfaces 24 within the channel Z6~ thereby creating a unitary structure. The amount of stress required to de-laminate the foam Z0 from the aluminum surfaces ZZ and 24 is greater than the foam itself, and thus, the foam will be pulled apart before it de-laminates.
Referring now to Figures 5 and 6, the manner of applying polyisocyanurate foam 20 over aluminum skin 1 will now be des~ribed. There are two way6 in whirh the foam 20 may be applied and these two ways are by pouring7 as indicated by Figure 5, and by spraying, as indicated by Figure ~. Whether the foam 20 is to be applied by pouring or spraying, the polyisocyanurate resin initi~lly applied will increase in volume by about thirty times in forming the polyisocyanurate foam 20. In addition, the amount of strength required must be determined and this determination i5~ in part, based upon the distance between bulkheads. It i5 sugge~ted that where the bulkheads 12 are separated by three feet, two inches of foam 20 be applied; that where the bulkheads 12 are separated by six feet5 three inches of foam 20 be applied and where the bulkheads 12 are separated by up to nine feet, four inches of foam oe applied.
Referring first to Figure 5, the method of pouring the polyisocyanurate resin to form the polylsocyanurate foam 20 will be described. In order to pour any resin to form a foam, a form must be created by closing char,nel 2 in order to contain the resin and resulting foam 20. ~e-cause of the shape of bulkheads 12 in an aircraft are gen-erally closed loop non-circular shapes and further, since the bulkheads 12 are not uniform in size or shape throuyh-out an airplane~ apparatus must be developed adaptable for 2031~2~
use on many different size and shaped bulkheads~ hs seen in Figure 5, this apparatus includes a thin sheet 36, such as commercially available heavy duty aluminum Foil or fi-berglass sheeting. and a flexible backing 3~. ~acking 38 may be a series of laterally expandable narrow strips hingedly affixed together and si~ed to fit against adjacent bulkheads 12. ~acking 38 may be designed to position foil 36 either against the in~ard -fating edges of the bulkheads lZ, as shown in Figure 5~ or within channel 2~ if the de-sired thickness of foam 20 i5 less than the thickness of channel 26. By making backing apparatus 38 adJustable and flexible, it may be used to pour the resin for many differ-ent shaped bulkhead lZ shapes.
When utiliziny the pouring method of Figure 5, the aluminum foil 36 remains in place after the foam 20 has formed from the poured resin. Foil 36 then acts to provide additional strength and further as a vapor barrier to pre-vent any moisture absorption by foam 20. In pouring the polyisocyanurate resin, a calculation of the volume en-closed in channel 2~ by the backing apparatus 38 and foil 3L is first made and the proper amount of resin is added to fill that volume. ~pproximately two pounds of resin per cubic foot of enclosed channel 26 volume may be used. If too much resin i5 added7 the resulting excess foam 20 will ~5 have to be removed by mechanical means, such as knives, saws or other cutters7 since no solvent is av~ilable to remove the foam without damaging the aluminum components 12 and 20. Each pour should be for a portion of each closed ~3~92~
laop channel 26 in order to control the application of the material. However, a plurality of oacking appar~tus ~ and foil 3~ may be set in place fQr por-tions in different chan-nels 26 and each may be poured at generally the same time.
It should be noted that it only requires a few minutes for the poured resin to form the foam.
Referring now to Figure ~, the polyisocyanurate re~in may be applied by using existing spraying apparatus 40. Such apparatus is commonly used to spray urethane foam resins in other applic~tions, such as wall insulation.
Because converting the polyisocyanurate resin to a solid foam requires a two step cataly~ation~ extreme care ;n the spraying application is required. The amount of resin sprayed in channel 26 is to be commensurate with the de-sired thickness of the resulting foam 20 and this can belearned by the person operating the spraying equipment 40 with practice. It may be necessary to clean the alu~inum surfaces within channel Z6 with a solvent to remove grease and dirt prior to spraying the foam. This is particularly true for retrofitting existing aircraft and may also be necessary for new air~rafts as well. This cleaning may also be necessary when pouring the resin~ as shown in ~ig-ure 5. One solv~nt which may be used is methylene chlo-ride.
In the sprayin~ applica~i~n, th~ resin sh~uld be applied to a single closed loop channel 26 at the same ~im~. Thr~ ~im~ r~qui -~d ~r ~h~ 1 iquid rr~sin ~ rn in~
the solid foam i5 sufficient to permit an entire channel to ~3~28 be sprayed during a commnn application ~gain, i-F excess foam i5 formed, it may be removed oy mechanical cutter means~ such as knives and saws.
The applied Foam 2~ and aluminum skin 1~ act as a composite unit and the resulting deflectian between skin lb7 during no pressurization, anci skin 16~ a during pressur~
ization, is substantially reduced7 as seen in Figure 4.
This precludes additional slip dislocations -From Dccurring in the metal due to maintaining the stresse~ well into the elastic region of the stress/strain curve. The principal function of the palyisocyanurate foam layer 20 is to add mechanical strength. Not only does foam Z0 act as a com-posite7 but the aluminum skin 16 is strengthened by having a support at every point throughout its surface area~
Should a crack occur in the -Future for some reason7 it is held to the foam Z0 and not allowed to buLk1e into the slipstream. Further, foam 20 acts a5 a block to prevent the escape of air and subsequent depressurization of the Fuselage 10 as a result o-F a small crack.
In addition to adding strength and preventing de-pressurizationl the foam 20 acts as an excellent noise and thermal insulator~ as well as acting as a barrier to fire and heat in the event of a crash and provide those extra few minutes for the passengers, during which time firecrews may be able to extinguish the fire. Further, if a crash or ditch occurs over water, the low density of foam Z07 acts as additiDnal flotation and could prevent the aircra-Ft from sinking.
2~31~8 In order to calculate the theoretical stiffness of the aluminum/polyisocyanurate composite ma-terial, reference i5 made to Figures 77 3, 9~ and 9BA These calculations are based upon a sprayinq application in which no backing alu-minum foil sheet 3~. The calculations for this type of abeam would provide better results.
Referring now to Figure 7 an axial loaded member 2R
having a length 1 and a wirth ~ is shown. When a foroe P
i5 applied in opposite directions from the two ends of mem-ber 28, member 28 expands by an amount S. The amount S maybe determined according to the equation:
S - P~ E
where E is Young~s Modulu~ for the material of member 28.
The stiffness of member 2~ may then be expressed as:
5tiffness = P/~ E~13 Figure 8 shows a member 30, which may be either a single material, such as aluminum, beam or a composite ma-terial, such as foam covered aluminum, beam. For member 30, the deflection S may be expressed by the equation:
~ - P~13/48 E~I
and the stiffness may be expressed as:
.~
' 2~3~92~
Stiffness = Pt~ = 48 E~Itl3 To determine the increase in strength ~ a foam and aluminum composite member 30 over an aluminum member alone, ane must determine the ~tiffness ratio, which i~:
Stiffness R~tio - (Stiffness)~=f(5tiffness)~"
or = E,~-IC-13 ~lJE~ I~L~13~=
where I is the Moment of Inertia of the particular beam.
Referring to Figure ~, the aluminum skin 1~ and foam 20, as constructed according to Figures 3 and 4, i5 6chematically shown as the composite beam 32. This compos-ite beam 32 has a width W and the thickness of the foam 20is t~r~ the thickness of the aluminum 16 i5 t~ and the neutral axis 34, i5 y. Since the width of the alumin~lm skin lh and foam 20 is the same~ the stiffness ratio can be expressed as:
Stiffness Ratio = EC I~tE~ :I Cl]
One can determine the Young's Modulu6 E for bath ~luminum and polyisocyanurate foam from available handbooks and find that:
E rfE~ n = 1~40 ~ince n = 40, the composite beam of Figure ~ may be shown as in Figure 9B. Thus, the neutral axis 34 may be found by the equation:
(wt.~ t .~r J 2 ~ t~ I ) + ( nw t~ L ) ~ t~.~ .L / Z ) s y ~ ( n~w-t/~ :L . + W ' t1") or y = ~t.,v2fZ + n~t~ Z ~ t.t.~t~l~f~n tf.~, + t.,.) ~2]
From Figure ~B, the moment of Inertia I can be cal-culated as:
Ic a ( W ' t.f 3 )JlZ + (w-t.~)C(t ~rJ 2)+t ~t 3, - y } 2 ~ (n~w~t ft L3 ) J 1 Z
-~ (n-W~t~.tL)~y-(t~ 2)}2 C3]
The moment of inertia for the aluminum skin 12 is:
I ft:L = W~ t~.t L ~ 12 C4]
From equation Cl~, the stiffness ratio then b~-comes:
Sti~fness Qatio s ~lJn)~(I,.JIf.~ 5 2't and a calculation of the stiffness ratio of Equation C5 using equations ~3] and L4] yields the following results:
- ' , `
2~3~92~
1~
T~E~LE I _ __ Thickness ~inches)Stiffness Rati~
Ql ~ninum F~am 0.04 O.S 100 1 .01 172 1 . 5347 1 2.~ 7~61 0 . 05 0 . 5 99 1 0 1 . 0~42 1.5 1912 2.0 41ZZ
0.06 0.~; 61 1.0 39Z
1 .51 172 2 . 02534 0.07 0.5 40 1 . 0258 1 . 5773 2.0 1677 0.08 O.S 2a 1~0 180 1 . 5539 2.01 172 , ,' 2~
Thus, when one applies two inches of polyisocyanurate fo~m over a forty mil thick aluminum skin 1~ the composite material has a stiffness 7461 times as great as the alumi-num skin lb alone. Accordingly, the amount of deflection of the skin of fuselage 10 with foam 20 applied will be significantly less and the met~l fatigue problem will be correspondingly reduced A
Claims (23)
1. In an aircraft body formed by a thin metal material having an interior side affixed to a plurality of bulkheads, the improvement of a non-flammable polymer foam material affixed to the interior side of said metal material between said bulkheads and to said bulkheads.
2. The invention according to claim 1 wherein said polymer foam material has a compressive and tensile strength sufficient to prevent said metal material from fracturing due to the expansion and contraction of said metal material resulting from the continual pressurization and de-pressurization of said aircraft.
3. The invention according to claim 2 wherein said polymer foam material is a closed cell material.
4. The invention according to claim 3 wherein said polymer foam material is an insulating material.
5. The invention according to claim 4 wherein said polymer foam material is polyisocyanurate foam.
6. The invention according to claim 1 wherein said polymer foam material is a closed cell material.
7. The invention according to claim 1 wherein said polymer foam material is an insulating material.
8. The invention according to claim 1 wherein said polymer foam material is polyisocyanurate foam.
9. The invention according to claim 8 wherein said polyisocyanurate foam is applied with a thickness of at least approximately two inches.
10. The invention according to claim 1 wherein said polymer foam material is applied with a thickness of at least approximately two inches.
11. The invention according to claim 1, 2, 3, 4, 5, 6, 7, 8, 9 or 10, further comprising a pliable material affixed to said foam material and on a side of said foam material remote from said metal material so that said foam material is sandwiched between said metal material and said pliable material.
12. The invention according to claim 11, wherein said metal material is an aircraft skin and said pliable material is a sheet of metal foil.
13. The invention according to claim 12, wherein said sheet of metal foil is an aluminum foil sheet.
14. A method of strengthening an aircraft skin, such skin having an interior side affixed over a plurality of spaced bulkheads, said method comprising the step of:
affixing a layer of a polymer foam material to the interior side of said skin between said bulkheads and to said bulkheads.
affixing a layer of a polymer foam material to the interior side of said skin between said bulkheads and to said bulkheads.
15. The method according to claim 14 wherein said foam material is selected to have a compressive and tensile strength sufficient to prevent said skin from fracturing due to the expansion and contraction of said skin resulting from the continual pressurization and de-pressurization of said aircraft.
16. The method according to claim 15 wherein said foam is a polyisocyanurate foam.
17. The method according to claim 15 wherein said step of affixing includes affixing said foam material with a thickness-of at least approximately two inches.
18. The method according to claim 15 wherein said step of affixing includes spraying materials on the interior side of said skin between adjacent bulkheads to cause said foam to thereafter form.
19. The method according to claim 18 wherein said foam is a polyisocyanurate foam.
20. The method according to claim 19 wherein said method further includes cleaning said skin interior surface and facing surfaces of said bulkheads.
21. The method according to claim 15 wherein said step of affixing includes placing a pliable material against the interior side of said bulkheads and pouring said polymer material to form said foam between the interior of said skin and said pliable material.
22. The method according to claim 21 wherein said foam is a polyisocyanurate foam.
23. The method according to claim 14 wherein said foam is a polyisocyanurate foam.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US456,533 | 1983-01-06 | ||
| US07/456,533 US5251849A (en) | 1989-12-26 | 1989-12-26 | Strain reduced airplane skin |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2031928A1 true CA2031928A1 (en) | 1991-06-27 |
Family
ID=23813139
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002031928A Abandoned CA2031928A1 (en) | 1989-12-26 | 1990-12-11 | Strain reduced airplane skin |
Country Status (10)
| Country | Link |
|---|---|
| US (1) | US5251849A (en) |
| EP (1) | EP0435650B1 (en) |
| JP (1) | JPH04218493A (en) |
| KR (1) | KR910011586A (en) |
| AT (1) | ATE109091T1 (en) |
| AU (2) | AU6811890A (en) |
| BR (1) | BR9006576A (en) |
| CA (1) | CA2031928A1 (en) |
| DE (1) | DE69011069T2 (en) |
| ES (1) | ES2063297T3 (en) |
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| DE102007003278B4 (en) * | 2007-01-23 | 2013-10-31 | Airbus Operations Gmbh | Hull structure component of an aircraft or spacecraft with a foam layer as thermal insulation |
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| DE102009060706B4 (en) * | 2009-12-29 | 2014-12-04 | Airbus Operations Gmbh | Method and device for producing a stiffening structure for an aircraft fuselage segment and a stiffening structure |
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-
1989
- 1989-12-26 US US07/456,533 patent/US5251849A/en not_active Expired - Lifetime
-
1990
- 1990-12-11 CA CA002031928A patent/CA2031928A1/en not_active Abandoned
- 1990-12-17 AU AU68118/90A patent/AU6811890A/en not_active Withdrawn
- 1990-12-21 AU AU68481/90A patent/AU637441B2/en not_active Ceased
- 1990-12-24 KR KR1019900021674A patent/KR910011586A/en not_active Application Discontinuation
- 1990-12-24 AT AT90314251T patent/ATE109091T1/en active
- 1990-12-24 ES ES90314251T patent/ES2063297T3/en not_active Expired - Lifetime
- 1990-12-24 DE DE69011069T patent/DE69011069T2/en not_active Expired - Fee Related
- 1990-12-24 EP EP90314251A patent/EP0435650B1/en not_active Expired - Lifetime
- 1990-12-26 BR BR909006576A patent/BR9006576A/en not_active IP Right Cessation
- 1990-12-26 JP JP2406733A patent/JPH04218493A/en active Pending
Also Published As
| Publication number | Publication date |
|---|---|
| BR9006576A (en) | 1991-10-01 |
| DE69011069T2 (en) | 1995-03-30 |
| EP0435650B1 (en) | 1994-07-27 |
| AU637441B2 (en) | 1993-05-27 |
| KR910011586A (en) | 1991-08-07 |
| JPH04218493A (en) | 1992-08-10 |
| AU6811890A (en) | 1991-07-04 |
| ES2063297T3 (en) | 1995-01-01 |
| AU6848190A (en) | 1991-07-04 |
| US5251849A (en) | 1993-10-12 |
| DE69011069D1 (en) | 1994-09-01 |
| ATE109091T1 (en) | 1994-08-15 |
| EP0435650A1 (en) | 1991-07-03 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request | ||
| FZDE | Discontinued |