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Aircraft floor panel developments at British Airways (1967–1973)
Aircraft floor panel developments at British Airways (1967-1973) K. B. ARMSTRONG
In 1967 testing started on balsa, pvc and aluminium alloy cores with aluminium alloy skins and a cost-effectiveness formula was developed to provide a basis for comparison. A specification for improved aluminium/balsa floors was produced and flight trials began with aluminium/aluminium honeycomb floors. Carbon fibre and later glass fibre came on the scene and a new specification was raised, based on more fundamental criteria. As a result over 500 cfrp/Nomex panels have been fitted in 747 aircraft and about 70 grp/Nomex panels. These are much lighter than earlier types of flooring and more cost-effective.
Rate of interest on capital
INTRODUCTION
K
The investigation of floor panel performance, with a view to selection of the best available product, began late in 1967. When the investigation started, the problem seemed to be one of resistance to indentation, as scrap panels showed core failures and delamination, usually beyond repair. Corrosion was and still is, a prime cause of failure of aluminium alloy panels fitted in wet areas. It would appear from observation of the inside faces of the skins that the adhesives used are sometimes corrosive in themselves. No data on indentation resistance could be obtained from research establishments and it was decided that testing would have to be carried out to compare available materials. It was appreciated that any information that could be obtained would have a wide applicability to anyone concerned with the indentation resistance of sandwich materials. The designers of aircraft, hovercraft, freight carrying pallets, road vehicles, railway carriages and ships were considered certain to benefit. As the cost of spares was the main reason for this investigation a cost-effectiveness formula was produced to provide the basis for comparison.
x --- Cash value of 1 Newton/annum for the airline or other operation concerned
COST-EFFECTI VENESS FORMULA
The formula suggested is a simple one: ee KL
+ xW=C L
where: p=
Prime cost of finished floor per square metre (ie including inserts, edge finishing and labour)
L=
Service life of panel in years
i
Development Engineer, Structures Group, British Airways Overseas Division, London Heathrow Airport, Hounslow, Middlesex, England
COMPOSITES. JU LY 1974
=
W:
Weight increase or decrease about a datum value
C = Total cost of 1 square metre of flooring per year The formula can be extended to include the manhour cost of removal over panel life/square metre and the cost of repair over panel life/square metre or, if weight is of no concern it can be reduced to ~eKL
L
Clearly for an aircraft like Concorde where the cash value weight is £56/Newton (£250/1b) the more expensive but lighter type of floor is inevitable, but even on subsonic aircraft their use is well justified. A I R C R A F T FLOORS, DESIGN AND MAINTENANCE CONSIDERATIONS
These need to be designed for their particular location ie toilet and galley areas, gangway, underseat and freight areas. Corrosion resistance is important in toilet and galley areas and indentation resistance needs to be greater in gangway, toilet and galley areas than under seats. Whenever corrodible materials are used particular attention is required to sealing at bolt holes and the use of sealed inserts at all bolt positions is recommended. Freight floors are subject to higher average loads and higher point loads and much greater impact loads and therefore need to be considered separately. The floor structure should be so designed that sandwich panels can drop straight in without requiring separate frames and should also be designed to ensure the use of the minimum number of different panels. In most modern aircraft almost every panel is different
165
and this creates a serious spares problem; for example the Boeing 747 has over 350 different panels, the 707 about 130 and the VC 10 has 75. Much greater.standardization could be achieved with considerable economy to manufacturer and operator. Floors need to be sufficiently stiff to avoid passengers feeling unsafe because of their deflection. It is also important, whatever type of floor is chosen, that the time from incipient to final failure should be fairly lengthy to give the operator the chance to change a defective panel at the next convenient maintenance check.
STA TIC DESIGN OF A FLOOR PANEL
This is usually carried out to a uniform pressure requirement although such a loading case only occurs if a freight door or hatch door is lost in flight. Unfortunately a freight door was lost in a recent incident and an adjacent part of the floor failed because it was not designed to resist cabin differential pressure [about 62 kN/m 2 (9 lbf/in2)]. Most passenger floors are designed to 27.6 kN/m2 (4 lbf/in2) and freight floors to 41.4 kN/m 2 (6 lbf/in2). This incident posed the question of providing blow-out panels in the floor to meet such an unlikely need. Data sheets exist for pressure loading but not for the point loading normally experienced. Bending moments under point loading can, however, be calculated if the formulae given on Page 135 of Ref. 1 are used. Optimum design of a panel under pressure loading can be worked out using the Hexcel Design Handbook 2 or Ref. 3. Having obtained the optimum strength panel for a given weight the next step is to ensure that its indentation resistance is adequate. The load and the size of indentor through which it is likely to be applied must be decided at this stage. The most common
O'?l
0.02 I
0.03 1
0.04 I
STA TIC INDENTATION TESTS
The results of a considerable number of tests are given in Fig.1. At the time these tests were done it was normal to use a 19 mm (0.75 in) diameter steel ball under a 1780 N (400 lb) load and to measure the permanent indentiation of the panel that resulted. The panel under test was supported on a steel plate, see Fig.2. This test really measures the compensation strength of the core and is favourable to balsa, which normally fails in shear but has good com-
Permanent Indentation-in 0.05 0'0(~ 0.07 i
I
L73skins
I
0.08
0-09
I
l
0-10 I
171 Balsa: 144,2 kg/m3191b/ft 3) averoge:6-35mm(O.25in)cor¢ O Balsa: 144.2 kg/m3(91b/ft 3) average:6.35 mm(O'25in)cor4 0 " 0 6 [Specification permits balsa to vary between 96 kg/m3 and P76kg/m3(61b/ft 3 to I IIb/ft 3) to give this average] X PVC: 134"6 kg/m3 (8"4 Ib/ft3): 9.5-3mm (0.375 in) core -0.05 + Aluminiu.m honeycomb: 121.7kg/m3(76 Ib/ft3):3.18 mm (O.125 in) celI,O.O25mm ( 0 ' 0 0 2 in) walls, 5052 alloy, 7.94mm (O,313in) core A PVC; 89.71 to IO8.9 kg/m3(5"6to 6'8 Ib/ft3): 7-11mm (0.28 in) core 0 . 0 4 "-~ PVC to BMS 4-13A: 9.53 mm (O.37in) core
Scatter bond : 3 pvc 134.6 kg/m 1'5C
mistake made by floor panel designers is that of using a lightweight core with heavy skins. As in road design, 'hard core' is necessary. Honeycomb cores of various types are superior to others because they are the only ones which produce a non-linear increase in indentation resistance with increasing weight. Fig. 10 indicates that a 12% increase in panel weight can produce a 60% increase in indentation resistance. This occurs because the buckling of the cell wall is a function of the square of the wall thickness. All other types of core have a roughly linear variation of strength witla core density. Honeycombs are also much superior in shear to other core types. For freight floors aluminium alloy honeycombs are superior to Nomex. Although Nomex is initially more resilient, once broken, it is completely useless at the point of damage. Aluminium honeycomb buckles under impact but does not break and retains a considerable amount of strength in the damaged state. A much dented VC 10 freight floor panel has survived 3½ years having been noticeably dented for most of that time. The important features of good skin materials are low weight, high modulus and high proof stress. Some ductility is essential for freight floor applications together with adequate scratch resistance.
{8"4 Ib/ft 3) L73 skins
1"25 E E
== I'O0
C
c ,.x
.u
... " ~
.~ 0.75
Scatter band: 3 pvc 89-71 kg/m
~ " "- ...
e-
0.03
(5"6 Ib/ft 3)
u3
U3
+-.,+. - O'O2
0.50
/El
I Cl~.~X
~'-+'--~X+
""--
A ""A --0.OI
0.25 Honeycomb : 3-18 mm (O.125 In ) cell, O'O25 mm (O.OO2in) walls, 5 0 5 2 alloy I
O
Fig.1
166
I
I
I
I
I
I
I
I
I
0.25 0-50 0-75 FOO 1.25 1-50 1.75 2.00 2.25 2.SO Permanent indentation depth (mm) du¢ to 1780 N (4OOIb) load on Igmm (0"75 in) diameter steel ball
Permanent indentation depth due to a load of 1780 N (400 Ib) with a 19 m m (0.75 in) diameter bell; (test piece fully supported)
COMPOSITES . J U L Y 1974
Clearly this is the maximum possible value. Specially conducted tests using a 203 mm (8 in) thick block of honeycomb with rods at 25.4 mm (1 in) intervals showed that Ec varied from the centre of the block to the surface. At the centre, where end effects were absent a value approaching the theoretical was obtained. Surface effects in balsa and honeycomb may be due to the following:
Loading device Centre to stopslipPoge o f boll
~
~ t _ 19 turn ( 0 " 7 5 in)diameter steel ba~landwic h
[ "
I~panel Steel plate
Fig.2
(a) Salsa (i) Some buckling of end fibres will be caused when cutting the wood;
J
Ball indentation test
pression properties. It is less than fair to honeycombs which are better under shear than compression. The mode of failure is a function of indentor size, core thickness and the ratio of shear properties to compression properties in a given core, see Figs 1 and 3. Fig.1 clearly shows the scatter of properties in both balsa and pvc. Although pvc is consistent in any one slice, it is foamed as 'loaff and different slices give different results due to density variation. The result of these early tests was BOAC Spec. 4BA-10,000. 5 This called for closer selection o f the balsa, to reduce scatter due to density variation and the use of L. 88 for the top skin, as tests with L. 73 had shown a clear improvement over L. 72 and this was assumed to be due to the higher proof stress of L. 73. Because the tests 6 later included edge supported tests which are more likely to suffer shear failure it was found that metal honeycombs were likely to be better than balsa. However, pending flight trials with metal honeycomb cores which have since proved successful, it was decided to stick with the aluminium/balsa already known to be good and likely to be better when supplied to the new specifications.
(ii)
Some deformation will occur in the adhesive layer;
(iii)
Under ball loading conditions the ends of the fibres will be rotated out of the vertical as deformation takes place and this will cause buckling at lower loads.
(b) Honeycomb (i) Some movement out of the vertical may occur at the edges during cutting; (ii)
Deformation of the adhesive layer;
(iii)
Ball loading will rotate cell walls out of the vertical and cause buckling at lower loads.
In both cases, (i) and (iii) above could not apply to pvc and its modulus is also less than that of the adhesive layer. Floor panels are always relatively thin and it is by no means certain that the theoretical value is the one to use for balsa or honeycomb.
0.25 i
0"50 i
50
THEORETICAL APPROACH TO INDENTATION RESISTANCE Because of the cost and quantity of testing required to provide the information for floor panel selection two theoretical methods were tried. 8,9 The first was later fbund to be in error, by the National Physical Laboratory, but consideration of it brought up points worth mentioning:
Core thickness- in 0.75 FO0 1.2 F50 i i i r C) 20'3 mm{ 0 8 in) diameter riot faced Indentor ® AvT.rogF_~tqct diamete.r~r : ~.=mmtu./~ inldiometer Dart at IOmm (0.39in)diameter riotriot indentor
1"75
4O ~50
| i
E
=~
1.25 ~
o 30
f
1 Membrane effects occur when indentation exceeds half skin thickness in depth; 2 Compression modulus of the core material is a governing factor in both formulae and accurate determination of this property can be difficult. The modulus for balsa is best obtained by wave velocity tests. For pvc the value can be fairly accurately obtained from compression tests. For other stiffer materials platen deformation and flatness can be so critical that only when allowance is made for these errors after the most accurate and careful testing can true values be obtainedA o For aluminium honeycomb the compression modulus is usually obtained as
2-00
1-00 ~ 2C
0"75 Indentor diameter Core thickness
I0
C>50 0.25
® 0
4 fshtmr fcom~sioa
20 Core thickr~ss-mm
30
Young's modulus x density of honeycomb density of aluminium
COMPOSITES . JULY 1974
Fig.3 Comparison o f core materials showing lines o f simultaneous failure in shear and compression
167
The work of Cox and Martin i I has been compared to test results by Raimondol. s who found that although the general trend of his results was in line with.the formula, calculated results were below those obtained by test. Further work is being carried out by Payen of RAE Farnborough. SPECIFICATION FOR LIGHTWEIGHTFLOORS
Soon after the new balsa specification came off the press British Airways were asked if any application for carbonfibre could be envisaged. The material seemed to have most of the properties required for a floor panel skin. except ductility and scratch resistance. For these reasons it has been used for passenger floors covered with carpet rather than for freight floors. The improvement aimed at by the 4BA-10,000 Spec s was really just an improvement on an existing material but was not based on any fundamental data. To put the work on a firmer basis it was decided to approach the Building Research Station and they provided details of loads applied by the foot in walking 7. Cox of the National Physical Laboratory provided a mathematical basis for calculating indentation resistance I l and this has since been published in Data Sheet form by the Engineering Sciences Data Unit 12. Using the Building Research data the indentation requirements for BOAC Specification E E - R 7 2 - 4 ( A ) 13 were formulated, suitable factors being included for fatigue and overload cases. It was also encouraging to find that the minimum indentation values of good quality aluminium/balsa panels, of known long life, came out to about the same as the values obtained by applying a factor of 1.5 to the walking loads supplied by the Building Research Station and confirmed by Brunel University 14 using a strain-gauged floor panel. A 10 mm (0.39 in) diameter blunt indenter was used for the tests. The diameter chosen was representative of heels at the time and was the only assumption made in producing the Specification. It was considered a reasonable choice because experience on the Britannia, which had separate ladies' and gentlemen's toilets, showed that although damage to the panel outside the 'Ladies' was more than the 'Gents', damage in the latter case was still significant. This is not surprising because a new heel on a male shoe is quite sharp and can have only line contact on impact with the floor. This view is confirmed by Harper 16. Indentation fatigue requirements were specified considering the performance of good existing materials. The CIBA Indentation fatigue test is empirical and comparative only and is sensitive to panel stiffness. Comparisons should be made only between panels of equal thickness
Table 1. CIBA Fatigue test results on balsa core/aluminium skin panels: the load was the same for each case Core thickness mm (in)
6.35 (0.25) 9.53 (0.375) 12.7
(0.5)
* w i t h cracking of the t o p skin
168
Time to failure h
7 2 ½"
intended for the same application. The results in Table 1 are from a series of CIBA Fatigue tests, and indicate the importance of the core thickness. To overcome the thickness problem and the difficulty presented by the fact that the indentation fatigue test is partly an impact test, a test on one spot was specified for 1 000 cycles at 70% of the specified static load, residual indentation strength to be measured afterwards. The figure of 1 000 cycles was obtained from a Douglas report which calculated this to be the likely number of cycles on one spot in a 5 year period. Binding strength had to meet design requirements. Design requirements also covered the shear load at bolt inserts." It is worth noting that the performance of inserts on test depends considerably on the pulling straps or test frame used. The method used by the aircraft manufacturer should be adopted in each case. To cater for loading both at and between floor beams two indentation tests were specified, one fully supported and one edge supported. It is necessary for the fully supported requirement to be higher than the edge supported case because no cushioning of the load due to panel deflection occurs at a floor beam. The food-cart wheel test was included because calculations at Brunel University 14 had shown this to be a severe case with the small hard wheels often used. The Specification has remained unaltered except that a deflection test has been added together with a creep test because of unfortunate experience with grp/Nomex panels. With these panels it is possible to reach strains sufficient to cause resin cracking unless the panel stiffness can be maintained above a critical level. Great effort has been put into solving this problem and the early difficulties have been overcome by one manufacturer. Any other design of grp would need to be carefully checked to ensure that an adequate life was likely to be obtained. Because of the higher cost of lightweight panels a life of 15 000 to 20 000 h is necessary to make them profitable at present prices. Care in use is also necessary with all lightweight panels when the carpet is removed during maintenance. Panels should be covered, preferably with rubber sheeting, plimsoils should be worn and either canvas toolbags or plastic toolboxes with well rounded edges should be used by mechanics. Also, the cutting of carpet and underlay should be carried out on a plywood sheet and not directly onto the panels. Careful handling is necessary between the manufacturing shop and stores and between stores and fitment to the aircraft. The above remarks are obviously drawn from experience but the amount of damage has been remarkably small. Problems mentioned have been met and are given as warnings so that potential hazards can be avoided. Fortunately repair is quite easy especially if damage is only to one skin. The core can be filled with 'KwikFfll' or the lighter EC 3524 and glass fibre patches can be applied with cold setting epoxy resins or the new 'iron-on' cfrp patches from Bristol Composite Materials Engineering Ltd can be used. One case of damage where two holes about 25.4 mm (1 in) in diameter were punched right through a cfrp panel has been successfully repaired and returned to service. Furthermore the panel was still safe as a floor in the damaged state. To date no problems have been found due. to corrosion.
COMPOSITES . JULY 1974
Although carbon fibre is potentially dangerous, trouble only occurs' when bare carbon is in electrical contact with bare aluminium. In practice most of the carbon is embedded in epoxy resin; facing skins having an insulating coating of glass fibre and all cut edges are painted with black enamel. The theoretical case is unlikely to occur under these conditions and initial fears appear to be unfounded. Service experience with 250 panels has not yet revealed any corrosion. Furthermore the bolt insert used with cfrp panels has been shown to be leakproof.
about 1 tonne of carbon fibre) of which 350 panels made by British Airways have been, or are being, fitted by Boeing to their 14th, 15th, 16th and 17th aircraft prior to delivery. These panels weigh only 26.0 N/m 2 (0.54 lb/ft 2) compared to Balsa 41.0 N/m 2 (0.87 lb/ft2). The Boeing 747 has a floor area of about 280 m 2 (3 000 ft 2) which means a saving of 408 kg (900 lb) if carbon-fibre is used throughout.
CORRELATION BETWEEN FLIGHT TESTING AND LABORA TORY R'ESULTS
Blunt indentor (IOrnrndk3rnctzr)l
Fig.4 shows a relation between static indentation and indentation fatigue or 'knobbly wheel' performance. The static indent test is shown in Fig.5 and the fatigue test in Fig.6. Fig.7 and Fig.8 show the relationships between static indentation results and service life and between indentation fatigue tests and service life. Both indicate that increasing indentation resistance results in increased service life. Aithough the scatter band limits are wide and crude and drawn as straight lines no flight result to date, which has failed by indentation, has fallen outside the predicted lines. As with any fatigue situation one would expect to reach a point, at the higher indentation strength levels, where failure never occurred at all. There is at present insufficient data to locate that point. However, no pane] fully meeting the Specification has so far failed below the minimum estimated line. It is encouraging that floors made from a range of materials to this Specific.don all
seem likely to give the desired result.
.
~ 9 . ~''
For flight experience to date see Tables 2 and 3. CONCL US/ONS "The success of the work is shown by the installation on
Boeing 747 aircraft of over 500 carbon fibre panels (using
Fig.5
Explodedview of the square panel test
Load (Ibf)on CIBA fatigue test for 2h life 50 75
25
I
IOO
I
I
500
c O.
~,2ooo
O
GRP Nornex
X
Corbon-flbre/Nomex
6OO ._c
500 ~ .g C
(~ AI/balso
E .g
-{- 747 Floors: titanium/pro aluminum/pvc
E
E 15OO to
4oo~ E
39'3 N/m2 (O'82 Ibf/ft 2)
300 ~
O
c 0
-,~ I 0 0 0 O o
~
~
J
(~)
2
~m2(O.541bf/ft2 ) cf/Nomex 31.1N/rn2(O'651bf/ff 2) /m2 (O48 Ibf/ft2) 9rp/Nomex
2 0 o '8
_o
U
•£ 5OO ~
IO0 ~u
-Ej 29.7 N/m2(O.62 Ibf/ft2)Ti/pvc 27.5N/m2(O.575 Ibf/ft2)AI/pvc I 30
1 50
8 I 70
I 90
LoadN on CIBA fatigue test for 2h life Fig.4 Comparison of square panel static indentation test core failing loads [9.53 mm (0.375 in) diameter pin] with CIBA indentation fatigue test loads for a life of 2 h
COMPOSITES . JULY 1974
169
Fig.6
Test rig for indentation fatigue tests on aircraft floor panels
i,
Aluminium skins and aluminium honeycomb
400
rn Z
15OC z
300 tO
Aluminium skins and
aluminium honeycomb
/
~
7.
0 O.
~Y~
.\~,"
0
g !
200
g ~ 0 0
. ~ o ~ / ' ~
c ~
500
L~O/"y 0"~// ~//
J
~G-ASGB underseat panel
J
fitted in gangway: failed at f l o o r beam (core compression)
~
._c u 0
7 4 7 original fit: T i / p v c
I
5000
I
I
I0000
20000 Service life h
Fig.7
170
Relation between static indent strength and service life
C O M P O S I T E S . J U L Y 1974
700
160 m Z <
600
40
i
O
tO
t) 20
5OC
Z
< _m u 40O
0
z
~ p
r~
IOO
u
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80 3OC 60 G-ASGB e
O
o~¢~~'¢
,
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y
.,
.
~
~ /4/original t=t.h/pvc
G-ASGB undcrseat panel fitted in gangway failed at floorbeam (core compression)
I
O
o _J
20 1
SOOO
Fig.8
40
I
IOOOO Service life h
I
15000
I
20000
Relation between load for a life of 2 h on the indentation fatigue test and service life
This is equivalent to six passengers. The longest lived carbon fibre panel, fitted in a VC 10, is going strong after 13 000 h and many early trials on the 747 have exceeded 8 000 h. The only panel to fail was fitted in a VC 10 and was to British Airways' underseat grade. It was fitted intentionally in a gangway position to give an accelerated failure. This panel lasted 7 500 h and Suffered a compression failure over a floor beam. Although an early grp/Nomex panel was removed due to resin crazing of the bottom skin at 4 500 h, flight testing of 70 panels of an improved version is being carried out. These have insufficient hours in service to make a judgement and have themselves been overtaken by a further development of grp/Nomex. This latest material is a definite improvement on the earlier versions and test results indicate a good service life. It is also heavier than a comparable carbon fibre panel and this has to be accounted for on a costeffective basis. This means that a grp panel must be cheaper because it is heavier but developments indicate that further weight reductions are probable. The cfrp panels are 14.0 N/m 2 (0.3 lb/ft 2) lighter than an aluminium/balsa panel of similar service life. Although the use of a l 0 mm (0.39 in) diameter indentor remains the one empirical aspect o f the test programme apart from the CIBA indentation fatigue test, both choices have proved correct in service as no false predictions have yet occurred in either case. There is on~ point o f suspicion o f the latter in that cfrp and grp skins on metal cores give low results on test. The author has no flight data to prove whether or not this occurs in service. It could be the case because both are relatively low modulus skins on high modulus cores. BAC Filton have shown that some skins
COMPOSITES . JU LY 1974
Table 2. Flight trial experience to 1 January 1 9 7 4 : a l u m i n i u m honeycomb cored panels w i t h a l u m i n i u m alloy skins
Aircraft registration
Load to failure in static indent test Hours flown
Remarks
(Boeing 707) G-APFD G-ARRB G-ARRC
117 kg (258 Ib) 117 kg (258 Ib) 131 kg (290 Ib)
9 013 6 423 10 146
Continuing Continuing Continuing
(VC 10) G-ASGD G-ASGB G-ASGF
131 kg (290 Ib) 117 kg (258 Ib) 188 kg (415 Ib)
6 162 3 286 2 141
Continuing Continuing Continuing
are incompatible with some cores. It may be that when core and skin fail simultaneously the optimum condition for indentation resistance at minimum weight has been found or at least the best combination of core density and skin thickness. Conversely if the skin carries a lot of load after core failure, before punching through, this could be a measure o f inefficiency from an indentation resistance point of view. The critical points for core and skin can be listed as follows: Skin
1 Deflection up to about half the skin thickness is ~lastic and igrimarily in bending
; 71
Table 3. Flight test experience at 1 January 1974: cfrp/Nomex panels, in flight installations
Airline
Aircraft
Registration
Flying Hours
Landings
Panel locations
British Airways (Overseas Division) (formerly BOAC)
VC 10
G-ASGF
13 400
3 320
Gangway galley area
747
G-AWNA
8 800
2 210
747
G-AWNB
8 690
2 037
No No No No
Trident 1
G-ARPH
4 500
4 650
Lufthansa
747
D-ABYD
Condor
747
D-ABYF
7 630
2 560
No 1 left main entry door No 2 left main entry door
United
747
N-4718U N-4719U
9 321 9 054
2 576 2 368
No 1 left main entry door No 1 left main entry door No 2 left main entry door
Air France
747
-VB
3 991
994
British Airways (European Division) (formerly BEA)
1 left 2 left 1 left 2 left
main main main main
entry entry entry entry
door door door door
Entrance gangway galley area
Panels removed at 7 700 due to accidental inmaintenance damage
Gangway
Airline totals: panels in service - 290; panel hours in use -- 630 000 (Except where stated, all tests are continuing)
2 Beyond half skin thickness membrane effects become significant
160
3 Beyond elastic limit (proof stress) permanent indentation
14o
OCCURS
4 If deformation is continued skin puncture will occur Core
1 Elastic deformation of core 2 Plastic deformation of core or buckling of fibres or cell walls The relative effects will depend on the relative moduli and strengths and could vary from failure of a brittle skin within the elastic range of the core to buckling of the core within the elastic range of the skin. Matching of skin to core is clearly important.
Corbon fibre "LBOACs!~¢: Fibrelom IB JEE-R72-4(A)
---~
Aluminlum/bolso BOACSpe¢:
C O
14
4BA- iO,O00
E
c
16
120
C
--- I00
q,.
"~ 80
8~
~ 6o
6 o
u
U
40
4
20 i
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2
3
4
5
Life, yeors Fig.9
Cost-effective comparison of a number of flooring materials
It would seem that these requirements have been met by cfrp/Nomex and the latest development of grp/Nomex both of which show an improved cost effectiveness over traditional materials even on subsonic aircraft, Fig.9. CFRP has proved itself in service to be as good as predicted. It remains to be seen whether grp will do likewise, but test results suggest that it will when produced to the same Specification.
future. This will place the design of indentation resistant sandwich materials on a firm footing and result in the more efficient use of sandwich panels under these conditions. To do this it may be necessary to try to measure the compression modulus of core materials at or near the surface of a laminate. The resistance of sandwich panels to impact is another useful field for study.
RECOMMENDA ?'IONS FOR FUR THER WORK
ACKNOWLEDGEMENTS
The relationship between the tentative data sheet and testing will, it is hoped, be fully explored by RAE in the near
This paper is published by permission of British Airways although the author wishes to state that where opinions are
172
COMPOSITES . JULY 1974
Square panel static indentation load IIO mmdiolHter pln)lbf I00 2.00 300 400 I
I
I
I
@ EI82 core
i.5
x EI83
E E
8
0-06
9
core
0'05
~: ~-c
.--q
0.045~
.u
10
U
0-03 .E
._C
® G i
l
11
o.o2 X i
O'OI
12
i
500 I000 1500 2000 Square panel static indentation load (IOmmdlometerpin) N
13 14
Fig.10
Effect of core density on i n d e n t a t i o n resistance
15 16
expressed they are his o w n a n d n o t necessarily those o f the British Airways Board. The assistance o f NPL, The Royal Air Force College, Cranwell, Brunel University, The Engineering Sciences Data Unit a n d the material Manufacturers is gratefully acknowledged.
REFERENCES 1 2 3
4 5 6 7
Timoshenko and Woinowski-Kreiger. 'Theory of plates and shells, 2nd edition' (McGraw-HiU 1959) 'Design handbook for honeycomb sandwich structures, TSB 123'. Hexcel Research and Development Dept., Dublin, California, USA (March 1970) Kuenzi, E. W. 'Minimum weight structural sandwich',US Forest Service Research Note FPL-086, Forest Products Laboratory, Madison, Wisconsin, USA (Revised October 1965) 'Design data for the preliminary selection of honeycomb energy absorption systems, TSB 122'. Hexcel Research and Development Dept., Dublin, California, USA (1964) Whittle, K. A. and Armstrong, K. B. BOA C Spec. 4BA-IO, O00, Issue No 5 (1 March 1974) Armstrong, K. B. 'Cost-effective design of indentation resistant sandwich panels', BOAC Report EEA. S. 4. 8644 (December 1969) 'The forces applied to the floor by the foot in walking', National Building Studies Research Paper, No 32, "Part 1, 'Walking on a level floor', (1961); Part 2, 'Walking on a slope', (1967); Part 3, 'Walking on stairs', (1967)
C O M P O S I T E S . J U L Y 1974
Vint, J. and Elgood, W. N. 'The deformation of a bloom plate resting on an elastic base when load is transmitted to the plate by means of a stanchion', Phil Mag S. 7 19 No 124 (January 1935) Murphy, G. 'Stresses and deflections in loaded rectangular plates on elastic foundations', Iowa Engineering Experiment Station Bulletin No 135 (1937) Cox, tt. L. 'Modulus of sandwich core materials under compression', J Royal Aeronautical Soc 75 (February 1971 ) pp 128 Martin, D. W., Picken, Susan M. and Hinde, B. T.
'Axisymmetric local loading of sandwich panels', DNAM 86, National Physical Laboratory (July 1970) Engineering Sciences Data Unit. 'Stresses in the core of a sandwich panel under local normal loading', Data Sheet No 73025 (October 1973) Armstrong, K. B. BOACSpec EE-R72-4(A) (Revised 1 February 1974) Roberts, M. Project report, Honours degree thesis,Brunel University (1971 ) Raimondo, J. V. 'The resistance of sandwich panels to lateral indentation', thesis submitted for the award of B. Sc (Hans), Royal Aircraft College, Cranwell (1971) Harper, F. C. 'The mechanics of walking', Research 15 (January 1962) pp 23-28
BIBLIOGRAPHY Allen, H. G. 'Analysis and design of structural sandwich panels', Pergamon (1969) 'Mechanical properties of Hexcel honeycomb materials', TSB 120, Hexcel Research and Development Dept., Dublin, California, USA (Revised 1971) Design data sheets, CIBA-GEIGY (UK) Ltd., Duxford, Cambridge, England 'Design curves for the preliminary selection of honeycomb structures', TSB 121, Hexcel Research and Development Dept., Dublin, California, USA (1970) Warring, R. H. 'Engineering properties of balsa', reprint from Wood (February 1966), available from Solarbo Ltd., Commerce Way, Lancing, Sussex, England Hoffman, G. A. 'Poisson's ratio for honeycomb cores', The Rand Corporation, California, USA (April 1957) Raville, M. E. 'Deflection and stresses in a uniformly loaded, simply supported rectangular sandwich plate', Forest Products Laboratory, Madison, Wisconsin, USA (December 1955) Owen, M. J. 'How fatigue affects glass and carbon fibre composites', Original Equipment Manufacture and Design (January 1972) Broutman, L. I. and Sahu, S. 'Progressive damage of a glass reinforced plastic during fatigue', 24th SPI Annual Technical Conference, Section 11-D (1969) pp 1-12 American MIL Standard 401/B. 'Sandwich construction and core materials', Department of Defense, Washington DC, USA (December 1967)
173
In 1967 testing started on balsa, pvc and aluminium alloy cores with aluminium alloy skins and a cost-effectiveness formula was developed to provide a basis for comparison. A specification for improved aluminium/balsa floors was produced and flight trials began with aluminium/aluminium honeycomb floors. Carbon fibre and later glass fibre came on the scene and a new specification was raised, based on more fundamental criteria. As a result over 500 cfrp/Nomex panels have been fitted in 747 aircraft and about 70 grp/Nomex panels. These are much lighter than earlier types of flooring and more cost-effective.
Rate of interest on capital
INTRODUCTION
K
The investigation of floor panel performance, with a view to selection of the best available product, began late in 1967. When the investigation started, the problem seemed to be one of resistance to indentation, as scrap panels showed core failures and delamination, usually beyond repair. Corrosion was and still is, a prime cause of failure of aluminium alloy panels fitted in wet areas. It would appear from observation of the inside faces of the skins that the adhesives used are sometimes corrosive in themselves. No data on indentation resistance could be obtained from research establishments and it was decided that testing would have to be carried out to compare available materials. It was appreciated that any information that could be obtained would have a wide applicability to anyone concerned with the indentation resistance of sandwich materials. The designers of aircraft, hovercraft, freight carrying pallets, road vehicles, railway carriages and ships were considered certain to benefit. As the cost of spares was the main reason for this investigation a cost-effectiveness formula was produced to provide the basis for comparison.
x --- Cash value of 1 Newton/annum for the airline or other operation concerned
COST-EFFECTI VENESS FORMULA
The formula suggested is a simple one: ee KL
+ xW=C L
where: p=
Prime cost of finished floor per square metre (ie including inserts, edge finishing and labour)
L=
Service life of panel in years
i
Development Engineer, Structures Group, British Airways Overseas Division, London Heathrow Airport, Hounslow, Middlesex, England
COMPOSITES. JU LY 1974
=
W:
Weight increase or decrease about a datum value
C = Total cost of 1 square metre of flooring per year The formula can be extended to include the manhour cost of removal over panel life/square metre and the cost of repair over panel life/square metre or, if weight is of no concern it can be reduced to ~eKL
L
Clearly for an aircraft like Concorde where the cash value weight is £56/Newton (£250/1b) the more expensive but lighter type of floor is inevitable, but even on subsonic aircraft their use is well justified. A I R C R A F T FLOORS, DESIGN AND MAINTENANCE CONSIDERATIONS
These need to be designed for their particular location ie toilet and galley areas, gangway, underseat and freight areas. Corrosion resistance is important in toilet and galley areas and indentation resistance needs to be greater in gangway, toilet and galley areas than under seats. Whenever corrodible materials are used particular attention is required to sealing at bolt holes and the use of sealed inserts at all bolt positions is recommended. Freight floors are subject to higher average loads and higher point loads and much greater impact loads and therefore need to be considered separately. The floor structure should be so designed that sandwich panels can drop straight in without requiring separate frames and should also be designed to ensure the use of the minimum number of different panels. In most modern aircraft almost every panel is different
165
and this creates a serious spares problem; for example the Boeing 747 has over 350 different panels, the 707 about 130 and the VC 10 has 75. Much greater.standardization could be achieved with considerable economy to manufacturer and operator. Floors need to be sufficiently stiff to avoid passengers feeling unsafe because of their deflection. It is also important, whatever type of floor is chosen, that the time from incipient to final failure should be fairly lengthy to give the operator the chance to change a defective panel at the next convenient maintenance check.
STA TIC DESIGN OF A FLOOR PANEL
This is usually carried out to a uniform pressure requirement although such a loading case only occurs if a freight door or hatch door is lost in flight. Unfortunately a freight door was lost in a recent incident and an adjacent part of the floor failed because it was not designed to resist cabin differential pressure [about 62 kN/m 2 (9 lbf/in2)]. Most passenger floors are designed to 27.6 kN/m2 (4 lbf/in2) and freight floors to 41.4 kN/m 2 (6 lbf/in2). This incident posed the question of providing blow-out panels in the floor to meet such an unlikely need. Data sheets exist for pressure loading but not for the point loading normally experienced. Bending moments under point loading can, however, be calculated if the formulae given on Page 135 of Ref. 1 are used. Optimum design of a panel under pressure loading can be worked out using the Hexcel Design Handbook 2 or Ref. 3. Having obtained the optimum strength panel for a given weight the next step is to ensure that its indentation resistance is adequate. The load and the size of indentor through which it is likely to be applied must be decided at this stage. The most common
O'?l
0.02 I
0.03 1
0.04 I
STA TIC INDENTATION TESTS
The results of a considerable number of tests are given in Fig.1. At the time these tests were done it was normal to use a 19 mm (0.75 in) diameter steel ball under a 1780 N (400 lb) load and to measure the permanent indentiation of the panel that resulted. The panel under test was supported on a steel plate, see Fig.2. This test really measures the compensation strength of the core and is favourable to balsa, which normally fails in shear but has good com-
Permanent Indentation-in 0.05 0'0(~ 0.07 i
I
L73skins
I
0.08
0-09
I
l
0-10 I
171 Balsa: 144,2 kg/m3191b/ft 3) averoge:6-35mm(O.25in)cor¢ O Balsa: 144.2 kg/m3(91b/ft 3) average:6.35 mm(O'25in)cor4 0 " 0 6 [Specification permits balsa to vary between 96 kg/m3 and P76kg/m3(61b/ft 3 to I IIb/ft 3) to give this average] X PVC: 134"6 kg/m3 (8"4 Ib/ft3): 9.5-3mm (0.375 in) core -0.05 + Aluminiu.m honeycomb: 121.7kg/m3(76 Ib/ft3):3.18 mm (O.125 in) celI,O.O25mm ( 0 ' 0 0 2 in) walls, 5052 alloy, 7.94mm (O,313in) core A PVC; 89.71 to IO8.9 kg/m3(5"6to 6'8 Ib/ft3): 7-11mm (0.28 in) core 0 . 0 4 "-~ PVC to BMS 4-13A: 9.53 mm (O.37in) core
Scatter bond : 3 pvc 134.6 kg/m 1'5C
mistake made by floor panel designers is that of using a lightweight core with heavy skins. As in road design, 'hard core' is necessary. Honeycomb cores of various types are superior to others because they are the only ones which produce a non-linear increase in indentation resistance with increasing weight. Fig. 10 indicates that a 12% increase in panel weight can produce a 60% increase in indentation resistance. This occurs because the buckling of the cell wall is a function of the square of the wall thickness. All other types of core have a roughly linear variation of strength witla core density. Honeycombs are also much superior in shear to other core types. For freight floors aluminium alloy honeycombs are superior to Nomex. Although Nomex is initially more resilient, once broken, it is completely useless at the point of damage. Aluminium honeycomb buckles under impact but does not break and retains a considerable amount of strength in the damaged state. A much dented VC 10 freight floor panel has survived 3½ years having been noticeably dented for most of that time. The important features of good skin materials are low weight, high modulus and high proof stress. Some ductility is essential for freight floor applications together with adequate scratch resistance.
{8"4 Ib/ft 3) L73 skins
1"25 E E
== I'O0
C
c ,.x
.u
... " ~
.~ 0.75
Scatter band: 3 pvc 89-71 kg/m
~ " "- ...
e-
0.03
(5"6 Ib/ft 3)
u3
U3
+-.,+. - O'O2
0.50
/El
I Cl~.~X
~'-+'--~X+
""--
A ""A --0.OI
0.25 Honeycomb : 3-18 mm (O.125 In ) cell, O'O25 mm (O.OO2in) walls, 5 0 5 2 alloy I
O
Fig.1
166
I
I
I
I
I
I
I
I
I
0.25 0-50 0-75 FOO 1.25 1-50 1.75 2.00 2.25 2.SO Permanent indentation depth (mm) du¢ to 1780 N (4OOIb) load on Igmm (0"75 in) diameter steel ball
Permanent indentation depth due to a load of 1780 N (400 Ib) with a 19 m m (0.75 in) diameter bell; (test piece fully supported)
COMPOSITES . J U L Y 1974
Clearly this is the maximum possible value. Specially conducted tests using a 203 mm (8 in) thick block of honeycomb with rods at 25.4 mm (1 in) intervals showed that Ec varied from the centre of the block to the surface. At the centre, where end effects were absent a value approaching the theoretical was obtained. Surface effects in balsa and honeycomb may be due to the following:
Loading device Centre to stopslipPoge o f boll
~
~ t _ 19 turn ( 0 " 7 5 in)diameter steel ba~landwic h
[ "
I~panel Steel plate
Fig.2
(a) Salsa (i) Some buckling of end fibres will be caused when cutting the wood;
J
Ball indentation test
pression properties. It is less than fair to honeycombs which are better under shear than compression. The mode of failure is a function of indentor size, core thickness and the ratio of shear properties to compression properties in a given core, see Figs 1 and 3. Fig.1 clearly shows the scatter of properties in both balsa and pvc. Although pvc is consistent in any one slice, it is foamed as 'loaff and different slices give different results due to density variation. The result of these early tests was BOAC Spec. 4BA-10,000. 5 This called for closer selection o f the balsa, to reduce scatter due to density variation and the use of L. 88 for the top skin, as tests with L. 73 had shown a clear improvement over L. 72 and this was assumed to be due to the higher proof stress of L. 73. Because the tests 6 later included edge supported tests which are more likely to suffer shear failure it was found that metal honeycombs were likely to be better than balsa. However, pending flight trials with metal honeycomb cores which have since proved successful, it was decided to stick with the aluminium/balsa already known to be good and likely to be better when supplied to the new specifications.
(ii)
Some deformation will occur in the adhesive layer;
(iii)
Under ball loading conditions the ends of the fibres will be rotated out of the vertical as deformation takes place and this will cause buckling at lower loads.
(b) Honeycomb (i) Some movement out of the vertical may occur at the edges during cutting; (ii)
Deformation of the adhesive layer;
(iii)
Ball loading will rotate cell walls out of the vertical and cause buckling at lower loads.
In both cases, (i) and (iii) above could not apply to pvc and its modulus is also less than that of the adhesive layer. Floor panels are always relatively thin and it is by no means certain that the theoretical value is the one to use for balsa or honeycomb.
0.25 i
0"50 i
50
THEORETICAL APPROACH TO INDENTATION RESISTANCE Because of the cost and quantity of testing required to provide the information for floor panel selection two theoretical methods were tried. 8,9 The first was later fbund to be in error, by the National Physical Laboratory, but consideration of it brought up points worth mentioning:
Core thickness- in 0.75 FO0 1.2 F50 i i i r C) 20'3 mm{ 0 8 in) diameter riot faced Indentor ® AvT.rogF_~tqct diamete.r~r : ~.=mmtu./~ inldiometer Dart at IOmm (0.39in)diameter riotriot indentor
1"75
4O ~50
| i
E
=~
1.25 ~
o 30
f
1 Membrane effects occur when indentation exceeds half skin thickness in depth; 2 Compression modulus of the core material is a governing factor in both formulae and accurate determination of this property can be difficult. The modulus for balsa is best obtained by wave velocity tests. For pvc the value can be fairly accurately obtained from compression tests. For other stiffer materials platen deformation and flatness can be so critical that only when allowance is made for these errors after the most accurate and careful testing can true values be obtainedA o For aluminium honeycomb the compression modulus is usually obtained as
2-00
1-00 ~ 2C
0"75 Indentor diameter Core thickness
I0
C>50 0.25
® 0
4 fshtmr fcom~sioa
20 Core thickr~ss-mm
30
Young's modulus x density of honeycomb density of aluminium
COMPOSITES . JULY 1974
Fig.3 Comparison o f core materials showing lines o f simultaneous failure in shear and compression
167
The work of Cox and Martin i I has been compared to test results by Raimondol. s who found that although the general trend of his results was in line with.the formula, calculated results were below those obtained by test. Further work is being carried out by Payen of RAE Farnborough. SPECIFICATION FOR LIGHTWEIGHTFLOORS
Soon after the new balsa specification came off the press British Airways were asked if any application for carbonfibre could be envisaged. The material seemed to have most of the properties required for a floor panel skin. except ductility and scratch resistance. For these reasons it has been used for passenger floors covered with carpet rather than for freight floors. The improvement aimed at by the 4BA-10,000 Spec s was really just an improvement on an existing material but was not based on any fundamental data. To put the work on a firmer basis it was decided to approach the Building Research Station and they provided details of loads applied by the foot in walking 7. Cox of the National Physical Laboratory provided a mathematical basis for calculating indentation resistance I l and this has since been published in Data Sheet form by the Engineering Sciences Data Unit 12. Using the Building Research data the indentation requirements for BOAC Specification E E - R 7 2 - 4 ( A ) 13 were formulated, suitable factors being included for fatigue and overload cases. It was also encouraging to find that the minimum indentation values of good quality aluminium/balsa panels, of known long life, came out to about the same as the values obtained by applying a factor of 1.5 to the walking loads supplied by the Building Research Station and confirmed by Brunel University 14 using a strain-gauged floor panel. A 10 mm (0.39 in) diameter blunt indenter was used for the tests. The diameter chosen was representative of heels at the time and was the only assumption made in producing the Specification. It was considered a reasonable choice because experience on the Britannia, which had separate ladies' and gentlemen's toilets, showed that although damage to the panel outside the 'Ladies' was more than the 'Gents', damage in the latter case was still significant. This is not surprising because a new heel on a male shoe is quite sharp and can have only line contact on impact with the floor. This view is confirmed by Harper 16. Indentation fatigue requirements were specified considering the performance of good existing materials. The CIBA Indentation fatigue test is empirical and comparative only and is sensitive to panel stiffness. Comparisons should be made only between panels of equal thickness
Table 1. CIBA Fatigue test results on balsa core/aluminium skin panels: the load was the same for each case Core thickness mm (in)
6.35 (0.25) 9.53 (0.375) 12.7
(0.5)
* w i t h cracking of the t o p skin
168
Time to failure h
7 2 ½"
intended for the same application. The results in Table 1 are from a series of CIBA Fatigue tests, and indicate the importance of the core thickness. To overcome the thickness problem and the difficulty presented by the fact that the indentation fatigue test is partly an impact test, a test on one spot was specified for 1 000 cycles at 70% of the specified static load, residual indentation strength to be measured afterwards. The figure of 1 000 cycles was obtained from a Douglas report which calculated this to be the likely number of cycles on one spot in a 5 year period. Binding strength had to meet design requirements. Design requirements also covered the shear load at bolt inserts." It is worth noting that the performance of inserts on test depends considerably on the pulling straps or test frame used. The method used by the aircraft manufacturer should be adopted in each case. To cater for loading both at and between floor beams two indentation tests were specified, one fully supported and one edge supported. It is necessary for the fully supported requirement to be higher than the edge supported case because no cushioning of the load due to panel deflection occurs at a floor beam. The food-cart wheel test was included because calculations at Brunel University 14 had shown this to be a severe case with the small hard wheels often used. The Specification has remained unaltered except that a deflection test has been added together with a creep test because of unfortunate experience with grp/Nomex panels. With these panels it is possible to reach strains sufficient to cause resin cracking unless the panel stiffness can be maintained above a critical level. Great effort has been put into solving this problem and the early difficulties have been overcome by one manufacturer. Any other design of grp would need to be carefully checked to ensure that an adequate life was likely to be obtained. Because of the higher cost of lightweight panels a life of 15 000 to 20 000 h is necessary to make them profitable at present prices. Care in use is also necessary with all lightweight panels when the carpet is removed during maintenance. Panels should be covered, preferably with rubber sheeting, plimsoils should be worn and either canvas toolbags or plastic toolboxes with well rounded edges should be used by mechanics. Also, the cutting of carpet and underlay should be carried out on a plywood sheet and not directly onto the panels. Careful handling is necessary between the manufacturing shop and stores and between stores and fitment to the aircraft. The above remarks are obviously drawn from experience but the amount of damage has been remarkably small. Problems mentioned have been met and are given as warnings so that potential hazards can be avoided. Fortunately repair is quite easy especially if damage is only to one skin. The core can be filled with 'KwikFfll' or the lighter EC 3524 and glass fibre patches can be applied with cold setting epoxy resins or the new 'iron-on' cfrp patches from Bristol Composite Materials Engineering Ltd can be used. One case of damage where two holes about 25.4 mm (1 in) in diameter were punched right through a cfrp panel has been successfully repaired and returned to service. Furthermore the panel was still safe as a floor in the damaged state. To date no problems have been found due. to corrosion.
COMPOSITES . JULY 1974
Although carbon fibre is potentially dangerous, trouble only occurs' when bare carbon is in electrical contact with bare aluminium. In practice most of the carbon is embedded in epoxy resin; facing skins having an insulating coating of glass fibre and all cut edges are painted with black enamel. The theoretical case is unlikely to occur under these conditions and initial fears appear to be unfounded. Service experience with 250 panels has not yet revealed any corrosion. Furthermore the bolt insert used with cfrp panels has been shown to be leakproof.
about 1 tonne of carbon fibre) of which 350 panels made by British Airways have been, or are being, fitted by Boeing to their 14th, 15th, 16th and 17th aircraft prior to delivery. These panels weigh only 26.0 N/m 2 (0.54 lb/ft 2) compared to Balsa 41.0 N/m 2 (0.87 lb/ft2). The Boeing 747 has a floor area of about 280 m 2 (3 000 ft 2) which means a saving of 408 kg (900 lb) if carbon-fibre is used throughout.
CORRELATION BETWEEN FLIGHT TESTING AND LABORA TORY R'ESULTS
Blunt indentor (IOrnrndk3rnctzr)l
Fig.4 shows a relation between static indentation and indentation fatigue or 'knobbly wheel' performance. The static indent test is shown in Fig.5 and the fatigue test in Fig.6. Fig.7 and Fig.8 show the relationships between static indentation results and service life and between indentation fatigue tests and service life. Both indicate that increasing indentation resistance results in increased service life. Aithough the scatter band limits are wide and crude and drawn as straight lines no flight result to date, which has failed by indentation, has fallen outside the predicted lines. As with any fatigue situation one would expect to reach a point, at the higher indentation strength levels, where failure never occurred at all. There is at present insufficient data to locate that point. However, no pane] fully meeting the Specification has so far failed below the minimum estimated line. It is encouraging that floors made from a range of materials to this Specific.don all
seem likely to give the desired result.
.
~ 9 . ~''
For flight experience to date see Tables 2 and 3. CONCL US/ONS "The success of the work is shown by the installation on
Boeing 747 aircraft of over 500 carbon fibre panels (using
Fig.5
Explodedview of the square panel test
Load (Ibf)on CIBA fatigue test for 2h life 50 75
25
I
IOO
I
I
500
c O.
~,2ooo
O
GRP Nornex
X
Corbon-flbre/Nomex
6OO ._c
500 ~ .g C
(~ AI/balso
E .g
-{- 747 Floors: titanium/pro aluminum/pvc
E
E 15OO to
4oo~ E
39'3 N/m2 (O'82 Ibf/ft 2)
300 ~
O
c 0
-,~ I 0 0 0 O o
~
~
J
(~)
2
~m2(O.541bf/ft2 ) cf/Nomex 31.1N/rn2(O'651bf/ff 2) /m2 (O48 Ibf/ft2) 9rp/Nomex
2 0 o '8
_o
U
•£ 5OO ~
IO0 ~u
-Ej 29.7 N/m2(O.62 Ibf/ft2)Ti/pvc 27.5N/m2(O.575 Ibf/ft2)AI/pvc I 30
1 50
8 I 70
I 90
LoadN on CIBA fatigue test for 2h life Fig.4 Comparison of square panel static indentation test core failing loads [9.53 mm (0.375 in) diameter pin] with CIBA indentation fatigue test loads for a life of 2 h
COMPOSITES . JULY 1974
169
Fig.6
Test rig for indentation fatigue tests on aircraft floor panels
i,
Aluminium skins and aluminium honeycomb
400
rn Z
15OC z
300 tO
Aluminium skins and
aluminium honeycomb
/
~
7.
0 O.
~Y~
.\~,"
0
g !
200
g ~ 0 0
. ~ o ~ / ' ~
c ~
500
L~O/"y 0"~// ~//
J
~G-ASGB underseat panel
J
fitted in gangway: failed at f l o o r beam (core compression)
~
._c u 0
7 4 7 original fit: T i / p v c
I
5000
I
I
I0000
20000 Service life h
Fig.7
170
Relation between static indent strength and service life
C O M P O S I T E S . J U L Y 1974
700
160 m Z <
600
40
i
O
tO
t) 20
5OC
Z
< _m u 40O
0
z
~ p
r~
IOO
u
•
tO
80 3OC 60 G-ASGB e
O
o~¢~~'¢
,
O
'
20C
°oI
o .J
y
.,
.
~
~ /4/original t=t.h/pvc
G-ASGB undcrseat panel fitted in gangway failed at floorbeam (core compression)
I
O
o _J
20 1
SOOO
Fig.8
40
I
IOOOO Service life h
I
15000
I
20000
Relation between load for a life of 2 h on the indentation fatigue test and service life
This is equivalent to six passengers. The longest lived carbon fibre panel, fitted in a VC 10, is going strong after 13 000 h and many early trials on the 747 have exceeded 8 000 h. The only panel to fail was fitted in a VC 10 and was to British Airways' underseat grade. It was fitted intentionally in a gangway position to give an accelerated failure. This panel lasted 7 500 h and Suffered a compression failure over a floor beam. Although an early grp/Nomex panel was removed due to resin crazing of the bottom skin at 4 500 h, flight testing of 70 panels of an improved version is being carried out. These have insufficient hours in service to make a judgement and have themselves been overtaken by a further development of grp/Nomex. This latest material is a definite improvement on the earlier versions and test results indicate a good service life. It is also heavier than a comparable carbon fibre panel and this has to be accounted for on a costeffective basis. This means that a grp panel must be cheaper because it is heavier but developments indicate that further weight reductions are probable. The cfrp panels are 14.0 N/m 2 (0.3 lb/ft 2) lighter than an aluminium/balsa panel of similar service life. Although the use of a l 0 mm (0.39 in) diameter indentor remains the one empirical aspect o f the test programme apart from the CIBA indentation fatigue test, both choices have proved correct in service as no false predictions have yet occurred in either case. There is on~ point o f suspicion o f the latter in that cfrp and grp skins on metal cores give low results on test. The author has no flight data to prove whether or not this occurs in service. It could be the case because both are relatively low modulus skins on high modulus cores. BAC Filton have shown that some skins
COMPOSITES . JU LY 1974
Table 2. Flight trial experience to 1 January 1 9 7 4 : a l u m i n i u m honeycomb cored panels w i t h a l u m i n i u m alloy skins
Aircraft registration
Load to failure in static indent test Hours flown
Remarks
(Boeing 707) G-APFD G-ARRB G-ARRC
117 kg (258 Ib) 117 kg (258 Ib) 131 kg (290 Ib)
9 013 6 423 10 146
Continuing Continuing Continuing
(VC 10) G-ASGD G-ASGB G-ASGF
131 kg (290 Ib) 117 kg (258 Ib) 188 kg (415 Ib)
6 162 3 286 2 141
Continuing Continuing Continuing
are incompatible with some cores. It may be that when core and skin fail simultaneously the optimum condition for indentation resistance at minimum weight has been found or at least the best combination of core density and skin thickness. Conversely if the skin carries a lot of load after core failure, before punching through, this could be a measure o f inefficiency from an indentation resistance point of view. The critical points for core and skin can be listed as follows: Skin
1 Deflection up to about half the skin thickness is ~lastic and igrimarily in bending
; 71
Table 3. Flight test experience at 1 January 1974: cfrp/Nomex panels, in flight installations
Airline
Aircraft
Registration
Flying Hours
Landings
Panel locations
British Airways (Overseas Division) (formerly BOAC)
VC 10
G-ASGF
13 400
3 320
Gangway galley area
747
G-AWNA
8 800
2 210
747
G-AWNB
8 690
2 037
No No No No
Trident 1
G-ARPH
4 500
4 650
Lufthansa
747
D-ABYD
Condor
747
D-ABYF
7 630
2 560
No 1 left main entry door No 2 left main entry door
United
747
N-4718U N-4719U
9 321 9 054
2 576 2 368
No 1 left main entry door No 1 left main entry door No 2 left main entry door
Air France
747
-VB
3 991
994
British Airways (European Division) (formerly BEA)
1 left 2 left 1 left 2 left
main main main main
entry entry entry entry
door door door door
Entrance gangway galley area
Panels removed at 7 700 due to accidental inmaintenance damage
Gangway
Airline totals: panels in service - 290; panel hours in use -- 630 000 (Except where stated, all tests are continuing)
2 Beyond half skin thickness membrane effects become significant
160
3 Beyond elastic limit (proof stress) permanent indentation
14o
OCCURS
4 If deformation is continued skin puncture will occur Core
1 Elastic deformation of core 2 Plastic deformation of core or buckling of fibres or cell walls The relative effects will depend on the relative moduli and strengths and could vary from failure of a brittle skin within the elastic range of the core to buckling of the core within the elastic range of the skin. Matching of skin to core is clearly important.
Corbon fibre "LBOACs!~¢: Fibrelom IB JEE-R72-4(A)
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Cost-effective comparison of a number of flooring materials
It would seem that these requirements have been met by cfrp/Nomex and the latest development of grp/Nomex both of which show an improved cost effectiveness over traditional materials even on subsonic aircraft, Fig.9. CFRP has proved itself in service to be as good as predicted. It remains to be seen whether grp will do likewise, but test results suggest that it will when produced to the same Specification.
future. This will place the design of indentation resistant sandwich materials on a firm footing and result in the more efficient use of sandwich panels under these conditions. To do this it may be necessary to try to measure the compression modulus of core materials at or near the surface of a laminate. The resistance of sandwich panels to impact is another useful field for study.
RECOMMENDA ?'IONS FOR FUR THER WORK
ACKNOWLEDGEMENTS
The relationship between the tentative data sheet and testing will, it is hoped, be fully explored by RAE in the near
This paper is published by permission of British Airways although the author wishes to state that where opinions are
172
COMPOSITES . JULY 1974
Square panel static indentation load IIO mmdiolHter pln)lbf I00 2.00 300 400 I
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I
I
@ EI82 core
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13 14
Fig.10
Effect of core density on i n d e n t a t i o n resistance
15 16
expressed they are his o w n a n d n o t necessarily those o f the British Airways Board. The assistance o f NPL, The Royal Air Force College, Cranwell, Brunel University, The Engineering Sciences Data Unit a n d the material Manufacturers is gratefully acknowledged.
REFERENCES 1 2 3
4 5 6 7
Timoshenko and Woinowski-Kreiger. 'Theory of plates and shells, 2nd edition' (McGraw-HiU 1959) 'Design handbook for honeycomb sandwich structures, TSB 123'. Hexcel Research and Development Dept., Dublin, California, USA (March 1970) Kuenzi, E. W. 'Minimum weight structural sandwich',US Forest Service Research Note FPL-086, Forest Products Laboratory, Madison, Wisconsin, USA (Revised October 1965) 'Design data for the preliminary selection of honeycomb energy absorption systems, TSB 122'. Hexcel Research and Development Dept., Dublin, California, USA (1964) Whittle, K. A. and Armstrong, K. B. BOA C Spec. 4BA-IO, O00, Issue No 5 (1 March 1974) Armstrong, K. B. 'Cost-effective design of indentation resistant sandwich panels', BOAC Report EEA. S. 4. 8644 (December 1969) 'The forces applied to the floor by the foot in walking', National Building Studies Research Paper, No 32, "Part 1, 'Walking on a level floor', (1961); Part 2, 'Walking on a slope', (1967); Part 3, 'Walking on stairs', (1967)
C O M P O S I T E S . J U L Y 1974
Vint, J. and Elgood, W. N. 'The deformation of a bloom plate resting on an elastic base when load is transmitted to the plate by means of a stanchion', Phil Mag S. 7 19 No 124 (January 1935) Murphy, G. 'Stresses and deflections in loaded rectangular plates on elastic foundations', Iowa Engineering Experiment Station Bulletin No 135 (1937) Cox, tt. L. 'Modulus of sandwich core materials under compression', J Royal Aeronautical Soc 75 (February 1971 ) pp 128 Martin, D. W., Picken, Susan M. and Hinde, B. T.
'Axisymmetric local loading of sandwich panels', DNAM 86, National Physical Laboratory (July 1970) Engineering Sciences Data Unit. 'Stresses in the core of a sandwich panel under local normal loading', Data Sheet No 73025 (October 1973) Armstrong, K. B. BOACSpec EE-R72-4(A) (Revised 1 February 1974) Roberts, M. Project report, Honours degree thesis,Brunel University (1971 ) Raimondo, J. V. 'The resistance of sandwich panels to lateral indentation', thesis submitted for the award of B. Sc (Hans), Royal Aircraft College, Cranwell (1971) Harper, F. C. 'The mechanics of walking', Research 15 (January 1962) pp 23-28
BIBLIOGRAPHY Allen, H. G. 'Analysis and design of structural sandwich panels', Pergamon (1969) 'Mechanical properties of Hexcel honeycomb materials', TSB 120, Hexcel Research and Development Dept., Dublin, California, USA (Revised 1971) Design data sheets, CIBA-GEIGY (UK) Ltd., Duxford, Cambridge, England 'Design curves for the preliminary selection of honeycomb structures', TSB 121, Hexcel Research and Development Dept., Dublin, California, USA (1970) Warring, R. H. 'Engineering properties of balsa', reprint from Wood (February 1966), available from Solarbo Ltd., Commerce Way, Lancing, Sussex, England Hoffman, G. A. 'Poisson's ratio for honeycomb cores', The Rand Corporation, California, USA (April 1957) Raville, M. E. 'Deflection and stresses in a uniformly loaded, simply supported rectangular sandwich plate', Forest Products Laboratory, Madison, Wisconsin, USA (December 1955) Owen, M. J. 'How fatigue affects glass and carbon fibre composites', Original Equipment Manufacture and Design (January 1972) Broutman, L. I. and Sahu, S. 'Progressive damage of a glass reinforced plastic during fatigue', 24th SPI Annual Technical Conference, Section 11-D (1969) pp 1-12 American MIL Standard 401/B. 'Sandwich construction and core materials', Department of Defense, Washington DC, USA (December 1967)
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