A brazed titanium honeycomb sandwich system for supersonic transport wing cover panels provides the most efficient structure spanwise, chordwise, and loadwise. Flutter testing shows that high wing stiffness is most efficient in a sandwich structure. This structure also provides good thermal insulation if liquid fuel is carried in direct contact with the wing structure in integral fuel tanks

Davis, R. A.

Elrod, S. D.

Lovell, D. T.

English

NASA Technical Reports Server

TITANIUM HONEYCOKE STRUCTURE 
By R. A. Davis, 5 .  D. Elrod, a d  D. T. Love11 
INTKODUCITON AND EACRGROL'ND 
During i n i t i a l  engineering design s tudies  f o r  a commercial supersonic 
t ransport ,  titanium w a s  shown t o  provide the lowest weight structure to 
carry a given peyload. A summary of t h i s  da ta  compared t o  aluminum and 
stainless steel  i s  shown i n  Figure 1. The lowest s t r u c t u r a l  weight has 
a favorable i m p a c t  on the operating economics of a commerciai transport  
airplane.  
pressure loads, P lu t te r ,  and range / payload developments, a mach. 2.7 
fixed wing titanium airplane evolved. 
As design refinements occurred based on aerodynamic. propulsion, 
The major impac t  of t h i s  f i n a l  configuration on s t ruc tures  and materials 
technology w a s  t o  require development of a titanium honeycomb sandwich 
system f o r  ving cover pane i s .  There were basical ly  three technicai reasons 
f o r  a saiidwich structure.  F i r s t ,  spanwise and chordwise and loadinns 
<ziguLe 2) were generaiiy low ana sandwich material provided the  most 
e f f i c i e n t  s t ructure .  Second, f l u t t e r  t e s t ing  showed tha t  high wing 
s t i f f n e s s  w a s  required and  a s s i n  sandwi-ch stx-vctire w a s  most e f f i c l en r .  
Third, l-lquid f u e l  w a s  carr ied i n  d i r ec t  contact. with the wing struccure 
i n  i n t e g r a l  f u e l  tanks and sandwich s t ruc ture  provided thermal insu la t ion  
f o r  the fuel.  
Figure 3 shows the  temperature p r o f i l e  f o r  the prototype supersonic transport .  
Generally the bas ic  s t ruc tu re  would be o?!erating a t  450 F with some loca l  
areas reaching peaks of 500 k' under spec ia l  f l i g h t  conditions. For the 
bas ic  wing sandwich material tne nasimum operating temperature would be 
450°F. 
of the  prototype a i r c r a f t .  
expanded t o  the power p lan t  pod and eai2ennage s t ruc ture  f o r  imprmcd s t r u c t u r a l  
efficiency. Additionally, res is tance to re la t ive ly  high scn ic  leve ls  (166 d 3 )  
was required i n  porcions of the empennage s t ruc ture  and honeycomb sandwich 
W ~ L  L i i i s  requireiarnc.  iiLanium rioneycono sariawicn w a s  not plannea t o r  the 
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Figure 4 shows the major s t r u c t u r a l  concepts planned €or construction 
Titanium honeycomb sandwich applications were also 
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most highly loaded center sec t ion  of the wing nor the  r e l a t ive ly  deep 
(up t o  11 inches) wedge s t ruc tu re  on both leading and t r a i l i n g  edges of 
wing and t a i l  s t ruc tu re  of the prototype airplane because process develop- 
ment could not  be carr ied out i n  t i m e  t o  meet the planned prototype 
manuf actbring schedule. 
PROCESS DEVELOPMENT 
The cr i ter ia  of concern i n  development of a brazed t i tanium system involved 
three major areas; (1) metal lurgical  compatibility, (2) manuf aeturing 
f e a s i b i l i t y ,  and (3) design v i ab i l i t y .  Process development, although 
concerned primarily with metal lurgical  parameters, had t o  consider con- 
current ly  both manufacturing and design aspects. 
The i n i t i a l  s t ep  w a s  t o  assess and se lec t ive ly  test brazing al loy systems 
f o r  compatibil i ty with titanium, f o r  processing parameters, and f o r  
preliminary s t rength properties.  
brazing a l loy  systems and t h e i r  per t inent  charac te r i s t ics .  As a r e su l t  
Table 1 shows a summary of the various 
of t h i c  a q q e s f i m p n t ,  tha  a l i r m i n i i m  I.lc0 5 l l o y c  y g y g  ~ s ~ = ~ ~ = ~  f o r  fcr:5:=r 
evaluation and subsequently aluminum al loy 3003 w a s  chosen as the bes t  a l lay.  
This choice w a s  based on brazing tempera ture  range, f o i l  ava i l ab i l i t y ,  
f la twise  t e n s i l e  s t rength and corrosion resistance.  
! 
A primary concern w a s  the formation of any embri t t l ing e f f e c t s  as a r e s u l t  
of the  formation of titanium aluminide (TiA1 ) during the brazing cycle. The 
brazing cycle temperature envelope w a s  developed t o  l i m i t  t itanium aluminide 
formation by r e s t r i c t i n g  the holding t i m e  above 1175'F t o  one hour maximum. 
With t h i s  control,  an aluminide layer of .0003 inches maximum is  formed. This 
t h i n  layer  has no noticeable e f f e c t  on s ta t ic  propert ies  of the  bas ic  titanium 
6Al-4V face sheet  alloy. 
20%, however, due t o  d i f f i c u l t i e s  i n  specimen preparation, the avai lable  data  
3 
Fatigue propert ies  may be reduced by approximately 
are inconclusive. Figure 5 shows the f i n a l  high temperature portion of the 
braze cycle envelope used f o r  process control,  
-2- 
Another major concern w a s  galvanic coupling e f f e c t s  resu l t ing  from exposing 
aluminum and ti tanium t o  a corrosive environment. From a theoretical 
standpoint based on s ing le  e lectrode poten t ia l s  of the two al loys,  aluminum 
should preferen t ia l ly  corrode when i n  contact with titanium. Extensive 
accelerated laboratory t e s t ing  and commercial a i r l ine f l e e t  exposure showed 
tha t  no galvanic accelerated corrosion "effects  occurred. 
an unprotected honeycomb sandwich panel i n s t a l l e d  on the mud f l a p  of a Boeing 
727 model a i rplane f o r  se rv ice  evaluation. 
a i r l i n e  service exposure have been car r ied  out with no corrosive a t tack  
occurring. 
Figure 6 shows 
Approximately three years of 
Although a l l  corrosion t e s t i n g  indicated tha t  galvanic acceleration would 
not  be a problem, a conservative approach w a s  taken by s t ipu la t ing  tha t  oon- 
perforated core would be used exclusively. 
cel l  cavity would he hermetically sealed and any progression cf moisture 
through a panel would have t o  progress a c e i l  a t  a t i m e .  
With t h i s  approach each individual 
In 1969 at  the  t i m e  of i n i t i a t i o n  of development of brazed titanium honeycomb 
sandwich, the state-of-the-art s t ruc tu ra l  sandwich sys tem which exis ted w a s  
Oilve~ ' v ra~ru  2ii L J - ~ L W  sca in ie s s  sceei .  Tnis sysrem had been used extensively 
on the B70 supersonic bomber. A comparison of the  s t a in l e s s  s teel  and t i tanium 
honeycomb sandwich processes is shown i n  Table 2. Several items erp, nore- 
worthy. F i r s t  the brazing temperature of the titanium system is  much lower 
and t h i s  f a c t  s impl i f ies  both tooling and heat source requirements. Secondly, 
the titanium system requires  no post brazlng thermal cycle. Finally,  a 
generous amount of aluminum braze al loy i s  used during the process t o  produce 
a .030" f i l l e t  a t  the core t o  face sheet  junction. 
thick one-quarter inch cel l  s i z e  and optimuui braze conditions, t h i s  amount 
of braze a l loy  provides enough s t rength t o  break the 3Al-2.5V core f o i l  during 
a flatwise t e n s i l e  test. 
- -  -.- 
With the standard .002" 
MECHANICAL PROPERTIES 
As process development progressed, f irm requirements were developed f o r  bas i c  
honeycomb panel strength.  The standard honeycomb core consisted of .002" 
tliick one-quarter inch cel l  produced from 3A1-2.5V f o i l  (density = 5 pounds 
per cubic foot) .  
mechanical propert ies  were developed f o r  the  bas ic  honeycomb panels from 
tests of f la twise  tension, flatwise compression, p l a t e  shear,  beam shear, and 
edgewise compression specimens, 
property data. 
C e l l  w a l l s  were corrugated f o r  added s t i f fness .  S t a t i c  
Figure 7 contains a summary of the mechanical 
In  addition, creep behavior w a s  of concern and a simple s ing le  cell  tubular 
test specimen w a s  developed fo r  evaluation. 
mechanical loading fo r  da ta  gathering. The data  were later confirmed using 
standard 2 inch square  honeycomb sandwich specimens. Figure 8 contains a 
summary of both s ingle  c e l l  and multiple c e l l  test specimens fo r  both 45OoF 
and 600°F. 
No perceptible creep occurred with any of the tes t  specimens. 
Tf i i s  specimen permitted easy 
A l l  test f ractures  tha t  occurred were of a stress rupture- nature. 
DESIGN DEVELOPMENT 
Basic panel s t rength requirements could be m e t  with f i v e  Foiind ?- ,-A+- 
f t i 2 - i ~ ~ ~ .  
used t o  fas ten  panels t o  spars  and r ibs ,  higher density core was used. 
high density core provided both increased resistance t o  e m i r s m e n t a l  effccis 
a t  panel edges as w e l l  as cap.ability t o  withstand fastener  loads. Table 3 
shows an approximate distribut%on of the  various types of core planned f o r  
use on the supersonic transport  prototype. 
density core are schematically shown i n  Figure 9. 
A i  load rransfer areas however, where mechanical fasteners were 
This 
Edge designs u t i l i z i n g  the higher 
One of the major developments involved u t i l i z a t i o n  of appropriate non- 
destruct ive t e s t ing  techniques t o  ascer ta in  tha t  brazed panel qual i ty  w a s  
acceptable. 
determine panel quali ty.  
100% t o  each panel t o  check f o r  the proper d is t r ibu t ion  of the aluminum 
braze alloy within the panel. Radiographic inspection w a s  used select ively 
cy &.teriiArlr LIU& L L U W ~  shear t i e  in t eg r i ty  and the extent of core crushing 
Figure 10 shows schematically the techniques u t i l i z e d  t o  
The eddy current scan technique would be applied 
c. . c- 
-4- 
. I  
i f  any. 
determine braze al loy f i l l e t  s i z e  and the location of any sheet-to-core- 
The Ultrasonic C-Scan technique w a s  used 1002 on each panel t o  
voids. 
f i l l e t  s ize .  
Figure 11 shows an ul t rasonic  trace representing three levels of 
1 
SCALE-UP 
Although many f i n e  ideas can be s a t i s f a c t o r i l y  demonstrated i n  the laboratory, 
the f i n a l  proof of acceptabi l i ty  f o r  aerospace s t r u c t u r a l  usage consists of 
f u l l  scale hardware fabr ica t ion  and tes t ing.  
of the prototype program, a major e f f o r t  w a s  devoted t o  building three foot 
by twenty foot panels representative of both wing and empennage designs. 
The philosophy behind t h i s  approach is  tha t  there are generally technical  
problems encountered with fabricat ion of f u l l  sca le  hardware that a re  not  
revealed i n  small .scale  fabrication. This f a c t  w a s  a l so  t rue  with t h i s  
Concurrent w i t h  the design phase 
alurninum brazed system. Considerable e f f o r t  w a s  expended i n  tooling 
development, temperature control,  r e t o r t  purging, panel res t ra int ,  core 
machining and non-destructive tes t ing .  
The s implif ied tooling concept used f o r  face sheet forming and panel brazing 
is  shown i n  Figure 12. This concept w a s  devised t o  minimize production cos ts  
of the process. 
s i d e  w a l l  r e ta iners  f o r  the braze f ix ture .  Figure 13 shows a successfully 
completed 20 foot wing panel. 
I n  f a c t  masonite s i d e  walls were found t o  be acceptable as 
A t  the t i m e  of the SST cancellation, the development and evaluation of the 
system w a s  e s sen t i a l ly  complete €or  the planned applications.  Design 
cr i ter ia  and propert ies  were established. 
shown tha t  ne i the r  corrosion nor creep rupture would be a problem f o r  the 
proposed applications.  
acceptmce cri teria and inspection methods were established as demonstrated 
by the successful fabr ica t ion  of 3' x 20' production wing panel. A t  the 
present tine, fu r the r  development work sponsored by the Department of Trans- 
portat ion is  underway t o  extend the process f o r  broader applications such 
as more highly loaded titructure, wedge configurations, and acoustic panels. 
Subsequent extensive t e s t ing  has 
Process and material specif icat ions and qua l i ty  
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Titanium honeycomb structure

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