The application of advanced composite materials to high performance structure frequently results in the desire to fabricate a structure from more than one composite system in order to tailor the composite material capabilities to the design requirements. The bi-composite transition provides a means of joining two different composite structural systems without the weight and complexity of mechanical attachments. The monolayer plies or combinations of plies of one composite system are interleaved with and bonded to the plies of the adjacent composite system, thereby providing a direct load transfer between the two composite structures

Dullea, K. C., Jr.

English

NASA Technical Reports Server

THE B I - C O M P O S I T E  T R A N S I T I O N  JOINT 
S D  7 2 - S H - 0 0 5 1  
K. C. DULLEA, JR. 
MAY 26, 1972 
Space Division 
North Arnencan Rockwell 
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https://ntrs.nasa.gov/search.jsp?R=19740022826 2020-03-17T15:13:39+00:00Z
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THE BI-COMPOSITE TRANS ITION JOINT 
By K. C. Dullea, Jr. 
I 
The application of advanced composite materials 
t o  h i g h  performance structure frequently results i n  
the desire to  fabricate a structure from more than 
one composite system i n  order t o  t a i lo r  the composite 
materi a1 capabi 1 i t i es  to the design requirements. 
The bi-composite transition provides a means of 
joining two different composite structural systems 
without the weight  and complexity of mechanical 
attachments. The monolayer pl  ies or combinations 
of plies of one composite system are interleaved w i t h  
and bonded to  the plies of the adjacent composite 
system, thereby providing a direct load transfer 
between the two composite structures. 
transition j o i n t  is sui table to general applications 
of composite materials and is not constrained to 
specific unique applications. 
The bi-composite 
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BI-COMPOSITE JOINT DEFINITION 
The term "advanced composite" is used to represent material combinations 
such as boron/al uminum, boron/epoxy or graphi te/epoxy which are composed of 
h i g h  strength and h i g h  st iffness fibrous material integrated i n t o  a matrfr 
t o  produce high performance structure. 
composite laminates i s  that they can be tailored t o  the required design loads 
A significant feature of advancer 
by lamina quantitive and cross plying build up. 
The several available advanced composite material systems provi de the 
composite designer w i t h  many different choices concerning design a? lowablp(r, 
and. mechanical properties. 
ables such as tension, compression and shearbstrength as a function o f  f iber  
volume percent as well as between material systems. Table 1 shows the 
There is  a considerable variance i n  design a1 low 
variance between material systems for boron based composites of 50 percent 
f iber  volumes and f o r  graphite based composites w i t h  55 percent fiber volume 
This data shows a cornpression.strength allowable for  boron/epoxy of 250 k s l ,  
investigation of the data reveals difference i n  tensile strength, tensile 
mdul us, compression modulus and temperature effects for  the noted compos 7 t e  
ma te r i  a? sys tems . 
I 
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1 ,  
I for graphite/epoxy o f  140 ksi and for boron/aluminum of 157 k s i .  Further I 
I 
I 
There are also choices from the mechanrical properties such as density, 
thermal expansion coefficient, thermal conductivity, minl'mum form radius, 
machineability, ply thickness and cost as shown i n  Table 2. T h i s  mechanical 
properties data is based on the same fiber volume percent as the design 
allowable data .  
than boron based composi tes. 
density difference the specific strengths are very close. 
minimum form radius f o r  boron composite is 0.50 inches while for the graphite 
composite i t  i s  only 0.06 inches. 
among the referenced mechanical properties, 
I t  can be seen that graphite based composites are less dense 
I 
The boron composites design a1 lowables are 
general ly  higher than those for graphi te  composites however because of t h e  
The tensile 
strength factor for graphite epoxy is 3.04 and for  boron/epoxy is 2.44. 
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Many other differences can Se observec' 
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Because o f  the large select ion of desi gq a1 1 owabl es and mechani cal 
propert ies available, i n  some design applications, i t  may become 
desireable t o  change f r o m  one advanced composite material system t o  a 
dif ferent one, i.e., boron/epoxy t o  graphite/epoxy. This approach i s  
a continuation o f  the design f l e x i b i l i t y  already being used i n  aerospace 
s t ructure advanced composite material system design applications. The 
bi-composite j o i n t  provides a means of j o i n i n g  two di f ferent advanced 
composite s t ruc tu ra l  systems t o  t a i l o r  the unique design allowable 
and mechani cal propert ies o f  each system t o  the design requirements of 
i n  te res  t. 
The bi-composite j o i n t  therefore i s  an interface between two 
d iss imi la r  materials. Advanced composite structures a re  norma? l y  
. fabr icated by means o f  tape layup where the p l i es  o f  material are 
stacked i n  a predetermined pat tern t o  provide the desired s t ruc tu ra l  
properties. 
s t ructure i s  fabricated f r o m  one composite material t o  take advantage 
o f  i t s  propert ies while the remainder o f  the structure i s  fabr icated 
o f  a d i f f e r e n t  cgmposite material t o  take advantage of other material 
The bi-composi t e  j o i n t -  i s  the j o in ing  c f  these two composi te 
sys'tems as shown i n  Figure 1. 
leaved w i th  p l i e s  o f  the second composite system and the voids are f i l l e c  
w i th  epoxy resin. 
t ransfer  o f  loads across the j o i n t .  
I n  using bi-composite j o i n t  concept, a por t ion of the 
t 
. propert ies. 
P1 ies o f  one composi t e  system are i n t e r -  
The bi-composite j o i n t  provides a smooth t rans i t i on  atid 
The s t ruc tu ra l  requirement i n  designing the bi-composite j o i n t  
i s  t o  bond each l a y e r  o f  the f i r s t  composite system t o  the correspondinc 
layers i n  the adjacent composite system. This i s  required because of  
the cross p l y ing  used w i  t h  composites. 
or ientat ion used i n  the i l l u s t r a t i o n  [0/0-/+45/T45/90/0], - the boron/epoxy 
outermost layers are arranged wi th  the f ibers  or iented i n  the longi tudinal  
For example, w i th  the layer  
d i rect ion.  
w i th  the outermost layers i n  the longi tudinal  d i rect ion.  This approach 
provides an ordered load t ransfer  across the j o i n t ,  i.e., tension t o  
Therefore, the graphi te/epoxy material must also be arrangea 
/ .  ,.-_, 
tension, shear t o  shear, etc. 
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REQUIREMENTS FOR BI-COMPOSITE JOINTS 
One requirement fo r  bi-composite j o in t s  involves the forming of 
a boron/epoxy stringer. The boron/epoxy composite sys tem requires a 
0.50 inch form radius which  is d i f f icu l t  to- layup and impractical to  
interface w i t h  the end f i t t ings.  One solution t o  the problem is t o  
use a bi-composite t ransi t ion j o i n t  and fabricate  the end of the 
stringer from graphite/epoxy which has a 0.06 inch form radius as 
shown i n  Figure 2. T h i s  approach permits forming the end of the 
stringer in to  a closeout f i t t i n g  t ha t  accepts mechanical attachments. 
Another bi-composi te j o i n t  requiremnt involves the efficiency 
of composite structu,res loaded i n  compression. Boronlepoxy has a 
twenty percent greater spec i f ic  compressi ve strength than graphi te/epoxy. 
However, as i n  the previous example, the boron/epoxy 0.50 inch form 
radius causes d i f f i cu l ty  i n  fabricating hat  section s t i f f ene r s  for  a 
s k i n  stringer panel. The bi-composite j o i n t  can be applied by us ing  
graphi te/epoxy 1 ayers arranged i n  [245] orientation draped over the 
entire ha t  surface w i t h  boron/epoxy layers used on the caps only i n  
the. [O] orientation as shown i n  Figure 3. The r e s u l t i n g  s t r inge r  has 
. the advantage of the h igher  speci f i  c compress i ve strength of boron/epoxy 
.and the smaller form radii  available by us ing  graphitelepoxy. 
Thermal isolat ion requirements provide a use f o r  the bi-composite 
joint. 
and advanced composi te pri M ry s t ructure ,  the s t ruc ture  coul d be 
fabricated from graphi te/epoxy o r  boron/aluminum t o  take advantage of 
lower material and fabrication costs w i t h  the thermal stand-offs 
fabricated from boron/polyimide t o  provide thermal is01 ation and higher 
temperature capabili ty as shown i n  Figure 4. T h i s  requirement fo r  the 
bi-composite j o i n t  permits the design of an a l l  composite s t ructure  where 
panels. 
In the case of reentry vehicle w i t h  external thermal protection 
F' ,t the only mechanical j o in t s  are the attaching of  the thermal protection 
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Figure 2, BI-Composite Structural Re1 nforcemen t 
Figure 3. Bi-Coffplosi te Thermal Isolatfon 
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Thermal d i s to r t ion  requirements also provide for 'appl ication 
of the bi-composite jo in t .  
such as antenna disks, solar panel structures and thrust structures 
which are subjected to large thermal gradients, b u t  where thermal 
deflection must be minimized, graphite based advanced composite 
systems should be used t o  take advantage of  the very low, approximately 
zero, coefficient of thermal expansion. However graphite composites 
have a f a i  r ly  h i g h  thermal conduction for  non-metal1 i cs , therefore 
the supports for  these items should be fabricated from boron composite 
materials t o  provide thermal isolation from the vehicle structure as 
shown i n  Figure 5* 
In designing advanced composi t e  structures 
One difficulty i n  applyinq boron fiber composite materials to 
aerospace structures i s  the d r i l l i n g  of holes for mechanical attach- 
ments. Successful use of mechanical attachments i n  composites has 
been. demonstrated by interleaving metal shims for local reinforcement. 
However, d r i l l i n g  holes through a boron/epoxy laminate interleaved 
w i t h  metal shims is a very d i f f icu l t  operation. Again the bi-composite 
j o i n t  provides a means for  designing around this difficulty by making 
a transition from boron/epoxy to graphite epoxy interleaved w i t h  
titanium shims thereby p rov id ing  a graphi te/epoxy and titanium 
structure for d r i l l i n g  as shown i n  Figure 6. T h i s  j o i n t  permits a boron/ 
epoxy composi te  structure w i t h  the r e s u l t i n g  h i g h  compressive s t r e n g t h  
w i t h o u t  the penalty i n  d r i l l i n g  and machining the boron material. 
Combined stiffness and thermal warp requirements present another 
use for the bi-composite joint. On large manipulator arms such as those 
proposed for orbital payload handling, st iffness is a d r i v i n g  function. 
However, form radius and thermal warp make graphite composite necessary 
for the [+45] layer. The longitudinal layers of the structure could be 
made from a boron composi,te and significantly increase manipulator arm 
stiffness w i t h  a resulting small increase i n  thermal warp. 
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Thermal Is01 at ion 
Grapn i te/Epoxy 
/-- Honeycomb Sandwi CII 
Figure 5. Bi-Cowosi t e  Thermal Expansion 
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TYPICAL BI-COMPOSITE JCINT APPLICATION 
Several Saturn S-I1 stage unpressurized structures were 
considered for  application of advanced composite materials i n  a 
recent NASA Contract NAS8-27278 w i t h  Marshall Space F1 i g h t  Center. 
These structures, shown i n  Figure 7 include the forward skirt, a f t  
skirt, thrust cone structure, inzerstage structure and the engine 
suppor t  longeron w i t h  the a f t  skirt being selected for advanced 
composite materials design application study. The composite 
applica%ions included boron/epoxy and graphite/epoxy skin stringer 
structures as we1 1 as boron/epoxy and graphi te/epoxy honeycomb sandwich 
as shown i n  Figure 8. 
The a f t  skirt i's a 33 f o o t  diameter cylinder a t  the mold l ine,  
There are 216 extruded 
located by the inner surface of the s k i n ,  approximately 7 feet  i n  
length and stiffened by three r i n g  frames. 
hat-section stiffeners,  two inches h i g h  and two inches across the 
crown, equally spaced around the shell perimeter, resulting i n  a 
stringer center t o  center spacing of 5.76 inches. 
The a f t  skirt is a load carrying shell structure, which transmits 
loads from the interstage dur ing  Saturn I/C boost and from the thrust 
cone d u r i n g  Saturn I1 boost t o  the a f t  flange of the LOX tank bolting 
r i n g .  
s tab i l i ty  for  the a f t  skirt d u r i n g  Saturn I/C boost and thrust cone 
.structural integrity dur ing  Saturn 11 boost. The maximum external 
body limit loads and associated distributed loads cr i t ical  for  a f t  
skirt design are 5265 pounds per inch compression and 1250 pounds 
per inch tension- The a f t  skirt differential pressure is 3 psi  and 
occurs a t  max qK. 
interstage end and -280 F a t  the LO2 tank' interface. 
Two of the r i n g  frames are dual purpose members providing shell 
The temperature extremes are 200 F on the 
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One of the composite application p o i n t  desi gns , boron/epoxy 
skin-stringer, used the bi-composite j o i n t  concept. Composites of 
boron/epoxy are viable structural candidates especially where stiffness 
t o  weight properties are desireable. 
a higher thermal coefficient of  expansion t h a n  graphite/epoxy.and is  
therefore more compatible fo r  bonding t o  metal structures and also pro- 
vi des better thermal is01 at ion properties t h a n  graphite/epoxy. Boron 
composites however, are more difficult  to  machine and tne prepreg 
requires a 1/2 inch form radius and is difficult  t o  drape. The boron/ 
epoxy stringer reinforced skin concept given i n  Figure 9 i s  based on a 
laminate consisting o f  nine plies .007 thick w i t h  a laminate configuration 
of [0/+45/0/90/0/T45/0]. 
the composite layers a t  the edges t o  provide for attachments. 
Boron/epoxy has the advantage of 
T i t a n i u m  reinforcement is  interleaved w i t h  
The minimum weight  boron/epoxy stringer has a dimension across 
the foo t  of 4.00 inches; a w i d t h  across the crown of 2.09 inches and 
a h e i g h t  from crown t o  foot  of 2.09 inches. The stringer spacing is 
6.50 inches which  results i n  48 stringers per quarter panel. The 
boron/epoxy stringer has eight plies of composite w i t h  a laminate 
configuration o f  [24!i2]s. The form radius is o f  course 1/2 inch. 
A bi-composite j o i n t  was used i n  this study t o  counteract the 
1/2 inch form radius a t  the stringer termination. 
skirt boron/epoxy stringer interfaces w i t h  a three bo1 t attachment 
to  adjacent stage hardware. One of the bol ts  is  located on the 
stringer center line and one each on either side o f  the stringer 
through a bonded f i t t i n g  as shown i n  Figure 10. A bi-composite 
t r ans i t i on  was used t o  change from boron/epoxy to  graphite/epoxy and 
the resulting 1/16 form radius. With the added layup flexibility a 
closeout type f i t t i n g  could be designed into the end of the stringer. 
The smaller form radius also provided a more efficient bond f o r  the 
adjacent molded chopped fiber f i t t i n g .  
Each end of the a f t  
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The bi-composite j o i n t  i s  appl icable- to  a wide 'variety of design 
applications as i l l u s t r a t e d  i n  the preceeding discussion. 
e f fec t i ve  so lut ion t o  many design problem areas such as thermal expansion 
and/or i s o l a t i o n  and t o  s a t i s f y  requirements f o r  varying s t ruc tu ra l  load/ 
strength requirements w i th in  a given s t ruc tu ra l  component. 
provides an ef fect ive way t o  a1 l ev ia te  manufacturing p roduc ib i l i t y  
constraints such as forming radius and machining operations associated 
w i th  the boron f iber  materials. 
It i s  an 
It also 
The inherent f l e x i b i l i t y  i n  fabr ica t ion  operations associated w i th  
the laminar p l y  lay-up technique o f  composite materials permits the 
inclusion o f  a bi-composite material t rans i t i on  sect ion wi th  very l i t t l e  
impact on the fabr icat ion operation. 
potent ia l  benef i ts t o  be gained make the bi-composite j o i n t  a very a t t rac t i ve  
concept f o r  sa t i s f y ing  spec i f i c  design requirements i n  composite material 
structures. 
This f a c t  i n  conjunction wi th  the 
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APPROVAL FOR 
SCIENTIFIC AND TECHNlCAL PRESENTATIONS AND PUBLICATIONS 
DATE May 25, 1972 T O  : Jules M. Niller, K/S 105, 3251 --- 
f??. 
q. FROM : Aero-space Technologist, 
i 
AGENCY CODE 1 13-32-01 -21 
SUBJECT: Matcrid enclosed consist of 0 Abstract a Rough draft, L- er] Final paper for 2 Article 
O Paper rzQ TA: entitled: 
Small Scale Explosive Seam welding 
I 
Authors (if not all Langley authors, indicate affiliation and describe nature of collaboration in par. 9’: 
827- L. J. Bement 
1. 
2. Meeting 
3. 
4. Sponsor: 
5. Status of marcrid: @ Data are in hand 0 Data will be available by -- 
Classification of material: Unclassified 0 Confidential 9 Secrct - 
Paper will be presented by: L. J Bement 
0 Publication 
Date and place of meeting: 
Rcspcctive authors’ teI. nos. ( I )  3823 ,z 3, 
Name: Symposi um on We1 ding , Bonding , and Fas teni ng 
May 30-31 , June 1 , 1972 - Wi 11 iamsburg , VA 
NASA, George Washington University, & Tne American Society for Metals 
Previous approvals: 
f3 None Rough draft ‘Jp L31lg;Cy a SpuXjJr 
From a Abstract 0 ,  a Hc‘idqtarterc 2 
6 .  Previous or future publicition,si of essentially same material :if PO NASA report wii! bs published. <rate wL;. . 
Author(s) . Title Status 
.1fpre~rn;eJ.  s f i t c  V;!IK:. :% .:x 
L.’J. Bement Small Scale Explosive Welding The Franklin Institute 
o f  Aluminum Philadelphfa, PA 
Sept. 1971 
8. Deadline date of action: June 30, 1972 
9. Additiolid comments (include statement regarding in-house rcwcw of paper for esrcrnal Fublication 1. 


The bi-composite transition joint

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