sorts of interesting design concepts previously not attempted with the t r a d i t i o n a l aerospace materials. Along with the positive aspects of composites, the limitations must also be assessed. Prior to failure, modern metallic aircraft structures have a capacity to redistribute loads by plastically deforming, whereas filamentary reinforced materials do not. Failure of composite structures under load is usually sudden. Examination of the stress-strain response of three typical aircraft structural materials will clarify the point, as shown in Figure 1. First, note that the response of the 143 Eglass fabric is linear (straight line) from 0 stress to f a i l u r e stress of 105,000 psi. The stress-strain relationship is the same at 80,000 psi as
Modern Composite Aircraft Technology By Hans D. Neuhert (EAA 631181 6051 Prado St.
Anaheim, CA 92807 and
Ralph W. Kiger 10201 Wembley Circle Westminster, CA This article, first in a series, deals with materials and predesign aspects of modern composite aircraft technology. We have tended to be a little more technical than one normally finds in SPORT A VIATION on the premise that education and information to you, if not abused,
strengthens the EAA movement.
in homebuilt and certified aircraft is certainly not new, having been used in cowls, tips, closures, landing gears, etc. for many years. However, the current trend of using composite materials in primary structure (wings, fuselages and empennages), as evidenced by the Vari-Eze, KR's and others, represents a substantial departure from previously accepted applications. These materials offer new avenues for efficiency, cost savings, fabrication ease and all TABLE 1
FIBER DESCRIPTION
BORON
During World War II, the need for new materials to replace scarce metals accelerated the research and development of plastics and reinforcements. In 1975, the total consumption of reinforced plastics in all markets reached approximately 1.7 billion pounds, with reinforced thermosetting materials accounting for about 85''f of the total. In the middle
1950's, NASA, the military services and i n d u s t r y recognized t h a t advanced composite materials would dominate in the 1980's due to inherent low weight, high strength and stiffness characteristics, whereas the strength-to-density and stiffnessto-density ratios of most metals fall in a closely grouped area. The emphasis has been to exploit a new class of reinforcements. These new fibers include tungsten substrate Boron, rayon and polyacrynitrile precursor graphite and inorganic Aramid fibers.
SELECTED ADVANCED COMPOSITE FIBER PROPERTIES
TRADE NAMES
4,5.6
PART I. MATERIALS BACKGROUND AND PRELIMINARY DESIGN
FIBER SUPPLIERS
4 8 MIL BORON
4.2 & 5. 7 MIL BORSIC HIGH STRENGTH GRAPHITE
HIGH MODULUS GRAPHITE
400,000
55
1290
FIBERON 4 AVCO
400,000
55
J305
450,000
34
S50
HERCULES
450,000
30
J45
MODMOR III
MORGANITE
450,000
30
J50
GREAT LAKES CARBON
400,000
49
190
HERCULES
350,000
55
J95
MODMOR 1
MORGANITE
350,000
55
J95
CELION GY-70
CELANESE
310,000
75
(100
KEVLAR 49
DU PONT
525,000
19
115
FIGURE 1 STRESS-STRAIN COMPARISON Of E-GLASS, 4130 STEEL, AND 2024-T3 ALUMINUM
STYLE 143 E-GLASS WARP DIRECTION
EPOXY RESIN
To date, NASA and the military services have funded the aerospace in-
dustry in excess of $400 million to bring this technology into full maturity. These advanced fibers are mentioned here so that you may be aware of their existence because, as increased demand lowers cost, their utilization by the homebuilder will become possible. Table 1 gives a physical property comparison of these materials. In the meantime, the only cost effective fiber available to homebuilders is fiber-glass. The use of fiber-glass construction 58 JULY 1976
.01
COST PER POUND
FIBERON 4 AVCO
UNION CARBIDE
TYPE HM-S
ARAMID
FIBER FIBER TENSILE MODULUS STRENGTH IPSI) (lO^SI)
TYPE A-S
THORNEL 300
FORTAFIL CG-5
ULTRA- HIGH MODULUS GRAPHITE
it is at 40,000 psi. At 105,000 psi, the
glass fabric fails without warning. In comparison, the a l u m i n u m and steel materials shown also have a linear elastic regime. However, at
.02
.03
.04
.05
.06
STRAIN, IN/IN
.07
.08
.09
.1
the proportional limit, the response of these materials changes into "plastic deformation". When one bends metal to a new position (i.e., bending an angle out of sheet stock), the material is taken into the "plastic" range. When the load is released, you
each glass composition, but is generally 2500' F. The molten glass is then drawn directly into fibers, or made into modules which are subsequently remelted and drawn into fibers. The fibers are drawn from a plati-
to bend up the angle from sheet stock, assume you are stressing the aluminum sheet to 52,000 psi and, as a result, have introduced a permanent set of .01 inch/inch strain. Upon load release, the material relaxes back
holes, gathered together and drawn or stretched m e c h a n i c a l l y to the proper dimensions. The number of
have permanent set, but the material still retains its properties up to the proportional limit. For example,
to zero stress. At any additional applied stress up to 52,000 psi, the material response is linear, just as it
was before being bent (fatigue aspects not being considered for this example). The fact that a unidirectional glass fabric does not have a "plastic deformation" range is not a bad feature; it only must be considered in the design process. Most EAA members recognize that
not all plans promoted in the classified section of SPORT AVIATION are the product of engineering trained
num alloy bushing containing many holes in its bottom. The molten glass in the bushing is gravity fed through
holes in the bushing determines the
number of filaments per strand, while
the glass temperature and the drawing speed determine filament diameter.
Although glass is strong, it is friable and subject to self-abrasive damage. To prevent this, a binder composed of starch and lubricating oil is applied to the filaments as they leave the bushing. The binder also aids in the further processing of this
yarn by preventing damage to the
filaments during various preparation steps, such as twisting, plying and weaving. The nomenclature of glass yarn
designers, nor have all prototype designs enjoyed the luxury of a thorough stress analysis and subsequent load test verifying the analysis. Since the EAA policy for homebuilders is "caveat emptor" (let the buyer beware), it is our purpose to educate
consists of an alphabetical and a numerical part. The first letter refers to glass composition (E for electrical grade, C for chemical, S for high strength, etc.); the second letter designates fiber form (C for continuous, S for staple); the third ( a n d fourth, when used) specifies the filament diameter (D= .00021 in., E = .00029 in., G= .00036 in., etc.). The
lightly; that the design/analysis process is difficult to perform, more so
is the glass yield in yards per pound.
those contemplating the use of composites in primary structure. We do this to demonstrate that aircraft structural design is not to be taken
in composite materials; and that you, the builder, ought to have some means of applying judgment to test the "caveat emptor" principle. Our educational approach to you will be via a series of articles dealing with preliminary design, a detailed design example, fabrication demonstra-
first part of the numerical portion
TABLE 2
FIBER DESCRIPTION
E-GLASS
FERRO
tion, and maintainability and repair.
In this first article, let us present
material, and perform a preliminary design tradeoff and optimization of
the T-18 outer wing.
FIBER-GLASS
Introduced in the late 1930's, fiberglass is made by mixing the various
ingredients together in dry form and melting the mixture in a refractory oven. The temperature varies for
fill for making 181 style cloth — a
typical aerospace grade. The code tells us that it is E glass, continuous, with 204 filaments per strand and .00029 in. filament diameter.
The yield is approximately 22,500 yds./lb. and three strands of ECE225
are plied together to form ECE2251/3. ECE225-4/3 would be made by first plying four ECE225-1/0 strands together, resulting in an ECE2254/0 yarn. Three sets of ECE225-4/0 plied together yield the ECE225-4/3 final yarn, which is used to make 184 style cloth.
E-glass is the predominate fiber
produced today. For special applications, typically aerospace, high
strength S-glass has been used. This fiber is marketed to two forms — aerospace (MIL-Spec) grade and commercial grade. The difference is primarily in the quality control requirements and means a 3-to-l price differential. All E-glass is of commercial grade, receiving nominal qual-
ity control during its manufacture. Shown in Table 2 are data for E and S-glass.
GLASS FABRIC CONSTRUCTION
The inherent properties of glass yarn are passed on to the fabric. Principally, these include high strength, flame, heat, weather and chemical resistance, and dimensional stability. FIBER GLASS PROPERTIES
FIBER TENSILE STRENGTH (PSI) (VIRGIN MONOFILAMENT)
500,000
FIBER MODULUS
(10* PSI)
COST PER POUND (YARN)
10.5
$.60
12.6
$8.00
12.6
$2.50
(STRAND TENSILE
STRENGTH = 285,000 PSI)
UNIGLASS INDUSTRIES
PITTSBURG PLATE & GLASS
redesign a typical outer wing of an
a brief background on composite materials, select and characterize a
OWENS-CORNING LIBBY-OWENS-FORD
The method to be utilized will be to
existing design (John Thorp's T-18) on an equivalent strength and stiffness basis, and finally a test-to-failure to show satisfactory compliance and correlation.
FIBER SUPPLIERS
singles plied together. For example, ECE225-1/3 is used in the warp and
S-GLASS (AEROSPACE GRADE) S-1014
FERRO
S
OWENS-CORNING
656,000 (STRAND TENSILE STRENGTH = 535,000 PSI)
S-GLASS (COMMERCIAL GRADE) S-994
FERRO
S2
OWENS-CORNING
The second series of numbers which resemble a fraction designate the number of single strands in the plied
yarn. The digit to the left of the slash shows the number of strands twisted together, and the digit to the right of the slash indicates the number of
650,000 (STRAND TENSILE STRENGTH = 500,000 PSI)
Glass fabrics have been engineered
for specific applications by designing the fabric construction to give
the required weighfarea, thickness
and breaking strength. The weave construction governs the drape characteristics and how the warp (long SPORT AVIATION 59
ber, before using any fabric without knowing its history, determine that the finish is compatible with the resin; otherwise, property degradation will occur with time due to moisture in the air.
directionl and fill (width direction) yarns are interlaced. For instance, a satin weave is used in producing u n i f o r m laminates with compound curvatures and deep draws. The basic weave patterns are plain, twill and satin. Variations of these i n c l u d e basket, crawfoot satin, eight h a r ness satin, leno, mock leno and unidirectional. Table 3 lists only a few of the more t h a n 250 designated weave styles available. Standard width is 38 in., with 44, 50 and 60 in. available.
EPOXY RESIN
Since t h e i r i n t r o d u c t i o n in the United States in the early 1950's, epoxy resins have gained wide acceptance in such diverse applications as surface coatings, transfer molding compounds, reinforced plastics and adhesives. This versatility is due to the combined factors of chemical structure, reaction mech-
FINISHES
It is recalled that during manu-
be varied from bisphenol A resins to other chemical types to impart specific characteristics to an epoxy composition. These variations include: Epoxy novolac, cycloaliphatic epoxy, brominated epoxy, aliphatic epoxy and special high functionality resins. Each of these special and distinct formulations were developed to satisfy the requirements of special applications. For example, epoxy novolac is most widely used in electrical components for elevated temperature performance, chemical resistance, stiffness and r e a c t i v i t y reasons. Epoxy novolac is very brittle and not useful for laminating in homeb u i l t aircraft applications. Bromi-
TABLE 3 DATA FOR SIX SELECTED FIBER GLASS FABRICS (E-GLASS)
CONSTRUCTION STYLE
WARP ENDS/INCH
FILL PICKS/INCH
YARN WARP
WEAVE
FABRIC THICKNESS (MILS)
FILL
TENSILE STRENGTH LB/INCH WIDTH WARP FILL
WEIGHT (OZ/SQYD)
STD ROLL LENGTH (YARDS)
COST PER YARD STD ROLL LENGTH (HEAT CLEANED VOLAN-A FINISH)
120
60
58
450 1/2
SAME
CROWFOOT SATIN
4
135
125
3.16
500
J.86AD
143
49
30
225 2/3
450 1/2
CROWFOOT SATIN
9
675
85
8.90
125
J.85AD
181
57
54
225 1/3
SAME
8 HARNESS SATIN
9
350
340
8.90
125
J1.13/YD
1543
49
30
150 2/2
450 1/2 CROWFOOT
9
675
85
8.65
500
J.85AD
1621
30
14
150 1/0
SAME
LENO
6.2
78
98
2.36
1000
S.245AD
7634
16
14
75 2/2
75 2/3
PLAIN
15
410
624
12.14
125
SU3AD
SATIN
MAJOR FABRIC WEAVERS: BURLINGTON INDUSTRIES, J.P. STEVENS, HEXEL/COAST MFG. UNITED MERCHANTS, AND CLARK-SCHWEBEL
facture a starch/oil sizing was applied to the fibers to prevent abrading. Because glass fabrics have many diversified uses, it is necessary to remove this sizing from the woven fabric and apply a special finish which will provide optimum performance for a specific application. For aircraft uses with epoxy resins, the interface between the fiber surface and the resin is chemically degraded in the presence of moisture. The moisture in the air will diffuse through the resin and attack this bond. The result is a rather dramatic reduction in strength properties. To overcome this problem, a number of standard and proprietary finishes have been developed by the weaving industry. These finishes are normally applied in solution form after the fabric has been heat cleaned and washed to remove the starch sizing. Finishes have been developed for melamine, silicone, polyester, phenolic, epoxy, polyimide and other resin systems. Epoxy and polyester-compatible finishes include Volan A
anism, wide formulation latitude and high performance inherent in epoxy resin technology. Epoxy resins are a class of low-molecular-weight resinous compounds generally characterized by the presence of the threemembered oxirane ring. Most epoxies produced are based on the reaction product of epichlorohydrin and bisphenol A. Thus it is bisphenol A epoxies that are usually referred to in discussions of epoxy properties. The physical structure of bisphenol A epoxy molecules imparts the desired characteristics. The rigid aromatic backbone connected by short glycidyl ether linkages promotes high strength and modulus properties, as well as increased thermal stability and good c h e m i c a l resistance.
A1100 ( a m i n o - t y p e silicone), UC Y4087 (glycidoxyprophy trimetho-
susceptibility to ultraviolet degradation, which manifests itself as yellowing and chalking. The result is a reduction in strength properties. The basic chemical structure can
( m e t h a c r y l a t e chromic c h l o r i d e ) ,
loxysilane) and Garan (vinyl-triet-
hoxy). There are others, but these
are the most widely used. Remem60 JULY 1976
Chemical polarity provided by hy-
droxyl and ether groups contributes toward the widely recognized adhesive qualities of epoxy resins. Thus, functionality promotes intimate wetting of metal, wood, fiber-glass and other polar substrates. This aromaticity does, however, result in one of the drawbacks of epoxy — their
nated epoxies are utilized where a high degree of fire retardency is desired. Cycloaliphatics provide good electrical arc resistance, and find applications such as electrical potting and encapsulations. Unsaturated polyesters (the type most familiar to homebuilders) have not been mentioned, primarily since epoxies, as a whole, provide 20 to 25^ increases in performance over the polyesters in the areas of strength, stiffness, stability, impact resistance and fatigue life. While formulations are available to bring a specific characteristic equal to or greater than a comparable epoxy, other desirable traits are reduced. As a whole, the epoxies are much more suited for p r i m a r y aircraft structure where maximum structural performance is desired at minimum weight. Polyesters are most ideally suited in secondary structural applications — i.e., wing tips, cowls, fairings, seats, etc. Most epoxies are designated by a number and several manufacturers offer essentially the same epoxy. For example, Shell E P O N " 828, Dow Chemical D.E.R. 331, Union Carbide 2774 and Ciba-Ciegy 6040 are all equivalent. The same holds true for curing agents, reactive dilutents, modifiers and other additives. With
literally hundreds of products on the
market, the situation is confusing,
even to those who have worked in the
resin field for many years. Criteria for selecting a particular epoxy include characteristics during
processing, strength, stability and
properties after cure. Pure bisphenol A resin by itself is not generally used for l a m i n a t i n g : it is b r i t t l e when
cured and has a high viscosity, giving poor wetting during layup. Resin distributors normally blend various
combinations of resins together, add
reactive diluents to control viscosity, mix in elastomeric modifiers to control toughness, solvents and cheap
extenders to reduce cost. Sometimes
fillers and pigments are used to reduce shrinkage, adjust thermal pro-
perties, improve chemical resistance and increase hardness (powdered
a l u m i n u m , for example). It's all like
a cake recipe, with countless variations for special uses and each with a different identification n u m b e r .
The result is a high degree of confusion, claims and counterclaims of what is better, and most often the f o r m u l a t i o n is proprietary to that p a r t i c u l a r resin supplier. All you see is a can with resin inside and a number on the label outside. Therefore, selection of a resin for a particular application is a complex pro-
cess, but should be done with great care.
EPOXY PROPERTIES AND CURING AGENTS Bisphenol A-type epoxies are all very similar. For comparison, let's examine Shell EPON 815, 820 and 830. The chemical structure of a typical molecule of the base resins is identical. Shell has added to the base Bis A, resin additives and reactive diluents to control viscosity and reactivity (pot life) with the curing agent. For Shell epoxies with increas-
ing numerical designation, the viscosity increases, the average molecu-
lar weight increases and elevated temperature s t a b i l i t y increases. EPON" 815 and 820 contain a monofunctional diluent, whereas 828 contains a bifunctional d i l u e n t . Incidentally, a diluent is a thinner which controls viscosity, but is chemically included in the resin chain reaction
4. EPON 1 Curing Agent T-l For longer pot lives and room temperature curing, one may use: 5. EPON" Curing Agent V-15 6. EPON' Curing Agent V-25
7. EPON Curing Agent V-40 All of the above curing agents require seven days at 77' F for complete cures.
Each of the above agents results
in a cured epoxy having slightly different f i n a l characteristics. They
were developed for large non-aerospace user(s), each with special requirements. Of the EPON" (as well as other manufacturers) resins cured
with the above, they all tend to have brittle fracture characteristics, which
is not desirable from a fatigue nor an impact point of view. These resins lack toughness, as evidenced by poor
peel strength performance. These are i m p o r t a n t c o n s i d e r a t i o n s for
homebuilts, since one expects to get utilization from the airframe over a
long period of time. Early attempts toward improving
toughness and peel strength have been successful, but only at the expense of other performance characteristics. Only recently, a new approach was taken, which is an elastomeric (as opposed to chemical) modification of the epoxies which maintains essentially all of the attributes TABLE 4
blending this resin for use by aircraft
homebuilders. We thank him for the data he has supplied, which is given in Tables 4 and 5. As one can recognize, the elastomeric modified resin
has comparable strength and stiffness, but the substantial increases
in shear strength and climbing drum peel are noteworthy. Chemical resistance is also comparable to the
standard epoxies. From what we can determine from the literature at this time, this "new" epoxy system is preferred for our homebuilt aircraft applications. MATERIALS CHARACTERIZATION
In order to successfully design an
aircraft structure, one must know the stress-strain behavior of the material. For our example, we have
chosen to use Style 143 fabric, which is sometimes known as "the poor
man's unidirectional weave". This material, with its highly different
MECHANICAL PROPERTY COMPARISON (CAST RESIN SAMPLES) TYPICAL BISPHENOL-A EPOXY SYSTEM
FLEXURAL STRENGTH
13,900 PSI
CUSTOM AIRCRAFT BUILDERS EPOXY SYSTEM
17,000 PSI
FLEXURAL MODULUS
4 . 4 X 10 PSI
4.4 X 105 PSI
TENSILE STRENGTH
11,400 PSI
10,000 PSI
SHEAR STRENGTH
4,760 PSI
6,100 PSI
CLIMBING DRUM PEEL STRENGTH
1 in-lb/in
54 in-lb/in
ULTIMATE ELONGATION
4.4%
9.0%
HEAT DISTORTION TEMPERATURE
270 f
248 F
5
TABLE 5 CHEMICAL RESISTANCE COMPARISON % WEIGHT GAIN OF CAST RESIN SAMPLES AFTER 30 DAYS IMMERSION)
at cure, as opposed to paint thinner
TYPICAL BISPHENOL-A EPOXY SYSTEM
which evaporates from the pigment. For EPON" 815, 816, 820, 826 and
828, Shell recommends seven curing agents for wet layup laminating and room temperature curing. For short pot lives and fast cure, one may use: 1. DIETHYLENETRIAMINE (DTA) 2. TRIETHYLENETETRAMINE (TETA) 3. EPON' ( Curing Agent U
of the epoxy, but which possesses the added benefit of improved toughness and excellent peel strength. These elastomeric epoxies are relatively new (1960's) and have not yet enjoyed widespread application. An associate, Ray Lambert of Northrop's Manufacturing Research and Development Department, has been independently evaluating and
CUSTOM AIRCRAFT BUILDERS EPOXY SYSTEM
DEIONIZED WATER
0.66%
0.63%
10% CAUSTIC SOLUTION (NaOH)
0.52%
0.52%
25% ACETIC ACID (HOAC)
9.30%
10.0%
METHANOL
7.24%
8.63%
TOLUENE
14.7%
2.06% SPORT AVIATION 61
properties in the warp and fill directions, allows us a better chance of
optimizing the layup. This basic material (143 E-glass/ epoxy) is then tested in tension and compression in the warp, as well as transverse to the warp direction. Inplane shear and Poisson's ratio are also determined by test. The results of these tests are shown in Figure 2 in the form of stress-strain curves. In composite materials, one must think in terms of strain, as opposed to m e t a l l i c materials where one thinks in terms of stress. In order to utilize composite materials properly, we now establish a design allowables criteria and some definition of terminology. Limit Loads: Most severe combination of loads and environment expected to encounter in service. Ultimate Loads: Limit loads times a factor of safety, usually 1.5 for aircraft. Margin of Safety: Additional safety factor above ultimate before actual failure. The M.S. must be equal to or greater than 0; limit loads to limit strain or ultimate loads to ultimate strain. Yield: Yield is the onset of inelastic behavior. Proportional Limit: The point on the stress-strain curve beyond which stress is no longer proportional to
will always be 1.5 times limit. Our margin of safety between limit and ultimate loads will always be conservative. Second, by designing at limit loads, we are using linear elasFIGURE 2
every time one mixes a batch of resin and applies it to a fabric, one experiences scatter in the property values obtained by test. Experience has shown that properly prepared
STRESS - STRAIN BEHAVIOR OF STYLE 143 E-GLASS/EPOXY RESIN
LEGEND O
AVE. ULTIMATE FAILURE
O
ELASTIC YIELD
•
DESIGN LIMIT
LONGITUDINAL
TENSION I———
strain.
Damage: An inconservative phenomena which may be observed by: a) Permanent Deformation bl Degradation of Modulus c) Reduction in Strength d) Loss of Environmental Resistance Criteria: We chose to design to limit loads using proportional limit strains, with damage to the composite not permitted at limit loads. In the determination of design properties, we actually work with strain and apply our factors to those values. To satisfy our damage criteria, we choose between either 2/3 of design ultimate strain or proportional limit, whichever is lower. A summary of these computations for in-plane tension, compression and shear is shown in Table 6. We have now defined the properties of one layer of Style 143 E-glass cured in epoxy. This is summarized in Table 7. This one layer of material we refer to as a "lamina". The next step is to define and predict how a "laminate" behaves. A laminate consists of any number of layers in any orientation relative to some point of reference. For wings, the span is usually the reference direction. In applying this criteria to our design example, we have done a number of things. First, by proper selection of limit strain, ultimate strain 62 JULY 1976
.004
.008
.012
.016
.020
.024
.028
STRAIN, IN/IN
tic properties. If we were to design using ultimate strain and loads, an elastic plastic analysis would be required. Also, fiber-glass tends to deflect more than metals due to lower modulus, and large deflection analysis (updating of stiffness matrix) is difficult to perform. Combining an elastic-plastic material with large deflection analysis is extremely difficult to do for complex structures. The only recourse is to go to finiteelement techniques and there are only a few computer programs in the country to handle that class of problem. Third, all available optimization techniques consider only linear
elastic materials. In summary, by
designing with limit strains to limit
loads, the design problem is solva-
ble by expedient techniques.
Since a new material is created
materials fall within an 0.8 to 1.2 distribution of test averages. This factor agrees remarkably well with statistical approaches. For example, our material has an average ultimate failure stress of 105,000 psi. For comparable fabricated laminates from different lots and batches, we
would expect that 95r^ of all tests would be greater t h a n 0.8 times 105,000 psi, or 84,000 psi. Therefore, our computed design test ultimate, based on statistical history, is 84,000 psi. If our structure fails
higher than that, then we have a built-in factor of safety.
In the optimization process, we
are concerned for orienting the least number of layers in the proper direction to achieve an acceptable structure at minimum weight. Since it is impractical to fabricate and test vir-
TABLE 6 DETERMINATION Of DESIGN LIMIT PROPERTIES FROM STRESS-STRAIN DATA (FIG. 2)
STRAIN
DIRECTION & LOADING
YOUNG'S MODULUS (M5% FIBER VOLUME
ULTIMATE STRENGTH (FROM FIG 2)
.8ULT STRENGTH (KSII
ULTIMATE STRAIN (FROM FIG 2l
YIELD STRAIN (FROM FIG 2) (IN/IN)
(IN/IN)
CALCULATED DESIGN LIMIT STRAIN
.
ultimo'*
'10* PSD
(KSII
4.23
105
84
.025000
.013340
56
443
69
60
-.016875
-.009000
40
1.82
10.2
8.16
.019000
.003000
,008100
5.4
1.80
33
26.4
-.026000
-.002700
-.009700
4.86
..002560
t. 011600
5.05
( .8 x 2/3 x .iroin)
' "t LONGITUDINAL IN TENSION '"c LONGITUDINAL IN COMPRESSION '», TRANSVERSE IN TENSION
CALCULATED DESIGN LIMIT STRESS ( E X ' 1
'"c TRANSVERSE IN COMPRESSION *.2 IN-PLANE SHEAR
6.3
'.86
.51
t. 017500-
•ULTIMATE SHEAR STRAIN IS DETERMINED AS THE INTERSECTION OF THE SHEAR STRESS-STRAIN CURVE AND 70% SLOPE OF THE INITIAL SHEAR MODULUS.
TABLE 7 STYLE 143 E-GLASS/EPOXY RESIN PROPERTIES USED FOR LIMIT LOAD DESIGN ANALYSIS ALLOWABLE LIMIT STRESS = MODULUS X DESIGN LIMIT STRAIN
STRESS
MODULUS
56,000 PSI
4 . 2 3 X 106 PSI
DIRECTION & LOADING LONGITUDINAL TENSION LONGITUDINAL COMPRESSION
40,000 PSI
IN-PLANE SHEAR
-.009000 IN/IN
6
.003000 IN/IN
6
-.002700 IN/IN
1 . 8 2 X 10 PSI
4880 PSI
TRANSVERSE COMPRESSION
.013340 IN/IN
6
4 . 4 3 X 10 PSI
5400 PSI
TRANSVERSE TENSION
STRAIN
1.80X 10 PSI 6
5050 PSI
.435 X 10 PSI
+ .011600 IN/IN
POISSON'S RATIO = .10
tually hundreds of orientation and layer combinations, we choose instead to compute these properties using the industry-wide "plane stressorthotropic material-lamination theory" method and the maximum strain failure criteria. To explain all that
"1 • y l °1
T
=
°66
12
2
(tension or compression)
22 2
(tension or compression)
C
C
V
(shear)
12 ,
where c
l * C2 ' T 12
are
c, , c- , Y,p
would fill another book. MIL-HDBK17A probably treats the subject as
The following will alert you to the computational complexity involved. In industry, we employ the computer
'onqitudinal, transverse and shear stress
are
longitudinal, transverse and shear strains
En/d -
Q
well as anyone.
22
E22/(l -
- v)2 v 21 ) "66
for expedience and accuracy.
The mathematical relationship of Hooke's Law for specially orthotropic materials is as follows:
°2
=
= G
12 •
The above holds true only in the direction of the fibers and transverse to the fibers. If a layer is placed at sone angle 8 relative to our reference (span), we nerforn a fourth-order tensor transform, giving us:
xy
where SPORT AVIATION 63
sin
cos 9
22
w i_ub
w T u-*j b i n
. 2 cos 2 9« * O cos 4 9 sin ft
= Qn sin 9 + 2 (Q 12
Q 12 = (On + 022 - 4Q66) sin2 cos2 9 + Q)2 (sin4 9 * cos4 9)
26
°22 '
2f)
12 - 2Q66>
- Q ie '
20
66>
s1
9+
"
5in 9
9 cos
The ahove relates the properties of one layer at some angle to our reference, where the x direction is spanwise and the y direction is chordwise. We do the above for every layer. The above equations are then incorporated into strain-displacement relationships. If we have totally confused you, and shown you that it becomes a computational nightmare if done by hand, then let's stop, be-
(0
9)
°66
12 * Q22 +
2f1
66>
° + (012 ' °22
up combinations. 64 JULY 1976
cos3 9
cause there is much more and it gets worse. Also, if you are impressed that an engineering approach to aircraft composite materials design is complicated and therefore beyond the scope of the average person, then our point is made. Do the design job right or not at all; improper design could be hazardous to your health. Results of three computations for our material are shown in Figure 3. Notice the elastic properties on the lower left of each curve are the lamina input properties obtained from Table 6. For the layup orientation shown in the title, the material will result in elastic properties shown in the lower right hand of each curve. The limit strengths of the layups are read from the curve, X (spanwise), y (chordwise).
FIG. 3. In-plane strength characteristics of style 143E — glass for 3 lay-
cos 9
In summary, we have taken actual test data for the chosen material, and carefully applied judgment and experience to its strain-stiffness characteristics to define a single layer which meets our design criteria. Then we mathematically transform and compute the performance of any combination of layers and orientations relative to our reference system (span-chord). In the optimization process, we search for the minimum number of layers at the correct angle to just satisfy the applied loads. Additional layers result in excess weight, extra cost and more labor. M I N I M U M WEIGHT DESIGN EXAMPLE Nearly every application of composite hardware to contemporary airframe structures has been through the replacement of a conventional metallic component with a composite one satisfying the same form, fit and functional requirements. These replacements have been justified on the basis of singular or synergistic advantages offered by composites through increased structural efficiency, weight savings, cost savings, fabricability, r e p a i r a b i l i t y , maintainability, etc. An example of such a trade study is presented here and also uses a conventional metallic component as a baseline for design purposes. This aluminum baseline provides geometric and structural parameters which pace the sizing of the composite concepts and
thus allow a comparison of each design on the basis of relative structural weight. The concepts presented are not strictly "optimum" designs, but more a combination of theoretically derived minimum requirements modified by practical considerations which allow for simplicity of construction. The net result, however, is an indication of the inherent flexibility of design and potential weight reduction associated with the judicious application of filamentary composites to a typical primary aircraft
component.
BASELINE METALLIC CONCEPT
The existing metallic component selected as the baseline for this exercise (with the approval of the originator) is the outer wing panel of the Thorp T-18 Sport. This segment of primary structure, the entire aircraft not withstanding, has been designed
with the homebuilder in mind and has sacrificed e v e r y t h i n g but integrity in order to provide construction simplicity. The wing skin is a one-piece wraparound sheet of 0.025 in. 2024-T3; all ribs are formed 0.032 in. 6061T4 sheet; the rear spar is formed 0.032 in. 2024-T3 sheet; the front spar has an 0.025 in. 2024-T3 web capped with 3/32 in. thick 2024-T4 extruded angles. The modified NACA 63j -412 airfoil is constant chord and constant depth. This type of simplicity is the trademark of the very popular Thorp T-18. OPTIMIZED CONFIGURATIONS
Attaining a m i n i m u m weight design for a given structural application involves a process of tradeoffs between the various portions of the structure. To accomplish these tradeoffs, some functional relationships must be established which identify weight changes in the various components resulting from changes in configuration. Once these relationships have been determined, the configuration can be systematically altered and a minimum weight design found.
The inter-relationships of ordinary
indeterminate structures are nonlinear, so some form of simplifications are necessary to allow solution of the problem in a reasonable time. These assumptions fall i n t o two classes — those required to yield solvable mathematical relationships, and the technique utilized to solve
the resulting equations. Other assumptions include linear elastic materials at constant temperature, the structural geometry and loads are known quantities and assumptions required based on fabrication
details (i.e., simply supported boundary conditions).
CONCEPTS I AND II — SKIN-STRINGER
For this design configuration, the literature indicates that a near minimum weight design is reached when the applied loads cause panel buckling of the skin and column buckling of the stringers simultaneously. This way, the skin carries its share of the load up to ultimate failure and is known as the simultaneous failure mode theory for compression covers. The upper cover is designed to be adequate in compression at the 5g load case, and the lower cover is designed to be adequate in compression at the 2.5 negative g load case. The respective covers are satisfactory in tension, determined by subsequent inspection, since the material is fully effective up to ultimate load. A series of mathematical expressions can be derived which relate the following parameters together: Skin material strength and stiffness; skin thickness; stringer material strength and stiffness; stringer area and moment of inertia; stringer f i x i t y ; rib material stiffness: rib thickness: rib spacing and end fixity. For a given set of loads, dimensions (span, chord and height), El and GJ stiffness requirements, and m i n i m u m gage requirements, a series of computations may be performed which tries numerous possibilities between skin thickness, number of stringers, area and moment of ine r t i a , rib spacing, rib thickness, then c o m p u t i n g a weight for that combination. A change is made and a new weight is found. This process, while not elegant, is continued until the combination which gives minimum weight is found. This process is repeated for each rib station from root to tip. The results are then collected together and a weight is computed based on the total volume of material used. By hand, this technique would be very tedious, but with high speed computers, mere seconds. Some of the limitations of the algorithm include 1) the user supplies the m a t e r i a l properties based on layup o r i e n t a t i o n , 2) each station
is optimized independent of the previous station and 3) stringer configuration must be determined independently. All this means is that we try different material layup combinations for skin, stringer and ribs
w h i l e we try to m i n i m i z e overall
weight. We search to find the opti-
mum layup orientations which give the lowest weight, then we "massage" the results slightly to make the results physically realizable. At one rib station, the results may indicate a stringer at 1.5 inch intervals and the next station might in-
66 JULY 1976
dicate a 1.6 inch interval. For constant chord wings, it is desirable to maintain some consistency of stringer spacing throughout the span, changing areas, moment of inertias and terminating some as we proceed outboard. A number of test computations were made, including an all aluminum version and an all fiber-glass version. Three different skin layup combinations were tried (see Figure 3) with unidirectional stringers and ±45 l( layup ribs. The results of this preliminary optimization search are shown in Table 8, Concepts I and II. CONCEPT III — HONEYCOMB SANDWICH PANELS
One of the most efficient design concepts available for flexural beam applications is honeycomb stabilized skin. By utilizing the shear carrying core to support thin facesheets, wing box covers of comparatively low weight can be achieved. The technique we employed here was to size and optimize each cover for the average load intensity. In this way, core thickness and facesheet thickness could be reduced (and weight saved) as the loads decreased going outboard. The procedure of finding the right combination of skin and core thicknesses is given in MIL-HDBK-23. The process is to iterate on facesheet stiffness, facesheet thickness and core depth, all as a function of layup orientation, using the design charts provided. Observations made during the calculations include: Panel stability against buckling was insensitive to core density, the skin thickness plays the predominant role; the weight tradeoff of panel size versus more ribs was not thoroughly explored, hut favored the configuration shown in Table 8. Foam core was not considered since a 1.8 Ib. density honeycomb was shown to be adequate and no appreciable weight savings would result. Besides, honeycomb is structurally superior for sandwich panels. For homebuilt application, foam core would be more adaptable to our fabrication methods. Of the configurations studied, the concept shown was found to be of minimum weight using fiber-glass composite materials. CONCEPT IV — FULL DEPTH FOAM
This configuration was of interest
to us since this is the design employed by Burt Rutan in the VariEze. Our optimization procedure was
one of hand calculation using the results of the other concepts. Much
to our surprise, this configuration is
heaviest. This is due to the large
volume of foam, light as it is. This design required the least amount of
fiber-glass, but the weight savings were lost due to large mass of 2 Ib. per cubic ft. full depth foam. Of the approximately 12 Ibs. total weight, half is in the foam. Results of our analysis for this type of design are shown also in Table 8. SUMMARY
In this first article, we've given you a background on the composite materials and a summary of a predesign effort using an existing design as a baseline. If the data shown was a little too technical, we regret any efforts to "snow you"; that was not our purpose. Composite materials for aircraft structures are here and now, and have earned a favorable reputation for their continued usage. Just as the introduction of a l u m i n u m required a new kind of t h i n k i n g while wood and fabric aircraft were being made, the same holds true now for composites. SUBSEQUENT ARTICLE
For the next article, we plan to choose one of the design concepts and perform further optimization, culminating with a final configuration, the results of which w i l l be passed on to you in the form of analysis and engineering drawings. ACKNOWLEDGMENT
We thank Dr. E. E. Sechler, Professor of A e r o n a u t i c s and Astronautics, California Institute of Technology, Pasadena, California for rev i e w i n g t h e draft a n d s u p p l y i n g meaningful comments.
ABOUT THE AUTHORS
Hans D. Neubert has been in the aerospace industry for 10 years, currently with TRW Defense and Space System Group, responsible for the utilization of composite materials on spacecraft structures. He was formerly employed with Convair in San Diego. He holds a Bachelor's degree in Mechanical Engineering and a Master's degree in Aerospace Engineering. He has 9 years experience w i t h composite materials, design
and analysis. Ralph W. Kiger has been in the aerospace industry for 9 years, currently, with Northrop Aircraft Division, responsible for the design and analysis dealing with maintenance and repair of composite aircraft structures. He was also formerly employed with Convair in San Diego. He holds
a Bachelor's degree in Aeronautical Engineering and a Master's degree in Aerospace Engineering. He has
8 years experience with composite
materials, design and analysis.
Modern Composite Aircraft Technology By Hans D. Neubert 6051 Prado St. Anaheim, IA 92807 and Ralph W. Kiger 10201 Wembley Circle Westminster, CA 92683 (All Rights Reserved by the Authors) PART II. OPTIMIZED OUTER WING PANEL DESIGN
N THE PREVIOUS article (Part I, July SA), we preIsented to you a brief overview of the materials technology of interest to the composite airplane designer/ builder, and a preliminary design summary of the T-18 outer wing panel in four different configurations. A brief review of the last article is prudent, hopefully to
answer some questions that may have been raised. Also, some limitation on the size of the first installment dictated restraint on our part to the quantity of material that could have been presented. In this second article, we plan to discuss the origin of the loads imposed on the wing during flight, resolve those loads into the wing structural elements, show how they are treated and proceed to an optimization of
the entire assembly. Since highly loaded joints are the nemesis of composite structural analysis, we are receding slightly from the initial cavalier approach taken of proceeding directly to fabrication and plan instead to first test a representative wing spar/sandwich cover joint section before committing the design to final form. The proposed
joint concept is shown in the later portions of this article. MATERIALS REVIEW
One recognizes that a great variety of different fiberglass weave styles are available and we normally would use 3 to 4 of them in combination. In doing so, a better strength-weight optimization will be realized. To illustrate, one would normally use 2 styles of unidirectional material having equivalent composite tensile and compressive strengths, but with different cured thicknesses. This permits one to taper out skin thickness more efficiently from root to tip. We would also normally use 2 styles of bidirectional weave having equal strengths in both directions for the same reasons. Our choice of Style 143 glass cloth was purely judgmental in order to reduce the complexity of the analysis. The E-glass composition is used herein since S glass (although more desirable from a strength point of view)
is not as readily available, and its cost penalty does not warrant its usage at this time. Dynel, Union Carbide's trade name, which is the same as Verel, Eastman Kodak's trade name, is a modacrylic fiber having '/.-ta of the linear elastic strength of fiber-glass, and is therefore not worthy of our consideration as a material for primary structural applica-
tions. It may, however, find favor and utilization for
secondary structural applications (i.e., cowls, fairings, covers, wheel pants, etc.). MATERIALS CHARACTERIZATION REVIEW
In the design of a laminate required to sustain a specified set of loads, the optimum combination of individual layers for minimum weight normally results in a laminate consisting of longitudinal plies and angle plies. Normally we design to "symmetrically balanced about the midplane laminate" configurations in order
to uncouple the membrane extensional and plate bending responses to the loads. That means "for every layer at a + H angle relative to the span we have a - H layer, and for every layer above the middle of the laminate thickness (mid-plane) there is one below". By constraining ourselves to a single material (i.e., Style 143 unidirectional cloth), the m i n i m u m number of plies is 2, but only if they run spanwise. Inboard, when layers are required at some angle to increase shear strength, then
the m i n i m u m number of layers is 4 [ + 0/-M-0/ + 0}. If angle and longitudinal plies are both used, the m i n i m u m is 6. If the "symmetric balanced laminate" rule is not
followed, the laminate is no longer specially orthotropic, but anisotropic. and becomes substantially more difficult to analyze since it is coupled. The utilization of anisotropic coupled laminates is only now being fully explored as a technique to increase high speed flutter, and also to control aerodynamic twist. A coupled laminate w i l l twist when bent, which leads to the possibility of a wing design which twists as a function of the bending moment, reducing the relative angle of attack as a result of bending loads (as might occur in an accelerated stall). We are, however, using the sandwich core as the mid-plane in the sandwich panel configuration: thus the laminate is still uncoupled: |0/+fl/-tf/Core/-#/ + #/0]. In our plan of constructing a wing panel for structural tests, we intend to use the new high impact/ superior toughness resin system, even though the strength characteristics presented on the 143 glass were based on MIL-HDBK-17 values for a typical bisphenol-A resin. Specimens to obtain test values of 143 with the improved resin are in work, but the complete results are not available at this time. Our approach is expected to be conservative, based on initial test data. PRELIMINARY DESIGN EXAMPLE REVIEW
After reviewing the results of Table 8 of the previous article (preliminary optimization for the 4 configurations), we expected you to conclude that using fiberglass composite in aircraft construction does not necessarily yield a lower weight design than contemporary a l u m i n u m construction. The conclusion that fiber-glass designs are. at best, equal to a l u m i n u m designs on the basis of design weight is a result familiar to us. Having gone through numerous aircraft designs of widely differing weight, size, speed and mission as a result of our occupation, similar comparative results have been SPORT AVIATION 55
obtained. The conclusion that leads one to consider composite airframes is not a lower weight (advanced composite materials excepted), but less tooling, less fabrication time, increased serviceability, better maintainability and repair, no corrosion and, normally, lower overall cost. Of the 4 configurations, the skin-stringer designs were optimized by using automated procedures. For aluminum, material properties from MIL-HDBK-5B were directly used; with the fiber-glass version, the properties of the skins, stringers and ribs, which are a function of the layup orientation, were determined and optimized independently, and then plugged into the minimum weight algorithm. Therefore, the overall minimum weight for the skin-stringer Concepts I and II is a combination of layup orientation optimization (which is an input quantity to the box beam analysis), and the box beam optimization, which is a tradeoff of the skin, stringers, spar and rib parameters. The same optimization problem of material and the structural configuration exists for the sandwich and full depth core designs. These were done by hand analysis, but are within a reasonable minimum weight tolerance for predesign tradeoff purposes. As a result of this work, we concluded that the sandwich cover design (Concept I I I ) is the most logical for further consideration. Among the payoffs is that a wet wing design is feasible, and that fabrication can be accomplished in a manner which gives a smooth exterior surface.
tact, suffer no irreversible damage and show no signs of permanent set. However, repeated application of limit load (usually 1,000 times or more) may lead to damage or failure due to low cycle fatigue characteristics of most materials. A factor of safety is normally introduced to specify the ultimate load factor. The value of 1.5 is traditional for aircraft structures (1.25 for unmanned spacecraft), and comes from the fact that the ratio of ultimate tensile strength to yield strength of typical aluminum alloys is 1.5. Basically, the factor of safety was established by the materials capability, and not some mysterious revelation. In the design of fittings, castings, and joints, higher factors of safety are used. For certificated composite aircraft, the value of the margin of safety for the structure is still a negotiable number. During the analysis of any particular component, a margin of safety value is computed. Given the stress (or load) a structural element is expected to sustain, a value is computed which compares the actual stress to the material's permitted stress (limit applied stress due to limit load compared to the material's yield stress) minus 1.0; thus, M.S. = (Allowable Stress/Applied Stress) - 1. Any value greater than 0 indicates that structural element has reserve capacity for additional load, provided other modes of failure do not occur first.
APPLIED AND DESIGN LOADS
For convenience and uniformity to the aerodynamicist, as well as the structural analyst, the aerodynamic force coefficients are usually given in non-dimensional form resolved about the aerodynamic center (the location on the wing chord about which the moment of the air forces are independent of the lift forces). These coefficients are C^ (coefficient of lift), CQ (coefficient of drag) and Cj^ (coefficient of moment). They are a result of the pressure distribution of the air acting on the airfoil (see Figure 1). It is not possible to predict with exact certainty the worst case load conditions which will be imposed upon the airplane structure. Our knowledge of aerodynamics, together with past experience, does enable us to limit the scope of the necessary investigations to a number of standard conditions. This leads one to construct the customary V-n diagram familiar to many homebuilders, where the gust load factor is combined with the various design speeds to yield a structural flight envelope. For the T-18 aircraft, limit load factor "n" is equal
The airplane will, during its lifetime, be subjected to an infinite number of load combinations. These loads are of two general types — aerodynamic flight loads and those which are a result of landing conditions. For wing structure, landing loads do not normally dictate the design. Flight loads arise from intentional maneuvers by the pilot or from "sharp-edged" gusts. During unaccelerated level flight, the lifting force on the airplane equals its weight. This results in a load factor n = 1 + (a/g) = 1. This load factor is defined as the multiplication factor by which the level flight aerodynamic forces are multiplied to obtain the equivalent static effect of dynamic forces acting during acceleration of the airplane. The value of the load factor is initially chosen by the aircraft designer based on the type of usage he expects the aircraft to encounter in service. The limit load factor is the maximum load factor which is to be expected in any normal maneuver. At the limit load, the airplane is expected to remain in-
FORCES ACTING ON WING
Air Pressure Loods
WING LIFT = C Sq
negative pressure
TWISTING MOMENT = C
Sqe
.DUCED DRAG = CdSq
Aerodynamic Center positive pressure
Dynamic Pressure ~ q = -y P V 5
= wing or«a
c
= chord
FIGURE 1 - AERODYNAMIC WING LOADS 56 SEPTEMBER 1976
to 5 at maximum gross weight. For these articles, we have assumed that the aerodynamic loads which result
cover also reacts to in-plane shear forces. Lower Cover — The lower cover is similar to the
the design maneuvering speed. The values cited in Table 8 (Part I) are derived as follows (see Figure 2, also, see page 62).
and the in-plane forces are tension. The in-plane shear forces are the same. Ribs — The ribs serve a number of functions. They
in wing bending and twisting occur simultaneously at
One must include the tip portion of the wing in determining the moment, shear and torque. By adding the 12inch wing tip to the structural box length and applying the equations, one will obtain the maximum values cited at the fittings of the outboard wing panel.
upper cover, except that air pressure loads are positive
act as transverse panel stiffeners for the compressiveloaded upper cover; they resist the secondary forces of wing bending (without ribs or spars, the upper and lower covers would try to come together as a result of the bending loads): they redistribute loads from aileron/flaps;
Aerodynamic Center
w = 28 Ib/in
t 1 ft t I ,1! I ! Wing Box
Wing Tip 12.0'
47.4"•
X -0
X = Any wing Station
FIGURE 2 - GEOMETRY OF DIMENSIONS FOR LOAD CALCULATIONS
RESOLUTION OF FORCES ON WING BOX STRUCTURAL ELEMENTS
These aerodynamic loads, resolved into equivalent externally applied moment, shear and torque, must now be reacted by the various structural elements of the wing structure. Refer to Figure 3, as well as the next few paragraphs to understand this transformation. Upper Cover — For wings in positive "n" bending,
the upper cover is subjected to a negative aerodynamic pressure gradient acting normal to the surface and, if
the cover is structural, must resist compressive in-plane loads. Therefore, the upper cover distributes air loads
to the substructure (ribs and spars), and reacts to compressive forces which are a result of wing bending due to the aerodynamic pressure of the air. In a clothcovered aircraft wing, the cover only distributes the air pressure loads to the substructure, but plays no role in resisting bending forces. Due to twisting forces as a result of aerodynamic shape of the wing section and concentrated loads at the fittings as a result of aileron deflections, the upper
and they react to and redistribute air loads.
Spars — Spars, together with the wing covers, react to bending loads, resist the vertical shear forces, and resist the torsion forces in both differential bending as well as in-plane shear.
DESIGN CONSIDERATIONS FOR STRUCTURAL ELEMENTS Since the design concept chosen for subsequent analy-
sis is the sandwich panel configuration, specific con-
siderations affecting the successful design will be addressed next. Structural sandwich is a layered construction formed by bonding two thin facings to a thick core. It is a type of stressed-skin construction in which the facings resist nearly all of the applied edgewise (inplane) loads and flatwise bending moments. The thinspaced facings provide nearly all of the bending rigidity to the construction. The core spaces the facings and transmits shear between them so that they are effective about a common neutral axis. The core also provides most of the shear rigidity of the sandwich construction. The sandwich is analogous to an I-beam, in which the flanges carry direct compression and tension loads, as do the SPORT AVIATION 57
_
Compress!ve Load N
Moment
"
Shear Load N
xy
due to Twisting Moment
xy
_ Twisting Moment 2 ( enclosed area)
Bending Load due to Moment
Secondary compressive load
ng flexure
Shear Load due to Twisting Moment
and aileron/flap reaction load
xy Twisting Moment
Note: Air pressure loads not shown for clarity.
FIGURE 3 LOADS ON WING BOX STRUCTURAL ELEMENTS
sandwich faces, and the web carries shear loads, as does the sandwich core. As a consequence of employing a lightweight core, design methods account for core shear deformation because of the low effective shear modulus of the core. The main difference in design procedures for sandwich structural elements, as compared to design procedures for homogeneous material, is the inclusion
of the effects of core shear properties on deflection,
buckling and stress for the sandwich. Because thin facings can be used to carry loads in a sandwich, prevention of local failure under edgewise direct or flatwise bending loads is necessary, just as prevention of local crippling of stringers is necessary in the design of sheet-stringer construction. Modes of
failure that may occur in the sandwich under edge load are shown in Figure 4. Shear crimping failure (Figure
58 SEPTEMBER 1976
4B) appears at first to be a local mode of failure, but is actually a form of general overall buckling in which the wavelength of the buckles is very small because of
low core shear modulus. The crimping of the sandwich
occurs suddenly and usually causes the core to fail in shear at the crimp, and it also may cause shear failure in the bond between the facing and the core. Refer to Table I for sandwich panel design requirements.
Crimping may also occur in cases where the overall buckle begins to appear, and then the crimp occurs suddenly because of severe local shear stresses at the ends
of the overall buckle. As soon as the crimp appears, the overall
buckle may
disappear. Therefore,
although
examination of the failed sandwich indicates crimping or shear instability, failure may have begun by overall buckling that finally caused crimping.
TABU
If the core is of cellular (honeycomb) or corrugated material, it is possible for the facings to buckle or dim-
1 DESIGN REQUIREMENTS FOR SANDWICH PANELS
ple into the spaces between core walls or corrugations, as shown in Figure 4C. Dimpling may be severe enough
so that permanent dimples remain after removal of load, and the amplitude of the dimples may be large
enough to cause the dimples to grow across the core cell walls and result in a w r i n k l i n g of the facings. Wrinkling, as shown in Figure 4D, may occur if a sandwich facing subjected to edgewise compression
The f.icinps shall be thick enough Co withstand the tensile, comprcssivc. and shear stiesscs induced by the dcstpi Ui.id.
buckles as a plate on an elastic foundation. The facing may buckle inward or outward, depending on the flatwise compressive strength of the core relative to the flatwise tensile strength of the bond between the facing and core. If the bond between facing and core is
The tore sh:Jl have sufficient strength to withstand the shear stresses induced by the design loads.
strong, facings can wrinkle and cause tension failure in
«*
the core. Thus, the w r i n k l i n g load depends upon the
The coie shall be thick cnoui'h and have sufficient shear modulus to prevent overall buckling of Ihe sandwich under load.
elasticity and strength of the foundation system; namely,
the core and the bond between facing and core. Since the facing is never perfectly flat, the wrinkling load will also depend upon the initial eccentricity of the facing
Younp's modulus of Ihc core and Ihc comprcssive strength of Ihc facing shall be svflxicnt to prevent wrinkling of the faces under the design load.
5
The core cells shall be small enough to prevent interccll buckling of the f.icinf.s under design loud.
O
The core shall have sufficient comprcssivc strength to resist crushing by design lojtls acting normal to the panel facings or by compressivc
or original waviness. The local modes of failure may occur in sandwich
panels under edgewise loads or normal loads. In addition to overall buckling and local modes of failure, sandwich is designed so that facings do not fail in tension, compression, shear, or combined stresses due to edgewise loads or normal loads, and cores and bonds do not fail in shear, flatwise tension or flatwise compression due to normal loads, as shown in items E through H of Figure 4. ANALYSIS METHODS AND FINAL OPTIMIZATION
stresses induced through flexure
Up to this point, we have summarized how the air forces on the wing surface are resolved into equivalent /
forces about the airfoil aerodynamic center, how those forces about the aerodynamic center are resolved into applied loads on the structural elements which comprise wing structures, and since we plan to design using
The sandwich structure shall have snOicient Rcxural and shear rigidity to prevent excessive delieetions under design load
A.
GENERAL BUCKLING
FIGURE 4 POSSIBLE FAILURE MODES FOR SANDWICH PANELS F.
D.
INTRACELL BUCKLING (DIMPLING)
Caused by insufficient par,tl thickness or insufficient
core shear rigidity
B.
TRANSVERSE SHEAR FAILURE
HONEYCOMB COIE
FACES BUCKIE INIO CO>E
Caused by insufficient core shear strength or panel
thickness
SHEAR CRIMPING
Applicable to ccllnl.ir cores only Occurs will) very thin facings and.l rrc core cells This cftccl may cause failure by prop.icatmg across adjacent cells thus inducing face wrinkling.
G.
FLEXURAL CRUSHING OF CORE
Sometimes occurs following, and as a consequence
of. general buckling Caused by low core shear modulus, or low adhesive shear strength.
C.
E. FACING FAILURE
FACE WRINKLING
Cauied by insufficient core flai*i« compresuve strength or excessive beam deflection.
H. LOCAL CRUSHING OF CORE
ADHESIVE BOND FAILUffE
TENSILE FAILURE IN FACING —A COtE COMPRESSION FAILURE
Initial failure may occur in either compression or
tension face. Caused by insufficient panel thickncii, facing thickness, or facing strength Caused by low core compression strength
Facmp buckle* a* a "plate on an clastic foundation." It may buckle inward or outward, depending on relative strength* of core in compression and adru sivc m fhtwiv: tension
SPORT AVIATION 59
* OPTIMJM CONFIRMATION LAYOUT
.
.1) fttfiwvMww F.«»
\/ X 3 *** •« *««
composite sandwich structure, a summary of the design requirements and possible failure modes. The moment of truth is at hand; "what is the optimum design configuration which satisfies all requirements, considering the tradeoffs between number of spars and ribs,
panel size and thickness, number of layers of glass cloth, orientation, etc., etc.?"
To answer this question, we have taken two approaches. First, we assumed various combinations of ribs and spars and computed the following: number of layers per facesheet, optimum orientation, sandwich core thickness for three distinct types of core, panel weight per square foot, and margins of safety to satisfy strength, stiffness, stability, and wrinkling. Secondly, we reviewed the literature to obtain classical solutions for minimum weight structures. Finally, we converged the two approaches together to yield what we believe to be the optimum design, based on m i n i m u m weight, for the constraints to which we worked. The results you see in the remainder of this article are a summary of somewhat numerous calculations done by hand as well as by computer.
There are those that might argue that all of this work
is a little too grandiose, and that all of this complexity
of analysis does not support the trend towards simplicity. The argument in favor of simplicity is valid, and is one the builder must make. Our intent is to present information which will show the impact of the various approaches. Consider even the tradeoff between the various composite design concepts shown in Table 8 of the first article which
were reasonably well optimized. Between Concepts II, III, and IV there is a 2\r'< difference in weight. If that weight penalty is translated into the empty structural
weight of a typical 1500 pound two seat homebuilt, a weight penalty of 100 pounds results, which could have been converted into payload (assuming you have some place to put the extra payload). On the other extreme, a brute force analysis approach, together with simplicity, may lead to two conclusions: excessive weight
with satisfactory structural margins, or acceptable weight but with the potential for premature failure. Murphy's law always seems to prevail.
Two references by Gerard were found which deal SPORT AVIATION 61
with the question of minimum weight analysis. Following the work of Gerard in optimizing a multicell box beam on the basis of structural weight, the baseline configuration (Concept IV) has been finalized as a 2 bay (3 spar) with one midspan rib box beam reflecting the overall planform and depth of the T-18 Sport outer wing panel. In order to arrive at the lightest substructure/cover combination, a parametric analysis of the compression cover (upper cover) was performed, a computer program referred to as SPADE (Sandwich PAnel DEsign) was used. This analysis procedure (using the methods found in MIL-HDBK-23) generates the lightest weight panel sufficient to carry the specified set of load intensities (axial and shear) and outputs the necessary design information. A sample run is included as Figure 5 for your assessment. The run was made using the Style 143 glass cloth, 3 types of core (2 Ib. foam, 4 Ib. foam, and 1.8 Ib. 1/8 cell aluminum honeycomb) for the loads corresponding to the 3(f< span point from the tip. The results indicate that an unbalanced layup is sufficient to carry the loads; 1 + 30/-30/ + 30I each facesheet, and we have therefore added one additional ply of material at -30 degrees in order to satisfy the midplane symmetry requirement, giving I30/-30/-30/30] at that station. One of the results out of the analysis which we thought was interesting was that on the inboard panels where the load is highest, the program showed that four plies were sufficient to meet strength, stiffness, and stability requirements, but 3 more plies were required to satisfy against wrinkling, using a 4 pound per cubic foot density foam. Given only properties of 3 cores to work with initially, the program reverted to satisfy the wrinkling criteria by adding additional layers of glass. The alternate technique is to increase the density of the foam since higher densities yield better flatwise tensile strengths. Subsequent computer runs using a 6 pound density foam yielded a more optimum panel; using less glass but heavier foam, the net weight per square foot was lower, however. The final overall configuration as derived from all this analysis is shown in Figure 6. With three spars, this wing does not offer the capability for retrofit onto existing T-18's, and we would not necessarily like to see that anyway until after the actual part is fabricated and tested. As mentioned earlier, getting the uniformly distributed loads out of the upper and lower covers and through attachment fittings is a troublesome design problem. To that end, we intend to complete the design of the fitting and to fabricate an element of the wing box to verify the load redistribution capability before completing the drawings or committing to the fabrication of the entire outer wing assembly. The design of the fitting we intend to assess is shown in Figure 7.
HGUH 7 SPAl JOINT TEST SKOMiN
'•^\^^
CORE DAMAGE
INCLUSIONS
MISLOCATED DETAILS
FIBER DEFECTS
SONICS (TAPPING)
A
A
C
B
B
C
D
THERMOGRAPHY
B
B
B
C
B
A
C
PENETRANTS
A*
A*
B*
D
D
D
D
LEGEND -
A= B= C= D =
GOOD CAPABILITY FAIR CAPABILITY POOR CAPABILITY NO CAPABILITY
NOTE -
* = EDGES OR SURFACES ONLY
Class 1: Dents, scars, scratches or erosion in the facings, not accompanied by a puncture or a fracture. Class 2: Punctures or fractures in one facing only, possibly accompanied by damage to the core, but without damage to the opposite facing. Class 3: Holes or damage extending completely through the sandwich, affecting both facings and the core. Class 4: Extensive damage requiring replacement of a complete sandwich part or parts.
DAMAGE TO FACING
DAMAGED MATERIAL REMOVtD »Y SANDING
REPAIR TECHNIQUES
Class 1 Repairs: Scars, scratches, surface abrasion or rain erosion may be repaired as follows. Apply one or more coats of resin, catalyzed to room temperature, to the abraded surface (number of coats depends upon the severity of the abrasion). Small fractures may be filled with a putty made from room temperature-setting resin and short glass fibers. Over this coated surface, apply a sheet of cellophane that extends 2 or 3 inches beyond the painted surface. After it is taped in place, work out all air bubbles and excessive resin with the hand or a rubber squeegee. The resin can then be allowed to cure at room temperature or, if necessary, the cure can be hastened by the use of infrared bulbs or hot sandbags. Occasionally, on small parts, the whole part can be put in an oven set at 100°C (212°F) to hasten the cure. After the resin has been cured, the cellophane is removed, the excess resin is sanded off and the whole repaired area is lightly sanded preparatory to refmishing. If the surface abrasions, scratches or scars are deep enough to seriously affect the strength of the facing (usually to more than the first ply of fabric), they should be repaired in the following manner. Sand the damaged
REPLACED LAMINATIONS ,SANDED TO CONTOUR
FIGURE 5 STEPS IN REPAIR OF CLASS 1 DAMAGES TO SANDWICH FACINGS
DAMAGE TO FACING AND CORE
DAMAGED CORE REMOVED FACING SCARFED FOR PATCH
area either by hand or with a flexible disk sander to a
smooth contour, as shown in Figure 5. Sand to a distance of at least 100 times the depth of material removed. Coat the sanded area with one coat of room temperature-setting resin and apply pieces of glass fabric
WET-LAMINATED PATCH
soaked in resin to a resin content of about 50 percent.
Lay these pieces of fabric in place in the sanded depression, as shown in Figure 5. Tape cellophane in place
over the repair and work out excess resin. After the
resin has cured, the surface of the repair is sanded down to the original surface of the facing. Class 2 Repairs: Damages that extend completely through one facing of the sandwich and into the core
require removal of the damaged core and replacement
of the damaged facing in such a manner that normal
58 APRIL 1977
CORE REPLACED
FIGURE 6
STEPS IN THE REPAIR OF CLASS 2 DAMAGES
stresses can be carried over the area. Figure 6 shows
one method for accomplishing this type of repair. The damaged portion is carefully trimmed out to a circular or oval shape and the core removed completely to the opposite facing. Caution must be exercised not to damage the opposite facing or to start delaminations between that facing and the core around the damage. The damaged facing around the trimmed hole is then scarfed back carefully by using a flexible disk sander, a belt or rotating pad sander, or by hand to a distance of at least 100 times the facing thickness. This scarfing operation must be done accurately to a uniform taper and usually takes a little practice before acceptable scarfs are obtained. Contour lines produced by the individual plies of fabric in the sanding operation can be used to judge the accuracy of the scarfed surface. WARNING
The sanding operation on laminates reinforced with glass fabric gives off a fine dust that may cause skin irritation. In addition, breathing an excessive amount of this dust may be injurious. Therefore, precautions as to skin and respiration protection must be observed.
The details of the scarfed method are shown in Figure
7. They consist of trimming out the damaged portion
and proceeding as with Class 2 damage, except that the
opposite side of the sandwich is provided with a temporary mold or block to hold the core in place during the first step. After the first facing repair is cured completely, the mold and the shim (which temporarily replaced the facing on the opposite side) are removed and the repair is completed by repeating the procedure used in the first step. Typical steps in making this type of repair are shown in Figure 7. By using the stepped-joint method shown in Figure 8. the damaged portion is trimmed as before to a round or oval shape, or to a rectangular or square shape (prefer having rounded corners). The thicknesses of the individual plies in the facings are determined, for choice of replacement fabrics, from the portion removed; the total overlap of the stepped joint is computed from the number of plies in one facing, minus one, times IVz inches. This overall size is then marked on the sandwich. The marking can be done with cellophane tape or by lightly scratching the surface. The outer layer of fabric only is then cut with a sharp knife or a specially prepared cutter along these lines. CAUTION
A piece of replacement core material (or a suitable substitute) equal in thickness to the original core material is cut to fit snugly in the trimmed hole. The glass fabric laminations for the facing repair are then prepared, with the largest piece being cut to the exact shape of the outside of the scarfed area. The smallest piece is cut so that it overlaps the scarfed area by its proportionate amount, depending on the number of plies in the repair, and the intermediate pieces are cut to have equal overlaps. A convenient means of preparing these pieces is to brush-spread the resin on the pieces of fabric and sandwich the spread fabric between two sheets of colored cellophane. The pieces are then cut to shape without the usual fraying at the edges. The resin content of the fabric should be about 50 percent. When all of the pieces are ready for assembly, the opening from which the damaged core was removed is coated on all sides and bottom with room temperaturesetting resin. The piece of core that is to be inserted is likewise coated on all sides, including top and bottom surface, and inserted in the hole. The pieces of fabric are then laid in place by first removing the cellophane sheet from one side of the fabric, placing the exposed fabric in position on the repair and then removing the second cellophane sheet. The whole area is then covered by a piece of cellophane and carefully worked down to remove as much excess resin and air as possible. Light pressure is applied to the cellophane by means of sandbags, taping (if the repair is on a convex area) or a vacuum blanket, if facilities permit. After the repair has cured, it is lightly sanded to contour it to the original shape and it is then ready for refinishing.
An alternate method that may be used for Class 2 repairs is the "stepped-joint" method, described under Class 3 repairs. Class 3 Repairs: Damages that are completely through the sandwich may be repaired by two methods: (1) The scarf-joint method (similar to that described for Class 2 damage); and (2) the stepped-joint method described later. The scarfed method is normally used on small punctures up to 3 or 4 inches in maximum dimension, and in facings made of thin fabrics (which are difficult to peel). The stepped-joint method is often employed on larger repairs to facings composed of thick
fabrics.
Do not cut through more than one layer. If the layer of fabric underneath is scratched, the strength of the repair will suffer. Using a knife blade, the outer fabric layer can be lifted and carefully peeled away from the layer underneath until the entire sheet is removed. This leaves a clean-cut step round the area. The process is then repeated, with the cut being made at a distance of 1V6 inches inside the original step, as shown in Figure 8. Each consecutive layer of glass fabric lamination shall be removed in this manner except the last one (bonded to the core), which is exposed for an area approximately l'/2 inches wide around the trimmed hole. This surface is then lightly sanded. A piece of core material of identical thickness to that in the sandwich is prepared of the same material (or an approved substitute) and of a size to provide a snug fit in the trimmed hole. Glass fabric sheets of appropriate thickness are then cut slightly too large (approximately an inch or two oversize) for each of the steps in the repair. A mold and shim combination is now prepared for the opposite side of the sandwich to preserve the contour while the first facing is being repaired. After the mold and shim have been temporarily secured in place by clamping, propping or lashing, the damaged area is ready for the first step in rebuilding. The replacement core piece is coated with resin on all edges and the top surface only, leaving the bottom surface (next to the temporary shim) uncoated. It is then inserted in place above the temporary shim. The glass fabric sheets for repairing the facing are now impregnated with resin to a resin content of about 50 percent, and the smallest one is laid in place over the replaced core. It is then trimmed with scissors to the exact shape of the trimmed hole. After the trimmed portion has been removed, successive plies of glass fabric are laid in place
and trimmed, just as was done for the first ply. An extra layer of 112 cloth is then applied over the repair and trimmed so that it laps about Vz inch over the undamaged facing. The area is then covered with cellophane, and the excess resin and air are worked out as described earlier. Class 3 Repairs (Access Not Available to Both Facings): When access is not available to the inner facing SPORT AVIATION 59
HOLE THRU SANDWICH
HOLE THRU SANDWICH
HOLE TRIMMED, CORE REMOVED, AND FACING "STEP-PEELED"
HOLE TRIMMED, CORE REMOVED AND FACINGS SCARFED
WET-LAMINATED PATCH
TEMPORARY BLOCK OR MOLD
TEMPORARY BLOCK OR MOLD
TEMPORARY SHIM CORE REPLACED
s
x
TEMPORARY SHIM
CORE REPLACED
COMPLETE BY SCARFING AND WET-LAMINATING OPPOSITE FACING IN A SIMILAR MANNER
FIGURE 8 FIGURE 7
DETAILS OF STEPPED-JOINT METHOD OF REPAIR OF CLASS 3 DAMAGES
STEPS IN THE SCARFED TYPE OF REPAIR TO CLASS 3 DAMAGES
of a sandwich panel, the following procedures are suggested: 1) Remove the damaged facings and core in the form of an oval area, using the general methods described on preceding page. 2) Fabricate glass fabric backing plate and drill necessary holes in plate and sandwich panel (Figure 9) for self-tapping screws to be used in applying pressure. 3) Clean the surface of the backing plate and the area around the cutout on the back of the repair. 4) Apply adhesive to the backing plate and position in place on the inner facing of the sandwich part. 5) Fabricate replacement core patch and place in the core cavity over the backing plate. 6) Apply cellophane sheet, plywood caul and self-tapping screws, as shown in Step 1 of Figure 9. 7) After the adhesive has cured, the screws and caul can be removed and the screw holes filled with a viscous mixture of epoxy resin. 8) The upper overlap facing is then bonded in place using screw pressure. Heat for curing the adhesive can be applied to the outer bond line by the use of heat blankets, hot sandbags or heat lamps, but it is difficult to obtain heat on the inner bond line, and therefore adhesives should be
used there that cure adequately at room conditions. Voids and Delaminations: Foam sandwiches sometimes develop voids or delaminations at the interface between the facing and the foamed-in-place core, caused by a blow from a blunt object. These voids may be repaired by injecting a small quantity of room temperaturecatalyzed resin by means of a hypodermic needle and syringe. Two 1/32-inch holes are drilled in the facing
at diametrically opposite sides of the void area. The resin is injected into one hole until it starts to extrude
from the other. Slight pressure is then applied by means
of a hot sandbag over the void to remove excess resin,
and to insure contact between the facing and the core while the resin is curing.
60 APRIL 1977
ENVIRONMENTAL CONSIDERATIONS
Durability: An acceptable composite structure, in addition to possessing specified initial physical properties, must be relatively unaffected by continued exposure to conditions imposed by severe service. Obviously, with composite combinations involving new and untried materials, it is impossible to provide the assurance of adequate service tests, and reliance must therefore be placed on tests conducted after artificial aging under conditions thought to be typical of severe service. Artificial aging tests are time-consuming, and often do not include all the possible combinations of exposure conditions that might be encountered in actual service. Weathering: The service history of sandwich of reinforced plastic materials combined with simulated weather aging has developed data showing wide scatter, but analysis of the information showed that service life reduced strength more than did shelf life. After 3 years exposure at several outdoor sites, the salt air exposure appeared to be the most severe and produced a somewhat eroded surface on laminates, with exposure of some glass fibers on the surface. A summary of U.S. Air Force service experience showed sandwich with reinforced plastic facings had no basic weakness except for rapid rain erosion. A variety of solvent materials can be successfully resisted by epoxy resin matrix composites for exposures of 30 days, except exposure to paint removers, which degrade strength drastically.
The most serious and difficult service problem is that of water entry into the sandwich cores. Water accumulation can result in damage to the panel from freezing, which can rupture the core or panel itself. Fluid
entry can also contribute to dynamic unbalance and
weight gain in such parts as control surfaces. In several instances, foam core materials have failed due to vibration and have accumulated moisture, with a resultant loss of the bond. In production, or during fabrication and
finishing operations, moisture may enter if panels are
immersed or sprayed with water or other fluids. Also, panels with adhesive bonds with poor moisture resistance, or bonds of a porous nature and skins or edges of laminated plastic or other porous materials, allow fluids to enter by transmission directly through the material, though the quantity depends on the degree of porosity. Difficulty may result even with a fully sealed panel because of one or more of the following conditions: * High levels of sonic vibration cause sealant breakdown * Structural deflection * Differential pressure * Inadequate overlap in bonded joint * External accidental damage to panel * Inadequate design for sealing * Thermal stresses due to steam cleaning cause sealing or bonded joints to fail * Aging and deterioration of sealing materials * Inadequate material properties The techniques and materials used to provide effective sealing must take the above conditions into consideration. The quality of the adhesive bond — i.e., degree of porosity and adhesion qualities — has a direct bearing on the sealing characteristics. Polysulfide and filled epoxy compounds are the most frequently used materials for sealing joints between edge members, relief cutouts and tool or vent holes. For laminated facings, the use of void-free processing techniques and nylon or epoxy sealer coatings is recommended. Erosion and Impact Hazards: Rain and Sand — Water (rain) and sand particle impact subject a material to combined tensile, compressive and shear loads at rates significantly greater than in static mechanical property tests. It appears that, in general, those materials in a given class having highest hardness and tensile or yield strength are the most resistant to solid or liquid
particle erosion. In addition, material properties such as ductility, modulus, shear and fatigue strength enter into the erosion process of a material. The amount of erosion can be related to particle size, quantity and velocity, and depends on the response of the material to dynamic impact and the time of exposure. Conventional paint finishes can generally be selected or designed so as to provide the desired level of erosion protection for most composite structures (other than leading edges), such as control surfaces, fixed trailing edges, doors, etc. Greater care is required to provide erosion protection for leading edge structures, as defined in Figure 10. Materials such as neoprene or polyurethane have proven satisfactory in the protection of fiber-glass structures against erosion. Impact Hazards — Through both flight and ground servicing environments, aircraft components can suffer damage due to foreign object impact (e.g., hailstones, runaway debris, dropped tools, personnel contact, etc.). A synergism of design features and material systems is necessary to provide adequate resistance to such hazards and is best accomplished at component inception, as opposed to afterthought. Judicious selection of material type and form, lamina stacking order, thickness and properties can mitigate the consequences of many common hazards. Impact damage typically manifests itself through delaminations or disbonds up to the point at which fibers fracture and the laminate (or facing) is completely penetrated. It is preferable then to have materials with high elongation and high interlaminar shear strength to absorb the energy induced by the impactor S-glass fibers, for instance, have been shown to have better impact strength than E-glass due to their higher strain-to-failure, but are much costlier. Rubber-
FILLED SCREW HOLES IN COVER PLATE .(HOLES USED IN BONDING COVER PLATE) SCREW HOLES UNDER COVER PLATE THROUGH PANEL AND BACKING PLATE
'OUTER EDGE Of COVER AND BACKING PLATE
WOOD OR SELF-TAPPING METAL SCREWS PLYWOOD CAUL SHEET
STEP 1 RELEASE FILM
BACKING PLATE (REPLACEMENT FACE SHEET), AND REPLACEMENT CORE BEING BONDED / INTO PLACE / REPLACEMENT CORE STEP 2 ........... TAPERED COVER PLATE (REPLACEMENT FACE SHEET)
ADHESIVE SQUEEZE-OUT NON-RIGID FOAM CUSHION
fl^-^r.-j )' -,,/,"•"-
FILM
BEING BONDED INTO PLACE / /
( TAPERED COVER; \PLATE OVERLAP \ ______I——
125 T
COMPLETED REPAIR FILLED SCREW HOLES
SECTIONS THROUGH A - A SHOWING STEPS IN MAKING REPAIR FIGURE 9
METHOD OF REPAIRING CLASS 3 DAMAGE WHENl PANEL IS ACCESSIBLE FROM ONLY ONE SIDE
SPORT AVIATION 61
ZONE Of MAXIMUM EROSION TANGENT TO AIRFOIL , AIRFOIL CHORD PLANE
There may not be enough historical evidence, experiments or tests available to draw a clear cut conclusion to questions as "What happens to the structural integrity of my composite aircraft primary structure if ... ?"
Moisture Absorption and Ultraviolet Degradation:
AIR FLOW ASSUMED PARALLEL TO AIRFOIL CHORD PLANE
FIGURE
10
EROSION PROTECTION AREA
modified epoxy resins are tougher than conventional nitrile epoxies, phenolics or polyesters, but their choice should not sacrifice interlaminar strength. Woven fiberglass epoxy fabrics possess inherently better resistance to delamination from impact than do impregnated unidirectional tapes due to the mechanically interlocked nature of the weave. Multidirectional^ oriented laminates resist damage from impact better than unidirectionally oriented laminates, since energy dissipation tends to follow the path of the stiffer constituent (fiber) of the composite and, hence, the more fibers in more directions the better. It stands to reason that, for a given impactor shape and energy, that a thicker laminate will suffer less damage than a thinner one, and thereby retain more of its original strength (up to the point of penetrating damage). As general practice, the primary load-carrying layers of a laminate (usually designated as the 0° direction) should be placed within the laminate (not on the extreme surfaces) so as to prevent damage to them due to more mundane hazards such as abrasion and scratches. Core selection should be based first on structural strength (and cost or fabricability) considerations and then that selection modified, if necessary, to provide adequate damage resistance to the sandwich. The core of a sandwich panel will absorb energy from impact by compressing (crushing) and flexing. Thus, it is best to have a thick, but compressible, core that has sufficient shear strength to carry flight loads, yet not impose a severe weight penalty. Continuous cores (e.g., foam, balsa) help reduce facing damage more than cellular cores (e.g., hexagonal honeycomb), since the latter provides only intermittent support at the nodes (cell walls) and the facing experiences high local bending strains between the nodes which can cause fiber failure at lower energies. The addition of extra layers of material (sometimes called "sacrificial plies") at the surface enhance the durability of composite laminates. Such concepts as elastomeric films (polyethylene or mylar), aluminum foil, ballistic nylon and Kevlar fabric/epoxy have been employed to reduce he deleterious effects of erosion, abrasion, impact and moisture penetration. Complete impact protection is, of course, impossible. The designer of composite aircraft should, consequently, take precautions to design for inspectability, accessibility, repairability, replaceability and avoid minimum gage designs where durability is a consideration.
ENVIRONMENTAL RELATED DESIGN FACTORS Throughout the course of these four articles, we've alluded to a number of factors which relate the initial design to long term effects. These effects, summarized in Table 3, are important if the aircraft structure is to be utilitarian over a long service life and retain its market value. The conditions or effects listed are understood by the composite technological community at various levels of competence and agreement. In a number of cases, the level of agreement is minimal, primarily due to the fact that little or no reliable data base exists. 62 APRIL 1977
There are two phenomena which were described in Part I, July SPORT AVIATION. The absorption of moisture from the air into the resin is a process that can be computed from the one-dimensional diffusion equation. For thin laminates, this process takes days to weeks, depending on the relative humidity-temperature environment and the chemical nature of the cured resin. The effect is to reduce the glass-transition temperature of the resin, which in turn reduces the temperature at which the composite is a viable structural material. In general, it is a process which cannot be prevented and, at best, can only be retarded. Moisture absorption is considered in the design process by using an empirical strength "knock down" factor or by using strength allowables generated at the "hot-humid" test condition. Ultraviolet radiation serves to sever the chemical chain comprising epoxy resin systems. This process can be prevented by protecting the structural composite by surface gel coats, paint, barrier films, additives to the resin, etc. Since UV degradation is a long term (years) process, and since it may be readily avoided, it is not considered directly in the design process except for a note on the drawings. Abrasion/Impact Damage: Invariably, some time during the lifetime of the aircraft structure, damage will occur. This article deals with this condition (as well as the erosion problem) and offers suggestions for its prevention and repair. Fatigue Life: In comparison to metals, which have fatigue-safe stress levels of one-third their ultimate static strength, composites have fatigue-safe stress levels from two-thirds to one-fourth their ultimate static strengths. This wide range is fiber orientation-to-load direction dependent. A unidirectional laminate will have substantially differing ultimate static strength to fatigue runout strength (107 cycles) ratios when loaded in the fiber direction and in the direction transverse to the fibers. In multi-oriented laminates, the presence of internal laminate shear influences the fatigue strength. To date, there is no good reliable analytical model which successfully predicts composite fatigue strength, and therefore one must rely heavily on test data taken at the specimen and component level. Hot and Cold Environment: Epoxy resin, being a thermosetting plastic, loses its strength and stiffness characteristics with increasing temperature. If the aircraft structure is expected to sustain design loads at the high temperature extreme (130 F to 160 F), then the structural analyst/designer must check that condition. The reduction of room temperature strength allowables at 150 F for ambient cured epoxies may be as much as 25%. To some extent, this substantial reduction in strength may be reduced by post-curing the structure at the high temperature extreme (say 160 F). This may be done with heat lamps, a box or tent connected to the wife's clothes dryer, etc. Another approach is to minimize the heating due to solar radiation. This is
why all composite sailplanes are white. In any event, the degree to which the hot environment design condition influences the design process depends on the expected service environment. In cold environments, the strength of fiber-glass epoxy composite laminates increases somewhat. At -40 F, this increase in strength would not normally be greater than 10% over the room temperature values. At the
cold condition, differential expansion between the composite and surrounding structure may cause problems.
Fiber-glass composites have thermal expansions of about 6 microinches per degree F. Wood, steel, and titanium match the composite thermal expansion quite closely; e.g., wood, 2 to 7 microinches/F; steel, 6 to 9 microinches/ F; titanium, 5 to 6 microinches/F. Materials which do not match as closely include: aluminum, 13 microinches/ F; acrylics (Plexiglass/Swedcast), 30 to 50 microinches/ F; and Dynel/epoxy, 40 to 50 microinches/F (estimated). Materials having different expansivities may be structurally joined provided the engagement is not too great and the thermal extremes not too large, since the induced stresses between the two materials is a function of AF.
Creep/Relaxation: A composite material could exhibit an appreciable amount of creep, depending on the state of stress and the temperature. Viscoelastic flow in the matrix, as well as internal flaw formation and Table 3. CONDITION OR EFFECT
MOISTURE ABSORPTION BY RESIN
growth, are normally the main sources of creep. Creep, then, can be observed as a measurable change in dimension, deflection or rigging. Except for unidirectional specimens loaded parallel (in-plane) to the fibers, considerable nonlinearity has been observed for multioriented composites. Creep for multi-oriented composites is a short-term process, where measurable changes can be detected within minutes to hours, depending on the stress level. In cases where the load is cyclic or reversible, relaxation back to the original state or position is observed. In cases where the load produces very high constant stress, creep rupture may occur. That is, if the applied load produces sufficiently high enough stresses, the creep will be sufficient to cause failure. A timedeflection history of an end-loaded cantilever beam might be a) initial static tip deflection due to the end load at time = 0 hours; b) continued tip deflection due to creep at the inboard end between time = 0 plus hours and time = XX hours; and c) at time = XX plus hours, the beam fails at the inboard end. For a beam where the stress is 90^ of the ultimate static strength, failure
Summary of Environmental and Configuration Related Design Factors
CONSEQUENCE
PLASTIZATION OF
RESIN, REDUCTION OF STRENGTH ALLOWABLES
HOW OBSERVED
NOT OBSERVED
ULTRAVIOLET DEGRADATION OF RESIN
ABRASION/IMPACT (HANGER RASH)
REDUCTION OF FATIGUE
STRENGTH
REDUCTION OF STRENGTH BUILT INTO DESIGN ALLOWABLES (I.E. KNOCKDOWN FACTOR)
DUE TO DISRUPTION OF BOND BTN FIBER & RESIN CHALKING OF SURFACE;
IMBRITTLEMENT OF RESIN LEADING TO MICRO-CRACKING
COMMENTS
HOW MINIMIZED
NOT A SERIOUS PROBLEM WHEN TAKEN INTO CONSIDERATION FOR HOMEBUILT DESIGNS
RESISTANT MATERIAL
U-V DEGRADATION OF RESIN IS A LONG TERM PROCESS, BUT MUST
COAT, ALYLIC ENAMEL PAINT U-V BARRIER FILMS, WAX ETC
PROCESS
EXTERIOR SURFACES COATED WITH U-V (I.E. POLYESTER SURFACE
LOSS OF AERODYNAMIC PERFORMANCE, POSSIBLE
PITS, DEEP SCRATCHES, DINGS; RESIN APPEARS
IMPACT RESISTANT RESIN, THIN METALLIC OVERLAYS,
STRUCTURAL FAILURE INITIATION POINT
MILKY AT IMPACT POINT
(EG. .005 ALUM OR CRES FOIL BONDED TO LEADING
BE MINIMIZED SINCE IT IS A NON-REVERSIBLE
PROTECTION CONTRIBUTES TO LOWER LONG TERM MAINTENANCE
EDGES OR OTHER DAMAGE PRONE AREAS)
LONG TERM FATIGUE LIFE
USEFUL LIFE OF
STRUCTURE REDUCED
MICROCRACKING,
VARIATIONS OF RETURN
ON ACCOUSTIC TEST (COIN TAPING)
CONSIDERATION OF FATIGUE ASPECTS DURING DESIGN/ ANALYSIS PHASE
FATIGUE OF FG/EPOXY LAMINATES VERY GOOD. FATIGUE OF LAMINATE
TO-FOAM CORE BOND AND FOAM CORE IN SHEAR UNKNOWN
HOT AND COLD ENVIRONMENTAL EXTREMES (-40F TO 160F)
CREEP/RELAXATION
LOSS OF STRENGTH PROPERTIES WHEN HOT. POTENTIAL BOND FAILURES DUE TO MATERIALS THERMAL EXPANSION MISMATCH
VISCOELASTIC FLOW IN RESIN 1 INTERNAL FLAW
FORMATION & GROWTH
MICROCRACKING; RESIN CRAZING AT LOCATIONS WITH THERMAL MISMATCH IN MATERIALS
STRIVE TO UTILIZE MATERIALS HAVING
EXPANSIONS CLOSE TO LAMINATE EXPANSION
(EG. 4 130 STEEL, WOOD,
DON'T OVERSTRESS STRUCTURE WHILE HOT.
OBSERVE AREAS
OF POSSIBLE CONCERN DURING PRE-FLIGHT
AND TITANIUM MATCHES FG/EPOXY CLOSELY)
CHANGES IN DIMENSION, RIGGING, AND SHAPE
DESIGN STRUCTURE SO THAT NORMALIZED
OCTAHEDRAL SHEAR STRESS /^/T^OFALL
LAYERS IN LAMINATE IS 2000 PSI OR LESS FOR 1G CONDITION. STORE IN
HIGH LOADS PRODUCING HIGH STRESSES, CAUSES, CHANGES IN SHAPE
SINCE RESIN IS A VISCOELASTIC MATERIAL. PHENOMENA IS TEMPERATURE DEPENDENT
COOL ENVIRENMENT.
RESIN DAMAGE
DEGRADATION OF STRUCTURAL INTEGRITY
SOFTENING/DISOLVEMENT OF RESIN
DO NOT USE EPOXY PAINT REMOVERS, CHLORINATED OILS, S. STRONG ACIDS ON EPOXY SURFACES
HYDRAULIC FLUID, SOLVENTS (MEK, ACETONE, ETC), FUEL AND OTHER CHEMICAL AGENTS NORMALLY FOUND IN
VICINITY OF AEROSPACE STRUCTURES DO NOT NORMALLY CAUSE PROBLEMS RAIN/HAIL EROSION
LOSS OF AERODYNAMIC PERFORMANCE
PITTING OF SURFACES
SCALE EFFECT
REDUCTION IN STRUCTURAL MARGINS OF SAFETY
PREMATURE FAILURE DURING LOAD TESTS
(SEE ABRASION/IMPACT)
TESTING IS NECESSARY AT (1) MATERIAL, (2) STRUCTURAL ELEMENT, (3) SUBCOMPONENT, AND (4) COMPONENT LEVELS.
BONDED JOINTS HAVE DEMONSTRATED SIGNIFICANT SCALE
EFFECTS. SCALE
EFFECT DIFFERS FOR
FATIGUE i STATIC
STRENGTH S. DIFFERS
WITH CONFIGURATION LAMINATE DISCONTINUITIES
STRESS CONCENTRATION
INITIATION OF DAMAGE
AVOID HOLES i CUTOUTS IF POSSIBLE. REINFORCE WITH ADDITIONAL MATERIAL (DOILIES)
FRACTURE BEHAVIOR CONTAINING DISCONTINUITY DEPENDS 1. ANISOTROPY 2. EFFECT OF BOUNDARIES 3. SIZE 4. ORIENTATION RELATIVE TO LAMINATE.
SPORT AVIATION 63
might occur in minutes to hours, depending on the material, orientation, environment, prior fatigue history, etc. Aircraft are not normally subjected to that type of stress level but, if they were, the load time would usually occur over a few seconds, not minutes or hours. But the combined effects of moisture, temperature, fatigue and creep are not understood at this time, and therefore the design process shall be intentionally conservative until these effects can be properly treated and judicial recommendations offered. Resin Damage: Typical liquid agents found at airports, such as aircraft fuels, hydraulic fluid, cleaning agents, etc. do not affect epoxy resins. However, strong acids (sulfuric nitric, aqua regea and others), chlorinated oils (methylene chloride) and epoxy paint removers can and will seriously damage the resin. These materials should not be used on resin matrix composites. If an aircraft requires repainting, then the old paint must be removed and stripped by mechanical means, such as sanding. Rain/Hail Erosion: Erosion of aerodynamic stagnation points can occur due to rain and/or hail impact. Since the impact energy is a function of velocity squared, the higher speed aircraft would be expected to receive severe damage. If the primary structure of the wing consisted of only the "D box", or if torsional integrity only depended on the exterior skin, then protection of the leading edge becomes paramount. This protection could be in the form of additional sacrificial layers of material, thin metallic overlays and/or neoprene boots. Electromagnetic Compatibility: When fuel is pumped from storage into an aircraft, static change can be either accumulated in, or dissipated from, the fuel as it moves through various elements in the fuel delivery system. Typically, charging occurs near surfaces where ionic materials provide an electrical double layer. The flowing fuel sweeps away charged ions of one polarity, leaving the opposite charges to flow to ground. Then, the charge in the fuel dissipates as the fuel is exposed to electrical paths to ground, and the charged ions combine with available opposite charges. Obviously, charge generation and charge dissipation occur simultaneously throughout the entire fuel delivery system. Local system characteristics determine if the net effect is to charge or discharge the fuel. Charge generation is a function of the area of the fuel-to-surface interface, properties of the fuel and surface, and the relative velocity of the fuel past the surface (flow rate). Charge generation can be either positive or negative, and may change from one part of the system to the next. Charge dissipation is principally a function of the electrical conductivity of the fuel (one conductivity unit = 1 CU = 1 picosemen/cm) and time. If the conductivity of the fuel is high, charge dissipation (or neutralization) occurs almost instantaneously and very little charged fuel may be found even a short distance from the chargegenerating surface. If the conductivity of the fuel is low, the fuel itself acts as an insulator and the charge remains for a considerable length of time.
Fuel with a high charging tendency, low electrical conductivity and processed at high flow rates can easily deliver enough electrical energy to an aircraft fuel tank
to produce incendiary sparking. The method by which that energy is handled as it enters the tank determines if an incendiary spark will occur. Arcing will not occur if the energy is dissipated uniformly at the rate it enters the tank. Arcing is likely if the energy is allowed to accumulate. If the accumulation of energy occurs in a conductive material that can give up significant amounts of energy instantaneously, the spark can contain a high enough concentration of energy to be incendiary. 64 APRIL 1977
The electrical conductivity of the tank must be considered in evaluating the ability of a particular design to safely dissipate expected levels of charge energy. Comparing idealized conductive and nonconductive tanks helps one understand the impact of tank conductivity on the dissipation of incoming charge energy, and to separate this effect from that of detail tank design. If the tank material is now considered perfectly insulating, the idealized spark conditions for the nonconductive tank may be deduced and several conclusions become apparent. Charged fuel in the idealized nonconductive tank has no path to the ground; hence, the fuel cannot discharge either in a controlled manner or by a spark. The tank/fuel combination is analogous to a charged capacitor. A path to ground must be available for sparking to occur. In the nontrivial case of an actual composite tank, ground paths of this type are probably common. They occur in the form of detail tank design such as systems, plumbing or splices to conductive structure. The sparking situation is created by detail tank design. Remember, a low fuel conductivity is required for an arcing voltage to develop between the fuel surface and ground in the conductive tank. However, this low fuel conductivity inhibits quick energy release. If a spark occurs, it is generally of low energy. The perfectly nonconductive tank, on the other hand, does not discharge the fuel. An accumulation of charge in a high conductivity fuel is possible. Incendiary sparking could be possible if a path-to-ground becomes available as the tank fills. Note, however, that the possibility of unbonded conductors is more likely for the nonconductive tank. Metal is typically used for splice plates, clips, etc. These items would naturally be grounded if they were installed in a metal tank. Special precautions might be necessary to assure grounding in a nonconductive tank. Also, the nonconductive tank would maintain the full incoming fuel charge level, exposing any charge collector to a higher charge level than would probably occur in the identical situation with a conductive tank. Fuel charging characteristics, fuel conductivity, fuel delivery equipment and fuel delivery rates all impact the sparking tendency within a given tank. Assuming, however, that all tanks will be exposed to a wide variation of these parameters, the following conclusions concerning idealized conductive and nonconductive tanks may be drawn. 1. Sparking in all tanks is primarily a function of detail tank design. 2. Extensive use of insulating materials may increase the possibility of a detail design deficiency. Many conductive clips, splices, fasteners, etc. will be used, and a conductive charge collector will be present if they are not adequately bonded. 3. The idealized nonconductive tank maintains the incoming charge in the fuel and presents this higher charge level to any sparking situation present. This characteristic is more critical if highly charged conductive fuel is delivered to the tank.
Factors affecting the consideration of actual composite fuel tanks are: 1. The extent to which the composite tank approaches the idealized nonconductive situation, -TS measured by the energy dissipating capability. 2. The necessity of special design considerations to assure bonding of all conductive materials used within the tank. 3. The capability to deliver highly charged and highly conductive fuel to the tank. To put all this into perspective, one must recognize
that a fiber-glass fuel tank is nonconductive, that the
design of the tank is very important, and that the grounding clip normally attached to the exhaust stack during
fueling may not offer any safety refuge for situations of a high charged fuel, low conductance fuel being pumped into the tank. Additional study and testing is required to clarify the problem and to assert constructive design guidelines. Electromagnetic Vulnerability: The vulnerability
in observing this type of discharge. However, these strokes are considered to be less severe than the cloudto-ground strokes, so 200.000A can be assumed to be
fiber-glass construction, this is still a controversial subject, since an inadequate understanding of the problem exists. For the time being, flight of primary structure fiber-glass aircraft in regions of lightning activity is not suggested. A natural lightning discharge initiates from a charge center in a cloud and propagates as a step leader toward another charge center or toward earth. Aircraft have triggered discharges, but usually the stroke is merely diverted from its path through the vehicle. An intense voltage exists between the step leader and the aircraft before attachment, and this causes streamers to form on all external surfaces of the aircraft. When one of these streamers contact the step leader, the aircraft is raised to the discharge potential and streamers extend off the opposite extremities of the structure. This step leader continues until another charge center or earth is contacted. An ionized path then exists. Conduction of a large surge current begins and the ionization wave travels back up the stroke path. This wave, referred to as the "return stroke", consists of a high-current surge with a fast wave front. The stroke may be considered stationary with the aircraft moving relative to it. This sweeps the ionized channel along the aircraft and permits contact at points aft of the attachment. This phenomenon is called "swept stroke" and is frequently accompanied by a "restrike". The restrike is a moderately high current surge caused by other charge centers within a cloud discharging down the ionized path. Note that the discharge only passes through the aircraft and is neither the termination or initiation point. Even large aircraft cannot store enough energy to appear as a sufficient charge center for lightning discharges to occur. Damage caused by lightning attachment to the metal skin of an aircraft is caused by the metal's inability to carry the discharge current without heating to a melting or vaporization temperature. Aluminum has good thermal and electrical conductivity, so aircraft skins have seldom shown severe damage. Since composites are less conductive, vulnerability will be greater, even to lower currents. Attachment studies on aircraft models indicate certain areas of high probability stroke contact. Actual strike data confirms that surfaces on the vehicle can be grouped into "zones", according to probability of direct attachment. From this, the zones can be defined as follows: * Zone 1 — Surfaces of the aircraft for which there is a high probability of direct lightning stroke attachment. * Zone 2 — Surfaces of the aircraft for which there is a high probability of a stroke being swept aft from a Zone
extremities (nose cone, wing/empennage tips, fairings, etc.). For an all fiber-glass composite aircraft, intuition might suggest that a nonconducting structure would survive, just as insulators are used on high voltage power lines. Not so; the total power (voltage times current) of a nominal strike is higher, the lightning strike phenomena is a dynamic (propagation) occurrence, and the characteristics of the material change in the presence of an ionized plasma. Severe damage has been reported (in laboratory tests) between metallic fasteners and the composite material, as might be in a typical structural joint. Fiber-glass sailplane damage (or nondamage) has not been adequately documented and reported, so as to lead one to prophesize blanket safety coverage for powered aircraft in all situations.
of composites to lightning strike is discussed here. For
1 attachment point. * Zone 3 — Areas other than those covered by Zones 1 and 2 which have a low probability of direct or swept stroke attachments. Data on the current amplitudes of a lightning strike is limited. Estimates have placed 200,000 amps as the realistic maximum for cloud-to-ground strokes, with this magnitude occurring only about 0.5% of the time. The 30,OOOA stroke is an average value. Cloud-to-cloud amplitudes are relatively unknown due to the difficulty
the worst case.
To our knowledge, the problems and test data associated with lightning strikes have been for aircraft
with aluminum construction and fiber-glass or graphite
SUMMARY
In this part of the series, we've attempted to surface a number of not frequently discussed issues on the use of composites for aircraft, as well as common techniques for their repair should damage occur. For the builders, as well as the designers, we hope this part of the series was informative. SUBSEQUENT ARTICLE
Fabrication of the outer wing static test article is proceeding on schedule. The final installment will be on the fabrication methods and test results of that component. REFERENCES
We've arranged our references in a bibliographical form for your convenience. Materials 1. Ferro Fiberglass Data Sheet, "Ferro S Glass Filament Winding Roving Prices", 7 January 1971. 2. Technical Bulletin, "Ferro S-1014 Ultra-High Strength Glass Fiber Products", FG-101. 3. Technical Bulletin, "Ferro S-1014 Ultra-High Strength Zero Twist Fabrics (with Data Sheets)", FG-102. 4. Owens-Corning Fiberglass Corporation Publication No. 5-ASP-6869, "We've been where you've been, we're going where j'ou're going". 5. 3M Data Sheet SP-RNB(1031)BE, "Scotchply Reinforced Plastics". 6. Owens-Corning Fiberglass Corporation Publication No. 5-ASP-6870, "Comparative Data; E, S and S2 Glass". 7. Burlington Glass Fabrics Company, "Industrial Woven Glass Fabrics (A Guide to Specifications and Prices)". 8. Owens-Corning Fiberglass Corporation Publication No. l-PL-6305, "An Introduction to FiberglassReinforced Plastics/Composites", March 1974. 9. Owens-Corning Fiberglass Corporation Publication No. 5-GT-5442-A, "Textile Fibers for Industry", March 1972. 10. J. H. Flickinger, "Elastomer Modified Epoxy Resins", 22nd SPI Technical Conference. 11. Dow Chemical Company Technical Bulletin, "High Peel Strength Liquid Epoxy Resin, SC-7575.02, for SPORT AVIATION 65
Improved Peel Strength, Tensile Shear Strength and Toughness in Epoxy Adhesives". 12. F. J. McGarry, "Crack Propagation in Fiber-Reinforced Plastic Composites", Fundamental Aspects of Fiber-Reinforced Plastic Composites, Wiley and Sons, 1968. 13. C. A. Harper (Editor-in-Chief), "Handbook of Plastics and Elastomers", McGraw-Hill, 1975. 14. Hexcel Aerospace Bulletin TSB120, "Mechanical Properties of Hexcel Honeycomb Properties", August 1971. 15. Hastings Plastics Company TNB 11-1-3, "Hathane 1680 Polyurethane Foam Data Sheet", 15 July 1973. 16. MIL-HDBK-17 and 17A, "Plastics for Aerospace Vehicles", Part 1, January 1971. 17. MIL-HDBK-5B, "Metallic Materials and Elements for Aerospace Vehicle Structures", September 1971. 18. Uniglass Industries, "Uniglass Fabrics Handbook", 1967. 19. "Modern Plastics Encyclopedia", Vol. 53, 10A, McGraw-Hill, 1976/1977. 20. Shell Chemical Company Technical Bulletins SC: 72-60, SC: 72-25, SC: 71-14, SC: 71-1, SC: 71-12. Design and Analysis 21. A. G. H. Dietz, "Composite Engineering Laminates", MIT Press, 1969.
22. MIL-HDBK-23, "Structural Sandwich Composites", December 1968. 23. P. Kuhn, "Stresses in Aircraft and Shell Structures", McGraw-Hill, 1956. 24. E. E. Sechler and L. G. Dunn, "Airplane Structural Analysis and Design", Wiley and Sons, 1942. 25. A. Gomza and P. Siede, NACA TN 1710, "Minimum Weight Design of Simply Supported Transversely Stiffened Plates Under Compression", Sept. 1948. 26. B. W. Rosen, NACA TN 3633, "Analysis of the Ultimate Strength and Optimum Proportion of Multiweb Wing Structures", March 1956. 27. E. E. Sechler, "Elasticity in Engineering", Wiley and Sons, 1952. 28. G. Gerard, "Minimum Weight Analysis of Compression Structures", NYU Press, 1956. 29. G. Gerard, "Optimum Number of Webs Required for a Multicell Box Under Bending", Journal of Aeronautical Sciences, January 1948, pp. 53-56. 30. W. C. Paulsen, "Finite Element Stress Analysis", Machine Design, Part 1, September 30, 1971, pp. 46-52; Part 2, October 14, 1971, pp. 146-150; Part 3, October 28, 1971, pp. 90-94. 31. K. H. Huebner, "Finite Element Method — Stress Analysis and Much More", Machine Design, January 10, 1974, pp. 92-99. 32. R. M. Jones, "Mechanics of Composite Materials", McGraw Hill, 1975. 33. D. H. Emero and L. Spunt, "Optimization of Multirib and Multiweb Wing Wing Box Structures Under Shear and Moment Loads", 4th AIAA/ASME Struc-
38. S. Timoshenko and S. Woinowsky-Krieger, "Theory of Plates and Shells", 4th Edition, McGraw-Hill, 1957. 39. S. A. Ambartsumyan, "Theory of Anisotropic Plates", Technomic Publishing Co., 1970. 40. J. E. Ashton and J. M. Whitney, "Theory of Laminated Plates", Technomic Publishing Co., 1970. 41. G. N. Savin, NASA TT F-607, "Stress Distribution Around Holes", November 1970. 42. S. G. Lekhnitskiy, FTD-HT-23-608-67 (AD 683218), "Anisotropic Plates", March 1968. Environmental 43. D. J. Birmingham, M.D., "Contact Dermatitis from Synthetic Resins", Regional SPI Technical Conference, NY Section, "Plastics in the Medical Sciences", September 1967. 44. J. R. Stahmenn, "Model Studies of Strike Probabilities to Selected Points on Aerospace Vehicles", Paper No. 700915, 1970 Lightning and Static Electricity Conference, 1970. 45. J. F. Shaeffer, "Aircraft Initiation of Lightning", AFAL-TR-72-325, 12-15 December 1972, pp. 192200. 46. J. A. Plumer, "Data from the Airlines Lightning Strikes", AFAL-TR-72-325, 12-15 December 1972, pp. 282-289. 47. N. Cianos and E. T. Pierce, Stanford Research Institute Technical Report 1, "A Ground Lightning Environment for Engineering Usage", August 1972. 48. SAE Committee AE-4 Special Task F, "Aerospace Recommended Practice: Lightning Effects Tests on Aerospace Vehicles and Hardware", 13 September 1975. 49. R. O. Brick, C. H. King and J. T. Quinlivan, "Coatings for Lightning Protection of Structural Reinforced Plastics", AFML-TR-70-303, Part II, February 1972. 50. C. D. Skouby, McDonnell Aircraft Company, "Electromagnetic Effects of Advanced Composites", January 1975. Other 51. J. E. Dougherty, Jr., "Certification Requirements for Civil Composite Commercial Airframe Structures", 3rd NASA/USAF Conference on Fibrous Composites in Flight Vehicle Design, Williamsburg, Virginia, 4-6 November 1975. 52. FAR, Part 23, Changes 1 through 15.
tures, Dynamics and Materials Conference. 34. F. R. Shanley, "Principles of Structural Design for Minimum Weight", Journal of Aeronautical Sciences, March 1949. 35. W. R. Micks, "Method of Estimating the Compressive Strength of Optimum Sheet-Stiffener Panels for Arbitrary Material Properties, Skin Thickness and Stiffener Shapes", Journal of Aeronautical Sciences, October 1953, pp. 705-715. 36. G. Gerard, "Minimum Weight Analysis of Orthotropic Plates Under Compressive Loading", Journal of Aeronautical Sciences, January 1960.
37. S. Timoshenko, "Theory of Elastic Stability", McGraw-Hill, 1956.
A VP-2 recently completed by D. H. D. Lowry (EAA
17348), 34 Bradgate Rd., Belleville, Ontario. It has since had a long turtle deck added to eliminate some rudder
buffet. The tires are 8.00 x 6.00 on 6.00 x 6.00 wheels. 66 APRIL 1977
