Embed Size (px)
Citation preview
I .
I t-
I .
i
I
LLYIL*CI OLI(ALIY FKaCTURE TOUGHNESS OF 2219-T87 ALUMINUM ALLOY
AT ROOM AND CRYOGENIC TEMPEMTURES
CARLM. CARMAN JOHN W. FORNEY JESSE M. KATLIN
GPO PRICE k CFSTI PRICE(S) S -
Microfiche (MF) * drn ff 853 July 85
Prepared f o r
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION L e w i s Research Center
August 1966
NASA PURCHASE ORDER C6860A
D i s t r i b u t i o n of t h i s report i s unl imi ted .
UNITED STATES ARMY FRANKFORD ARSENAL
PHILADELPHIA, PA.
NOT ICE
This report was prepared as an account of Government sponsored work. Nei ther the United S t a t e s , nor the Nat iona l Aeronatucs and Space Adminis t ra t ion (NASA), nor any person a c t i n g on behalf of NASA
A.) Makes any warranty o r r e p r e s e n t a t i o n , expressed o r impl ied , with r e spec t t o the accuracy, completeness o r u se fu lness of t he i n f o r m t i o n contained i n t h i s r e p o r t , o r t h a t t he use of any information, appara tus , method, o r process d i sc losed i n t h i s r e p o r t may not in f l r in .ge provate ly owned r i g h t s ; o r
B.) Assumes any l i a b i l i t i e s wi th r e s p e c t t o the use o f , o r f o r damages r e s u l t i n g from the use of any informat ion , appa ra tus , method o r process d isc losed i n t h i s r e p o r t .
A s used above, "person a c t i n g on behalf of NASA" inc ludes any employee o r con t r ac to r of NASA, o r employee of such c o n t r a c t o r , t o the e x t e n t t h a t such employee o r con t r ac to r of NASA, o r employee of such con- t r a c t o r prepares , d i sseminates , o r provides access t o any informat ion pursuant t o h i s employment o r con t r ac t wi.th NASA o r h i s employment with such con t r ac to r .
D ISP O SIT ION INSTRUCT IONS
Destroy t h i s r epor t when i t i s no longer needed. the o r i g i n a t o r .
Do n o t r e t u r n i t t o
The f ind ings i n t h i s r e p o r t a r e no t t o be cons t rued as an o f f i c i a l De- partment of the Army p o s i t i o n un le s s so des igna ted by o t h e r au thor ized documents.
t
? ”
.
NASA FA Report R-1821
PLANE STRAIN FRACTURE TOUGHNESS OF
AT ROOM AND CRYOGENIC TEMPERATURES 2219-T87 ALUMINUM ALLOY
CARL M . CARMAN JOHN W. FORNEY JESSE M. KATLIN
Prepared f o r
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Lewis Research Center
August 1966
NASA Purchase Order C6860A
Technical Management NASA Lewis Research Center
Chemical Rocket Div is ion Cleveland, Ohio
Richard M. Johnson and Gordon T . Smith
Pitman-Dunn Research Labora tor ies FRANKFORD ARSENAL
Ph i l ade lph ia , Pa. 19137
Carl M. Carman, John W. Forney, and Je s se M. K a t l i n
ABSTRACT
The t e n s i l e p rope r t i e s and plane s t r a i n f r a c t u r e toughness of 1 / 2 and 1 inch t h i c k 2219-T87 aluminum a l l o y have been determined as a func t ion of t e s t i n g a t temperatures from room t o -423" F. The t en - s i l e and y i e l d s t r eng ths of t h i s m a t e r i a l show a gradual i nc rease as the t e s t i n g temperature i s decreased t o -423" F while t he e longa t ion and reduct ion of a rea remain e s s e n t i a l l y unchanged.
The plane s t r a i n f r a c t u r e toughness of t h i s m a t e r i a l i s r e l a - t i v e l y i n s e n s i t i v e t o t e s t i n g temperature and shows only a s l i g h t i n - c rease with decreasing t e s t i n g temperature . Specimens machined so t h a t the crack propagat ion i s perpendicular t o the r o l l i n g plane show somewhat h igher values of plane s t r a i n f r a c t u r e toughness than when crack propagation i s para l le l t o the r o l l i n g plane. The plane s t r a i n f r a c t u r e toughness of t he 1 inch t h i c k 2219-T87 aluminum a l l o y w a s somewhat lower than t h a t of the 1 / 2 inch t h i c k p l a t e .
I l l u s t r a t i v e examples a r e presented us ing these parameters i n design.
ii
TABLE OF CONTENTS Page .
.
nOSSARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1
INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
PROGRAMOBJECTIVE . . . . . . . . . . . . . . . . . . . . . . . . 2
.TER. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
PROGRAMAPPROACH . . . . . . . . . . . . . . . . . . . . . . . . . 3
Measurement of Plane S t r a i n Frac ture Toughness . . . . . . . 4
Tens i le Tes t ing . . . . . . . . . . . . . . . . . . . . . . . 10
EXPERIMENTAL TECHNIQUE . . . . . . . . . . . . . . . . . . . . . . 10
EXPERIMENTAL RESULTS AND DISCUSSION . . . . . . . . . . . . . . . 11
One-half Inch Thick 2219-T87 Aluminum Alloy P l a t e . . . . . . 11
Engineering Tens i l e P r o p e r t i e s . . . . . . . . . . . . . 11 Plane S t r a i n F rac tu re Toughness . . . . . . . . . . . . 14
One Inch Thick 2219-T87 Aluminum Alloy P l a t e . . . . . . . . 22
Engineering Tens i l e P r o p e r t i e s . . . . . . . . . . . . . . 22 Plane S t r a i n F rac tu re Toughness . . . . . . . . . . . . 22
DESIGN CONSIDERATIONS . . . . . . . . . . . . . . . . . . . . . . 30
CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
REFEREXCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33
DISTRIBUTION . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
L i s t of Tables Table
I . Chemical Composition of 2219-T87 Aluminum Alloy . . . . . 3
I1 . Value of f ( a /d ) a s a Function of a / d . . . . . . . . . . . 8
111 . Tens i l e P r o p e r t i e s of 1 / 2 i nch t h i c k 2219-T87 Aluminum Alloy P l a t e . . . . . . . . . . . . . . . . . . . . . . . 14
iii
L i s t of Tables (Cont'd) Tab l e Page
IV. Plane S t r a i n F rac tu re Toughness as a Function of Specimen S ize and Temperature f o r One-half Inch Thick 2219-T87 Aluminum Alloy . . . . . . . . . . . . . . 1 7
V . Plane S t r a i n F rac tu re Toughness of One-half Inch Thick 2219-T87 Aluminum Alloy P l a t e (Crack Propagat ion P a r a l l e l t o Rol l ing p l ane ) . . . . . . . . . . . . . . . . 20
V I . Tens i le P r o p e r t i e s of One Inch Thick 2219-T87 Aluminum Alloy P l a t e . . . . . . . . . . . . . . . . . . . . . . . 22
VII. Frac ture Toughness of One Inch Thick 2219-T87 Aluminum Alloy P l a t e (Crack Growth P a r a l l e l t o Ro l l ing p lane) . . 27
VIII. Frac ture Toughness of One Inch Thick 2219-T87 Aluminum Alloy P l a t e (Crack Growth Perpendicular t o Ro l l ing P lane) 29
L i s t of I l l u s t r a t i o n s Figure
1. Notched Bend Specimen . . . . . . . . . . . . . . . . . . 6
2. Method of Loading Notched Bend Specimens i n I n s t r o n Tes t ing Machine . . . . . . . . . . . . . . . . . . . . . 7
3. Plane S t r a i n F rac tu re Toughness of 2014-T6 Aluminum Alloy as a Function of Beam Depth a t Room Temperature . . . . . 9
4 . Small Round Tens i le Specimen. . . . . . . . . . . . . . . 10
5 . Schematic of Cryos ta t and Associated Apparatus used f o r Tes t s Conducted a t -423" F. . . . . . . . . . . . . . . . 1 2
6. Tensi le P r o p e r t i e s of 1 / 2 inch t h i c k P l a t e of 2219-T87 Aluminum Alloy 8 s a Funct ion of Tes t ing Temperature . . . 13
7. Plane S t r a i n F rac tu re Toughness of 2219-T87 Aluminum Alloy a s a Funct ion of Beam D e p t h , f o r Room Temperature T e s t s . . . . . . . . . . . . . . . . . . . . . . . . . . 15
8 . Plane S t r a i n F rac tu re Toughness of 2219-T87 Aluminum Alloy as a Funct ion of Beam Depth, f o r -110" F Tests. . . 15
9. Plane S t r a i n Frac ture Toughness of 2 2 1 9 - ~ 8 7 Aluminum Alloy as a Funct ion of Beam Depth, f o r -320" F Tests. . . 16
i v
L i s t of I l l u s t r a t i o n s (Cont'd) F igure Page
10. Plane S t r a i n F rac tu re Toughness of 2219-T87 Aluminum Alloy as a Function of BeamDepth, f o r -423" F Tests. . . 16
11. O r i e n t a t i o n of Test Specimens wi th Respect t o P l a t e . . . 19
12. Plane S t r a i n F rac tu re Toughness of 1 /2 Inch Thick 2219-T87 Aluminum Alloy P l a t e a s a Funct ion of Tes t ing Temperature (Crack growth p a r a l l e l t o r o l l i n g p lane) . . . . . . . . . 19
13. Lower Bound Plane S t r a i n Frac ture Toughness of 1 / 2 Inch Thick 2219-T87 Aluminum Alloy P l a t e as a Funct ion of Tes t ing Temperature (Crack growth p a r a l l e l t o r o l l i n g p l ane ) . . . . . . . . . . . . . . . . . . . . . . . . . . 21
14. Lower Bound Plane S t r a i n Frac ture Toughness of 1 /2 Inch Thick 2219-T87 Aluminum Alloy P l a t e as a Funct ion of Tes t ing Temperature (Crack growth perpendicular t o r o l l i n g plane),. . . . . . . . . . . . . . . . . . . . . . 21
15. F rac tu re Surfaces of Ben Specimens showing Delaminations f o r Cracks Propagat ing Perpendicular t o the Ro l l ing Plane of t h e P l a t e . . . . . . . . . . . . . . . . . . . . . . . 23
16. T e n s i l e P r o p e r t i e s of One Inch Thick P l a t e of 2219-T87 Aluminum Alloy as a Function of Tes t ing Temperature . . . 25
17. Plane S t r a i n F rac tu re Toughness of One Inch Thick 2219-T87 Aluminum Alloy P l a t e a s a Funct ion of Tes t ing Temperature (Crack growth p a r a l l e l t o r o l l i n g p lane) . . . . . . . . . 26
18. Plane S t r a i n F rac tu re Toughness of One Inch Thick 2219-T87 Aluminum Alloy P l a t e as a Funct ion of Tes t ing Temperature (Crack growth perpendicular t o r o l l i n g plane) . . . . . . 28
19. V a r i a t i o n of Cr i t i ca l Crack Depth f o r I n s t a b i l i t y as a Funct ion of t h e Gross Sec t ion S t r e s s f o r Seve ra l K I ~ V a l u e s . . . . . . . . . . . . . . . . . . . . . . . . . . 31
.
V
GLOSSARY
a
A
B
d
E
tr
-%c
K
KIc
M
Q
r Y s
T
Y
€
U
OYS
w
- Notch depth o r 1 / 2 c rack l eng th
- Area
- Specimen th ickness
- B e a m depth
- Young's modulus
- S t r a i n energy r e l e a s e r a t e
- S t r a i n energy r e l e a s e r a t e f o r onse t of f a s t f r a c t u r e i n oepning mode type of f r a c t u r e
- Parameter desc r ib ing the l o c a l e l e v a t i o n of the e l a s t i c s t r e s s f i e l d ahead of a c rack
- C r i t i c a l va lue of above parameter f o r onse t of f a s t f r a c t u r e i n opening mode type of f r a c t u r e
- Bending moment of beam
- Form f a c t o r
- Radius of p l a s t i c s t r a i n zone
- Surface t ens ion
- Poisson ' s r a t i o
- S t r a i n energy
- Gross s e c t i o n s t r e s s
- 0.2% o f f s e t y i e l d s t r e s s
- Work func t ion composed of su r face t e n s i o n and p l a s t i c deformation
Subsc r ip t s
C - C r i t i c a l va lue of a parameter
I - F i r s t , o r opening, mode of f r a c t u r e
.
v i
PLANE STRAIN FRACTURE TOUGHNESS OF 2219-T87 &UMINLTM ALLOY AT ROOM AND CRYOGENIC TEMPERATURES
Carl M. Carman, John W. Forney, and J e s s e M. K a t l i n
SUMMARY
Inasmuch as f u t u r e upper s tage rocke t s w i l l use l i q u i d hydrogen as a f u e l , t h e p r o p e l l a n t tanks w i l l be requi red t o opera te a t -423" F. It is t h e r e f o r e d e s i r a b l e t o employ materials f o r t hese s t r u c t u r e s which w i l l possess very high s t rength- to-dens i ty r a t i o s and material p r o p e r t i e s which w i l l be s a t i s f a c t o r y a t the minimum ope ra t ing temperatures .
It appears that th ree types of a l l o y s o f f e r t he promise of achiev ing the requi red s t r e n g t h i n combination with adequate f r a c t u r e toughness a t -423" F. This r e p o r t p re sen t s engineer ing des ign d a t a f o r one such ma- t e r i a l - 2219-T87 aluminum a l l o y .
The s u i t a b i l i t y of 2219-T87 aluminum a l l o y f o r cryogenic tankage a p p l i c a t i o n s has been s tud ied by determining the mechanical and f r a c t u r e p r o p e r t i e s of t h e material a t t e s t i n g temperatures ranging from room t o -423" F. Small round t e n s i l e specimens were developed t o measure the t e n s i l e p r o p e r t i e s over t he range of t e s t i n g temperatures. Plane s t r a i n f r a c t u r e toughness measurements were a l s o made a t t hese temperatures us- i n g t h e !!pop-in" technique with a small notched bend specimen.
S p e c i a l techniques u t i l i z i n g the s p e c i f i c h e a t of vapor i za t ion of l i q u i d helium were developed t o t e s t t h e specimens a t -423" F. The t en - s i l e and y i e l d s t r e n g t h s of the 2219-T87 aluminum a l l o y show a gradual i n c r e a s e as t h e t e s t i n g temperature i s decreased. The e longa t ion and r educ t ion of area remain e s s e n t i a l l y unchanged.
The plane s t r a i n f r a c t u r e toughness of t h i s material i s r e l a t i v e l y i n s e n s i t i v e t o t e s t i n g temperature, showing only a a l i g h t i nc rease wi th dec reas ing t e s t i n g temperature. Tes ts performed on specimens wi th c rack propagat ion perpendicular t o the r o l l i n g plane show somewhat h igher v a l - ues of plane s t r a i n f r a c t u r e toughness than t e s t s performed on specimens wi th c rack propagat ion p a r a l l e l t o t he r o l l i n g plane. The plane s t r a i n f r a c t u r e toughness of the one inch t h i c k p l a t e w a s somewhat lower than t h a t of t he 1 / 2 inch p l a t e .
The da ta a r e summarized i n terms of a par t - through d e f e c t which w i l l be s t a b l e a t v a r i o u s ope ra t ing temperatures and s t r e s s l e v e l s .
1
IXTRQDUCTLON
The majori ty of l i qu id - fue led rocket boos te r tanks used i n the p a s t have been cons t ruc ted of ma te r i a l s which have room temperature y i e l d s t r eng th - to -dens i ty r a t i o s genera l ly no t exceeding 650,000 inches. These rocket boos te rs use l i q u i d oxygen as the f u e l ox id i ze r and the ope ra t ing temperature of t he l i q u i d oxygen tanks i s -297" F. A t t h i s temperature, the maximum y i e l d s t r eng th - to -dens i ty r a t i o of these ma- t e r i a l s i s approximately 850,000.in. It i s . p l a n n e d t o f u e l f u t u r e rocke t boos t e r s with l i q u i d hydrogen and the tanks would t h e r e f o r e have t o func t ion a t -423" F. Before l i q u i d hydrogen can be used , however, t h e engineer ing, mechanical, and f r a c t u r e p r o p e r t i e s must be determined a t -423" F fo r candidate materials,
It would be d e s i r a b l e t o employ materials wi th t h e h ighes t s t r e n g t h - to -dens i ty r a t i o s t o reduce launching weight, provided t h e material prop- e r t i e s a t the opera t ing temperature were s a t i s f a c t o r y . Data presented a t the 1960 ASTM Symposium on Low Temperature P r o p e r t i e s of High S t r eng th A i r c r a f t and Missile Alloys1* showed t h a t the s t r eng th - to -dens i ty r a t i o of 850,000 inches could be exceeded s u b s t a n t i a l l y by s e v e r a l m a t e r i a l s a t cryogenic temperatures , namely, cold worked s t a b l e a u s t e n i t i c and meta- s t a b l e s t a i n l e s s s t e e l s , annealed alpha t i t a n i u m a l l o y s , and c e r t a i n a l u - minum a l l o y s of the copper-bear ing s e r i e s .
The p resen ta t ion of t hese da t a was concerned most ly with the en- g inee r ing mechanical p r o p e r t i e s a t -423" F, and only a l imi t ed amount of t he work covered f r a c t u r e toughness i n v e s t i g a t i o n s . Based on these papers , the Lewis Research Center of t h e Na t iona l Aeronaut ica l and Space Adminis t ra t ion, Cleveland, Ohio, sponsored i n v e s t i g a t i o n s con- cern ing the f r a c t u r e toughness c h a r a c t e r i s t i c s a t t e s t i n g temperatures a s low as -423" F t o Frovtde b a s i c design da ta f o r the t i t a n i u m and aluminum a l l o y s .
A repor t2 has been publ ished r e c e n t l y by Frankford Arsenal which desc r ibes the plane s t r a i n f r a c t u r e toughness c h a r a c t e r i s t i c s of 5A1- 2.5Sn ELI t i t an ium a l l o y a t room and cryogenic tempera tures . The p resen t r e p o r t covers a s imi la r type i n v e s t i g a t i o n us ing 2219-T87 aluminum a l l o y .
PROGRAM OBJECTIVES
The object of t h i s program i s t o determine the b a s i c engineer ing design parameters of 2219-T87 aluminum a l l o y a t 70", -110", -320", and -423" F .
*See REFERENCES
2
MATERIAL
Nominal composition of the 2219-T87 aluminum a l l o y i s given i n -. Table I. The material w a s suppl ied as 1 /2 inch and 1 inch t h i c k r o l l e d
p l a t e .
TABU I. Chemical Composition of 2219-T87 Aluminum Alloy
Cow onen t
Si Fe cu Mn Mg Zn T i v Z r
Pe rcen t
0.20 max 0.30 max 6.3 0.30 0.02 max 0.10 max 0.06 0.10 0 -18
PROGRAM APPROACH
The ob jec t ive of t h i s programmay be m e t by determing the engineer- i n g tezlsile p r c p e r t i e s and t h e f r a c t u r e toughness c h a r a c t e r i s t i c s of t he m a t e r i a l over the s e r i e s of t e s t i n g temperatures . The s t r e n g t h l i m i t a - t i o n s due t o b r i t t l e f r a c t u r e may bes t be i n v e s t i g a t e d by the a p p l i c a t i o n of t h e G r i f f i t h - I r w i n f r a c t u r e mechanics approach. G r i f f i t h 3 s t a t e d t h a t f o r an i d e a l l y b r i t t l e m a t e r i a l such as g l a s s , t he s t r a i n energy r e l eased p e r u n i t c rack ex tens ion , a t i n s t a b i l i t y , w a s equa l t o twice the su r face t e n s i o n p e r u n i t c rack extension:
For more d u c t i l e materials, however, t he e f f e c t of t he work absorbFd by n l = - t i c r --.a defo rmt lo i ; c ~ t t he crack t i p had t o be accounted f o r . Irwin* proposed a work func t ion be s u b s t i t u t e d f o r the s u r f a c e t e n s i o n t e r m and, a t i n s t a b i l i t y , Equation 1 becomes
where o= work func t ion composed of two terms: (1) su r face f r e e energy, and ( 2 ) p l a s t i c deformation.
3
i n the s p e c i a l case of a through-tho-thickness crack i n an i n f i n - i t e l y large p l a t e , I n g l i s 5 has solved the s t r e s s a n a l y s i s f o r dw, giv- ing r i s e to Equation 3 : dA
= b - - d w - n cr2 a dA E (3)
I r w i n 6 proposed t h a t the events a t the lead ing edge of a c rack may be described i n terms of a parameter , K , which i s a func t ion of t he l o c a l e l eva t ion of the e l a s t i c s t r e s s f i e l d ahead of the c rack . Crack propagation w i l l take p lace when the s t r e s s i n t e n s i t y , K , reaches a c r i t i c a l value, Kc o r KIC ( f r a c t u r e toughness) , depending on t h e s t a t e of s t r e s s . Therefore , i t may be shown t h a t
Kc2 = E bc (plane s t r e s s ) ( 4 ) I K I ~ ~ = E &rc (plane s t r a i n )
(1 - rL> Since t h i s program i s pr imar i ly concerned wi th the plane s t r a i n
f r a c t u r e toughness, Equation 5 i s a p p l i c a b l e .
(5)
Measurement of Plane S t r a i n F rac tu re Toughness
The measurement of the plane s t r a i n f r a c t u r e toughness i s compli- ca ted by the requirement t h a t t he p l a s t i c zone s i z e a t the c rack t i p be q u i t e s m a l l r e l a t i v e t o the specimen c r o s s s e c t i o n . I n the p a s t t h i s necess i t a t ed the use of r e l a t i v e l y l a r g e specimens.
H i s t o r i c a l l y , t he c i r c u m f e r e n t i a l l y notched round specimen has been the most popular specimen. However, f o r t he lower s t r eng th -h igh toughness m a t e r i a l s , the major diameter necessary t o prevent y i e l d i n g of the net s e c t i o n i s q u i t e l a rge . The requirements f o r t he s u r f a c e - flawed specimen also make the use of l a r g e specimens mandatory.
The slow notched bend t e s t has been used ex tens ive ly t o measure the plane s t r a i n f r a c t u r e toughness of m a t e r i a l s . It has the advan- tage t h a t , when r e l a t i v e l y l a rge specimens a r e r equ i r ed , t he loads needed t o break the specimens may be kep t low by i n c r e a s i n g the span of the specimens. specimens with only the maximum load o r a simple load d e f l e c t i o n curve being recorded. Modern ins t rumenta t ion has r e s u l t e d i n t h e use of smal le r s i ze notched bend specimens t o o b t a i n r e l i a b l e plane s t r a i n f r a c t u r e toughness d a t a .
Early work was p r i m a r i l y confined t o q u i t e l a rge
Recently, Boyle, Su l l ivan , and K r a f f t 7 developed a technique f o r measuring the plane s t r a i n f r a c t u r e toughness us ing sha rp ly notched
s h e e t specimens. They observed t h a t t he i n i t i a l b u r s t of c rack growth from the notch o r f a t i g u e c rack occurred under plane s t r a i n condi t ions . This technique c o n s i s t s of determining the load -de f l ec t ion curve of t he specimen. de t ec t ed as a n i n f l e c t i o n i n the load-def lec t ion curve.
The i n i t i a l b u r s t of crack ex tens ion , o r "pop-in," may be
S ize i s an important f a c t o r i n s e l e c t i n g specimens f o r cryogenic t e s t i n g . Large specimens r e q u i r e excess ive amounts of c o s t l y cryogenic l i q u i d s t o cool them. Consequently, i t i s d e s i r a b l e t o use as small a specimen as poss ib l e and s t i l l be c o n s i s t e n t wi th the experimental con- d i t i o n f o r ob ta in ing v a l i d f r a c t u r e toughness measurements. Therefore , the small notched bend specimen was s e l e c t e d f o r t h i s s tudy . This speci- men o f f e r s t he advantage of small s i z e and minimum breaking load. u s ing the "po?-'id' technique, it may be poss ib l e t o f u r t h e r reduce the specimen s i z e .
By
I n a t tempt ing t o aeasure t h e plane s t r a i n f r a c t u r e toughness of low s t r eng th -h igh toughness m a t e r i a l s , i t i s necessary t o determine the mini- mum s i z e specimen which w i l l g ive v a l i d plane s t r a i n f r a c t u r e toughness va lues . depth was made us ing 1 /4 inch t h i c k 2014-T6 aluminum a l l o y , a d i f f e r e n t material. This ma te r i a l was used i n a previous program and a plane s t r a i n f r a c t u r e toughness value ( K I ~ ) of 32,000 p s i a t room t e m - p e r a t u r e had been e s t a b l i s h e d . t a b l i s h e d by performing t e s t s us ing c i r c u m f e r e n t i a l l y notched-fat igue cracked round specimens and pop-in tes ts of cen te r cracked s h e e t s and s i n g l e edge notched specimens.
A pre l iminary survey of the e f f e c t s of beam depth and notch
The v a l i d i t y of t h i s KIc value was e s -
The genera l conf igura t ion of the notched bend specimen i s shown i n F igure 1, Three Seam depths (1 /2 , 3 / 4 , and 1 inch) and two crack depths (10 and 20 percent ) were used f o r t hese tests. The specimens were ma- chined wi th a notch roo t r ad ius of 0.001 inch , maximum. Previous work9 has shown t h a t even f o r s o f t e r aluminum a l l o y s , notch sharpness may a f - f e c t t h e x a s u r e d va lue of plane s t r a i n f r a c t u r e toughness. Consequently, a l l of t he notched bend ba r s were precracked i n f a t i g u e p r i o r t o t e s t i n g . A small, h igh , e longat ion s t r a in gage w a s cemented a t t h e t i p of t h e crack t o d e t e c t t h e pop-in. The specimens were t e s t e d i n th ree -po in t l oad ing u s i n g a n I n s t r o n t e s t i n g machine. A schematic of t he loading arrangement i s shown i n Figure 2. The output of the s t r a i n gage w a s fed i n t o a s t r a i n gage preampl i f ie r and a p l o t of load vs specimen d e f l e c t i o n w a s ob ta ined on t h e x-y recorder . The pop-in was de tec ted as a n i n f l e c - t i o n i n the load-def l e c t i o n curve .
The plane s t r a i n f r a c t u r e toughness was c a l c u l a t e d from t h e load a t pop- in and i n i t i a l crack length , using a s t r e s s a n a l y s i s developed by Bueckner.10 H i s express ion i s
6M 31z f ( a /d ) - K1 - (d - a)
The v a l u e of f ( a / d ) i s given i n Table 11,
5
m
m
t
U-
6
-L
-
1 7
-I -I w 0 0 U 0 J
ct
ln
w K 0 fn
s
1
0 d
7
TABLE I1 Valce of f ( a / d ) a s a Function of a /d
a / d 0.05 0.10 0.20 0.30 0.40 0.50 0 .GO f (a /d) 0.36 0.49 0.60 0.66 0.69 0.72 O"73
The room temperature plane s t r a i n f r a c t u r e toughness da t a f o r 2014-T6 aluminum a l l o y s a r e p l o t t e d as a func t ion of the beam depth i n Figure 3. It w i l l be observed t h a t t he maximum value of K I ~ i s less than the value of 32,000 p s i Under these cond i t ions , a p l a s t i c zone co r rec t ion may be requi red . The r ad ius o f the p l a s t i c zone may be ca l cu la t ed from 6
The value of rys i s added t o the measured va lue of a. The new value of a This process i s repea ted u n t i l the change i n KIc i s very small . The s e r i e s rap id ly converges s o t h a t only th ree o r fou r c a l c u l a t i o n s are needed t o produce a n e g l i g i b l e change i n K I ~ . These va lues a r e p l o t t e d i n Figure 3. I n s t u d i e s of t4e e f f e c t of specimen s i z e on the measured va lues of plane s t r a i n f r a c t u r e toughness of 7075-T6 aluminum a l l o y s , Boyle, Su l l ivan , and K r a f f t 7 have shown t h a t a minimum specimen s i z e e x i s t s fo r v a l i d measurements of the plane s t r a i n f r a c t u r e toughness. Tes t s of l a r g e r specimen s i z e s do not apprec i ab ly a l t e r t he K I ~ va lues . These data f o r 2014-T6 aluminum a l l o y show t h e same t r end as t h a t re- ported fo r t he 7075-T6 aluminum a l l o y . beam depth should be the minimum specimen s i z e .
i s s u b s t i t u t e d i n Equation 6 and KIc i s ca l cu la t ed aga in .
By analogy, then , t he 3 /4 inch
Having e s t a b l i s h e d v a l i d plane s t r a i n f r a c t u r e toughness va lues f o r t he 2014-T6 aluminum, i t i s poss ib l e t o desc r ibe the necessary experimental condi t ions f o r accu ra t e plane s t r a i n f r a c t u r e toughness measurements. It was shown7 t h a t , f o r s a t i s f a c t o r y pop-in measure- ment, the specimen th ickness should be e q u a l t o a t l e a s t fou r t i m e s the rad ius of the p l a s t i c zone s i z e . p l a s t i c zone s i z e gave a value of 0.0387 inch . The 0.25 inch t h i c k specimen, t h e r e f o r e , w a s more than adequate t o meet t h i s requirement.
C a l c u l a t i o n of the r a d i u s of the
From the preceeding d i scuss ion i t w a s decided t h a t a specimen depth of 0.750 inch w a s necessary f o r a c c u r a t e p lane s t r a i n f r a c t u r e toughness measurements. K r a f f t , i n h i s work on s h e e t specimens, has recommended t h a t the specimen width be a t l e a s t 20 times the r a d i u s of the p l a s t i c zone s i z e , By analogy, t he beam depth of t he specimen should conform t o t h i s requirement. The re fo re , a minimum beam depth of 0.774 inch w a s needed.
8
I -
o c cn Q J
c 3 0 I k 3
n
+-
I 0 rr)
h 0
9
Gross l l has s t a t e d t h a t the l i m i t of a p p l i c a b i l i t y of the specimen w i l l be reached i f the nominal s t r e s s a t the crack t i p reaches the y i e l d s t r e s s of the m a t e r i a l , as s t a t e d i n
Solu t ion of Equation 8 f o r d-a gave a value of 0.481 inch. With a 20 percent notch depth, t h i s would give a beam depth of 0.601 inch . This c r i t e r i o n i s somewhat l e s s conserva t ive than 20 t imes the p l a s t i c zone s i z e .
Tens i le Tes t ing
A s mentioned previous ly , l a r g e r consumptions of cryogenic coo lan t s and the mechanical l i m i t a t i o n s of the c r y o s t a t n e c e s s i t a t e d the use of small specimens, cons i s t en t with obta in ing v a l i d d a t a , f o r K l c measure- ments. ?his a l s o holds t r u e f o r the de te rmina t ion of t he engineer ing t e n s i l e p rope r t i e s . Therefore , a smal l round t e n s i l e specimen, 0.160 inch i n diameter (Figure 4 ) w a s used f o r t hese s t u d i e s . Small s t r a i n gages were used t o determine the s t r a i n of the specimen.
I- 2'1
Figure 4 . Small Round Tens i l e Specimen
EXPERIMENTAL TECHNIQUE
The experimental arrangement used f o r loading the notched bend specimens has been descr ibed previous ly . Specimens t e s t e d a t ambient temperature were broken i n a i r a t approximately 70" F. T e s t s a t -110"
10
c
and -320" F were conducted by immersing the compression column spec i - men 2nd a s soc ia t ed appara tus i n a mixture of dry i c e and acetone and i n l i q u i d n i t rogen , r e s p e c t i v e l y .
Since i t was not p r a c t i c a l , from a s a f e t y s t andpo in t , t o use l i q - u i d hydrogen i n the labora tory , a system had t o be developed whereby t e s t i n g could be accomplished a t -423" F. The method devised u t i l i z e d t h e evapora t ing gas of l i q u i d helium (-452" F) as t h e coolan t . The co ld gas passed over an e l e c t r i c a l r e s i s t a n c e h e a t e r c o n t r o l l e d so t h a t t he emerging stream of hea ted gas maintained the tes t specimen a t l i q u i d hydrogen temperature. A schematic of the c r y o s t a t and a s soc ia t ed appa- r a t u s used is shown i n F igure 5.
A b r i e f d e s c r i p t i o n of the operat ion of t h e c r y o s t a t fol lows. Af- t e r p l ac ing tbe specimen i n pos i t i on i n the ho lde r , making a l l e l e c t r i c a l connect ions, and s e a l i n g the c ryos t a t t o t he t e s t i n g machine, l i q u i d n i t r o g e n is introduced i n t o the outer dewar u n t i l the proper l e v e l i s obtained. This s e r v e s as a s h i e l d f o r t he l i q u i d system. Liquid n i t r o g e n i s then slowly introduced through i n l e t the inner dewar and allowed t o b o i l . The bo i l -o f f gas then drops down i n t o cup 11 as shown by @ The gas then t r a v e l s up through the opening a t t h e 9 o t t o m of t he cold f i n g e r @@over the h e a t e r and the rmis to r , over the specimen, and out t h e exhaust . Excess p res - su re i s b l ed o f f by exhaust va lve 0. Cooling wi th l i q u i d n i t rogen i s cont inued u n t i l the specimen temperature i s approximately -250" F. The system i s now purged w i t h helium gas, and l i q u i d helium i s then t r a n s - f e r r e d i n t o the i n n e r ( a rea @) dewar. The pa th of t h e co ld helium gas is the same as t h a t of t h e n i t rogen gas. However, when the temper- a t u r e of t he cold helium gas stream i s below -423" F , t he h e a t e r c o i l i s energ ized t o cond i t ion the gas stream t o l i q u i d hydrogen temperature. The temperature of t h e specimen i s confirmed by means of a d i f f e r e n t i a l thermocouple. The specimen i s maintained a t temperature f o r 10 minutes p r i o r t o t e s t i n g .
EXPERIMENTAL RESULTS AND DISCUSSION
One-half Inch Thick 2219-T87 Aluminum Alloy P l a t e
Engineer ing Tens i l e P r o p e r t i e s
The engineer ing t e n s i l e p rope r t i e s of t h i s material, as d e t e r - mined us ing the 0.160 inch diameter t e n s i l e specimens, are p l o t t e d i n F i g u r e 6 as a func t ion of t e s t i n g temperature, and the va lues are tabu- l a t e d i n Table 111,
"Circled numbers r e f e r t o Figure 5 .
11
Neg : 3 6.23 1 .SO 719/AMC. 65
1. 2. 3. 4. 5. 6. 7. 8. 9.
10. 11.
rt 4
I \ I L I
Moveable Cross head S t r a i n Bar Exhaust - He lium Exhaust - Nitrogen & Helium Vacuum Inner Glass Dewar F l a sk Outer Glass Dewar F l a sk Liquid Nitrogen Glass Cold Finger H e l i u m Gas Downflow Area Outer Cup - Cold Finger
12 . Wire Mesh Cage f o r Heater
13. Specimen Holder 14. Beam Specimen 15. Heater , Thermistor & Specimen
16. Liquid and/or Gaseous Helium 17. S t e e l Compression Tube 18. Liquid Helium I n l e t 19. L iquid Ni t rogen I n l e t 20. Amphenole P l u g f o r Instrumenta-
t i o n Lead Wires
& Thermistor
Thermocouple
Figure 5. Schematic of Cryos ta t and Assoc ia ted Apparatus used f o r Tests Conducted a t -423" F
12
l -
100
90
80
70 5 .)
Y * $0 a I- v)
50
40 \
bl
w LONGITUDINAL * TRANSVERSE
TENSILE STRENGTH
I I
REDUCTION OF AREA
-400 -300 -200 -100 0 loo TEMPERATURE (OF)
Figure 6 . Tensile Properties of 1/2 inch thick Plate of 2219-T87 Aluminum Alloy as a Function of Testing Temperature
13
Tens i l e P r o p e r t i e s of
Tes t
(OF) Di rec t ion Temp
Ambient Longi tudina l Ambient Longi tudina 1 Ambient Transverse Ambient Transverse
-110 Longi tudina l -110 Longi tudina l - 110 Transverse -110 Transverse
-320 -320 -3 20 Transverse
Long i tud i n a 1 Long i t ud i n a 1
-423 Longi tud i n a 1 -423 Lsngi tudina l -423 Transverse -423 Transverse
The s t r e n g t h
TABLE 111. 1/2 inch t h i c k 2219-T87 Aluminum Alloy
St renpth ( p s i ) Y i e l d a t
0.20% Offse t
57,300 5,6,500
54 , 400 55 , 100
60 , 400 59,500 58,000 58 , 500 67,500 66 , 300 64 , 500 77,500 72,800 67,000 66 , 500
Tens i l e
70 , 000 69,000 66 , 700 67,400
73 , 500 73 , 800 72 , 300 72 , 100
83 , 800 83,800 81,900
10 2 , 000 96 , 100 96 , 100 92,500
Elongat ion (%)
16.4 15. .. 1 12.9 13.9
15.6 14.7 12.9 13.8
15.6 16.4 14.7
20.4 17.3 17.3 15.6
P l a t e
Reduction of Area
(%)
29.9 29.9 23.5 20.1
27.9 28.9 24.5 19.1
28.9 28.9 22.5
25.5 27.9 23.5 26.5
prope r t i e s of t h i s material are not g r e a t l y sen- s i t i v e t o the t e s t i n g temperature. However, a gradual i nc rease i n y i e l d and t e n s i l e s t r e n g t h s d id occur upon decreas ing the t e s t i n g temperature t o -423" F. The e longat ion and r educ t ion of a r e a were no t a f f e c t e d by reducing the t e s t i n g temperature These va lues are i n c l o s e agreement with those repor ted by Ti f fany . 15.
Plane S t r a i n F rac tu re Toughness
I n rccordance wi th the concepts advanced ear l ie r , it i s neces- s a r y t o determine the minimum specimen s i z e t o o b t a i n v a l i d plane s t r a i n f r a c t u r e toughness va lues . A s e r i e s of notched bend b a r s , having a 20 percent notch wi th depths varying from 1/2 t o 1-1/2 i nches , was machined from the 1/2 inch t h i c k 2219-T87 aluminum a l l o y p l a t e . The experimen- t a l l y determined plane s t r a i n f r a c t u r e toughness va lues are p l o t t e d as a func t ion of specimen depth i n F igu res 7, 8, 9, and 10, and the va lues a r e tabula ted i n Table IV,
14
- LONGITUDINAL
A-& TRANSVERSE
40
-\
30 .- m x
k-
X Y 20
IO
Figure 7 . Plane Strain Fracture Toughness of 2219-T87 Aluminum Alloy as a Function of Beam Depth, for Room Temperature Tests
Figure 8 . Plane Strain Fracture Toughness of 2219-T87 Aluminum Alloy as a Function of BeamDepth, for -110" F Tests
15
I 12 I
d (inch)
Figure 9. Plane S t r a i n F rac tu re Toughness of 2219-T87 Aluminum Alloy as a Function of B e a m Depth, f o r -320" F T e s t s
.- I j4 I 12
d ( inch)
Figure 10. Plane S t r a i n F rac tu re Toughness of 2219-T87 Aluminum Alloy a s a Funct ion of B e a m Depth,for -423" F Tests
16
Test Temp 0
70
-110
-320
-423
TABLE 1V. Plane Strain Fracture Toughness as a Function of
Specimen Size and Temperature for One-half Inch Thick 2219-T87 Aluminum Alloy
Specimen Depth (in.)
3 I4 3 I4
1 1
1-1/2 1-112 1-112
11 2 If2 112 112
3 /4 3 14
1 1 1 1
112 112 11 2 112
3/4 3 I4 3 /4 314
1 1 1 1
112 1/ 2 1/2 I! 2
1 1 1 1 1 1
Direct ion
Longitudinal Transverse
Longitudina 1 Transverse
Longitudinal Longitudinal Longitudinal
Longitudinal Long i t ud ina 1 Transverse Transverse
Longitudinal Transverse
Longitudinal Longitudinal Transverse Transverse
Longitudinal Longitudinal Transverse Transverse
Longitudinal Longitudinal Transverse Transverse
Longitudinal Longitudinal Transverse Transverse
Long i tud ina 1 Longitudinal Transverse Trar?sverse
Longitudinal Longitudinal Longitudinal Transverse Transverse Transverse
Load 0
1835 14 20
3450 2865
5300 5690 5 740
1120 10 90 9 10 9 60
2050 1715
3460 3 640 28 90 3080
1090 1100 900 956
2290 2350 1880 1730
4125 4080 33 70 25 90
1118 1240 1050 1090
2800 3 200 2930 2850 3175 2750
Initial Crack
Length (in.)
0.1696 0.1716
0.2119 0.2156
0.3754 0.3756 0.3626
0.1025 0.1132 0.1180 0 -1097
0.1672 0.1607
0.2104 0.2103 0.2157 0.2106
0.1178 0.1175 0.1236 0.1176
0.1705 0.1634 0.1662 0.1642
0.2140 0.2166 0.2100 0.2722
0.1148 0.1187 0.1205 9. I122
0.3178 0.2759 0.2904 0 -3026 0.2167 0.2750
KIc (psi &>
31,200 24,600
3 7,000 30,900
34,800 37,200 36,400
33,800 34,500 29,300 29,500
34,900 28,400
37,600 38,300 30,900 32,800
35,800 36,200 30,100 30,700
39,300 39,400 31,800 30,800
44,900 43,700 35,600 32,800
35,800 40,600 34,400 3 4 , io0
40,500 41,200 39,300 39,600 34,300 35,200
Average KIc
Jpsi JL';;I>
36,100
34,200
29,400
37,900
31,800
36,000
30,400
39,400
31,300
44,300
34,200
38,200
34,300
40,300
36,300
Fcl?svFag t h e discuss ion of t he da t a f o r the 2014-T6 aluminum a l l o y , i t may be concluded t h a t the one inch deep beam i s s u f f i c i e n t l y l a r g e . fo r t h i s m a t e r i a l . It was f e l t t h a t room temperature da t a f o r l ong i tud ina l specimens would provide adequate information s i n c e the toughness i n the o the r d i r e c t i o n s i s usua l ly l e s s and the y i e l d s t r e n g t h inc reases wi th decreas ing t e s t temperature. These condi t ions would r e - q u i r e a smaller minimum s i z e specimen f o r t he cryogenic temperatures . Data obtained a t -320" F show t h a t a one inch beam depth was needed e v e n ' a t t h i s temperature; t h e r e f o r e , a one inch beam depth w a s s e l e c t e d f o r a l l t e s t temperatures.
Poss ib l e v a r i a t i o n of the plane s t r a i n f r a c t u r e toughness wi th r e spec t t o crack o r i e n t a t i o n i s of practical importance t o the des igner . I n t h i c k p l a t e s i t i s poss ib l e t o study the plane s t r a i n f r a c t u r e tough- ness of the ma te r i a l with r e spec t t o an iso t ropy . The o r i e n t a t i o n of t he var ious specimens i s shown i n Figure 11. I n t h i s f i g u r e , the "L" and "T" designat ions give the o r i e n t a t i o n of the specimens r e l a t i v e t o the r o l l i n g d i r e c t i o n of the p l a t e ; i n the "S" s e r i e s , the d i r e c t i o n of '
crack propagation i s p a r a l l e l t o the r o l l i n g p lane , while specimens i n the "D" s e r i e s a r e s o o r i en ted t h a t the pa th of c rack propagat ion i s perpendicular t o the r o l l i n g plane.
I n Figure 12, the plane s t r a i n f r a c t u r e toughnesses f o r one inch beam depth LS and TS specimens a r e p l o t t e d as a func t ion of t e s t i n g temperature, and the data a r e given i n Table V. The plane s t r a i n f r a c - t u r e toughness of t h i s ma te r i a l i s r e l a t i v e l y i n s e n s i t i v e t o t e s t i n g temperature. These da t a show a t rend s imilar t o t h a t repor ted by T i f f a n y l l ( s o l i d p o i n t s ) .
In order t o compare the e f f e c t s of c rack o r i e n t a t i o n on the plane s t r a i n f r a c t u r e toughness of the 1 / 2 inch t h i c k p l a t e of 2219-T87 aluminum a l l o y , only beam depths of 1 / 2 i nch could be used. Lower bound plane s t r a i n f r a c t u r e toughness values* obta ined f o r the LS and TS s e r i e s using 1 / 2 inch beam depth specimens a r e p l o t t e d i n Figure 13 a s a func t ion of t e s t i n g temperature . These d a t a show a gene ra l de- p re s s ion of the measured plane s t r a i n f r a c t u r e toughness a t t e s t i n g temperatures above -423" F. The da ta p o i n t s f o r -423" F a r e i n agree- ment wi th those obtained us ing l a rge specimens.
S imi l a r da t a f o r t he LD and TD series a r e shown i n F igure 14. Comparison of F igures 13 and 14 shows t h a t t he lower bound plane s t r a i n f r a c t u r e toughness f o r c racks propagat ing pe rpend icu la r t o the r o l l i n g plane of the p l a t e i s g r e a t e r than f o r c r acks propagat ing p a r a l l e l t o t he r o l l i n g plane. us ing sur face- f lawed specimens
*Plane s t r a i n f r a c t u r e toughness va lues c a l c u l a t e d us ing data obta ined
These va lues a r e always less than the t r u e Plane
S imi l a r behavior has been observed by Tiffany13
from specimens which a r e too small have been a r b i t r a r i l y def ined as lower bound va lues . s t r a i n f r a c t u r e toughness.
18
R O L L I N G D I R E C T I O N
Figure 11. Orientation of Test Specimens with Respect to Plate
F
TESTING TEYPERATURE~F)
OPEN POINTS FROM FRANKFORD ARSENAL M LONGITUDINAL
TRANSVERSE SOLID POINTS FROM TIFFANY (12 *PART THROUGH SHEET @NOTCHED ROUND
Figure 12. Plane Strain Fracture Toughness of 1/2 Inch Thick 2219-T87 Aluminum Alloy Plate as a Function of Testing Temperature (Crack growth parallel t o ro l l ing plane)
19
It i s be l ieved t h a t t he h igher energy abso rp t ion f o r c racks propagating perpendicular t o the r o l l i n g plane w a s due t o delaminat ion of t h e plate. These delaminat ions are q u i t e apparent upon examination of t he f r a c t u r e su r face (Figure 1 5 ) .
TABLE V. Plane S t r a i n Frac ture Toughness of
One-half Inch Thick 2219-T87 Alumipum Alloy P l a t e (CrackPropagaSionParallel t o Rol l ing Plane)
Test Temp (OF)
70
-110
-320
-423
Direc t ion
Longi tudina l Transverse
Longitudina 1 Longi tudina l
Transverse Trans ve r s e
Longi tudinal Longi tudina l L ongi tudina 1
Transverse Transverse Transverse
Long i t u d i na 1 Long i tudina 1 Longi tud ina 1 Transverse Transverse Transverse
Load 0 3450 28 65
3460 3 640
2890 3080
4125 4105 4080
3320 3370 2590
2800 3 200 2930 2850 3175 2 750
20
I n i t i a l Crack Length
( in . )
0.2119 0.2156
0.2104 0.2103
0.2157 0.2106
0.2140 0.2150 0.2166
0.2170 0.2100 0.2722
0.3178 0 -2759 0.2904 0.3026 0.2167 0.2750
5 C (psi K)
3 7,000 30,900
37,600 38,300
30 , 900 32,800
44,900
43 , 700
36,000 35,600 32,800
40,500 41,200 39,300 39,600 34,300 35,200
44 , 100
TEMPERATURE CF)
e 30 6 J; 20
- P
0
IO
A I
I I
F i g u r e 14. Lower Bound Plane S t r a i n F rac tu re Toughness of 1 / 2 Inch Thick 2219-T87 Aluminum Alloy P l a t e as a Funct ion of Tes t ing Temperature (Crack growth perpendicular t o r o l l i n g plane)
2 1
One Inch Thick 2219-T87 Aluninum Alloy P l a t e
Engineer ing Tens i l e Proper t i e s
The engineer ing t e n s i l e p r o p e r t i e s of t h i s m a t e r i a l a r e p l o t t e d i n Figure 1 6 as a func t ion of t e s t i n g temperature , and the data a r e summarized i n Table VI. E s s e n t i a l l y the same comments a r e p e r t i - nent regarding these data as f o r the da ta from the one-half inch p l a t e .
TABLE V I . Tens i l e P rope r t i e s of One Inch Thick 2219-T87 Aluminum Alloy P l a t e
Test S t r e c g t h ( p s i ) Reduction Temp Yie ld a t Elongat ion of Area
Direc t ion 0.20% O f f s e t Tens i l e (%) (%) 0 Ambient Longi tudinal 54 , 200 66,200 14.7 28.9 Amb i e n t Long i t ud ina 1 54 , 500 67,200 17.3 30.9
Ambient Transverse 53 , 800 67,500 12 .7 Ambient Transverse 55,000 66 , 700 12.9 21 .1
-110 Longitudinal 58,000 70 , 600 14.7 26.5 -110 Longitudinal 57,000 70 , 100 14.2 27.9 -110 Transverse 56,700 71 , 200 11.1 16.7 -110 Transverse 54,800 7 1 , 200 11.1 19.1
-320 Longi tudinal 65 , 500 81 , 600 16.4 28.9 -320 Longi tudinal 67,000 82,000 18.2 30.9
-320 Transverse 66 , 000 83 , 050 12.9 24.5 -3 20 Transverse 66 , 100 83,900 12.9 21 .1
-423 Longi tudinal 7 1 , 200 95,000 22.2 28.9 -423 Long i t ud ina 1 70 , 600 92 , 700 27.9
-423 Transverse 76 , 200 96 , 100 16.9 24.5 -423 Trans ve r s e 68,000 96,000 16.4 1 9 . 1
Plane S t r a i n F rac tu re Toughness
The plane s t r a i n f r a c t u r e toughness of t h i s p l a t e was deter- mined using a s m a l l bend specinen which w a s one-half inch t h i c k by one inch deep. The values of plane s t r a i n f r a c t u r e toughness d e t e r - mined using specimens o r i e n t e d i n the LS and TS d i r e c t i o n s a r e shown as a func t ion of t e s t i n g temperature i n F igure 1 7 and a r e t a b u l a t e d i n Table VII.
22
~ _________~
Neg : 36.23 1. S 1565/AMC. 65
L a m i n a t i on s
Lami n a t i ons
D i r e c t i o n o f C r a c k Growth
R o l l i n g D i r e c t i o n -
Figure 15. F rac tu re Surfaces of Bend Specimens showing Delaminations f o r Cracks Propagat ing Perpendicular t o t h e Ro l l ing Plane of t h e P l a t e
23
IOC
9c
8(
T a Y u
cn 7c u) w I- cn a
60
X
40 \
Figure 16.
REDUCTION OF AREA
0 IO0 -400 -300 -200 -100 TEMPERATURE (OF)
Tens i le P r o p e r t i e s of One Inch Thick P l a t e of 2219-T87 Aluminum Alloy as a Function of Tes t ing Temperature
25
m
I 'c)
.A $ 2
26
TABLE VII.
(Crack Growth P a r a l l e l t o Ro l l ing P lane) F rac tu re Toughness of One Inch Thick 2219-T87 Aluminum Alloy P l a t e
Tes t Temp 0
70
-110
-320
-423
f r a c t u r e
Di rec t ion
Longi tudina l Longi tudina l L ongi tud i n a 1 Transverse Transverse Transverse
Longi tudina l L ongi tud i n a 1 Longi tudina l Transverse Transverse Transverse
Longi tudina l Longi tudina l Longi t ud ina 1 Transverse Transverse Transverse
Longi tudina 1 Longi tudina l Longi tudina l Transverse Transve r se Transverse
Load 0 2875 2 640 2610 2485 23 65 2340
2 730 2550 2650 2300 2470 2420
3175 3 150 3400 2840 2820 2720
3 700 3590 3875 3410 3300 3 225
These da t a show a gradual
I n i t i a l Crack Length
( in . )
0.2566 0.2753 0.2718 0 2523 0.2589 0.2643
0,2561 0.2855 0 2529 0.2632 0.2547 0.2519
0.2707 0.2777 0,2519 0.2651 0.2624 0.2516
0.2615 0.2603 0.2437 0.2340 0.2580 0.2752
K I C
( p s i &Z) 35,700 33 , 200 32,600 29 100 28,400 28 500
32,600 33,000 31,300 27,800 29,300 28 700
39,300 39,900 40,100 34 , 800 34,400 29,400
44,800 37,100 44,500 37,700 34,100 40 , 100
t r end of i nc reas ing plane s t r a i n toughness wi th decreas ing t e s t i n g temperature below -110" F.
The gene ra l l e v e l of plane s t r a i n f r a c t u r e toughness repor ted here i s somewhat lower than t h a t of the 1 / 2 i nch t h i c k p l a t e . of K I ~ are i n c l o s e r agreement wi th those r epor t ed by Ti f fany . How- e v e r , t h i s w s u l d 3e a n t i c i p a t e d s ince T i f f a n y ' s specimens were taken from r e l a t i v e l y heavy p l a t e , i .e . , up t o 2-1/2 inches i n th ickness .
These va lues
The d a t a obtained f o r the LD and TD s e r i e s of specimens a r e shown as a func t ion of t e s t i n g temperature i n F igure 18 and a r e tabu- l a t e d i n Table VIII. The l e v e l of plane s t r a i n f r a c t u r e toughness determined wi th these specimens was approximately 30 percent h igher
27
(d
m (d
3
t 3
3
3
I 2 -
LL. e w a 3 I-
D Q D L Z N W I n I W c
0 0 rr)
I
0 0 * I
b
E-c c .rl w ( d o Lc i r c * 2
28
than for comparable specimens taken in the LS and TS orientation. This effect may be attributed to delaminations in the short transverse di- re c tion.
TABLE VIII.
(Crack Growth Perpendicular to Rolling Plane) Fracture Toughness of One Inch Thick 2219-T87 Aluminum Alloy Plate
Test Temp 0 70
-110
-320
-423
Initial Load Crack Length KIc
Direction 0 (in.) (psi K) L ong i t ud ina 1 3 930 0.2703 48,400 Long i tud ina 1 40 75 0.2639 49 , 600 Longitudinal 3550 0.2248 38 , 300 Transverse 3410 0.2706 42 , 700
Transverse 3050 0.2473 37,900 Transverse 3975 0.2682 49 , 100
Long i t udina 1 4110 0.2515 48,400 Longitudinal 4800 0.2211 51,400 Longitudinal 3 760 0.2165 39 , 700 Transverse 35 70 0.2489 41,400 Transverse 3 240 0.2555 38 , 600 Transverse 3500 0.2516 41,300
Longi tud ina 1 4190 0.2077 43 , 100 Long i tud ina 1 40 10 0.2148 44 , 900 Longitudinal 43 70 0.2165 46,500 Transverse 3 600 0.2350 40 , 300 Transverse 3690 0.2426 42,100 Transverse 3680 0.2425 42,000
Longitudinal 4 700 0.2460 54 , 600
Longitudinal 4700 0.2480 54,500 Transverse 41 75 0.2566 49,400 Trans ve r se 4 740 0.2593 56,600 Transverse 43 10 0.2596 51,600
Long i tud ina 1 4500 0.2591 55,000
29
DESIGN CONSIDERATIONS
Of the many cons idera t ions t o which the des igner must devote a t - t e n t i o n , those discussed i n t h i s r epor t a r e the y i e l d s t r e s s and plane s t r a i n f r a c t u r e toughness.
The data f o r the y i e l d s t r e s s a t the va r ious t e s t temperatures have been t abu la t ed and are se l f -explana tory . I n F igure 19, the c rack depth f o r i n s t a b i l i t y i s p l o t t e d as a func t ion of gross s e c t i o n s t r e s s f o r a panel conta in ing a s e m i e l l i p t i c a l crack. Examination of t hese curves shows t h a t plane s t r a i n i n s t a b i l i t y w i l l be achieved before a c rack could develop through the p l a t e t h i ckness . Consequently, des ign cons idera t ion must be d i r e c t e d toward the maximum s i z e of d e f e c t which is s t a b l e a t the opera t ing stress. An example of the use of t h i s f i g u r e as an a i d t o engineer ing design fol lows.
I f a s t r u c t u r e were t o be f a b r i c a t e d of the one inch t h i c k 2219-T87 aluminum a l l o y f o r s e r v i c e a t -320" F , t he K I ~ i n the d i r e c t i o n of i n - t e r e s t would be determined from Table V I 1 1 (e.g., LD K I ~ = 45,000 p s i G). If the design s t r e s s f o r t h i s s t r u c t u r e were approximately 80 percent of the y i e l d s t r e n g t h (53,000 p s i ) , then , as determined from Figure 19 , the maximum al lowable depth of d e f e c t would be 0.24 inch. I f t he s t r u c t u r e were f a b r i c a t e d from t h i s material f o r use a t -423" F and the crack were propagat ing perpendicular t o bo th the plane of t h e p l a t e and r o l l i n g d i r e c t i o n (TD), t h e K I ~ would be 52,500 p s i & If t h e minimum crack depth which would be de t ec t ed by nondes t ruc t ive means were 0.200 inch , t he i n F igure 19 would be 0.200 inch and the max- imum allowable s t r e s s a t which t h i s s t r u c t u r e could be opera ted would be approximately 67,000 p s i . These va lues , however, do no t inc lude any allowance f o r a s a f e t y f a c t o r .
CONCLUSIONS
It may be concluded t h a t
1. The t e n s i l e p r o p e r t i e s of 2219-T8 aluminum a l l o y are r e l a - t i v e l y i n s e n s i t i v e t o t e s t i n g temperature . The y i e l d s t r e n g t h v a r i e s from approximately 57,300 p s i a t room temperature t o 75,000 p s i a t -423" F.
2. The plane s t r a i n f r a c t u r e toughness of 2219-T87 aluminum a l - loy i s a l so r e l a t i v e l y i n s e n s i t i v e t o t e s t i n g temperature . s t r a i n f r a c t u r e tou hness v a r i e s from 36,100 p s i a t room tempera- t u r e t o 40,300 p s i $m a t -423" F.
The plane
30
0 q
In t
0 t
cn f?
0 n
8
0 v
.In
- g
cn ' 9
-0 3 0
ro
a, 5 w 0
c 0 rl U
2 El Fr m m m h U 4 rl rl P a i m ua, (03 c 4 H a
3 $4 0 0
W H %4
a @ a , U
? sa, ucn m & $ 4
5 4 na,
" 2 4 a i m o m 4 a , u$4 -rl u & c n u
d w o 0 4
U d o oa, .rl cn U a i m r l m $40
3 0
a\ 4
a, & El 80 .rl Fr
31
3 . The plane s t r a i n f r a c t u r e toughness of 2219-T87 aluminum a l - loy is s e n s i t i v e t o the o r i e n t a t i o n of the specimen i n the p l a t e . Those specimens o r i en ted s o t h a t the crack propagat ion is perpendicular t o the r o l l i n g plane show higher va lues of plane s t r a i n f r a c t u r e toughness. This e f f e c t has been a sc r ibed t o delaminat ion i n the p l a t e .
4 . The plane s t r a i n f r a c t u r e toughness of t he one inch t h i c k 2219-T87 aluminum a l l o y p l a t e i s somewhat lower than t h e value obtained f o r the one-half inch t h i c k p l a t e .
32
MFEFLENCES
1. G. B . Epsy, M . H . Jones , and W. F. Brown, Jr., "Factors Inf luenc ing F rac tu re Toughness of Sheet Alloys f o r Use i n Lightweight Cryogenic Tankage," Spec ia l Technical Pub l i ca t ion No. 302, ASTM; Symposium on Evalua t ion of Metallic Mate r i a l s i n Design f o r Low Temperature Serv ice , 1961.
2. C. M. Carman, J . W. Forney, and J . M. K a t l i n , "Plane S t r a i n Frac- t u r e Toughness of 5A1-2.5Sn ELI Ti tanium a t Room and Cryogenic Temperatures, Frankford Arsenal Report R-1796 (NASA CR-54296), Apr 66.
3. A. A. G r i f f i t h , "The Phenomena of Rupture and Flow i n Sol ids ," Ph i losoph ica l Trans of t he Royal Socie ty of London, Vol 221, 1920.
4 . G. R. I rwin , "Fracture Dynamics," ASM, F rac tu r ing of Metals, pp 147-166, 1947.
5. C. E. I n g l i s , "S t resses i n a P l a t e Due t o the Presence of Cracks and Sharp Corners ,I1 Proceedings of t he I n s t i t u t e of Naval Archi- t e c t s , Vol 60, 1913.
6. G. R. I rwin , "Relat ion of Crack Toughness t o P r a c t i c a l Applicat ions," Welding Journa l Research Supplement, Nov 1962.
7. R. W. Boyle, A. M. Su l l i van , and J . M. K r a f f t , "Determination of P lane S t r a i n F rac tu re Toughness wi th Sharply Notched Sheets," Welding Journa l , Vol 41, No. 9 , Research Supplement 428-s t o 432-s, 1962.
8. A . M. S u l l i v a n , "New Specimen Design f o r Plane S t r a i n F rac tu re Toughness Tests ,'I Materials Research and Standards, J a n 1964.
9. C. M. Carman, D. F. Armiento, and H. Markus, "Plane S t r a i n Frac- t u r e Toughness of High-strength Aluminum Alloys," ASME Paper No. 64-WA/Met-ll, Dec 1964.
10. H. F. Bueckner, I n t e r n a l Reports of the General E lec t r ic Company, Schoectady, N. Y.
11. 3. Gross, 2 . E. Scrawley, and W. F. Brown, Jr., "St ress I n t e n s i t y Fac to r f o r a S ingle Edge Notched Tension Specimen by Boundary Co l loca t ion of a S t r e s s Function," Lewis Research Center , NASA Report TN-D2603 , J a n 65.
12. P. M. Lorenz and C. F. T i f fany , "Fracture Toughness and S u b c r i t i c a l Flaw Growth C h a r a c t e r i s t i c s of Sa tu rn SI-C Tankage Materials," Boeing Document No. D2-22802 (Contract WAS8-5608) , Apr 1964.
33
DISTRIBUTION
Nat iona l Aeronautics & Space Adm Lewis Research Center 2100 Brookpark Rd Cleveland , Ohio 44135
1 At tn : Contract ing O f f i c e r MS 500-210
8 Attn: Liquid Rocket Technology B r MS 500-209
1 Attn: Tech Report Control Of f i ce MS 5-5
1 At tn : Technology U t i l i z a t i o n Off . MS 3-16
2 At tn : AFSC Lia ison Of f i ce MS 4-1
2 At tn : Library
NASA, F l i g h t Research Center At tn : L ibrary P . 0. Box 273 Edwards, C a l i f . 93523
NASA, Langley Research Center Attn: L ibrary Langley S t a t i o n Hampton, Va. 23365
NASA Manned Spacecraf t Center At tn : L ibrary Hous ton, Texas 7 700 1
NASA, Geo. C. Marsha l l Space F l i g h t
Hun t sv i l l e , A l a . 35812 Center
1 Attn: Library
1 Attn: R. N. Eilerman, M-P & VE-PS
1 Attn: Of f i ce , R e l i a b i l i t y & Q u a l i t y NASA, Western Operat ions Assurance, MS 500-203 At tn : L ib ra ry
2 At tn : Richard Kemp, MS 49-1
1 At tn : W. F. Brown, MS 105-1
1 At tn : J. L. Shannon, MS 105-1
1 At tn : J . E * . Srawley, MS 105-1
Nat iona l Aeronautics & Space Adm
Washington, D. C . 20546 At tn : Code RV-2 (2)
S c i e n t i f i c & Technical Information
Attn: NASA Representa t ive , Code CRT P . 0. Box 5700 Bethesda, Md. 20014
F a c i l i t y ( 6)
NASA Ames Research Center A t t n : Library Moffe t t F i e l d , C a l i f . 94035
NASA, Goddard Space Fl i .ght Center At tn : Library Greenbel t , Md. 20771
150 P i c 0 Boulevard Santa Monica, C a l i f . 90406
NASA, John F. Kennedy Space Center A t t n : Libra ry Cocoa Beach, F l a . 32931
J e t Propuls ion Laboratory 4800 Oak Grove Drive Pasadena, C a l i f . 91103
1 At tn : L ib ra ry
1 At tn : L . D . J a f f e , Ch, Mat'lks Sec
AFFTC (FTAT - 2) Edwards AFB, C a l i f . 93523
Wright -Pa t te rson A i r Force Base Ohio, 45433
1 At tn : AFML(MAAE)
1 At tn : AFML(MAAM)
34
I -
Commanding General Aeronaut ica l Systems Divis ion Wright-Pat terson AFB , Ohio 45433
1 Attn: WCLPRX
1 At tn : M r , Geo. C. Young Appl ica t ions Laboratory Code ASRECEM-1
Commander AF B a l l i s t i c Missile Div (ARDC) At tn : WDSOT P , 0. Box 262 Inglewood, C a l i f .
RTD (RTNP) B o l l i n g A i r Force Base Washington, D. C. 20332
Arnold Engineer ing Development C t r At tn: AEOm A i r Force Systems Command Tullahoma, Tenn. 37389
A i r Force Systems Command SCLT/Capt S . W. Bowen Andrews A i r Force Base Washington, D. C . 20332
A.F. Rocket P r o p e l l a n t Laboratory Edwards , C a l i f . 93523
1 At tn : AFRPL (RPC)
1 At tn : AFRPL (PRM)
1 At tn : AFRPL (RPR)
O f f i c e of Research Analyses (OAR) Attn : RRRT Holloman AE'B, New Mexico 88330
AF O f f i c e of S c i e n t i f i c Research At tn : SREP, D r . J. F. Masi Washington, D. C.
Commander A i r Un ive r s i ty L ib ra ry , USAF Maxwe 11 AFB , A l a .
Hq, U.S . Army Materiel Command Washington, D. C. 20315
1 Attn: AMCRD-RS-CM-M
1 At tn : AMCRD-RC
Hq, U.S . Army Materiel Research Agency At tn : AMXMR-OPT Watertown, Mass. 02172
Hq, U . S . Army Miss i l e Command Redstone Arsena l , A l a . 35808
1 At tn : Chief , Document Sec t ion Redstone S c i e n t i f i c I n f Center
5 Attn: M r . C. H. Martens
Commanding O f f i c e r U . S . Army Research Off ice-Durham Box CM, Duke S t a t i o n Durham, N . C . 27706
1 Attn: Technical Information Div
1 At tn : M r . Davis, Metallurgy ti Ceramics
Commanding O f f i c e r Ammunition Procurement & Supply Agency At tn : SWAP-Mat'ls Engineering J o l i e r , Ill. 60436
Commanding General U.S. Army Tank-Automotive Center U .S. Army Mobi l i t y Command Warren, Mich. 48089
2 Attn: SMOTA-RCM.l
1 Attn: SMOTA-RCS
Hq, U . S . Army Weapons Command At tn : Laboratory Rock I s l a n d , Ill. 61202
Commanding O f f i c e r Harry Diamond Labora to r i e s At tn : AMXDO-TIE Washington, D. C. 20438
35
H q , U . 3 . Army Munitions Command Dover, N . J. 07801
Commanding O f f i c e r B a l l i s t i c Research Laboratory Aberdeen Proving Ground, Md. 21005
1 At tn : Technical Information Div 1 Attn: AMXBR-1
1 At tn : AMSMU-S, D r . J.V.R.Kauffman 1 At tn : ORDBG-BLI
1 Attn: AMSMU-I, M r . R.M. Schwartz
1 At tn : AMSMU-E, M r . C . H . S t a l e y
1 At tn : AMSMU-LA, USAF L i a i s o n Off .
1 At tn : AMSMU-LC, CDC L ia i son Off
1 At tn : AMSMU-LM, USMC L i a i s o n Off .
commanding O f f i c e r P ica t inny Arsenal Attn: M r . J . Matlack
Dover, N. J . 07801 P l a s t i c s & Packaging Lab.
Commanding O f f i c e r Rock Is land Arsenal At tn : M r . Robt. Shaw, Laboratory Rock I s l and , Ill. 61202
Commanding O f f i c e r Spr ing f i e ld Armory Attn: M r . E . Abbe Spr ing f i e ld , Mass. 01101
Commanding O f f i c e r Watertown Arsenal Attn: Technical Information S c Watertown, Mass. 02172
E
Commanding O f f i c e r At tn : STEAP-DS-TU, M r . W . P l e s s Aberdeen Proving Ground, Md. 21005
Commanding Of f i ce r ( 2) At tn : Technical Library , Bldg 313 Aberdeen Proving Ground, Md. 21005
U.S. Army Coating & Chemical Lab. Attn: D r . C . P i c k e t t Aberdeen Proving Ground, Md. 21005
Commanding O f f i c e r Army Tra in ing Command School U .S. Army Combat Development Command Attn: Combat Support Group Aberdeen Proving Ground, Md. 21005
Commanding General Engineering R&D Laboratory U.S. Army Mobi l i ty Command F o r t Be lvo i r , Va. 22060
Bureau of Naval Weapons Department of the Navy Washington, D. C. 20360
1 At tn : DLI-3
1 Attn: RMMP-2
1 At tn : RMMP-4
1 At tn : RRRE-6
1 At tn : Ad3, Technica l L ib ra ry
1 At tn : Res6 ( M r . Matesky)
1 At tn : ReW3a
Of f i ce of Chief S c i e n t i s t Department of t he Navy At tn : M r . H . B e r n s t e i n Washington 25, D. C.
Commanding O f f i c e r U . S . Naval P r o p e l l a n t P l a n t At tn : Research & Development Dept Ind ian Head, Md. 20640
Commanding O f f i c e r Lake C i t y Army Ammunition P l a n t Independence, Mo . 64050
U .S . Naval Research Laboratory Washington, D. C. 20390 U .S . Naval Pos tgraduate School
1 Attn: D r . J. M. K r a f f t (Mechanics) Monterey, C a l i f . 93900
Superintendent
Naval Academy
1 Attn: Dr. P. King, Code 6000
1 Attn: R. W. Carhar t
1 Attn: D r . G . R . I rwin, Code 6200
1 At tn : M r . J . A. Kies, Code 6210
1 At tn : D r . B . F. Brown, M e t a l l u r g i c a l Div
5 At tn : Mrs. Baster, Code 2027
Commander U.S. Naval Missile Center Attn: Technical L ib ra ry P o i n t Mugu, C a l i f . 93041
D i r e c t o r S p e c i a l P r o j e c t s Of f i ce Department of t h e Navy Washington, D. C. 20360
U .S . Naval Weapons Laboratory Dahlgren, V a . 22448
Commanding O f f i c e r Of f i ce of Naval Research 1030 E. Green S t . Pasadena, C a l i f . 91101
Cormanding O f f i c e r U.S. Naval Underwater Ordnance S t a t i o n Attn: W. W. Ba r t l e t t Newport, Rhode I s l and , 02844
Defense Documentation Center (20) Cameron S t a t i o n Alexandria , Va 22314
Off ice of the Di rec to r of Defense
At tn : Dr . H.W,Schulz, (Chem Technology)
Washington, D. C. 20301
Research & Engineering
Of f i ce of A s s t . D i r ec to r
Defense Metals Information Center B a t t e l l e Memorial I n s t i t u t e At tn : M r . Webster Hodge 505 King Ave. Columbus, Ohio
2 A t t n : D r . H. Romine, Warhead & U .S. Atomic Energy Commission Terminal B a l l i s t i c s Lab. Attn: Of f i ce of Technical Information
Extens ion 1 At tn : Technica l L ib ra ry
Commander U.S. Naval Ordnance Laboratory S i l v e r Spring, Md. 20910 Attn: M r . J. A. Bennett
1 At tn : L ib ra ry Washington, D. C. 20234
P.O. Box 62 Oak Ridge, Tenn. 37831
Nat iona l Bureau of Standards
M e t a 1 l u r gy Div .
1 Attn: E. L. Criscuolo U .S. Department of the I n t e r i o r
Commander (Code 753) Attn: M . M e Dol inar , U.S. Naval Ordnance T e s t S t a t i o n At tn : Technical L ib ra ry China Lake , C a l i f . 93557 4800 Forbes Avenue
Bureau of Mines
Reports L i b r a r i a n Ex p 1 os ive s Re s e a r c h Cent e r
P i t t s b u r g h , Pa. 15213
37
Aerojet-General Corporat ion 11711 S. Woodruff Ave. Downey, C a l i f . 90241
At tn : F. M. West, Ch L i b r a r i a n
Aerojet-General Corporat ion P. 0. Box 296 Azusa, C a l i f . 91703
At tn : E . J . Morgan, L i b r a r i a n
Aero j e t -Genera 1 Corporation P. 0 . Box 1947 Sacramento, C a l i f . 95809
1 At tn : Tech Library 2484-2015A
1 At tn : C . E . Hartbower
Ae r onu t r on i c Div i s i o n Ph i l co Corporation Ford Road Newport Beach, C a l i f . 92600
Attn: D r . L . H. L inder , Mgr Technical Information D e p t
Aeroprojects , Inc . 310 E. Rosedale Ave. West Ches te r , Pa. 19380
Attn: C . D . McKinney
Aerospace Corporat ion P . 0. Box 95085 Los Angeles, C a l i f . 90045
At tn : Library-Documents
Aluminum Company of America AECOA Research Labora to r i e s New Kensington, Pa.
At tn : M r . J . G . Kaufmann
ARO, Inc. Arnold Engineering Development C t r
B a t t e l l e Memorial I n s t i t u t e 505 King Ave. Columbus , Ohio 43201
1 At tn : Reports L ib ra ry , Rm 6A
1 Attn: G . T . Hahn
B e l l Aerosystems Box 1 Buffa lo , New York 14205
Attn: T . Reinhardt
Bethelehem S t e e l Company Homer Research Labora to r i e s Bethlehem, Pa.
Attn: M r . J . S c o t t
The Boeing Company Aerospace Div is ion P.O. Box 3707 S e a t t l e , Washington 98124
1 At tn : Ruth E . Peerenboom (1190)
3 At tn : C . W . T i f f any
1 At tn : Pau l Lorentz
Chemical P ropu l s ion Information Agency A p p l i e d Phys ics Laboratory 8621 Georgia Ave. S i l v e r Spr ing , Md. 20910
Douglas A i r c r a f t Company, Inc . Santa Monica Div i s ion 3000 Ocean Pa rk Blvd Santa Monica, C a l i f . 90405
1 At tn : M r . J . L . Waisman
1 Attn: M r . G . V . Bennet t , M i s s i l e s Arnold Air Force S t a t i o n , Tenn. 37389 & Space Systems Div.
Attn: D r . B. H. Goether t Chief S c i e n t i s t 2 At tn : B . V . Whiteson, Missiles
& Space Systems Div A t l a n t i c Research Corporat ion S h i r l e y Highway & E d s a l l Rd. E. I. duPont de Nemours & Company Alexandria , Va . 223 14 E a s t e r n Labora tory
At tn : Secur i ty Of f i ce f o r L ib ra ry Gibbstown, N . J . 08027 At tn : M r s . Al ice R . Steward
38
Esso Research and Engineering Co. Spec ia l P r o j e c t s Uni t P. 0. Box 8 Spokane, Wash. Linden , N . J. 07036 Attn: J. B. Herr
Kaiser Aluminum & Chemical Corp. Dept of M e t a l l u r g i c a l Research
1 Attn: M r . D. L. Baeder
1 Attn: V. E. Anderson
1 Attn: D. R. Sherman
FMC Corporat ion Chemical Research & Development C t r P. 0. Box 8 P r ince ton , N . J . 08540
Attn: Secur i ty O f f i c e r
General Dynamics/Astronautics P. 0. Box 1128 San Diego, C a l i f . 92112
At tn : Library & Information Serv (128-00)
Lockheed-California Company Burbank, C a l i f .
Attn: R. 0 . E. E a r l
Lockheed Missiles & Space Company Sunnyvale , Cal i f .
Attn: Richard E. Lewis
Lockheed Propuls ion Company P . 0. Box 111 Redlands, C a l i f . 92374
Attn: M i s s Belle Berlad, L i b r a r i a n
Marquardt Corporat ion 16555 Sa t icoy S t . Box 2013 - South Annex Van Nuys, C a l i f . 91404
General Dynamics/Convair Martin-Marie t t a Corporat ion San Diego, C a l i f . Middle River
Attn: W. E. Wi tze l Bal t imore , Md. Materials Research G r Attn: RIAS - M r . C. E. Thomas
General E l e c t r i c Co. Apollo Support Department P.O. Box 2500 Daytona Beach, F l a . 32015
At tn : C. Day
Hercules Powder Company Allegany B a l l i s t i c s Laboratory P -0. Box 210 Cumberland, Hd. 21501
At tn : L ib ra ry
I n s t i t u t e f o r Defnese Analyses 400 Army-Navy Drive Ar l ing ton , Va. 22202
At tn : C l a s s i f i e d L ib ra ry
I n t e r n a t i o n a l Nickel Co. 117 W a l l S t . New York, N.Y.
A t tn : M r . A. Graae
Mart in-Mariet ta Corporat ion Denver Div is ion Denver, Colo. 80201
Attn: F. R. Schwartzberg
Minnesota Mining & Manufacturing Co. 900 Bush Avenue S t . Pau l , Minn. 55106
V i a : H. C. Zeman, Secur i ty Adminis t ra tor ( f o r Code 0013, R&D)
North American Avia t ion , Inc. Space & I n f . S y s t e m Divis ion 12214 Lakewood Blvd Downey, C a l i f . 90242
Attn: Technical Information Center D/096-722 (AJO1)
Ordnance Engineering Associates , Inc . 1030 E. North Ave. DesPlaines , Ill. 60016
At tn : M r . A. D. Kafadar
39
Propuls ion Engineering Div (D .55-11) Thiokol Chemical Corporat ion Lockheed Missiles & Space Company 1111 Lockheed Way Sunnyvale, C a l i f . 94087
Reynolds Metals Company Metals Research Div is ion 4th & Canal S t s . Richmond , V a .
Attn: Mr. R. Zinkham
Rocket Research Corporation 520 S. Por t land S t . S e a t t l e , Washington 98108
Rocket dyne 6633 Canoga Ave. Canoga Park, C a l i f . 91304
Attn: Library Dept 596-306
Rohm & Haas Company Restone Arsenal Research Div Hun t sv i l l e , Ala 35808
Attn: L ib ra r i an
Space Technology Laboratory , Inc . 1 Space Park Redondo Beach, C a l i f . 90200
Attn: STL Technical L ib ra ry Documents Acquis i t ions
Thiokol Chemical Corporat ion Alpha Divis ion, Huntsv i l le P l a n t Hun t sv i l l e , Ala. 35800
Attn : Technical D i rec to r
Tbiokol Chemical Corporat ion Brunswick Ga.
Attn: Mr. Craig
Thiokol Chemical Corporation Elkton Divis ion Elkton, Md. 21921
Attn: L ib ra r i an
Thiokol Chemical Corporation React ion Motors Div. Denvi l le , N . J. 07834
Attn : L ib ra r i an
Rocket Operations Center P. 0. Box 1640 Ogden, Utah ; 84401
Attn: L i b r a r i a n
Thiokol Chemical Corporat ion Wasatch Div is ion P . 0. Box 523 Brigham Ci ty , Utah 84302
Attn: L ib ra ry Sec t ion
40
Titanium Metal Corporat ion of America 233 Broadway New York 7 , N . Y .
At tn: Ward Winkler
TRW Incorporated 23555 Euclid Avenue Cleveland, Ohio 44117
Attn : E. A . Steigerwald
United A i r c r a f t Corporat ion Corporat ion L ib ra ry 400 Main S t r e e t East Har t ford , Conn. 06118
A t t n : M r . Daivd Rix
United A i r c r a f t Corporat ion P r a t t & Whitney F l o r i d a R&D Center P . 0. Box 2691 W. Palm Beach, F l a . 33402
A t t n : Libra ry
United A i r c r a f t Corporat ion United Technology Center P 0. Box 358 Sunnyvale, C a l i f . 94088
Attn : L i b r a r i a n
United S t a t e s S t e e l Co. Ap p 1 i e d Re s e a r c h Lab o ra to ry Monroevi l le , Pa.
At tn : M r . J . Hodge
Vanadium Alloys S t e e l Co. La t robe , Pa .
At tn : D r . G . Roberts
Westinghouse Research Labora tor ies Beulah Road, Churchi l l Boro P i t t s b u r g h , Penna. 15235
1 Attn: G. 0. Sankey
1 At tn : J. H. B i t le r #
C a l i f o r n i a I n s t i t u t e of Technology 1201 E , C a l i f o r n i a Blvd Pasadena, C a l i f .
At tn: S e c u r i t y O f f i c e r
Carnegie I n s t i t u t e of Technology Department of C i v i l Engineering P i t t s b u r g h , Pa .
At tn : Robert B . Anderson
Department of Metallurgy C a t h o 1 i c Univers i ty Washington, D. C .
At tn : D r . E. P . K l i e r
U n i v e r s i t y of Chicago Co 1 lege of Engineering D e p t of Material Engineering Chicago C i r c l e , Box 4348 Chicago, Ill. 60680
At tn : Prof . W . Rostoker
P r o f . H . H. Johnson Dep t of Engineering Mechanics
And Materials Thurston H a l l Corne 11 Univers i ty I t h a c a , N . Y.
IIT Research I n s t i t u t e Techno logy Center Chicago, Ill. 60616
1 A t t n : C. K. Hersh Chemistry Div .
1 A t t n : K. E , Hofer
Department of Mechanics Lehigh Univers i ty Bethlehem, Pa.
A t t n : D r . Paul C. P a r i s
Department of Metal lurgy Lehigh Univers i ty Bethlehem, Pa .
Attn: Prof . Wayne K r a f t
Purdue U n i v e r s i t y L a f a y e t t e , Ind. 47907
Attn: M. J. Zuarow
D r . Earnest Chi l ton Stanford Research I n s t i t u t e Menlo Park, C a l i f .
Syracuse Univers i ty Research I n s t i t u t e Department of Metallurgy Syracuse, N. Y .
1 Attn: H. W. Liu
1 Attn: Volker Weiss
P r o f . Samuel Carpenter Swart hmo re C o 1 l e ge Swarthmore, Pa .
Univers i ty of Denver Denver Research I n s t i t u t e P . 0. Box 10127 Denver, Colorado 80210
Attn: S e c u r i t y Off ice
P r o f . H. T . Corten College of Engineering Univers i ty of I l l i n o i s Urbana, Ill.
Reproduction Branch FRANKFORD ARSENAL Date P r i n t e d : 9/27/66
41
UNCUS SIF XED
ORIGINATIN G A C r l V l T Y (Corporate author)
FRANKFORD ARSENAL P h i l a d e l p h i a P a . 19137 2 a R E P O R T S E C U R I T Y c L A S S I F I C A T I O N
Unc las s i f i e d
(SMUFA L3300) 2 6 G R O U P
N J A
AMCNS 5900.21.11603
NASA Purchase Order C6860A b. PROJECT NO.
0.
August 1966 ,
d.
1 2
I R-1821
Sb. p T H C R R PORT NO(S) (Any o t h o r n u m k n rh4t may bo 4arlmod 14 nmrJ
NASA CR-54297 I
10. A V A I L ABILITY/LIMITATION NOTICIIS
Dis t r ibu t ion of t h i s report is unlimited.
I 1 * SUCCL LMLNTARY N O T W (2. SCONSORINO MILITARY ACTIVITY
NASA, Lewis Research Center
IS, ABSTRACT
The t e n s i l e propert ies and plane s t r a i n f racture toughness of 1/2 and 1 inch th ick 2219-T87 aluminum al loy have been determined as a function of t e s t i n g a t temperatures €rom room t o -423' F. The t e n s i l e and y i e l d s t rengths of t h i s ma- t e r i a l ahow a gradual increase as the t e s t i n g temperature is decreased t o -423" F while the elongation and reduction of area remain e s s e n t i a l l y unchanged.
The plane s t r a i n f rac ture toughness of t h i s material is r e l a t i v e l y insensi t ive t o t e s t i n g temperature and shows only a s l i g h t increase with decreasing t e s t i n g temperature, Specimns machined so tha t the crack propagation is perpendicular t o the r o l l i n g plane show somewhat higher values o€ plane s t r a i n f rac ture toughnes than when crack propagation is p a r a l l e l t o the r o l l i n g plene. The plane s t r a i n f r a c t u r e toughness of the 1 inch th ick 2219-T87 aluminum a l loy was somewhat lower than t h a t of the 1/2 inch th ick plate.
I l l u e t r a t i v a examples a re presented OB using these parameitere i n design.
____. ---____ S e c u r i t y Classif icat ion
14. K E Y WORDS
_ _ _ ~ ~ - __
INSTWCTION!
I. ORIGINATING ACTIVITY Enter the name and address ,f the contractor, subcontractor, grantee, Department of De- ‘ense ectivity or other organization (corporate author) issuing he report. !a. REPORT SECURITY CLASSIFICATION Enter the over- 311 security classification of the report. Indicate whether ‘Restricted Data” is included ance with appropriate security regulations. !b. GROUP: Automatic downgrading is specified in DoD Di- ect ive 5200.10 and Armed Forces Industrial Manual. Enter he group number. Also, when applicable, show that optional narkings have been used for Group 3 and Group 4 a s author- zed .
1. REPORT TITLE: Enter the complete report t i t le in all :apital letters. T i t l e s i n all cases should be unclassified. f a meaningful t i t le cannot be selected without classifica- ion, show ti t le classification in all capi ta ls i n parenthesis mmediately following the title. 1. DESCRIPTIVE NOTES If appropriate, enter the type of eport, e. g., interim, progress, summary, annual, or final. ;ive the inclusive dates when a specific reporting period is :overed. i. AUTHOR(S): Enter the name($ of author(s) a s shown on ir in the report. f military, show rank and branch of service. The name of he principal author is an absolute minimum requirement, ). REPORT DATE Enter the date of the report a s day, nonth, year; or month, year. If more*than one date appears m the report, u s e date of publication. ‘a. TOTAL NUMBER O F PAGES T h e total page count ihould follow normal pagination procedures, i. e., enter the lumber of pages containing information ‘b. NUMBER OF REFERENCES Enter the total number of eferences cited in the report. la. CONTRACT OR GRANT NUMBER. If appropriate, enter he applicable number of the contract or grant under which he report was writ ten Ib, &, & 813. PROJECT NUMBER: Enter the appropriate nilitary department identification, such a s project number, ,ubproject number, system numbers, task number, etc. ‘a. ORIGINATOR’S REPORT NUMBER(S): Enter the offi- :ial report number by which the document will b e identified ind controlled by t h e originating activity. Th i s number must le unique to this report. Ib. OTHER REPORT NUMBER(S): If t he report has been issigned any other report numbers (either b y the originator ir b y the sponsor), also enter t h i s number(s).
Marking is to b e in accord-
Enter l a s t name, first name, middle initial.
R O L E
L I N K A L I N K B L I N K C
R O L E R O L E W T W T W T
10. AVAILABILITY/LIMITATION NOTICES Enter any lim- i tations on further dissemination of the report, other than those imposed by security classification, using standard statements such as:
(1)
(2)
(3)
“Qualified requesters may obtain copies of t h i s report from D D C ”
“Foreign announcement and dissemination of this report by DDC is not authorized.”
“U. S. Government agencies may obtain copies of t h i s report directly from DDC. Other qualified DDC use r s shall request through
t ,
(4) “U. S. military agencies may obtain copies of this report directly from DDC Other qualified use r s shall request through
I ,
(5) “All distribution of th i s report is controlled. Qual- if ied DDC users shall request through
3 ,
If the report h a s been furnished to the Office of Technical Services, Department of Commerce, for s a l e to the public, indi c a t e t h i s fact and enter the price, i f known. 11. SUPPLEMENTARY NOTES: Use for additional explana- tory notes.
12. SPONSORING MILITARY ACTIVITY Enter the name of the departmental project office o r laboratory sponsoring ( p a p ing for) t h e research and development. Include address.
13. ABSTRACT; Enter an abstract giving a brief and factual summary of the document indicative of the report, even though i t may a l so appear elsewhere in the body of the technical re- port. shal l be attached.
ports b e unclassified. Each pa rapaph of the abstract shall end with an indication of the military security classification of t he information in the paragraph, represented a s (TS), (s), (C), or (U).
How- ever, the suggested length i s from 150 to 225 words. 14. KEY WORDS: Key words are technically meaningful terms or short phrases that characterize a report and may be used as index entr ies for cataloging the report. se lected so that no security classification i s required. Iden- f iers , such a s equipment model designation, trade name, -nili- tary project code name, geocraphic location, may be used a s key words but will b e followed by an indication of technical context, optional.
If additional space i s required, a continuation sheet
It is highly desirable that the abstract of c lass i f ied re -
There i s no limitation on the length of the abstract .
Key words must be
T h e assignment of l inks, rules, and weights i s
Security Classification ~~ _____
