: t I i

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o ?7 -1 8 ,t 0;3CONCEPTS FOR IMPROVING TURBINE DISK INTEGRITY

Albert K aufman

NASA Lewis Research Center

SUMMARY

The trend toward higher turbine-blade tip speeds and inlet gas temperatures makes "'

It increasingly difficult to design reliable turbine disks that can satisfy the life and per-

formance requirements of advanced commercial aircraft engines. Containment devices

to protect vital areas such as the passenger cabin, the fuel lines, and the fuel tanks

against high-energy disk fragments would impose a severe performance penalty on the

engine. The approach taken in this study waz to use advanced disk structural concepts

to improve the cyclic lives and reliability of turbine disks. Analytical studies were con-

ducted under NASA contracts by the General Electric Company and Pratt & Whitney Air-

craft to evaluate bore-entry disks as potential replacements for the existing first-stage

turbine disks in the CF6-50 and JTSD--17 engines. Results of low-cycle fatigue, burst,fracture mechanics, and fragment energy analyses are summarized for the advanced

disk designs and the existing disk designs with both conventional and advanced disk ma-

terials. Other disk concepts such as composite, laminated, link, multibore, multidisk,and spline disks were also evaluated for the CF6-50 engine.

INTRODUCTION

A disk burst is one of the most catastrophic failures possible in an aircraft engine.

Flight failures of disks in commercial airliners have caused fires, rupture of fuel tanks,

penetration of passenger cabins, wing damage, ingestion of disk fragments by other en-gines, and aircraft control problems (ref. 1).

Aircraft engine companies generally endeavor to use conservative design practices

and modern quality control procedures in producing turbine disks. Itowever, failures

occur because of design errors, undetected manufacturing defects, uncontrollable oper-

ating factors, errors in engine maintenance and assembly, and failure of other engine

components. To attempt to design turbine disks to preclude failure from any of these

causes would result in prohibitively low allowable stresses. Contahlment devices to

protect vital areas of the aircraft against high-energy disk fragments would impost

severe performance penalties on the engine.

The approach t_,,Xenin this program was directed toward lmp_ving turbine disk re-

liability by using more adv_mced structural concepts to increase low-cycle fatigue life.

to impede crack propagation, and to reduce fragment energies that could be generated

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in the event of a disk failure. This paper reports the results of NASA-sponsored ana- I

lyUcal studies by the General Electric Company and Pratt & Whitney Aircraft (refs. 2 1

and 3) to evaluate bore-entry disks as potential replacements for the existing first-stage i

turbine disks in the CF6-50 and JTSD-17 engines, respectively; these engines were

selected because of their extensive use in commercial passenger aircraft. Other con- Icepts such as composite, laminated, and multidtsk designs were also studied for the op- ierating conditions of the CF6-50 engines.

The bore-entry disks were compared with the existing disks (henceforth called the

"standard disks") on the basis of cycles to crack initiation and overspeed capability for "_'initially unflawed disks and on the basis of cycles required to propagate initial flaws to

failure. Comparisons were also made of th_ available kinetic energies of possible burstfragments. All of these comparisons were also made for the standard disk with the ma-

terial of the bore-entry disk so that improvements resulting from changes in material

properties could be distinguished from those resulting from structural design changes.

DISK CONCEPTS

CFa-50 Turbine Disk Designs

The standard disk and the disk concepts considered as potential replacements are

illustrated in figure 1. The standard disk (fig. 1(a)) is machined from an inconel 718

(Inc-718) forging. Local bosses on both sides of the disk provide reinforcement around

the bolt holes to increase the low-cycle fatigue life at the hole rims. Cooling air from

the compressor is channeled through the shaft, cools the disk bore, is pumped upradially between the stage 1 and 2 rotors, cools the aft side of the disk between the bolt

holes and rim, and then enters the blades th,Dugh openings in the dovetails.

The bore-entry disk (fig. 1(b)) is a two-part disk of integral construction. The two

disk halves are connected by radial webs for channeling coolant up the center of the disk

from the bcre to the blades. Among the advantages of the bore-entry concept are im-

proved cooling effectiveness, reduced axial thermal gradients, and increased resis-

tance to crack propagation in the axial direction. One of the main attractions of the

bore-entry concept for the CF6 program was that it lent itself to a redundant construc-

tion where the disk would be overdcsigncd so that if half was failing, the undamaged disk

half would be able to assume a larger portion of the load and sustain the damaged D_rt;

however, this would require a substantial tncrcase in total disk weight. The integral

bore-entry disk would be fabricated from a single-piece forging of Ren_ 95 alloy with the

material between disk halves removed by electrochemical machining.

The composite disk (fig. 1 (e)) uses high-strength filament or wire hoops to provide

most of the load-carrying ability of the disk except at the dovetail attachments. The

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!._ hoops would have to be pretensioned in order to assure an even load distribution amongi the filaments; this could be accomplished by filament winding, by interference fitting,

_: or by the selection of filament and matrix materials so that the desired hoop pretension

• would be applied by differential thermal expansion under engine operating conditions.

. In the laminated design (fig. l(d)), a disk is constructed by bolting together a large

number of sheet-metal laminates. A stepwise variation in thickness provides more

:_ laminates at the rim and bore but leaves gaps between laminates in the web region, h,

,: the link design (fig. l(e)) a disk is constructed of pinned sheet-metal link segments.

Both the laminate and link concepts are directed toward low-cost fabrication, isolation

:: of propagating cracks, and generation of small burst fragments rather than toward ira-

: _i:i proving disk life.

'ii The multibore disk (fig. l(f)) separates the highly stressed bore region into a num-ber of circumferential ribs in order to prevent a crack or flaw at the bore from propa-

gating axially. At the ends of the ribs, the tangential stresses due to centrifugal loading

, would be less and, therefore, the crack propagation rate should be slower than at the

:! bore of the standard disk.

: The purpose of the multidisk design (fig. 1(g)) is to obtain improved disk cooling

and to provide for a redundant construction by transference of loads from a failed disk

member to the undamaged ones through the bolts. The spline disk (fig. l(h)) is essen-

_: tially a two-piece design where the members are coupled through splines on their center

faces. In order to counter the tendency of each disk half to straighten out due to the

_ lack of axial symmetry, the splines would have to be radially interlocked through pins.

. The mechanical coupling of the multidisk and spline designs prevents cracks in one disk

' : member from propagating to another.

_:-:'" These concepts are described in more detail in reference 2._.

JT8D-17 £urbine Disk Designs

The standard disk shown in figure 2 (a) is machined from a Waspaloy forging. Cool-

! ing air is bled from the combustion chamber liner and discharged at high velocity: i

through nozzles toward the front side of the disk near the rim. The cooling air is de-

livered to the blades through angl( _ holes at the disk rim. These holes result in cllip-

• ttcal exit openings with high stress concentrations, these are the limiting low-cycle

fatigue locations.

A split-bonded, bore-entry concept was selected as a possible replacement for the

_ standard disk. As with the integral bore-entry disk (fig. l(b)) for the CF6-50 turbine,

cooling air would be introduced at the bore, would be pumped up radi.111y through than-

• _!. nels formed by radial webs, and would enter the blades through openings in the bases.

i-li_ The two halves of the bonded bore-entry disk would be fabricated from separate forgings

425

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: l

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of Astroloy and diffusion brazed together at the center surfaces of the radial webs.

Dovetail broaching and final machining operations would be performed on the bonded disk

assembly. _l_e emphasis in the design of the bonded bore-entry disk was on improving

the cyclic life without providing redundarcy or increasing the disk weight.

DESIGN CONDITIONS

Design properties of the materials for the standard and bore-entry disks are pre-Ira.,

sented in table I. The simplified flight cycles used for the cyclic heat transfer and

stress analyses are shown in figure 3 for the CF6-50 engine and in figure 4 for the

JTSD-17 engine. The flight cycle shown in figure 4 was the cycle used in the original

design of the first-stage turbine disk for the JT8D-17 engine. The analytical methodsarc discussed in references 2 to 4.

DISCUSSION OF RE_ULTS

Preliminary Analyses of CF6-50 Disk Concepts

The results of preliminary analyses of the seven candidate design disk concepts are

summarized in table II. Two of the designs, the laminated and link disks, proved to

have excessive mechanical stresses and to be unsuitable for the CF6 operating condi-

tions. The multibore design exhibited high transient thermal stresses in the region

above the bore rims; therefore, the desired benefit of this design in retarding the prop-

agation of rib flaws was not fully realized. Analysis of the multidisk design under var-ious failure conditions revealed that the bolts could not contain a failed outer disk and

that a crack in a center disk would reach critical length before the load cx)uld be redis-

tributed to the undamaged members.

Only the bore-entry, composite, and splint disks appeared suitable for the CF6-50

turbine disk applications. From the standpoint of strength-to-density ratio, the compo-

site disk was the most promising c_ncept. Itowever, the composite design is furthest

removed from the current state,-of-the-art of fabrication and material processing tech-nology of any of the concepts c_nsidered. Because of the considerable fabrication de-

velopment that would be required, the composite disk was not further considered. The

spline disk presented special problems in analysis because the load distribution among

the splincs is dependent on the fabrication tolerances and it is not readily apparent how

the loading would be redistributed should one disk-half fail. _111cintegral construction

of the bore-entry disk gives more assurance that the loading due to a failed disk mem-

ber would be more evenly redistributed on the undamaged member. The intcgr_fl bore-

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I, i

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entry concept was, therefore, selected for more detailed study to replace the CF6-50standard disk•

Analyses of CF6-50 Standard :rod Be re,- Entry Disks

i

The rim and bore average temperature responses during the flight c vc'le of the st_m-

dard and bore-enttT disks are shown in figlare 5. Average effective stresses are also

. indicated at the start and end of takeoff, climb, cruise, and thrust reversal on descent.

•' In both disks the maximum rim and bore temperatures occurred at the end of takeoff _md..:. climb, respectively; the maximum stresses also occurred in the bore at the end of

' climb.

Bore temperatures in the bore-entry disk are only slightly lower thm_ bore

temperatures in the standard disk since the bore is cooled in both cases. Rim tempera-

tures were somewhat higher in the bore-entry disk because the coolant picks up some

heat fxx)m the center faces of the disk, whereas the coolant only comes into contact withthe sides of the standard disk near the rim.

Figure 6 shows the predicted cyclic lives to crack initiation in the initially unflawed

standard and bore-entry disks. The limiting fatigue life of .q0 000 cycles in the !no-718

standard disk was at the aft dovetail post rabbet, where the side plate is fastened to the

disk. This location was not further considered in the study because fragment generation

due to failure would be limited to the dovetail post and adjacent blades. The next most

critical location in the Inc-718 standard disk was at the bore with a predicted crack ini-

tiation time of 63 000 cycles. The Initial FAA certified life of the first-stage tuflJine

disk was 7800 cycles based on one-third of the minimum design life fox" the original de-

sign cycle, which was somewhat different fxx)m the simplified cycle used in this study,

this FAA approved life is subject to increase as the result of gt_)m_(l tests of three fleet

leader engines.

Calculated crack initiation lives for the Reng, 95 standard mad bore-entry disks were

over 100 000 cycles. Since the crack Initiation analyses were based on minimum guar-

anteed material plx)perties, it is evident that even the standard disk is vet5" conserva-

tively designed provided the design conditions are not exceeded and the disks al_, ini-

tially unflawed.

qqle cyclic lives for cracks propagating from initial sere!elliptical surface flaws

0.635 centimeter (0.250 in.) by 0. 211 centimeter (0.08.q in.) to critic'fl crack size :lr(,

shown in figure 7 for the most critical locations in the three disk._. .Manuf:tcturing flaw._

of this size should be readily detectal)lc by medea1 nondest rue'tire (,vnlu:dion tcclmiquc.,.

ltowevcr, in the pa_t, large defects in turbine disks have occasionally t,st'aped (h,t(,ction

fllt_ugh hum:m ern)r and have c.msod problems in sorer, military cngint,s in flight.

The most critical locations for flaws were at the dovetail slot bottom in the Inc=71_

•) -*

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. i : ,

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8tancard disk and at the bore in the llen_, 95 standard and boro-entry disks. Although

file bore-entry disk showed _tn impl_avement in the minimum crack propagation life of

mor,, titan :t00 lmr,,t,n! _ls ,.,,mlmt_,d wtlh lilt' lnt.-TIg standard disk, part of this increase

wall due to the superior strength propertl(,s of the lien6 95 :alloy. If the effect of differ-ent materials was eliminated by comp:lz_ng the bore-flawed boro-entl.'y and lien6 95

standard disks, the Improvement in crack propagation life resulting solely flv0m the

structural ch,'mge was 136 percent.

The crack propagation lives given in figure 7 for the Inc-718 standard disk with a

dovetail slot bottom flaw and the bore-entry disk with a bore flaw are only 5 and 20 per- --

cent of the I,'AA certified life of the disk. llowever, the probability of such large flaws

occurring at eritie,'d locations trod passing modern inspection procedures is statistically

remote. Of greater significance is that a substanU,'d improvement in the crack propaga-

tion life is added inburance against sudden catastrophic failure due to unforeseen design,

manufacturing, maintenance, or operating problems. The overspeed burst margins of

the boro-entry disk were 18 and It percent greater than for the Inc..718 and Rcn6 95

standard disks, respectively.

The redundmlt construction of the bort.,-entry disk resulted in an increase in weight

of 66 percent over the standard disk. This extra weight is equivalent to an increase of

0.29 percent in installed specific fuel consumption (8FC) for an average DC10-30 air-

craft flight.

The extra disk weight could also be added to the standard disk design to reduce the

centrifug,'d stresses due to the blade loads, llowever, this mechanical stress reduction

would probably be offset by the increased transient thermal stresses resulting flx)m the

slower thermal response of the bulkier disk. Also, a heavier standard disk would lack

the redundancy of the bore-entry disk and would generate even Mgher fragment ener_esfrom a burst disk•

Some possible fragment patterns resulting from manufacturing flaws are illustrated

in table III. The available kinetic energies that would be generated from these failures

are also indicated. The highest energy fragments are caused by failures initiating at

and pl_apagating radially from the bore, as shown by the 120° disk and blade fragment

pattern for the standard di_k in table 1If. Ilowever, the redtmdant construction oi the

integral bore.entry disk would enable the undamaged member to contain such a failed

part. The only possibility of a sebnnent separating in this wa.__would be if the radial

[allure ptx_pagated through a web to tht: opposite disk face; however, this i8 highly un-

likely because the total tMckness [or all the webs is only 20 percent of the bore circum-

ference and, as one web started failing, its load would be transferred to adjacent webs.

1lie most likely mode of fragment generation is a rim fra_mlent resulting f_m dcfcets

or crack initiation sites at the dovetail slot bottom or bolt hole rim. lSa._t,d on _pin pit

experience, the rim-initiated crack wouM result in the loss of three dovci:dl po,_ts and

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four blades, as shown in table III. 'll;e fragment {mergy of the bore-entry disk rim

fragment was only about 10 percent of tim 120° disk H(,_lnt'lll thai x_a._ _;HHllntt,d to I)c

generated from a bore defcc,t In tilt, stand_;rd disk.

Analyses of ,l'l'Sl ;- 17 Hl:uld_lr_l :rod I{ort-- l<nl ry i)i;;k,:

The average temperature responses for the JTS;D-17 turbint, disks in l'igurc 8 sllow

consistently lower bore :mtl rim tcmper:_tu rcs throughout th(, cyt.l¢, in tilt, bore-entry

disk as compared with the st_mda;_l disk. The lower tempcraturt.s in lilt, bor(_-(,ntry

disk were the result of Its superior cooling effectiveness m_d the u:;c of cooling air bled

from the compressor midstage. M_tximum temperatures and strt,sst,s ovcurred at the

end of takeoff and climb, respectively.

Predicted cyclic lives for the initially mfflawed st_mdard and bore-entry disks are

presented in figure 9. Tim FAA-certified life of the Waspaloy st{mdard disk is 16 000

cycles based on the limiting low-cycle fatigue life at the exit of the cooling air hole.

These results indicate an Improvement in the cyclic crack initiation life of the Astroloy

bore-._try disk of 88 percent over the Waspaloy standard disk {rod 67 percent over the

Astroloy standard disk. 'l_e most critical location in tim bore-entry disk was in the

bore region at the entrance to the cooling air channel.

Defects and manufacturing flaws in the JT8D-17 turbine disks weft: considered for

the critical locations hldicatcd in figure 10. Subsurface flaws of 0. 119 centimeter

(0.047 in.) in diameter were assumed in the bore and web regions for all three disks;

this diameter was selected because It is at the threshold of detectability by ultrasonic

inspection. The web flaws shown in figure 10 were at thc radius of mm,_imum raditdstress in the standard disks and at tim radius of mlLximum _hxitd stress at the bond sur-

face in the bore-entry disks. The surface flaws at the disk rim or bore were assumed

to be 0.081 centimeter (0.032 in. ) ha length.

The most critical location in the Waspaloy st:re(lard disk for "l flaw was "d the exit

of the cooling air hole with a predicted crack propagation lift, of 290(I cy(.lcs. Substi-

tuting Astroloy properties for the Wasptfloy reduced the calcul'dcd crack plx)pagatlon

life to 1150 cycles because of the lower ductility, llowevt, r, there are indit'ations that

ff the crack prep.lgat!on data had Included llold-timc (,ffcct,_, lilt, cr'l_.k prop;tgatlon life

of the Astroloy stm_dard disk would h._vc bccn superior to that of tilt, W:lspaloy shmd:lrd

disk. This would 'also mc.m tlmt the values given in figure 10 for th(. bore-entry disk

are too low.

The calculated Improvement in the, minimum cra(.k l)l_l_:ll_alion |if_, of Ih_, bert,-

entry disk over the W:lspaloy st:re(lard disk was 121 i)crcent. 'Mits Iml_rovt'ment is ._lg-

nlficantin im'reasingthccap.lbilItyof tilt,disk to survivt,tmconti_)l]:d)h,I:l_.t_r._ih:_I

might resultinc.ltaslrophi,'l'_dlur(,of {.()nv(,ntlon:dly(k,slR_cddisk,_.'rh{,r(,w:is'l

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: slight reduction in the overspeed burst margin of the bort_.entry disk n._ compared with,i_ the st_mdard disk because the overall disk weight was kept const,mt and thai portion of It

due to the radial webs was of smqll structur,d importance.

": A s,abstantlal reduction in fragment ener_,5' is shox_al in t:dfle II1 for tht, .l'l'sl)- t7

bonded bore-entry disk even though it was not designed for rt,dundmwy. This impn_ve-

ment would result from the confinement of the fragmentation from a bore flaw to one

:_ disk h_f; the other half would probably experience failure at the rim from the increased

blade loading.

CONCLUDING REMARKS..,[

i: Some advanced turbine-disk structural concepts have been an_dytlcally studied as

potential replacements for the existing first-stage turbine disks in the CI.'(;-50 and

_: JT8D-17 engines. An integral bore-entry design was selected for more detailed evalua-

i_ tion for the CF6-50 engine as a result of preliminary analyses of seven disk concepts in-, including composite, laminated, _-mdmultidisk designs. The integral bore-entry turbine

: disk was designed to improve disk life and to prevent high-enerk_' fragmentation by using

redundant construction at the ex'pense of an increase in disk weight.,L

_: A split-bonded, bore-entry design was selected for evaluation for the JT8D-17 en-

' gine. This bore-entry disk was designed to improve disk life without redun(hmce or ,an

increase in disk weight.

Cyclic thermal, stress, and fracture mechanics analyses of tl_c bor_entrv and

:_ standard disks demonstrated that substantial improvements in the cyclic lives of both": initially tmflawed and flawed disks could be achieved with the Imrc_-entrv disk designs.

_, The benefits of the advanced disk designs are influenced by differences in design philos-

,7 ophy, disk cooling method, fabrication procedure, and engine operating characteristics.

REFERENCES

-_ 1. National Transportation Safety Boarct Special Study: Tud)ine-Engine Rotor DiscFailures, 1975.

2.Barack, W. N. : and Domas, P. A. : An Improved Tud)h](, Disk l)t, sigll 1o ln('rcasc

• ' Reliability of Aircraft Jet Engines. NASA CR-1350:13 (I{_oAE(,. (;eneral Electric

Co.), 1976.

3. Alver, A. S. ; and Wong, J. K. : Improved Turl)inc l)isk i)csign to hwrca._c lk,li'l-

_ bility of Aircraft Jet Engines. NASA CR-13.t985 _PWA-5:',29), 1_7_;.

, 4. Kauhmm, Albert: Adwmce(I Turbine Disk l)eslg_ls to Im'reasc Reliability of Aircraft

Engines. NASA TM X-71804, 1976.

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, I J

TAIII,EI.o DI':_I(INPllOPl.;ll'lIF:l(H 'IIlillINlPl-:l-,MATI llIAl._

lll-apcrty I '[,'lJ- rill l'llJ_l/ll' ,t 'l_ll- 17 I'lll_llll'

U|tlm,te ti,nHl|t, Htl'¢,n_h, N/urn 2AJ 29,1K 12ii {li_{l l_l) (lllll l_ I lltm l:l'lUIIO

_t Hll K l Ill (IIII) I l,tl(IUll I Ill(IIHI l J7 flint

Ylehl utrc_ngthlo. 2 purccnt offsctl. N/t.m J m,A| "19,l I( llll Illlll l Ill OIll) _1, IIIIII ll; Illlll

At 811 K jl.l.t(Hill I I II (llll} %1, II(lll N7 Illlll

Elongation at failure, p;n.ent.At294 K 2u _.5 2 | l!_

At 811 K :|o _. 5 21 11 3

1000-tlour rupture strt.nKth at 8(;7 I_, N,'(.m2 t,8 UIIU Ill:} Illlll 71' I.., ,_I i,Hi

Stress range for crack initiation in i0 uou t,vch,,_ _l iiu/i ll_lIIiIU _,_')tltltl 'IN,')II|lllat 811. K (minimum stress, zcm|. N/cm 2

,. :l/2CHt/cal strips tntt_sity {actor at 894 K. N/cm 9'1 l)iJ0 ti_ 011D .l;; (i()(1 .l .|,t_ if|ill

aEstlm ated.

TABLE II. - RESUL'I_ OF PRELI,MINAitY ANALYSES Of,' CV*,-3,, DL'_ht't)Nt't l"l'_

Disk concepts Attv'antagcs |)i_:ttt_ :0)l.:tl_t,._

Bore entr_ Ht_tmd0zzc.**,imprvvcd thermal IDt't't':t.v.t'(I v.t'lKht to prt)vZdt, It_

rt.sponsc, longcr lift. dund:mt dt._igll

('omposltc lit_dut't_ _trt.ss lt,vt.ls, ]ongcl. Limited m:tt*.,i'Jal p._._lbllitit,-

cyclic lift, l:d) rit':Itlon dt._ C]Olm1_.nl It_

tlui I't'¢l

Lamtnatt_l l_cdtmd:mc_ low [ragtncnt l-.'._t,csst_ t, _l_,t,lt_hl higJ: ,_tt't,._l._

tq|Cl'l_'. |O%_ t'O_l :it h'Jh_ _Uid I.,]t huh,- tJll.l'tn iI

mis1T_III'_CS ht't_t't'l) l:ll}lill;lJt'_

IJnk l_t,durtdmlcy low frngrnctlt I-'xt'cs.,.l_c link _,trt,._t ,- titlh-

tmerg3. 1o%1cost ¢'tlllIt)s_,;il di._k h,l,t't'_,'lltCtiUI:UI I h,;t_ ;lRt'

Multihorc Ribs t)rc_,tqll, ;LXI;LI llo_,t llil_hI r;ol_.lt,llt th,.llll:ll -IIt,-_t,,.

propag.,iUonat hort, ;it rib OUtt'l' Iliillllt'tl' I................................

Mu|tl(Llsk Improved lhcrm:)l rt._pt)nst,, irlct't.iist,tl v_t.ilzht boll_ ,._ouhl

SO_II_.'rt, tlulltl_lt'_, Jail If oilier di_i_. {:lllt',t hi) l-;.I

.rhl|t I1 l/i/it, r dlq. l;tlh,d

Spllxtc II,.dtmd_uzr_, lOllgt, I lilt, lrlt.r(,;t..._,d ,,_('lt;ht tcz pt,v, ldl' l t_

tlund:tlllill'_[I.11i dllhHdt t,,

:ill:tl% ,'t' hu,,I _}illl _ Hit ,q}( I.IlJl'd

Ili._l'.

O0000005-TS E11

TAI_LE I11 - I, ItA(,MI,._ 1 I ,%r It(,lt,_ _l! '1 I'ltltlNl" DL'iI., l)t._lih_

lll.l_ iI(,NJl_n 1' r_ll_lllt,flt l)atlt, rTI A_allldJh, klneth' enellly

,I

Iqk_

llllll;ll ll,l_t --_,-- ""',

('FI;-50 stluldant dlKk E/ "_ _\l_tt" ! 172 SO0

t

CFli-5U Integral L,_rL_ -_ I _ I _ ( I f 1|11 [)tiLl

entry dl_k lnltl;fl lla_ ;

l.ttial tl:lw - -_

J'F{_D- 1T _tazldatxl disk (_(_ _i20 ° 1,78 600

Initial _ _ o "Aml

Fn,ant h; lte;H' h:fll

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OOOOOOO5-TSE12

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::J ill !l ,'- /ll_ WEBS'_, t ;:f,"

IT H _ ,l i, l,

i L , ,,', ,i1,i_'_i,:/" " ," /'_\' BORE _ '_'_,-INLET_:.:...., ..'.Ik,f',SUPPOR.;,,.,,._//TURNING

_.{-';_:'.",.:-,'..:I....A- _'"'J {(I_I_,;¢ VANESVIEWA-A

(a) Standard disk. (b) Integral bore-entry disk.

K

°

' ".I_!,' t _

•'- _M...L;.LL_ ',. '

COMPOSITE

FIBROUS/ I'_,,,, ..... .LOOPS'_ _', , ,, ' , . ,,,

.... i;,j ....

p

r .

_ii (c) Compos£te dLsk. (d) Laminated d£sk.

:-_ Figure i.- CF6-50 flrst-stage tnrblne disk designs,433

_,_o. _,,.,.,_.:_:_........... ,:..-,.:.... ..,_-_,--, ...... .,". ,, u ., ,,o _, ,, • o u r_.,(0

00000005-TSE 13

l II\ I I

l !

LNK LNK ISEbMENTS \ SEGMENTS

CIRCUMFERENTIAL /RIBS -,. \

',/' \,\

>"

(e) Link disk (typically a disk would (f) Multibore disk.

contain 20 to 40 layers each clocked

axially relative to the next).

Figure i.- Continued.

434

i_;_,,'-_"t.,--'- _... _- .......................... ._.,_ ., . - ,:-.,- ............ t" • ...... m ....... #'-" + - "" "

O0000005-TS E14

SPLINES/

'I

' t__III i, _IIP I

HALVES

(g) Multidisk. (h) Spline disk.

Figure I.- Concluded.

4 '}.5

O0000005-TSF01

.._ A

//_/'/

['__/,.//_ COOLINGAIR

I_!_i_ FLOWPATH L_".__._..._L;I...... '_ _--"" A--'-I\-BONDED COOLING

SURFACE AIRFLOWPATH

............ Jr --

SECTION A-A

(a) Standard disk. (b) Bonded bore-entry disk.

Figure 2.- JTSD-17 first-stage turbine disk designs.

436

O0000005-TSF02

+ i ,; 'il t

1 i 1 |, ,

/- TAKEOFF

_-, I. 2[ i' / CLIMB DESCENT

,, , I_I03

ALTITUOE, ;M

0LL+ .... ],

181X)Emmmmml_[_ i i

K 000[;' h , ,! , I

'_ mmmse_°'z°I_ II

0 20 40 60 _I00120

FLIGHTTIME. MIN

•, Figure 3.- CF6-50 simplified engine cycle.

,-TAKEOFF- / CUMB DESCENT

'019 "-_ I

" 'K iINLrT MACH NO. '0 CRUi l %,_1

AmTUDe.l_ °3_ !" 'I+]i l\i

• OLr i i ;* J i_ *_l

lll_j o ov.

K 600" J , +, ,, l....,,------.P_J

+" 12Xl#

mGlm SPED. 10_-'---'_-_ {

0 5 10 15 20 _ 20 35

RIGHT TIME, MIN

Figure 4,- JT8D-17 simplified engine cycle,

437i.

,_ _ ,.,-..._........................................ :.-_:................. : - . . -.:.._:_i_i,_:.-:._+:.+-.+_ .:.

": " +'..... ".........." " OOOO0005-+-SFO'T3

9001 EFFECTIVEkNIcm_:S'_ESS'3O 40

800-35 39

•_ 7oo- LTo "'_._u.r

,q

-77 1441)

/!500// _ RIM It_

//1/// .... BORE CooL 1 I 1 I I I 1 1 I 7% o 2o3o ,o ,o oo:o 8o9o_oo11o

(a) INC-718 standard disk.

EFFECTIVES_TRESS.900- kNIcm'_

M 32

8OO _,_

= 7oo /L,3 "'"'7_'..

_ /r,_ --..._N_11 "'-A-72 .

• ,___'L__L L-_--2°R_I L , , , _,0 lO _ 30 40 50 60 lO 80 90 lO0 ]lO

FLIGHTTIME,MIN

(b) Bore-entry disk.

Figure 5.- CF6-50 turbine dlsk temperature response.

438

_!_...: .-- _......... -- .... : .. : :_ i._ ._ . .... ,._- , .

00000005-TSF04

>I00000 >I00000 >IO0000

CYCLES--,_"j CYCLES-_ CYCLES-, -

V _ |FAA-APPROVED I "_ CALCULATED _--_I_>I00000 "- ' _I00000,_ I LIFEIS?800CYCLES. I" >100000 I1'.1['CYCLES"I I SUBJECTTOINCREASE| CYCLES rl:ll cYCLEs _

I I AS RESULTOFFLEET | l!lI

.... , >tooooo/ -- I"-T_oooooo_uuu / CYCLESJ CYCLESCYCLES-_

(a) Sta_ard disk (b) Standard disk (c) Integral bore-entry(INC-718). (Ren_ 95). disk (Ren_ 95).

Figure 6.- Crack initiation lives of CF6-50first-stage turbine disk designs.

380 I155 ,

''1809 ''/161 CYCLES--'__11',1I', _ o26

CYCLES YCLES IIiii C CLE

/ II \,.-ie3

CYCLES--' CYCLES-' "_CYCLESBURSTSPEED.PERCENTOFMAXIMUM

• TAKEOFFSPED- 126 1N 149WEIGHTCHANGE.

PERCENTOFSTANDARD 0 +66DISKWEIGHT" "'"

(a) Standard disk (b) Standard disk (c) Integral bore-entry(INC-718). (Ren_ 95). disk (Ren6 95).

Figure 7.- Crack propagation lives of CF6-50 first-stage turbine disk designs,

439

O0000005-TSF05

35000

ocYc cYc(FAA-CERTIFIEDLIFE) _ ]8000CYCLES_¢:_

elm.,

--->100000 ->100000 /t_->100000CYCLES

/'J CYCLES ./'"J CYCLES ,

(a) Standard disk (Waspaloy). (b) Standard disk (c) Bonded bore-entry(Astroloy). disk (Astroloy).

Figure 9,- Crack initiation lives of JT8D-17 flrst-stage turbine disk designs.

1150

CYCLES--x_ ] CYCLES_,,C_ _9

?_9C/C_3_ _Y 000 >100000CYCLE65000 000 CYCLES

CYCLES-'I_',,_'I CYCLES-"I_",___ 6900__.__'_..__.J __.L. CYCLES--L_-__e, ,--3/000CYCLES

BURSTSPEED,PERCENTOFMAXIMUMTAKEOffSPEED: 136 136 1)3

• WEIGHTCHANGE.PERCENTOFSTANDARDDISKWEIGHT: .... 4 0

(a) Standard disk (b) Standard disk (c) Bonded bore-entry(Waspaloy), (Astroloy). disk (Astroloy).

Figure 10.- Crack propagation lives of JT8D-17 flrst-stage turbinedisk designs.

441

'-- _--'--_"_ _" I nl .......................... """ IT lilllll ill ir" _ _ ,i=,,i :" _ ...... ". .... _" '" ° "':" ."'i -_"_', .... . _ -_-"_':'_:w :_':'L :"':-"-±.

O0000005-TSF07