CF6 jet engine performance improvement: New fan

~.4...tg:(/. ,$ .. ~~~~~- "o;;;$1!"%¥'R!htp;;a ___ ~~~-*¥?~-!Q;J4~'T"!IJ,*:-"},'{- f'¥¥)#F¥ffl4AL-~M;;t .-:¥¥i-.-ef'f.,,,-, ¥#**~ ,~~'!~~~~,_,,""?",t"'~,:,'-" ~"-'-"=';';i--.i" "; =~'f'''- <."!, . ..-~--,, -.~ ---'----

r

TABLE OF CONTENTS Section SUMMARY 1 1.0 2.0 INTRODUCTION 3.0 INITIAL NEW FAN DESIGN AND TESTING 3.1 Early Design Studies 5 3.2 Previous Engine Development Tests 3.3 Design Considerations 4.0 FAN ROTOR PHOTOELASTIC TEST 24 4.1 Rotating Stress-Freezing Test Facility 24 4.2 Test Configuration 24 4.3 Instrumentation 4.4 Test Procedure 4.5 Test Results and Discussion 5.0 FAN BLADE BENCH FATIGUE TEST 52 5.1 Test Setup and Test Configurations 52 5.2 Instrumentation 52 5.3 Test Procedure 55 5.4 Test Results and Comparison to Predictions 59 6.0 ENGINE CROSSWIND TEST 65 6.1 Crosswind Test Facility p,nd Configurations 6.2 Instrumentation 6.3 Test Procedure 68 6.4 Test Results and Discussion 74 7.0 ENGINE PERFORMANCE TEST 7.1 Engine Test Facility 7.2 Engine Test Configurations 98 7.3 Instrumentation and Data Reducticn 99 7.4 Test Procedure and History 7.5 Discussion of Results 7.6 Application of Results 123 8.0 ENGINE ACOUSTIC TEST 127 8.1 Engine Acoustic Test Facility 127 ·1 8.2 Test Configurations 8.3 Instrumentation - Acoustic 8.4 Test Procedure and Data Reduction 8.5 Test History 8.6 Test Results and Discussion 137 v PRECEDING PAGE BLANK NOT FILM£D

J

TABLE OF CONTENTS (Concluded) Page Section 9.0 ENGINE POWER MANAGEMENT TEST 159 9.1 Engine Test Facilities 9.2 Test Configuration 159 9.3 Instrumentation 161 9.4 Test Results and Discussion 165 10.0 ENGINE CYCLIC ENDURANCE TEST 177 10.1 Test Facility, Configuration, and Instrumentation 177 10.2 Test Description 10.3 Test Results and Discussion ! ~ 11.0 PRODUCTION ENGINE AND AIRCRAFT FLIGHT PERFORMANCE TESTS 184 ~, 12.0 ECONOMIC ASSESSMENT 186 P 13.0 SUMMARY OF RESULTS 189 APPENDIX A - QUALITY ASSURANCE APPENDIX B - NOMENCLATURE 195 APPENDIX C - REFERENCES DISTRIBUTION 198 vi j & ; '¥- _ .•• ¥- 1.0 SUMMARY Aa part of the NASA-sponsored Engine Component Improvement (ECI) Progr_.

a nev fan package has been developed to reduce fuel consumption of current CF6 turbofan engines for today's vide-bodied commercial aircraft. The new ran package consists of • modified fan blade. reduced fan tip clearance due to a fan case stiffener. and a smooth casing tip shroud (microballoon epoxy in open-cell aluminum honeycomb). The new CF6 Fan Program included full scale engine and component testing and monitoring of aircraft flight tests. Full scale CF6-50 engine testing included back-to-back performance and acoustic tests. a power management test. a crosswind test, and a cyclic endurance test.

Component tests consisted of a model fan rotor photoelastic stress test and a full-sizp. fan blade bench fatigue test.

Back-to-back sea level and simulated altitude engine performance tests demonstrated the predicted improvement in altitude cruise specific fuel con- sumption (sfc) of 1.8% for the improved fan compared to the original fan.

Subsequent sea level production engine and aircraft flight tests confirmed this cruise sfc improvement. Based on this demonstrated cruise sfc improve- ment of 1.8%. a 2.0% block fuel saving per aircraft is projected for a new CF6 engine with the improved fan for the longest U.S. domest~~ and interna- tional missions. The improved fan offers an annual fuel savings per aireraft up to 1.37 million liters (0.36 million gallons).

Back-to-back engine acoustic tests established that the improved and orig- inal fans will have comparable community noise exposure. The FAA has accepted the acoustic equivalency of the improved and original CF6 fans. The improved fan has a significant reduct inn in multiple pure tones (buzz saw noise) com- pared to the original fan which should significantly reduce aircraft passenger compartment noise levels during aircraft tak~off and initial climbout.

Power management tests of the CF6-50 engine with the improved fan defined the fan speed/engine thrust rel:t>ionship for the DC-IO-30. 8747-·200, and A300B aircraft. Full scale fan nozzle thrust and flow coefficients were determined t from instrumented engine ground tests and correlated with aircraft flight tests.

Several component and engine tests were conducted to confirm the struc- tural integrity of the improved fan. Component photoelastic and blade bench fatigue tests demonstrated that the stresses and fatigue margins of the im- proved fan blade are similar to the ~riginal CF6 fan blade. Engine crosswind testing demonstrated that the improved fan blade has similar crosswind/distor- tion characteristics to the original blade. The improved fan operated suc- cessfully without exceeding vibratory stress limits with both the DC-lO-30 and Boeing 747-200 inlets at allowable takeoff crosswinds up to 35 knots. A previous bird ingestion engine test demonstrated that the improved fan blade is as rugged as the original CF6 fan blade. Fan tip rub button tests indi- cated that the fan case stiffener provided a significant improvement in fan casing roundness compared to the unstiffened case. This permits reduced oper- ating fan tip clearances and improved fan efficiency.

The CF6-50 enaine cyclic endurance te.t demon.trated a ba.ic lifi capa- bility of the improved fan blade and the fan case .tiffen.r in over 1000 .imu- lated fliaht cycle. without any .ian of di.tr.... A .eparate blade/.hroud interaction rub test indicated no evidance of blade and ca.ing interaction due to heavy rub. into the .mooth microballoon ca.ina tip .hroud material.

Aircraft fliaht te.t. with the improved fan w.re con.idered. The i.- proved fan ha. been certifi.d by the FAA for use in the CF6-5OC2/E2 .nain •• and i. now in commercial .ervic. on the Boeina 747-200. Douala. DC-lo-30. and Airbu. Indu.trie A300B aircraft. The improved fan will al.o be incorporated in the CF6-6D2C and -61 enaine •• and provide. a ba.i. for performance improve- ment in the CF6-80 and -32 enaine.. The improved fan will make a very .ub- stantial contribution to reduced fuel consumption in commercial aviation through the balance of the century.

. e 2.0 INTRODUCTION National enerlY demand has outpaced domestic supply creating an increased U.S. depe~dence on foreign oil. This increased dependence was dramatized by the OPEC oil embargo in the winter of 1973 to 1974. In addition, the embargo triggered a rapid rise in the cost of fuel which, along with the potential of further increases, brought about a changing economic circumstance with regard to the use of energy. These events, of course, were felt in the air transport industry as well as other forma of transportation. AI a result of these ex- periences, the Government. with the support of the aviation industry, has initiated programs aimed at both the supply and demand aspects of the problem.

The supply problem is being investigated by looking at increasing fuel avail- ability from such sources as coal and oil shale. Efforts are currently under- way to develop engine combustor and fuel systems that will accept fuels with broader specifications.

Reduced fuel consumption is the other approach to deal with the overall problem. A long-range effort to reduce fuel consumption is to evolve new technology which will permit development of a more energy efficient turbofan or the use a different propulsive cycle, such as a turboprop. Although studies have indicted large reductions in fuel usage are possible (e.g., 15% to 40%), any significant impact of this approach is at least 15 years away. In the near term, the only practical propulsion approach is to improve the fuel efficiency of current engines. Examin~tion of .his approach has in- dicated that a 5% fuel reduction goal starting in th~ 1980 to 1982 time period is feasible for a current commercial engines. These engines will continue to be significant fuel users for the next 15 to 20 years.

Accordingly, NASA is sponsoring the Aircraft Energy Efficient (ACEE) Program (based on a congressional re~uest) which is directed at reduced fuel consumption of commercial air transports. The Engine Component Improvement (ECl) Program is the element of the ACEE Program directed at reducing fuel consumption of current commercial aircraft engines. The ECI Program consists of two parts: engine diagnostics and performance improvement. The engine diagnostics effort is to provide information to identify the sources and causes of engine deterioration. The performance improvement effort is di- rected at developing engine components having performance improvement and retention characteristics which can be incorporated into new production and existing engines.

The performance improvement effort was initiated with a feasibility anal- ysis which identified performance improvement concepts and then assessed the technical and economic merits of these concepts. This assessment included a determination of airline acceptability, the probability of introducing the concepts into production by the 1980 to 1982 time period, and their retrofit potential. The study was conducted in cooperation with Boeing and Douglas aircraft companies and American and United Airlines, and is reported in Refer- ence 1.

c- ',.

I j ;- In the fea.ibility analy.i., the new Cr6 fan performance improvement con- cept wa •• elected for development and evaluation becau.e of it. fuel .avinl' potential and attractive airline payback period. The objective of the new fan proar .. wa. to develop technololY and to verify the predicted fuel .avina. by enline around te.t.. lnititally, the new fan concept con.i.ted of an improved CF6 fan blade, reduced fan tip clearance. due to a new fan ca.e .tiffener, and a revi.ed fan operatina line change by increa.inl the fan nOllle area. An i.provement in crui.e .fc of about 1.8% wa. e.ti.ated due to the new fan on the Cr6-S0 enline.

The new fan proar .. wa. a 20-month effort that included model, component, and full .cale enline te.ting. Hodel fan rotor photoela.tic .tre •• te.t. and full-.ize fan blade bench fatigue te.t. were conducted. CF6-S0 enline te.ting included performance, acou.tic, power manalement, croa.wind, and cyclic endur- ance teata of the i.proved fan, and compari.on of re.ult. with the orilinal CF6 production fan. Additional GE-funded CF6-S0 engine te.t., includinl bird inge.tion and performance and aircraft flilht telta, were al.o conducted and are reported herein. An overview of the new fan program i. pre.ented in Refer- ence 2.

t - _ h.

14M I~nnd--- _ -'""""_~-n -_.,:..-

; 3.0 INITIAL NEW rAN DESIGN AND TESTING Initially, the new Cr6 fan concept con.i.ted of improved fan blade aero- dyn-.ic de.iln, a 1.50 .. (0.060 in.) reduction in fan tip clearance due to a new fan ca.e .tiff.ner and optimi.ation of thQ fan crui.e operatina line.

TOlether. the •• improvement it ... offered a potential reduction in Cr6-50 en- aine .fc at crui •• of 1.8% (Reference 1). Thi. potential .fc reduction i.

achieved vith • mode.t veilht increa.e of 13 kl (29 1b) and a forward center of aravity .hift of 0.8 em (0.3 in.). A maintenance co.t reduction (lower DOC) i. projected, re.ultina from the lover turbine aa. temperature. that accompany the improved enaine performance. An improvement in crul.e .fc of about 1.6% va. e.timated for the nev fan on the Cr6-6 engine.

The improvement in aerodynamic performance of the fan ha. been achieved laraely by way of a more forwrad throat location (more camber in the forward portion of airfoil) and aft location of the part-.pan .hroud, re.ultina in an improvement in the entire chordwi.e and .panvi.e efficiency. Reduced fan tip clearances, further improving performance, are achieved by atiffenina the fan

I

caae. Optimhation of the cycle 11 achieved by adju.ting the fan operating , line ao that it paa.ea through the region of peak efficiency. The total pre- dicted improvement in fan efficiency waa 4.7% at cruiae-equivalent power which is projected to be a 2.7% sfc improvement at sea level static crui.e equiva- lent power. Thia tranGlatea into the altitude cruiae afc improvement of 1.8%.

l.l EARLY DESIGN STUDIES Since the CF6 engine powers the DC-IO-IO, DC-IO-lO B747, and AlOOB air- t craft, it will continue to be A significant fuel user for the next 15 to 20 year.. In mid-1975, a GE atudy was initiated to con.ider an aerodynamic re- design of the fan blade for the CF6-6 and CF6-50 engines. It was the objec- tive of this study to further fan technology development to significantly im- prove the performance of these tw~ high bypa.a ratio engine. without compro- misiog .tructural integrity, maintenance co.t, or reliability. The study in- dicated that the CF6 fan could potentially be improved up to 2.0 point. in fan efficiency if the blockage due to the part .pan ahroud could be reduced. The detailed fan aerodynamic and mechanical designs were eatabl!shed; initial bench teatina for blade vibratory frequencies, nodal patterna, and .tresa dis- tributiona wa. performed; fan .tress and aerodynamic mapping was done on an actual engine; and initial engine performance improvement was evaluated throuah CF6-50 back-to-back te.ting. In .ddition, rotating rig bird inges- tion test. were conducted to e.tabli.h blade/bird strike integrity.

A cro •• section of the CF6-50 engine showina the proposed fan performance improvements is presented in Figure 1. The CF6-50 engine has s full diameter, single-stage fan and three low pre.sure compressor stages which boost the flow into the core engine (Figure 2). These booster stages allow higher thrust C'I REDUCED TIP CLEARANCE""O. 060 in.

TRIM

FA~ CASE STIFFENER J FAN NOZZLE

~

~

~~·i r t- -- ~ .-

( - ~- -~- .-. "1" .... ~ UIPROVED BLADES ..........

- [ ~

- J - ~i

~ ~ - \ \ I )1 1\ , ____

:Dr- O"'tl C:P

,.e>

r-FYt ~u; Figure 1. CF6-50 Engine: Initial Fan i>erformance Improvement Concept.

-

-R -p"- '&",,4 ~.",{ =:i:!'"_= 4- - __ ~ #=4ifE-, - ~-~-~~~-- Figure 2. CF6-50 Engine; Fan and Three-stage Low Prossure Compressor.

vith minimum over.ll enlin •• nv~lope and v.ilht. Th. low.r thru.t. CF6-6 .n- ,ine. Pi,ure 3, con.i.t. of a full di ... t_r •• inal.-.t ... fan follov.d by a .inal.-.t .... low pr ••• ure co.pre •• or to boOit the flow into the cor •• n,ine.

3.1.1 ran AerodYh~.ic Oft.i,n An an.1ytical .tudy of the detai1.d a.rodyn .. ic flow charact.ri.tic. v •• .. de of the CF6 fan v.r.u. the fir.t .t.,e with p.rt-'pan .hroud of a tvo- .t .. _ NASA fan. fbi •• tudy indicated area. of potential efficiency i.prove- aent .ianifieantly Ir.at.r than initially foreca.t. Procedure. WIre developed which permitted the ....... nt of .panvi.e blocka._ .ff.et. on airfoil .urfaee.

(.uch a. the part-'pan .hroud). Thi ......... nt r.v ... ~.d that thro.t marlin.

in the vicinity of the .hroud Vere more .ubcritical than initially con.id_r.d.

A co.parativ. a •• e ••• ent of the NASA part-.pan .hroud fan blad. r_v.a1.d that throat margin8 in the shroud vicinity were not .ubcritica1. It wa. deduced that the NASA d •• ian obtain.d the throat marlin in two way.. Pir.t, the pa.- .a,e throat va. forward relative to the current CF6 fan. A more forward throat location val achi.v.d in the r.d •• ian by puttinl more c .. b.r in the forw.rd portion of the airfoil. S.cond, the part-.p.n Ihroud wa. locat.d aft toward th~ trailinl ed,_ of the .irfoil r •• u1tin. in reduc.d throat blnck ....

In turn, the lead ina .d.e of thft .hroud oper.te. in • r.aion of lov.r M.ch number flow vithin the p....... R •• ov.l of the .ubcritic.l thro.t. fro. the r.d •• ianed fnn elimin.ted the l.rle r.di.l flow .hift .nd permitt.d the bl.de to oper.te ., de.ianed. The entire .p.nvi.e .fficiency level incre ••• d vith little effect on the .hroud w.k.. Hovin. the p •••••• throat forw.rd. by put- tinl more c.mber forw.rd in the .irfoil, unlo.ded the tr.ilin. ed.e .nd re- .ulted in more .ffectiv. c.aber. A photolr.ph of the orilin.l production .nd the improved CF6 r.n bl.d. i. pr •• ented in Filure 4.

In addition to the .bove red •• i,n con.ideration., .ddition.l c .. ber w •• put into the .irfoil to r.i.e the peak efficiency at ?re •• ure r.tio. ~orre .pondinl to the current CF6-S0 crui.e operatin, level throulhout the .peed r.nle.

t 3.1.2 F.n Mechanical De.i,n The mo.t obviou. ch.nge in the ~chanic.1 de.iln of the improved per- formance fan bl.de i. the movement of the part-.pan .hroud .ft on the airfoil.

A co.pari.on of the p.rt-'p.n .hraud. for the orilinal .nd the improved f.n b1.de. i •• hown in Filure S. Some .ma1l but import.nt difference. between the two part-'pan .hroud de.ian •• re (1) a 41- pr~ •• ure-f.ce anile with the en,ine .xi. for the rede.ian ver.u •• 39- .na1e for the orilin.1 de.i,n; thi. ch.nle corre.pond. with the .talRer anile ch.nae of the airfoil at it •• hroud 'ec- tion, and (2) a 1arler radius blendinl to .irfoil for improved .upport and incre •• ed .tiffne •• for the rede.ian.

I

i -----------.

I

i

t

~ I , Fan-- ...... ~

I

--------_.- Figure J. CF6-6 Ena1ne: Fan and Single-Stage Low Pre •• ure Comprellor.

__________________ ~~ ____ ~ ____ ~ .. ~ ......... tl~.-." ........ ________________________ _ .=...."..... ms ~ ..

LE TE PART-SPAN SHROUD MOVED AFT TE TE

~

PART-SPAN SIIROUD HOVED AFT O R) GT AL II LN) E Id l'RO ED BLADE Figure 4. CF6 Ori g inal Production nnd Im r r ove d Fan Blades.

Engine Centerline Original Blade ,'7 //

//

// // /,~/'

,,~

",....;t"/ /1 //1 ,/ / I /' ( Trailing Edge Improved Blade Figure 5. CF6 Fan Blade Part-Span Shroud Comparison.

Other features of the shroud design are n~arly identical. luch as span location. weight. thickness. and cross-sectional streamline shape.

The original and improved fan blades have identical design chord and thickness. as shown in Figure 6. However. the orientation angle of the chord- line (complement of stagger angle) of the improved fan blade is more open over the inner portion of the span and about 1.5· more closed in the outer tip por- tion, as shown in Figure 7. During the fan performance tests reported herein (Section 7.0). the blade part-span shroud interlocks were subsequently modi- fied (restaggered) to close the running blade stagger angle by about 1.5· at the part-span shroud to fine-tune the fan-engine match.

Camber angle is larger for the redesigned airfoil of the improved fan (as shown in Figure 8) than for the original design. The dovetail is identical for the redesign to facilitate interchangeability by sets and thereby take advan- tage of the improved performance with minimum hardware change.

3.1.3 Fan Case Stiffener Ring A clearance reduction betw~en the fan blade tip and casing tip shroud provides potential for further performance improvement. The radial clearance reduction potential relative to original production configuration is 1.5 mrn (0.060 in.). based on observation of revenue service hardware and analysis, provided fan casing roundness could be improved. This clearance reduction has a theoretical payoff of 1.1 points in fan efficiency. However, in order to take advantage of this payoff, it was concluded that the fan case must be stiffened to raise the critical interaction frequencies of the fan rotor and fan case above the maximum operating fan speed. These critical interaction points occur at the intersection of the case harmonics with the respective rotor backward traveling wave harmonics as shown in Figure 9. Only the backward-traveling rotor waves are critical in that an excitation rub by the case is a backwards excitation on the rotor relative to its travel direction.

The original production configuration avoided these phenomena via sufficient clearances to eliminate rubbing and/or negate the coupling from buildups be- tween the rotor and case.

Figure 9 also shows the case resonant frequencies when the stiffener ring is added. The interaction points are well beyond the maximum fan speed oper- ating range so the fan clearances can be closed down 1.5 mm (0.06 in.) with confidence. These results and the stiffening ring shown in Figure 10 are largely the result of experimental investigations. In parallel with the ex- perimental programs, analyt ical mode ling was also acco'.npl ished.

An analytical model of the fan ca~e structural elements was extended to include a wing and tail engine inlet adapter (Figure 11). Boundary conditions which are defined at the frame strut locations include variable axial stiff- ness based on the fan case mode shape in order to obtain correlation with com- ponent test. Good correlation was obtained between analysis and test for the ~-------r-------r-------r------,-------,-----~ ..

e Ql 210 --I II u II ..

I ~ 5

!.

'0" ..

'i e ..

"," 0 ..

..

.c .c tJ 180 '" '" ..

Ul ~ ..

.0: ..

u ...

ir-

" 0 .c '" 500 600 700 800 900 1000 Radius, mm Figure 6. CF6 Original and Improved Fan Blades, Chord and Percent Thickness Radial Distribution.

.. , ...

~~~~~~~" - : 80 ~--------~--------~------~--------~--------~~------~ ....

70 ~-+----+------+----t--+-----; ..

~ 60 t----~-'.,..~ 1 _

~ 50 -ti~inal~~-~J+-----+---i-

~ 40 1_

.~ 30 ~---.+----- __ f-----+---~~ 1_

.~ '01 I ~ ~

~ 20 g I-t- Improved Fan _ Eo<_

~I Ql ~

10 r----------r--------~----------r_------- I;---------~r_---- :-

I ~

o ~ ______ ~ ________ ~ ______ ~ ______ ~j~ ________ ~ _____ 6~

500 600 700 800 900 1000 1100 Radius, mm Figure 7. CF6 Original and Improved Fan Blades: Orientation Angle of Chord Line Radial Distribution.

-

r-~----------t-----------~-------"--- .... ~"--------~----------~----------, ,

"

--,--------+----------+----------+----------+-----------+.--------~ , , ,

\

rtJ <J.)

\ -.... 1-

OJ ell

, I I

"" <J.)

..,

'" I 1

" 1-+-----+- I -

<J.)

.....

",_ ~Improved Fan I 0.1

ell c:: «

~ ]1 ;:1-

f-+-

'----rl ...... ~' ... ~ ""I Q11

<J.)

"" .0

e ............ "6i ~~ I

C',l u ~~--o-r-i-g-i-n-a-lTF-a-n--------~~~ ............... __ ...... II ~ I -- '" ~

...... ~I

l-~~~,,~ 0

~~----------~----------~-----------r_------~~~- ...... __ I --

II i~

500 600 700 800 900 lO~O 1100 Radiu!l. mm

I

Figure 8.

CF6 Original and Improved Fall Blades: Camber Angle Radial Distribution.

~"-- _____ ...:...... ____ ...... _.ilL.·7Tirl'fiii- Ii, iii'I"IIiT1tM"'i' .:·ZIII-.',,'· IIIM.".'I1'·- m"IooM_"_~ ___ ---- ----- --"

~ .. 40,°°

-50E Aid Un. "2-1

~ t---r---f---+--

~ 0 Stltt.n.r I

8: 6 'Ir. ealln,,' 1

8 3200 --' --.

~ I I

! ~----I----+---+----I---~

Oil I

! 24001----i---4--_-+ __ -!-)

.e

CU .:!

~--+--~,~--~- ..

CII' Reson.~equencl ..

CO n .7 Q) a: n .6 "C 1600

Z

Co n .5 tn c

o

.-

..

& 800F=~~~~~~rt~---~~~_+----+_--_+--~

a I

e 1

Q. - -.-...;--i---~____l

~ I

~ I

> O~---~~~_~~~--~---~- ___ ~ __ ~_~

o 2000 4000

6000 8000 Fan Speed, rpm Figure 9. CF6-50 Campbell Diagram, DC-lO Tail Engine.

-fJ2f~

T

-' .-

I

l' 1 c:=.:-! -':_ i --:-~ » , , Figure 10. CF6-50 Fan Case Stiffener.

..

j ~ Stiffener II 1/

-.~~ t L L-L

J

Wing ~ Stiffenor /I

~LtLL~

II

Tail Figure 11. Engine Inlet Adapter.

. "mrn-irm _

wing inlet and duct inlet adapters for an unstiffened case configuration. For a stiffened fan case configuration, the correlation was not as consistent be- tween analysis and teat for the wing inlet.

The orignial CF6 fan casing tip shroud was open cell aluminum honeycomb.

Puring the performance testing, the tip shrouds were modified by installing microballoon epoxy in the honeycomb and grinding it smooth.

3.1.4 Fan Operating Line Cycle optimization studies had indicated that the fan operating line of the improved fan would have to be lowered for improved efficiency at cruise operation. This was to be accomplished by trimming the fan nozzle as shown in Figure 12. The performance tests demonstrated that the improvements for an increase in fan exit area are not sufficient to warrant a nozzle area change esee Section 7.7).

3.2 PREVIOUS ENGINE DEVELOPMENT TESTS TWo key requirements for an operationally serviceable fan are the ability to withstand FAA prescribed bird ingestion damage and to function over the full operating range with acceptable vibration stre.s levels. Analytical de- .ign aasessment for these two items is very difficult and empirical data are required. Because of the significant development cost and the importance of thes. two considerations, it was decided to perform an early screening before proceeding into the full fan development program. The first of these efforts was a fan stre,s mapping test which was performed on the first set of blades prior to the start of the NASA-sponsored program. The seclmd test was a bird ingestion test that was conducted in parallel during the early phase of the NASA-sponsored program. To provide a comprehensive overview of the fan devel- opment, these tests are also documented in this report.

3.2.1 Engine Performance and Fan Happing Test The fan mapping test was conducted with • CF6-50 engine equipped with the improved fan blades. This engine had an adjustable fan nozzle to control the fan operating line. Aerodynamic instrumentation was utilized forward and aft of the fan .tage to Qeasure pressures, airflow rates, and temperatures.

Strain ga~es were also installed at key locations on the fan blade to meaaure vibratory stress on the airfoil, shroud and shank/dovetail. The fan wa. oper- ated through the entire speed range at various operating lines up to and in- cluding full stall. From this testing, it was possible to define a complete fan map, including the stall line. Stresses measured on all the various oper- ating lines were very similar to stresses measured on the original CF6 fan blade. Aeromechanical performance of the improved fan blade was very similar to the original blade. Two resonant points occurred on both blades at 2500 and 3350 rpm. Vibratory stresses even in resonance are low. The vibratory stress of the fan blade in stall was less than 100% of the infinite life limits. The stall margin of the improved fan blade was within 1% of the stall margin on the original fan blade.

current

ed

D\a~eter

\ocreas

"rea \-OYler

opera\\n9 \-\n.

3.2.2 Bird Inae.tion Enaine Te.t A bird inae.tion te.t with the improved fan blade. va. conducted to .i ..

ulate typical bird in,e.tion incident. that can occur durin, ,round roll or fir.t .eament climb. Condition •• iaulatin, thi. condition vere .et up by u.ina • vind tunnel ahead of the en,ine to produce a vind velocity of 54.9 m/.ec (180 ft/.ec). An aircraft inlet wa. in.tliled on the CP6-50 ensine.

Three ~~.4 cm (lO in.) diameter aun. vere u.ed to inject a total of ei,ht 0.68 ka :1.5 lb) .eaaull. into the enaine inlet. Two of the bird, .truck tho 'pinnel, four of the bird •• truck the fan blade at the .plitter, and two of the bird •• truck the outer panel. The ,un. vere placed fir enou,h ahead ot the enaine to Illov the bird to take I natural po.ition vith wina' .pread prior to .trikin, the fin blade. Tne purpo.e of thi •• etup wa. to .imul.te an encounter with a flock of bird •• After .uch an inle.tion, the engine i. required to demon.trate at lea.t 75% of takeoff thru.t for S minute.. The 75% thru.t wa. demon.trated for 5 minute., and a hilher thru.t .ettinl of over 91% thru.t WI. demon.trated addi- tionally. Damlae to the fan blade. wa. limited mo.tly to the outer panel .trike.. Bendina of the fan blade in-board of the .hroud Wa' very .mall.

There wa. no dama,e to the .pinner, the root of the fan blade, or the boo.ter from the bird. thlt .truck the .pinner. Unbalance due to thl bird .trike. wa.

low. There were no crack. or ~ral .. ntation of fan blade. from the bird .trike.

Baaed on the favorab Ie experi.~lce, the improved fan blade wa. concluded to be a. rUlaed or better than the':i,inal CF6 fan blade and no de.ian nodification WI. required prior to proceedina throuah further development.

3.3 DESIGN CONSIDERATIONS 3.3.1 In.tallatiun The fan CUI! /Jl:iffener rina wa. initially in.tdled on IlDOckup to idti!!\- tify in.ta1 !ation interferenc~ problem. with the aircraft equipment in.talled in this area of the fan ca,@. ~odificltion. to enline buildup hlrdware Ind .upportina bracket. were identified and fabricated to complete the mockup.

The primar~' interference probletllti were with the electrical harne .. rout ina and modifit:91tion of bucket. and several lead. to provide .ome added lenath.

The hydraulic .y.tem required rework involvina tube red •• ian, lenathened hOlel, and bracket redelign, none of which were exten.ive. Some repo.itionina of the f~el i~let hOle and throttle cable rout ina wa. made without rework.

Filure t3 .hOWi a typical inltall,tion in the area of the fan ca.e Itiffener.

The imp;:.)ved perfOrllllhC4! fan blade haa been deailned to operate under the II. condi t; it:,nl .. " t~e .'trilinal fan blade. Steady-Itate and vibratory .tre ..

levels thro.t!t.houl the blade, part-.pan .hroud, shank, and dovetail are very __ ~ .... ,-- ___ ...:..... ____ .;... ... ~.J ...a;.;;;;Il'liIIn.'.· .Wli·.~n",slllll·· •••. ________ " ~~ .•.. " .. ___ .,., - - .

.imilar for the improved fan blade a. for the oriainal fan blade. The .ame .. rain of .afety ha. been built into the improved blade a. exi.t. for the oriainal blade. The .. nufacturina proce •• u.ed for the improved blade i.

the .... proce •• a. that u.ed for the orieinal blade. ~lade .. terial. and the coat ina' u.ed in the dovetail and the part'~.pan .hroud at vear poinu are the .... for both bladel.

Factory endurance telting hal been comple~ed n" ~he new fan blade limu- latina airline fliaht cyclel. Factory telting of 2034 cycle. and approxi- mately 1000 houri ha. been completed. All POltvilU.1 and fluorelcent pene- trant in.~ction. of the improved fan blade after enline te.tinl have .hovn no dh~r',", Three flilht te.t proarama of the improved fan have been co.pleted.

The.e flight te.t. were conducted by Airbu. Induatrie on the AJOOa, by Doulla.

on the DC-Io-30, and by Boeinl on the 747-200 aircraft. In all of these pro- Irams, the mechanical performance of the fan blade wa. equivalent to the oria- inal fan blade with no di.tre •• of any kind.

3.3.3 Reliability The improved fan blade ha. been des lined to achieve a total u.eful life with repairs of at least 30,000 hour. or 30,000 flight cycle •• whichever occur. fir.t. This ia the same technical objective used in the design of the oriainal fan blade.

All the maneuver load cnnditions, .uch as yaw and pitch vetocity, have heen used in analyzing the new fan blade. Loads imposed on the blade under these maneuver load conditions are very similar to the loadll impoled on the original fan blade. Stresses for such conditions are comparable between the two delilns. The improved fan blade is not limiting in low cycle fatilue life. None of the conditions in the factory testing, such as crollwind telt- ing or fan mappinl, showed any measured stresses a. high as the enduranc~ limit of the fan blade material. There have never been any fatigue failurel in the orilinal CF6 ran blade in more than 10 million flight hour. of exp~r ience. aa.ed on comparable strela levell between the new blade and the orig- inal deaign, this lame level of reliability i. expected for the improved fan blad!,;.

3.3.4 Maintainability The .nae maintainability ·'·;ture. have been designed into the improved performance fan blade a. exist for the original fan blade. It i. pos.ible to install or remove a .iRlle fan blade from the diak without affecting any of the adjacent blade.. This can be done when the engine ia inatalled in the aircraft by removing the Ipinner and di.a~.embling the fan blade retainer •.

The fan blade edge thick"e.s for the improved blade i& the ~ame a& that on the original ian blade. The experience gained to date show I comparable tolerance to forei,n object da .. ,e •• uch ~. bird. and •• all object •• that are tyrically in.e.t.d. It i. exp~c~ed that larger object •• uch a. tire tread and hand tool. can be in,e.ted by the nev blade without doing extensive da.a.e. Thi' ru,gednes, hal been previou,ly d •• on'trated on the oriainal fan blade. The fan blades are pairftd by lIo .. nt weight to fad litate the rtt.,lac .. ent "f a d .... ed blade without requirin, trill bal~~ce of the enaine. The wear life of the blade coat in •• i. expected to be the '''' for the i.proved blade a. the oriainal blade based on c~lculated bearing stre.,e. and the experience Rained in the factory and flight te.t proRr ....

Iff 4.0 FAN ROTOR PHOTOELASTIC TEST The principal objective of the photoelastic test was to obtain a detailed stress analysis of the fan blades and disk, especially in areas where stresses are considered likely to be limiting. The photoelastic method has the great- est value in analyzing locations of complex stress concentration which are difficult to analyze by other means.

4.1 ROTATING STRESS-FREEZING TEST FACILITY The Rotating Stress-Freazing Test Facility consists of an environmental chamber with vacuum, heating, and cooling capabilities and a variable speed motor drive for rotating model tests. The facility is capable of operation at pressures from ambient to 1 mfu of Hg and temperatures from 20· to 180· C.

The chamber has a full opening door with an observation panel. The chamber floor has a removable bottom plate for drilling and tapping, and installation of various supports. A photograph of the improved fan photoelastic model installed in the test facility is shown in Figure 13.

Because of the large size and weight of the model, a special rugged bearing support structure was used for the test, and large radiation screens were added to aid in maintaining a uniform model temperature in a vacuum. A slip ring assembly was mounted on the shaft to permit monitoring of thermo- couples in representative thick and thin parts of the model itself for tem- perature control.

For the photoelastic analysis of model slices, several polariscopes were used. Other conventional lab and shop equipment was used in manufac- turing the models, slicing, polishing, etc.

4.2 TEST CONFIGURATION Test hardware included a 0.6 scale model of the CF6-50 fan disk and a full set of improved fan blades. Three types of blade shanks were used: full shank (no pocket), half-pocket shank (original design), and full pocket shank (Figure 14). All a irfoi 1 sect ions and part-span shrouds were ident ical on all blades. As shown in Figure 15, the blades were held in place by means of retainers made of metal and nylon in such a way that their weight matched the properly scaled weight of the prOduction design retainers.

The fan disk was a scaled model of the production CF6 fan disk, except that the flange on the aft ends of the posts was omitted. It was mounted in the test rig so that it was ~antilevered from the aft flange.

,- " ' t'tt,?&:; - ,kj'zaarithrzirot"ro's Fi gu r 1 L Pho t o l as t ic Iud 1 of Imp r ov'd F;) n H u nl ed i n Rotatin g St ress - Fr ezi ng T s t Fac ilit y .

O n, r. f', ,-., .

,I 1- ;,"

, " Full Pocket Full Shank Half-Pocket Shank Shank Figure 14. Photoelastic Models of Fan Bladed With Three Types of Blade Shanks: Full Shank, Half-Pocket Shank, and Full Pocket Shank.

b J C.

.. .. ..

, .

'-,~" '-i$F~*""~' ,...,..'""""'M ..... --~...,. """"'~QI, ,'i!';g:·'''''-'''''''''''.--*'''''it''''-¥?F'-, ""M""""..,,_, ; !

In order to prevent "shingling" of the blade part-span shrouds, the shroud contact faces of alternate blades were built up with an epoxy paste material, as shown in Figure 16. This modification was found necessary after the first attempt to stress-freeze the model had to be aborted due to shroud shingling. The second attempt ran successfully after the blades were annealed and modified as shown.

4.3 INSTRUMENTATION Model instrumentation consisted of five thermocouples mounted at the following locations: l. Tip of blade located in Disk Slot l.

located 2. Tip of blade in Disk Slot 3.

3. Shroud of blade located in Disk Slot 3.

4. Inside rim post near disk 00.

Ins ide disk hub 19 mm (0.75 in. ) from !D.

5.

These thermocouples were used to monitor temperatures in the extreme locations (thinnest-outermost at the blade tips and thickest-innermost at the disk hub) to assure that model temperature remained uniform within approxi- mately 3° C.

4.4 TEST PROCEDURE Stress-freezing is possible because of the diphase behavior of many polymeric solids. The molecules of these materials are held together by pri- mary and secondary bonds. When such a material is subject to loads at room tem~erature, both sets of molecular bonds are in effect; however, as the temper4ture is increased to a certain level, the secondary bonds relax and the primary bonds must carry the entire applied load. At this temperature, relatively large, but still elastic, deformations occur. When polarized light is passed through the solid, double refraction occurs with the rela- tive retardation of the two light components (or fringe order) being linearly proportional to the principal stress difference. When the temperature is again lowered to room temperature with the loads still applied, the secondary molecular bonds re-form and lock in the deformations, as well as the above photoelastic effect, just as it existed at the stress-freezing temperature.

Since this "thawing" and "freezing" occur on a molecular scale, thin slices may be cut out of the solid model for analysis without disturbing the defor- mation or the double refrhction properties. Epoxy resins, such as Araldite, are commonly used for photoelastic stress-freezing.

: OJ C OJ > OJ \.J 0..

o OJ "-' o ~ OJ .....

> 0.

:J OJ U) o Epoxy models of compressor and turbine rotor parts, such as bladed disks, may have rotational stresses frozen into them by means of a specially designed oven. When mou~ted in such a facility, the model is first turned slowly by the variable spe~d motor as the temperature is slowly increased. Temperature gradients are carefully minimized by monitoring thermocouples mounted in a thick and a thin section of a dummy plastic part and adjusting the rate of heating so as to limit the temperature differential. When the stress-freezing temperature is nearly reached, the oven is evacuated to eliminate air loads and the motor speed is brought up to the desired rpm to reach a preselected deformation and stress level. After stabilization, the oven temperature i.

slowly reduced, again with continuous monitoring and control based on the thermocouple readings from the dummy part. When the temperature dropped below the stress-freezing temperature, the vacuum is slowly released and the motor is stopped.

After the oven reaches room temperature, the model is removed and anal- ysis begins. Thin slices are cut along planes where stresses are desired, and fringe orders and stress directions are determined in a polariscope.

Depending on the size and shape of the slices, a diffused-light polariscope or a magnifying slice analysis polariscope or a polarizing microscope may be used for this analysis. Micrometer measurement of slice thickness is also required. Since rotating models usually have a geometric pattern that is re- peated many times, several sets of slices may be cut with different orienta- tions using identical model areas in order to completely define the state of stress, and/or duplicate slices may be cut in identical areas for confirma- tion of results. Stresses in the slices are calculated by conventional photo- elnstic techniques with the aid of a "fringe constant" determined from a bend- ing or tensile calibration bar.

The rotating stress freezing test was conducted according to the cycle shown in Figure 17. After conclusion of the stress-freezing cycle, the optical effects of the frozen stresses were viewed with polarized light. For example, overall fringe patterns of the three types of blade shanks tested are shown in Figure 18 as viewed in a diffused-light polariscope with cross polarizer and analyzer.

The fan blades were sliced und analyzed in four areas (Figure 19): 1. Part-span shroud region (70% of span) 2. Airfoil - at Section G-G, 271 mm (10.66 in.) above the dovetail base (40% of span) 3. Airfoil root (airfoil - platform fillet tangency point) 4. Dovetail and shank.

Most of the analyses were done on four blades, one full shank (Blade A2), two half-pocket shanks (Blades B20 and B23), and one full pocket "!hllnk (Blade CI9). All of these blades were analyzed in the dl)'i,t,,:Ii.l and shank, but not all blades were analyzed in the other three areas.

Stress Freezing Temperature - ~-;-~--- About 120° C (248° F) Rate of heating and cooling controlled by thermocouples in a dummy plastic part so as to limit the temperature differential to 3° C between a representative thick and thin section of the test model Ambient I m .,-..-;:1:.,;;6:..;:3;......;;;.rA- ""-- ___ -I Test rpm : 50 rpm

(IJ 0 J----------:------ .----------

....

::I en en (IJ Ambient ....

c;l; ....

(IJ

~

..c: u I I -"--"-'-_._---- 30 (Typical) Time. hr Figure 17. Rotating Stress-Freezing Cycle for Fan Model.

-----~

.. 2

F' i g ur ' l H. OV l' rLlII Ik lrk F i l' l d F'r z' n S trL 'SS F rin g' Patl ' rns in I II Tllr .(. TypL's IH:ld Sh a n ks .

I

j I· !

A typical slicing diagram is shown in Figure 19 for the shank, airfoil root, and upper airfoil for blade B20. Except for the airfoil root slices,

I

all slices (including shroud slices) were oriented perpendicular to the blade midchord and in a radial direction. The airfoil root slices were oriented perpendicular to the platform. as indicated in Figure 19, to conform to the usual convention for orientation of strain gages in blade end-effects testing.

Shroud slices from Blade B20 located in Disk Slot 28 are shown in Figure 20.

Isochromatic fringe patterns of a typical set of shank/dovetail slices from Blade 23 are shown in Figure 21.

4.5 TEST RESULTS AND DISCUSSION Stress data in the part-span shroud member adjacent to the airfoil fil- let region are shown on the slice profile diagrams of Figure 22. The maximum shroud photoelastic (compression) stress (concave side of airfoil and on the radially outboard surface) of -521 MFa (-76,000 psi) in Slice 4 agrees well in location with peak values found in instrumented engine tests and also with finite element analysis. but it is about 50% higher in magnitude. The maxi- mum shroud photoelastic (tension) stress (convex side of airfoil and on the radially inboard surface) of 501 MFa (73.000 psi) in Slice 10 is similarly about 50% higher than both the theoretical analysis and strain gage experi- mental resultb. Figure 23 shows these photoelastic data compared with the finite element analysis of this blade and the similarity of the data distri- bution trends can be noted. The data for the shroud region on the concave side of the airfoil, which is in tension (the radiallv 'nb~ard surface). are in very good agreement with the finite element analysLD.

The spanwise airfoil stress distribution data at a section 270.8 mm (10.66 in.) above the dovetail base (reference Section G-G. Figure 19), are shown in Figure 24 for Blade A2 (full-shank blade). Blades S20 (half-pocket shank blade) and C19 (full pocket shank blade) were also sliced and photo- elastically analyzed at this section; those data were almost identical to that shown. Figure 24 shows the photoelastic r~sults compared with the fin- , ite element analysis of the blade. The data trends between the photoelastic test and the theoretical analysis are in excellent agreement. though the maximum photoelastic stress on the convex surface (at 0.4 fraction of airfoil chord location) of 568 MPa (82,500 psi) is about 25% higher in magnitude than the theoretical value. Figure 25 shows the spanwise stress data at Section G-G from Blade C19 along with theoretical "twisted blade program" analyses of this blade with (a) shrouds "free" and (b) shrouds "locked up".

The airfoil root stress data distribution from Blade B20 is shown in Figure.26 along with the results of the finite element analysis of the blade.

Agreement between these data is good; maximum stress in the airfoil root re- gion does not exceed 400 MFa (58.000 psi). Figure 27 shows these same photo- elastic data along with results from end-effects testing of a full scale metal blade. The photoelastic data and the metal blade strain gage data.

from the pure radial pull test for which the load conditions are similar (but obviously not identical), are for the most part in good agreement. The - 4¥t¥~ 4?- j1 M '"' M., &3,,--1 C!:: " I 0.2101 ..

('0.6~ 1"<") I i~ -$- Iw III I:' i- TO OO~[U r L BAS[ 'I ; ..

I [liCrNe ) \ 'LAOC Figure 19. Slicing Diagram for Blade B-20 for Dovetail and Shank, Airfoil Root, and Upper Airfoil.

(a) Sli ces R assembled with Airfoil Leading Edge a nd Shr o ud Tips.

• (h) Sli c Arr ng d a nd u mb r d ons c utiv ly " I' fr om Fo rw a rd t o Aft .

FI! ':u ," '() , Sh n l ud. J I ' s I r' III 131 ad ' O:!O ,

D

; 1, S l lc ',,1'> Il [ /trough .J (r 'I 'r"lh ' " Figur l ' I ) ) I .

1'1)( [ /I : r,l l'lls I"~ S h:lI\k - 1JuVL'l.rI I I'ltu l, j' I. ls t /l ' /);1 r rll ' td) l ;,;ul' llrOm,l l j " Fr/n ~:' 1',I [[I ' rl1 ~ , r"rw;trd I nk - in )' , "l 11.1l1 - I'uI'kd Sl ldl1k H l.l d Ilil ,

L

S I [n :s " thr ug h P r -I L' II ' n ' > Fi g u r - I ) Fi g ur ' I. Ph g n l phs II S h"llk - D Vl' lal! I'hll t U~ ' I ,,, lt (Dark F i~ ' l J) I so ' I! nlIIa l I ' Frln ' l' I' 'll. ' n S f Forw ~ lFJ L Ilk - In ' AI t . II, I ~ - Pock -l S hank U I i HI ' U j ( ' o nelll Il 'd •

I

Upper Airfoil Part-Span Shroud ',1 !; I

~ 3

Lower A.:lIAllY l~e=~It'_

Airfoil "'l

~=~n~a% ~HW.! isg~ • 0.10

1M 1I.:lIIllY ~JTeOAIIC - E!

e

'" ...

'" ..

'" ~ !

~ !

-188 u u ..

.77 (-27) (.11) = ...

99 B -501 (u) -268 (-73) .186. (.39) S

~5'fJ (-27) -2611

(.39) I ----- .. : (5~~ (5~~

5 6

WIALlY l_eOI'll)_ lJ~rDIl CuT lilit _ a. Slices 2 through 7 (See Figure 20a. for location) Figure 22. Airfoil Shroud Region Surface Stresb Distribution, 3785 RPM.

_,·,""',,_:..- ___ ....:. __ .. '2_' _t"'iIii' 'ioI.;,.b~'t.·ilttll1i1ibilli'.17ii9riJIiiiIii'iii'.W.'iii'''iIi:_I-."11111.1' 1I.1iII'i!!tllilllltr __ t ......... H' .. ,~~--~- ~-' -----,-- -" ....

...

e III i ~ u >- - - - -- - (33 )

•

14' (20)

\ ('

~O I~'J

"

h. Slices 8 through 12 (See Figure 20a. for location) Fi~ure 22. Airfoil Shroud Reg;on Surface Stress Distribution, 3785 RPM (Concluded) .

i r: I: I: jl 80 x l; 40') 60 ~O ....

30 c 200 ....

Ii ....

'a ..0 20 ....

II ~ - III .

.c GI II ., ... 112 ..

GI III 0 0 ...

..

'a til 6 -10 ."

-100 ...

;:I .c -20 ...

.c GI til -200 c -30 ...

GI III!

c C ....

~ -40 tIC -300 c ~ -50 -400 -60 -70 -500 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 Fraction of Airfoil Chord from Shroud Leading Edge Concave Side Convex Side of Airfoil of Airfoil Shroud Radially Outboard Surface o~----------o ~---U Shroud Radially Inboard Surface e:...-----------~ 0----0 Finite Element Analysis 0 _______ 0 0 ____ 0 Figure 23. CF6 Improved Fan Blade: Airfoil Part-Span Shroud Region Stress Distribution, 3785 RPM at Sea Level.

.. _____________ ..:.. ___ ... _.;.. ___ iliiiiillii"·iIiii'*IIIi~ .. ' iIIIi".~Ii';ji,.·-Iiii .. iii"ii .. ii·iI·h;..II"·iI· 111"';".'"11' .am-mIllil"··IiII~ .. g .. n .. IiII. __ .. ___ ~ __ -~"~------"---"- --- -- ~ Full Shank Blade (No Pocket) 600 r----------r----------.---------~----------~----------

~ t

i /.0-j'O", ~

f ___ //-6 --t- '<\\~----I-----~

..:I 1/ 0 .,,~.:._Q::.._o--~.::Lk::- E-o

.

//U "X, CD CD 400 ~

//):1' I --"

~ ~ I '[y' , Ul I • -,

· t- t\'_--e

---- ~I Concave Convex

Photoelastic, Airfoil Surface 0---0

e __ e

_---e

Finite Element Analysis

Photoelastic, Airfoil i [J--._. _._. -0 I

o ~ ____ --J,.. ____ ...J ____ ___" _____ IL_ ___ ___l 0

o 0.2 0.4 0.6 0.8

1.0 Fraction of AirfOil Chord Figure 24. CF6 Improved Fan Blade A2, Spanwise Stress in Airfoil at Section G-G, 0.2708 m (10.66 Inches) Above Dovetail Base, 3785 RPM at Sea Level.

Full Pocket Shank Blade 500 .----------r----------~--------~----------~--------~ 70 x 10

~ I "",e- - --'. Shrouds

'0 ./" "Locked Up"

: ~~-.o;>,c;..~.:~~~<

60 N.

I

400 ~ ------+---~:7 ~.cy.......... "'.~\ 'Q.., = ....

tU O/~, .,,'" ___ .----. .~\' .0

-

~ ~ /' /.,0' /. o.

P;' /e

d' .... ~ / ~D"'""'"

d' Shrouds /' .

. /

200 I--.p / •

\ Concave convex\ \

<J ./

\

0-----<> 20

/ /. ~:~;!~:; B!:=:u::o~~;~~ed uPq~, --~o

e---.

I Analysis; Shrouds "Free"

Photoelastic, Spanwise Stress o-.-.-.-.-.~

(Averaged Through Airfoil Thickness)

I

o ! I I I

l

o

f

I

1.0 , 0.2 0.4 0.6 0.8 Fraction of Airfoil Chord Figure 25. CF6 Improved Fan Blade Cl9, Spanwise Stress in Airfoil at Section G-G 0.2708 m (10.66 Inches) Above Dovetail Base, 3785 RPM at Sea Level.

I

Half-Pocket Shank Blade

t

'0 r&l s.:

bI) i

N ~----I---- .

40 ~ .......

.0

...... - .... 1""'1

.,,"

30 rn rn Q) +> til 20 1""'1 'M o 'H 10 ~ Q) r:: oM bI) o r:: o ~ COllcnvc Convex Photoclustic, Airfoil Surface U 6

0----.0

Finite Element Analysis e-----_e

• • -10

I I

-100 o 0.2 0.4 0.6 0.8

•

Fraction of Airfoil Chord Figure 26. CF6 Improved Fan Blade 820, 3785 RPM at Sea Level.

90 x Trailing Edge Half-Pocket Shank Blade I ~., ", \ g:, 'C 1/ \ ~

tlC ,. - \

'1"1 70

=

I \ 'C as Q,) ..:I

I \

\

I 60

,--...

/ I -0.

1;5- \

/

I / \

L';t-

~ .

as Il.

~ II

'1"1

::a =

.......

--0...

.0

-

III ....

9 \

III Q,) 30 200 ,..

'-

: -+-''0._-1- ~I

// III I III -+-> ~

til '-, -0"> Q,)

,..

\ .... 'I(. / -+-> '1"1 \ til

0 1>/ .......... --'-"/

'e- .

\ ....

'H ,..

.~ '1"1 -\ ~ 'H ,..

Q,) '1"1 ct .~

=

~I

/f-: Q) tlC r::: r::: .~ r.:l tlC

= r.:l

fj--

-10 '.

Concave Convex -100 6,

Photoelastic, Airfoil Surface 6 0----.Q

-20 e _____ • Engine Blade, Strain Gage

•

• -200 -300 o 0.2 0.4 0.6 0.8 Fraction of Airfoil Chord Figure 27. CF6-50 Improved Fan Blade B20, Engine Airfoil Root Stress Distribution, 3785 RFM at Sea Level.

plotted photoelastic stresses were taken at a height above the platform cor- responding to the strain gage locations and were oriented perpendicular to the platform, as were the strain gages. The true maximum stress for each slice was not plotted, because it generally did not exceed the value at the strain gage location by more th~n 10%. Blade C19 was also analyzed in the airfoil root region and those data are compared to Blade B20 in Figure 28; again, agreement between these data is excellent.

For the four blades analyzed, stresses were determined from shank slices cut as shown in Figure 19. Stresses obtained included: (1) the maximum dove- tail fillet stresses, (2) maximum pocket fillet stresses (for Blades B20 and CI9), (3) surface stresses at a reference section 21.6 mm (0.86 in.) above the dovetail base on both (concave and convex side) surfaces (all discussions are given in terms of full-size blade geometry rather than 0.6X model geome- try), and (4) shank centerline stresses at a reference section 38.1 mm (1.50 in.) above the dovetail base. Stresses were also read at a section 15.8 mm (0.624 in.) above the dovetail base, but this value was always close to the peak value in the dovetail fillet. The 15.8 mm (0.624 in.) and 21.8 mm (0.86 in.) dimensions corresponded to locations which were used for strain gage placement in end-effects testing of metal blades, while the 38.1 mm (1.50 in.)

dimension was chosen as a convenient reference location near the base but out of the stress concentration. It serves to define the location of a nominal stress when stress concentrations are desired or to relate photoelastic re- sults to finite-element results. Figures 29, 30, and 31 are plots of stress ratios, or stress concentrations, for the dovetail fillets. These concentra- tions are defined as the ratio of maximum dovetail fillet stress to the ref- erence shank centerline stress at 38.1 mm (1.50 in.) above the dovetail base and are given in this form for generalized application to different design cond it ions.

All stresses shown on the various figures are presented as the actual values for a full size titanium 6-4 blade rotating at 3785 rpm. Equations used to calculate the stresses from the photoelastic fringe orders are given below.

The model stress was calculated from fn (1)

am .. -

t where am • model stress (MPa) f • fringe constant for the model material (MPa/m) n • fringe order t • slice thickness (m) ,..------ ..------~-----"T'""'----__,.~ ....

/ \ ~ / \ '0

:0

'0 50 x 10 rIQ

/ • .t6-~\ ~

I I \\ C ~ 'f'4 ----.-+---- , I \\ ;:: '0 (II 6 I / \\ -; ill .,;J .... /, \\t,

'O..~ \'

~" \'

---+-- 'Q....- .... ,,/ \

--(Y ~

....

'f'4

\\

o

\

t 100

'f'4 « ill c 'f'4 \ bIl C -~.-----+-------+--- r.:I 0 o Concave Photoelastic, Airfoil Surface (Blade

B20) 6-----6

-10

Photoelastic I Airfoil Surface (Blade C19) & ... --.Aa

e----.

I -100 o 0.2 0.1 0.6 0.8 1.0 Fraction of Airfoil Chord Figure 28. CF6 Improved Fan Blades 820 and Cl9. Engine Airfoil Root Strpss Distribution, 3785 RPM at Sea Level.

.# .

.P

:: 4 r---,---+-----t-----+-----+---- r;/ •

.. ~ n I ~

w ""'\ i

g ~ \ I

~ ~

~ ~ \ 0.. 1-

... ~ ~-O- '0.. ~

~ 3 i -- " 0"" I .... ~ }J --;:

~.! ........, I '0- -"" ~

~

III III C> ..

+' {Il 2 t---- QI III DOVetail 'f'4 Fillet ~ as fJ;

t

Concave Side Convex Side 0.01111 • (0.124 111.)

of Airfoil of Airfoil

Dovetail Fillet 0-----0

6 6

[lovetail Base 0.4 0.2 0.6 0.8 1.0 Fraction of Chord Figure 29. Dovetail Fillet Stress Concentration Kt for Full Shank Blade A2. 3785 RPM at Sea Level.

t

4 ~ ----+------- tIQ& -= ....

~ ~ \ .

Dovetail Ba8e

. "

~ ' .....

o b ....... •

/ ~ .......... O 0 ......... 0 ;'

8 • /0

~, - ~/ ~

-""0... _--:::-::::- _" +

t ...... ~

•• •

•

I 6\ 6 6

Pocket Forward Edge ~ + Pocket Aft Ed.e -~----1

Concave Side Convex Side \ of Airfoil of Airfoil Dovetail Fillet, Blade B20 ~j CJ 0------0 Dovetail Fillet, Blade B23 ... .....------ •• e-----e o L- ______ ~ ________ -L ______ ~ _________ ~~ ______ ~ o 0.2 0.4 0.6 0.8 1.0 Fraction of Chord Figur~ 30. Dovetail Fillet Stress Concentration K for Half-Pocket t Shank Blades B20 and B23, 3785 RPM at Sea Level.

...

Ii

.. GI

tao "0 ~ IIQ ~i IIQ

"

tao

-

c:: 4 tao --.-~~-- ---~---- Dovetail • I 'f'4 ...

....

~ Fillet

•

Ii 'f'4 ;; ..

co co 0

"

~ ,;

I

I ~ foo ~

-

•

0.011' •

'~

(0.'14 &11.)

\

\

~,

*

Dovetail B.se

',0

, 1------- , 2 , "'Q.. ....

.... -.0

.... - ---0-

o -0---0- -a-

t I Pocket Aft Edge __ .1

l'-Pocket Forward Edge -.. I -., Concave Side Convex Side of Airfoil of Airfoil Dovetail Fillet 0---- .... 0 0-------0 o o 0.2 0.4 0.6 0.8 1.0 Fraction of Chord Figure 31. Dovetail Fillet Stress Concentration K for Full Pock~l t Shank Blade C19, 3785 RPM at Sea Level.

_______ ........ _ .. 1 r....;o;- .. .:::..IIiI·~.· "'·~.··Ii'·I11- 1I •• rlllr.· .I:I!,..-.;.'.l:l:""'''''- _______ - .- ---- _.

@,; w· ;: The prototype .tre •• was calculated from where a • engine part .tres. (MPa) p Yp • material density of engine part <g/m ) Y • den.ity of model material (81m3) m wp - r~tational tpeed (3785 rpm) of engine parts (rad/sec) wm • rotational .peed (163 rpm) of model (rad/.ec) , - characteristic length of engine part (m) p 1m - characterittic length of model (m) The geometric scale factor i.

''1\

(3) -. 0.6 tp Equation (2) for the prototyp~ .trest eventually reduce. to n (4) a p • k t (MPa) at 3785 rpm ~~ere k i. a con.tant factor, the value of which depends upon the material, physical properties, and units used to expres. the other quantitie. in the equation.

Model to prototype strain ratio (tm/t ) wa. calculated from p tm (5) f -.

. (~)

(~: ) 2 (:;~

(:; ) k

(~) tp

where Young'. modulus of engine part p E • Young'. modulus of model at critical temperature

Em-

-,.~ -- ~ ______________________ ~ ________________ ~~ .... -~--_"~-.~~-~,.~-~ .. ~~~.· .... ~7~j~~~- ____ ~- ___________________________ ~ "f For thi. test. Equation (5) reduce. to em _. 1.20 cp Poillon'l ratio for the model iIO.S.

Some unfor.een probleml were encountered in thi. te.t, mo.t of which were cau.ed in one way or another by the unulual large .ize of the photo- el.ltic model, and .ome of which warrant further inveltigation.

For example, the problem of part-'pan Ihroud "shinglinlt" ~as noted, AnJ corrective action wa. taken which .olved the problem. Yet, the shroud. did not lock up a. expected by theoretical analy.i. of blade untwilt and as ob- .erved in engine telt. In the ~hotoela8tic te.t, all Ihrouda were ob.erved to be f'Jlly cloaed (and precUlTled locked up) at the streas-freezing (or criti- cal) temperature; but .a the chamber temperature was alowly decrpM~ed. lome •• all gape appeared and moved around among the blades. This impliP8 a 108s of lockup and .ome undesirable relative motion of the blades. Photoelastic epanwi.e airfoil .tre.se. at reference Section G-G. 270.8 mm (10.66 in.)

above thr dovetail ba.e, were compared with theoretical results fer both "locked" and "unlocked" ahrouda <lee Figure 25). It should be noted that the ph~toelaetic atreaa resulte fell midway between the two, poseibly indi- cating a 10 •• of that lockup.

For this atre.s-freeze telt, the choice was made to rotate the model at .n rpm corre'ponding approximately to a 1:1 strain ratio with the full- size engine component. This wal done primarily to achieve realistic blade stressec, especially thOle due to untwist. and presumably to achieve realis- tic .hroud lockup. Sev~ral undesirable side-effects of this choice affected the accuracy of the re.ult.: • Fringe orders were low, magnifying the r~lative value of any re- .idual .tres.es or reading inaccuracies. This was partially cv~r come by taking relatively thick slices.

• Gravity load effects during rotation were relatively large compared to the centrifugal load. The gravity force allo produced ai/rev cycle which is considered undesirable during a stress-freeze.

• Hodel defects in such thingll as dovetail pressure-face fitup would be expected to show up in an exaggerated manner.

Results .hown in Figures 29, 30, and 31 show successively reduced con- centration factor. with increased pocket .ize. However, this does not neces- .arily mean that the peak stress ctf'cr~ases, since the "nominal" strell! 18 increa. ing at the same time. There does appear to be some .: ield ing of the dovetail fillet by the proximity of the pocket, but this benefit may b~ off- set locally by the flow of stress around the end. of the pockets.

5.0 FAN BLADE BENCH FATIGUE TEST Objectives of the improved fan blade bench fatigue tests were to demon- strate that (I) the design has no stress risers at the airfoil edges, around the part-span shroud, in the airfoil root platform shank dovetail area, or elsewhere; (2) the manufacturing processes produce a fan blade whose fatigue strength is in accordance with the material specifications; and (3) adequate design margin exists between measured stress levels in engine testing, such that no field fatigue problems would be anticipated.

5. I TEST SETUP AND TEST CONFIGURATIONS A bench test setup was utilized for these tests. An inner panel section of the fan blade was clamped at the dovetail and supported as a cantilever beam as sho~ in Figure 32. The fan blades were excited using a siren which produced air pulses by air flowing through a slotted disk rotating at con- trolled variable speeds. Pulse frequency was adjusted to coincide with the test parts fundamental vibretory mode frequency by varying the rotational speed of the slott~rl disk. Test component tip amplituu~ w~s measured through a calibrated microscope. Siren frequency and input were held constant during the fatigue test. Component failure was determined by a decrease in blade amplitude and resonant frequency.

Fan blade forgings and finished blades were used in this test. Test bars were machined from production forgings in the areas shown in Figure 33.

Two notched specimens were m&chined from the dovetail area of the forging for low cycle fatigue testing, and two smooth specimens were machined from the midspan shroud area of the forging for high cycle fatigue testing. All test bars were unpeened.

Production fan blades were cut 127 mm (5 in.) below the midspan shroud.

The lower blade half, consisting of the dovetail and the inner panel, was used for the inner panel test as-is. The upper blade half, consisting of the outer panel, midspan shroud and the inner panel stub below the midspan, was used for the outer panel test and the shroud test. For the outer panel test, a "kirksite" block was cast around the inner panel stub up to the top of the shroud. For the shroud test, a 3imilar block was cast around the convex side of the airfoil. The blocks \'ere mach ined square and used for clamping during the test.

5.2 INSTRUMENTATION No instrumentation was appl ied to the notched low cycle fatigue test bars nor to the smooth high cycle fatigue test bars. These tests were con- ducted in the Materials Laboratory where applied axial loads for low cycle fatigue tests and bending moments for high cycle fatigue tests are accurately measured.

. j . --::""--# ... ~~ .... '-,' ' ..

XaaCJHta . ~t .. A4iiieC2, LJMk,a, jj i _,It '1

$$2"_".##, anOCSJLS saNto. iUSQpj§l., Cycle Fatigue Bars Outer Panel Inner Panel Fatigue Bars

I~l ~ Low Cycle

L~y Dovetail

Figure 31. Location of Test Bars Taken From Forging.

- , - ~ .,." " .~-;- - For th~ inner panel fgtigue tests, eight strain gages were used. all on the concave side of the blade and shank at the leading edge. as shown in Figure 34. The gages were positioned to be in the area of highest stress.

Locating this area was based on previous dynamic stress distribution testing and on failure initiation from previous fatigue tests on similar fan blades.

Four gages were used on the airfoil. The first airfoil gage was right at the fillet tangency point. The next three gages were spaced 25.4 mm (1 in.)

apart. The four shank gages were located at the edge of the corner radius tangency point. The spacing of these gages was as close to the adjacent gage as possible. Grid sizes of all gages used were 1.6 mm and 0.8 mm.

Five strain gages were used on the outer panel fatigue tests. as shown in Figure 35. Selection of these locations was based on the same previous fatigue test discussed above. The gages were always located on the fillet tangency point and perpendicular to the tangency line. Failure initiation was always very close to the strain gage.

Eight strain gages were used on the con~ave shroud fatigue tests &s shown in Figure 36. Gages were placed back-to-back on the top and bottom sides of the shroud. The fillet tangency point was consistently used as the locator from the edge or airfoil. It was not possible to drive the shroud to an ampl itude that would produce failure. so it is not known how the instru- mentation corresponded to the fatigue InItIation points. However, previous tests on similar fan blades indicated that the middle gage in the group of three was the initiation point.

5.3 TEST PROCEDURE For the low cycle fatigue bars taken from the shank section of the for~ ing, an axial load to produce a stress level of 62,055 N/cm (90,000 Ib/in. ) was applied to the notched bars. The load was then cycled from 0 to 62,055 2 2 N/cm (90,000 lb/in. ) for the entire test until the bar failed.

A moment was applied to the first smooth bar, taken from the shroud sec- tion of the forging. This stress was cycled from a compressive to a tensile 2 2 stress at the 37,922 N/cm (55,000 Ib/in. ) level. Since the first bar tested did not fail in 10.6 x 10 cycles, the second bar was loaded to give 2 2 a stress level of 44,817 N/cm (65,000 Ib/in. ). The second bar sustained 15.5 x 10 cycles without failure.

The inner panel fatigue testing was accomplished by clamping the cutoff inner panel of the blade on the dovetail pressure face. The blade was then vibrated with an air siren at its first flexural frequency. The highest read- ing strain gage was on the airfoil. Previous tests with a similar blade had shown the maximum stress to be on the shank; that blade th~ugh had a larger pocket in the shank and that difference probably explains the difference in stress distribution. The siren air pressure was then increased until the 2 2 highest reading strain gage read 46,450 N/cm (67,368 Ib/in. ). The blade did not fail after 5 x 10 cyclps were accrued. 1be stress level was then 2 2 increased to 49,988 N/cm (72,500 lb/in. ). When the part still did not , f.

Concave Side Strain Gages 1 in.

Figure 34.

Inner Panel Instrumentation.

~ ~ ~ ~ . -" ,-"'. , .. . " . , ~- ~ -..-.....; ........... i.~~"' .... ~=-, _ .--.... .. "'"""';._.,.. .............. _-...... ........ '"'~~.....- ...... __ """" .. _~:.-_~ __ ... ~ .. __ .. '*'- ....... __ .. , ... _.".~ ,jU.. .¥'iH'¥!,~;i. , .. / ¥ .. ··,,··;;:;;::W; 4 W?-4.Q.

Concave Convel( Side Stra1n Gages Bide Kirkslte Kl rksi te

L_. __ _

------- -- .. - ----'> Figure 35.

Outer Panel Instrumentation.

Top Side Strain Gare-

..- j 6 _------

r----..:::::::..--.--

Conc:ave Klrkdte Figure 36. Concave Side Shroud Instrumentation.

_ -"L-~"-H~'W='* p' .....

2 2) fail after 10 cycles, the level wag increased to 53,436 N/cm (77,500 lb/in.

for another 10 cycles - again withl')ut failure. Finally. the level was 2 2 4 increased to 56.883 N/cm (82,500 lb/in. ) and failure occurred prior to 10 cycl~s. The same procedure was followed on another inner panel specimen, 2 2 except the test was start~d at 49,988 N/cm (72,500 lb/in. ) for the first 2 2 5 x 10 cycles and failure finally occurred at 60,330 N/cm (87,500 Ib/in. ) in under 10 cycles.

The outer panel fatigue testing was conducted by clamping the kirksite block. which had been cast around the inner panel below the part-span shroud.

The blade was driven in its first flexural mode of vibration b~ an air siren.

The maximum strain gage reading of 46,540 N/cm (69,500 lb/in. ) was attained for 5 x 10 cycles without failure. The stress level was then increased to 2 6 49,988 N/cm (72,500 lb/in.) for 10 cycles without failure, then to 53,436 2 2 6 N/cm (77,500 Ib/ir. ) for 10 cycles without failure. Finally, at 56,883 2 4 N/cm (82,500 lb/ ir .. ), fa ilure occurred in under 10 cycles. Another outer panel specimen was tested. The initial stress level of 49,988 N/cm (72,500 2 6 lb/in. ) for 5 x 10 cvcles was completed without failure; the part failed at 2 4 a level of 53,436 N/cm (72,500 lb/in. ) in under 10 cycles.

TIle concave side shroud fatigue test was conducted by clamping the kirk- site block that had been poure~ around the convex side shroud and airfoil and driving the shroud with an air siren at its first flexural natural frequency.

The stress level on the highest reading strain gage was 46,540 N/cm (67,500 lb/in. ). At this level, 5 x 10 cycles were accrued without failure. The 2 7 amplitude was then increased to 49,988 N/cm (72,500 lb/in.2), where 10 cy- cles were accrued without failure. Since the siren waR incapable of driving the shroud to a higher amplitude, the testing was terminated.

5.4 TEST RESULTS AND COMPARISON TO PREDICTIONS Results of the round bar low cycle fatigue test are shown in the room temperature fatigue diagram in Figure 37. Both bars were tested at 62,055 2 2 2) N/cm (90,000 Ih/in.2) peak stress or 31,027 N/cm (45,000 lb/in. alter- nating stress. The first test bar failed in 15,528 cycles. The second bar failed in 23,192 cycles. The average of these two test points would be 19,360 cycles. This compares with a material spPcification average life of 20,000 cycles, or a -30 standard deviation life of 6000 cycles. These re- sults are in good agreement with the average life and the material low cycle fatigue characteristics appear to meet all life requirements required of the fan blade.

High cycle round bar test results are shown on the Goodman diagram in 2 2) Figure 38. The first test specimen waF run at 37,922 N/cm (55,000 Ib/in.

alternating stress. This bar did not fail after 10 cycles. The second bar was tested at 44,817 N/cm (65,000 lb/in.) and it did not fail after 15 x 10 cycles. These results compare well to the average materials specifica- tions high cycle fatigue strength at an "A" ratio (alternating stress divided by mean stress) of infinity. The average high cycle fatigue material strength ..

Ti 6-4 Fordn.a Room T.mper.ture Stre.e/Cyclee to Fail Fatiaue Diaara., K • 1.5. Ao • 1.0 t Axial-Axial Loadin;. f • 20 Cycle. per Minute, (0.33 HZ) Low Stre .. Ground Specimen S'Jrface Finhh Solid Line· Average Haterial Propertie. of Notched Bar (K • 1.5) t Oa.hed Line· Averwae H1nu. 30 Deviation Unit.

N 100 x 10

-

e 60 N

.

....

80 c

-

~ ...

....

,Q

. SO

-

•

" • .

t 60

40 •

...

,

•

II) II

"

..

...

g II)

-

... "

30 ...

~ """IIit:- g U ...

Q/ ...

II) 40

-

" ~

u Q/ ... .... ,

"

- V.

Q/ IE ...

"

' ... Q/

20 :r- ~ ....

... DC

"T-~

c: ca ...

c: ...

..

Q/ I

... ~

... ..

Q/ < ...

" < , I I .2 .3 4 10 10 2 5 10 2 5 10 2 Cycle. to Failure Figure 37. Fatigue Characteristics of Titanium LCF Specimens Made from Fan Blade Forgings in Dovetail Area.

-~ .L...i1rc r .- trt #4- -¥- Ti-6Al-4V Forging Room Temperature Goodman Fatigue Diagram Solid Line - Average Material Properties of Smooth Bar Dashed Line - Average Minus 30 Deviation Units Mean Stre~q, 1b 50 100 150 200 x 103 o , , I 140 x 10

-

- 120

N .

C ....

N ,Q a

100 -

CJ -

.... """

..

Z en .!IG CIl C1) ..

...

en

- 4J

en tn C1) eo 4J C tn ....

~~

-

40 4J eo ro s: )G:.._ .....

."", c ...

4J ..........

C1) ell 4J ..........

s: ~

100 -

..........

C1) <: 4J ...........

""" 20 <: .. [ .......................

-

......

I .............

o

I

o

o

20 40 100 120 140 60 80 2 Mean Stress. kN/cm I Figure 38. Goodman Diagram for Titanium HeF Specimens Made from Fa~ Blade Forgings in the Midspan Shroud Area.

f'-I 2 2) is 44,128 N/cm (64,000 Ib/in. and the -30 standard deviation strength is 37,233 N/cm (54,000 1b/in. ).

Inner panel fatigue te.t re.ult. are .hown in Pilure 39 and are repre- lented by the circlel. The level of the firlt telt wal initiated at 46,540 2 2) N/cm (67,500 Ib/in. and increa.ed in increment. of 3450 N/cm (5,000 lbl 2 2 2) 1b/in. until the part finally failed at 56,883 N/cm (82,500 lb/in. ). The 2 2) second lpecimen wa. tested at an initial 49,988 N/cm (72,500 Ib/in. and 2 2) allo i~creaeed in 3450 N/cm (5000 Ib/in. increment. until failure at 60,330 2 2 N/cm (87,500 Ib/in. ). The ... relulte compare well with the materiale S-N 2 and a curve that Ihowe an average runout of 51,000 N/cm -30 standard devia- 2 2 tion .trength of 43,440 N/cm (63,000 lb/in. ).

Outer panel fatigue telt re.ulte on £lnilhed bladee are aho Ihown in Filure 39 and are repreeented by the trianllel. The eame levels were tested and increased incrementally the .ame a. for the inner panel telt. The first 2 2) teet epecimen failed at 56,883 N/cm (82,500 Ib/in. and the lecond at 2 2 53,436 N/cm (77,500 Ib/in. ). The outer panel points plot the lame al the inner panel on the Goodman diagram in Figure 40.

Part-.pan Ihroud fatigue teet relult. are Ihown in Figure 39 by the rec- 2 (67,368 2 ) tanlular Iymboll. A level of 46,540 N/cm Ib/in. wae run without failing the .hroud. The telt wal terminated without experiencing a Ihroud fa- tigue crack becaule of the limitations of the air siren.

Bench fatigue telt results with both the round bar test specimene and the finished airfoil demonstrated that the improved fan blade delign is equal in fatigue margin to the current CF6 fan blade. The current CF6 fan blade has never experienced a fatigue failure in over 12 million flight houri. The fatigue testing demonstrated that the improved fan blade has no high strels concentrations to degrade the fatigue strength of the design. A substantial margin exists between measured engine atres.es and the fatigue capability of the design.

.~ Tl-Ml-4V BlAd" Room T"mpt'ratut(' Stn'flR/Cvch!'R to Fa! 1 Fati~u .. 111ajtum Gla." 8 ... d r.'"n .. d 8PndlnR-n .. n,lln~ L.'ad In" A •• ' 0-- ----------, 60~---------------------~----------------------- 0-- J

I A'· .... Il ..... h-rlul rrul"'rt 1". /,1

.. ,,-- 0---

~o ~

~ -+-

..

..

• ....... -...

'" • ..

..

" ..

" ..

--------------r--

--

.

'" c I -3- rro""rll ... - -'- /'0 ..

!

~ .. 40 ..

1 ~ :< Olnn~r POI ... I :: ..

• ..

6out~t 1'.",,1 !

< ..

~ /6hr o ud ,.

-.",,/

I

I JII L _____ ..... __ i..... __ --I-:-_ ----------- ....... -- ----~---'--- ...

IO~ :I Ill'; 10 Figure 39.

Fatigue Characteristics of Fan Blade Inner and Outer Panels and Part-Span Shroud.

- ....

Ti-6Al-4V Blade RI Fatigue Limit Diagram Gla.. Bead Peened Estimated Condition: 6-12 N. No. 98 Bead. 125% Solld Line • Average Materlal Propertle8 Dashed Line • Average Mlnus 3c Deviation Unlts 140 x 10

-

• Bendin~ Alternating Stress Only

-

120 N N

._---

~ .

--

C

•

~ ~ .....

.t:l ~

-

~ ..

..

CI'l en CI'l en -.---- ~- GI 60 >------ GI ...

...

..

-

...

rn ~ tIC ,~ 1:10 c , C of'4 ~ ~

-

c ~, 'g j GI

=

~ co c

-

of'4 c

I

~

.. "

...

Ie

"-

C

"

... ~

...

GI

.. GI

-

"'- ..

r-I

'"

r-I -< <

"

o

I I

o 0 20 40 60 80 100 120 140 Meon Stress, N/cm Figure 40. Goodman Diagram for Improved Fan Blade Inner and Outer Panels and Part-Span Shroud.

. __________ .... -IiiI· ____ .;.. ___ ."IiII-~Ii' .. ··.··.·II1·i...iI· 1IIii.·II.L.....I··.~.·~.·-.- ... ~ .. --_ ...... _____ ._ .. _._. _. _____ ~_

r

6.0 ENGINE CROSSWIND TEST

-

Objective. of the cro •• vind te.t vere to eval~ate the perfor .. nce of the improved fin blade on a CF6-S0 en,ine ~nder the infl~ence of inlet di.tortion ind~ced by 90· cro •• vind. Ind to demon.trlte the .eromechanicil operation of the i.proved fin blade ~.in, both the D~ala. DC-10-30 and the loein. 747 in- let. at ero •• vind velocitie. ~p to 64.8 to 74.1 km/hr (35 to 40 kn).

6.1 CROSSWIND TEST FACILITY AND CONFIGURATIONS the te.t VI. cond~cted in the outdoor cro •• wind te.t facility .hown in Fiaure 41. A brid,e-type .tr~ct~r~ .~pport. I turntable thru.t fra .. for overhead enaine mo~ntina. Cro •• wind. ~p to 185 ~/hr (100 kn) for a .t.tion- ary enline centerline and up to 157 km/hr (85 kn) It anIle. between O· Ind 90· can be created by the thirteen 5664 .3/min (200.000 efm) two-.ta,e Ixial flow variable pitch f.n. driven by 150 kw electric .otor. in thi. flcility.

Autoaatic d.ta handlin, in thp Idj.cent control blockhou •• c.n be proce.,ed on .n Evendale time-.harinl computer.

Te.t vehicle. for the DC-IO-30 inlet and 8747 inlet te.t. were CF6-S0 en- gine.. The fan confiJUr.tion w •• the .Ime for eleh te.t Ind eon.i.ted of the followin,: • Improved f.n bl.de. with f.n c •• e .tiffener to permit tip cle.r.nce to be .et It 2.92 .. (0.115 in.)

• Open cell aluminum honeycomb f.n e •• inl tip .hroud.

• Inlet r.ke .y.tem which repl.ced the .t.ndard .pinner with a .plit .pinner to provide .upport for the rake •.

6.2 INSTRUMENTATION 6.2.1 Strain G.,e.

The .train I.,e. u.ed fo~ .... urina f.n bl.de .tre •• e. were 1.6 mm (1/16 I in.) grid. dynamic type. with 350 ohm re.i.tan~e. Str.in .iln.l. were routed out through a forward .lip ring and di.pl.yed on o.eillo.cop~ ••• well a. re-

I

corded on .. anetie t.pe •• Four fan blade. were In.trumented with. total of 16 .train g'le. at four

t

different locltion.. The.e bl.de. were in.talled in Di.k Slot. 1. 10. 19 and i , 29. Figure 42 i ••• thematic of blade .tr.in Pie loc.tion.. Theae gage lo- ; cation. were cho.en to in.ure a rel.tively high level of re.ponse to all an- ticipated vibratory mode •.

'1 ~ , C

"

' ''; ;3 C/J C/J o 1..0 U ClJ C ' ''; C "'-l ' ..

!, I ~ !

'" If

'-...-SD6

A

SDI

j

CON AVE

CO~VEX

STACKING

AXIS

SECT A-A

SHK

Disk Blade Slot Gases I I SDI, SHK. 5D6, 5DA SHK, 5D6, SD1, SDA SHK, 5nl. SD6. SDA SilK, 5))6, SD1, SDA Figure 42. Fan Blade Strain Ga~(' Instrumentatiou.

.... ~ -,.". ""_ '" li, ""- , '" ~ -" - . -~- , - - - .. - "' ~ . .

6.2.2 Fan Inlet Aerodynamic Instrumentation Fan inlet aerodynamic distortion w~s monitored and measured in the cir- cumferential and cadial directions by steady-state pressure and temperature probes mounted on rakes as illustrated in Figure 43.

The steady-state instrumentation was located slightly forward of the fan and consisted of six rakes spaced circumferentially clockwise (aft looking forward) at 38·,98·, 158°, 218·, 278°, and 338°. Each rake contained six total pressure probes, six static pressure taps, and five temperpture probes.

There were also two boundary layer rakes located at 55· and 270·. Each boundary layer rake contained nine total pressure probes located off the wall at immersions of 5.1 mm (0.2 in.), 10.2 mm {0.4 in.}, 15.3 mm (0.6 in.), 20.4 mm (0.8 in.), 25.4 mm (1.0 in.), 38.1 mn (1.5 in.), 50.8 mm (2.0 in.), 63.5 mm (2.S in.) and 76.2 mm (3.0 in.).

There were also two dynamic pressure transducers in the fan discharge wall. The test instrumentation which was uSt~d to monitor engine operation 1S shown in Figure 44.

6.3 TEST PROCEDURE The engine with the DC-IO-30 or B747 inlet was installed in the test stand and set up with the engine oriented with the crosswind from the left (aft look- ing forward) ~t an angle of 90° with the engine axis. Instrumentation was con- nected. The following test sequence was performed: 1. Perform normal prefire checks, including .i leak check.

2. Start the engine and stabilize for 5 minutes at ground idle.

3. Perform the mechanical checkout and break-in run.

4. Set fan speed at flight idle, and slowly turn on the required number of facility fans, one at a time, to obtain the desired crosswind velocity.

5. Slowly accelerate engine to the desired corrected fan speed, sta- bilize 3 minutes, and take data. If maximum fan speed cannot be reached because of fan stalls or high vibrations, back off 100 rpm, stabilize 3 minutes, and record data.

6. Re,~ord the fan speed at which the inlet flow reattached; and define the inlet separation, inlet distortion, and engine stability for each crosswind condition.

The engine with the DC-IO-30 inlet was tested for about 35 hours in order to complete the following crosswind test points: Fi g lln :' 43 . Filn In] L' l In~tnllnlnt;lti.'n i"n l" Cr "ss \vind TL'St .

..."

o !I')f) SH R 1 ~; ,; n '_ ' I ~ N TA T IO ~ -~ Y ~ ~ : ~ ~ ii :· sic '";' \ :- i: \c ,; o RU OE S IS AU ~ O T O R S TA ClS (l ~ TO lA L ) l1 ' '' ~J': IC S I C O ~ l D ~ II (' f VA/ . i:S I ~; S TAT ORS I AS O 2 ([2 TOT AL ; ;: il r);A HI C SI c .;s £~C If OF V , ~ ; E; IN : ;rA:l ll ~ 1 J ,-,SO ~ ( b TO TAL) 1 lJY)l A. '1I C 'Ie 0 , 2 tJC V 'S ( 2 1 OR J B/l RIS ES

-

TIc' s O~ f O RI.'ARD IIAT NEW nw AfT IN NEIl SE.> ,t :. ! ~ ; li' T P .AI :£S 21 !'I t:T Al .ic's - 8 AIR TI C'S 2 DYI'IA.'!I': SIc's DYN A.'1I C S IC C~ EA CH )e B lADE S ( 16 TO lA l) O', SA' lI C S IC OS EA CH B UDES ( 6 TOTAL) lW W •. ': 1:C ~ / C ' S " , J FA.~ S HA fT ' ;-" 1» 2 RE r r,\~l. 'LAA J ROS ETTES SI C ' S ~ ~ RlPoC

c

-, ( .J ~ £U:II£IIT) ' ........

0-

~ - ~ - . --

c:.

~- -- STA"D~RD TEST '~STP.UJl[~I~ T1IIIUST ~L fAN DISC P .ARCE Kl~rTE !UOS! [ R u , SC HAR C~ KU l lTE ---" ~; ".

"t.

T •• P (' f'ROUI) TU lfAnIlUL ":JrD DI!ECTrow • nLOtlTr Fi g ur e 44.

CF6- S0 Engine - Instrum e ntation f or Crosswind Test.

CtQllKiDd [I~ili~!(l) No. of Fans/Average Crosa- Fan SI!eed wind Veloci~!1 kmLhr rpm % 0/0 4/33.9 5/42.6 7/60.0 9/65.8 Ground Idle X Flight Idle X X X X X 2500 72.8 X X X X X X 2746 80.0 X 3350 97.6 X X X X X X(2) 3670 106.0 X X X X 3783 1l0.2 X X(2) X X X 3900 113.6 X (1) Figure 45 identifies the facility fans in operation (2) Not completed because of facility faHure 6.3.1 DC-10-30 Inlet - Baseline Test Procedure The baseline t~st was run with an ambient tailwind gusting up to 9.6 km/hr (6 mph). The engine was orerated from ground idle to 3900 rpm corrected fan speed. The following steaay-state speeds were set and held for a period of 5 minutes: Flight Idle

•

rpm

•

2746 rpm

•

3350 rpm

•

3670 rpm

•

3783 rpm

•

3900 rpm

•

6.3.2 DC-lO-30 Inlet - 90· Crosswind Test Procedure With tour facility fans on, the engine was operated from ground idle to 3900 rpm corrected fan speed. The following steady-state speeds were set and held for a period of 5 minutes: • Fligh t Id le • 2500 rpm _ _ r. ,,-.~~ ~ .:- ~ 111';-, T - Io.i,.of.- ,1 .11'''' ... , tI!iio1fj;~ ~. ~...:!; '" ~ , ~"""",~~""", DC-10-)0 In~et rests Average Crosswind Velocity No. of Fans knots k~hr Fans 1n 02eration 4 18.3 4,5,9,10 33.9 5 42.6 23.0 4,5,7,9,10 6 27.5 51.0 3,4,5,7,9,10 32.4 3,4,5,7,9,10,11 7 60.0 65.8 35.5 3,4,5,6,7,8,9,10,11 10 72.5 39.1 2,3,4,5,6,7,8,9,10,11 13 92.3 49.8 All

I

!

Looking Upstream Figure 45. Identification of Facility Fans for Crosswind Test.

• 2746 rpm.

• 3350 rpm • 3670 rpm • 3900 rpm This procedure was repeated, except for the 2746 rpm, with five and seven facility fans on. With nine facility fans on, the engine was operated from ground idle to 3580 rpm physical speed. The following steady-state speeds were set and held for a period of 5 minutes: • Fligh t Id '-e • 2500 rpm • 3350 rpm While taking data at the 3350 rpm point, two facility fans failed result- ing in foreign object damage to the test fan blades which terminated the test.

New facility fans were installed prior to the engine testing with the B747 inlet. The new facility fans generated less crosswind velocity than the origi- nal fans. Four facility fans generated 33.9 km/hr (18.3 kn) while the DC-10-30 inlet was being tested, whereas four facility fans only generated 29.6 km/hr (16 kn) when the B747 inlet was be ing tested. This explains why there wal: some difference between the two tests in the velocities used even though the same number of facility fans was used.

The engine with the B747 inlet was tested for about 15 hours in order to determine inlet separation, inlet distortion, and engine stability for the fol- lowing crosswind test points: Crosswind FacilitI(l) No. of Fans/Average Cross- Fan S:eeed wind Velocity, km/hr rpm 0/0 4/29.6 5/35.2 7/40.8 9/59.3 12/79.7 Ground Idle X X X X X X 2500 72.8 X X X X X X 81.6 2800 X X X 3350 97.6 X X X 3600 104.9 X X 3898 113.6 X (1) Figure 45 identifies the facility fans in operation ~ . , ~ _ ~_ ~ ~ . _ ,_ ~ _ ~ J ',"""" .S . 3 ¥- _.2 __ ,· .43#- Lt 11

£iI ~.l

6.3.3 B747 Inlet - Baseline Test Procedure The baseline test was run with an ambient tailwind gusting up to 6.4 km/ hr (4 mph). Th~ engine was operated from ground idle to 3898 rpm physical fan speed. The following steady-state speeds were set and held long enough to collect stress data: • 2500 rpm • 3350 rpm • 3600 rpm 6.3.4 H747 Inlet - 90 Crosswind Test Procedure With sufficient facility fans to supply 29.6 km/hr (16 kn) crosswind, the engine was operated from ground idle ,:,) 3550 rpm physical fan speed. The fol- lowing steady-state speeds were set 8tld held long enough to collect stress data: • 2500 rpm • 3350 rpm • 3530 rpm With 35.2 km/hr (19 kn) crosswind. the engine was operated from ground idle to 3400 rpm physical fan speed. Steady-state speeds of 2500 rpm and 3350 rpm were set and held long enough to collect stress data.

The fan speed was limited to 2800 rpm for the remainder of the test. This fan speed is sufficient to provide breakaway power levels required to achieve forward aircraft speed for initiating takeoff and inlet cleanup. This is con- sistent with the rolling takeoffs practiced in airline service.

With 40.8 km/hr (22 kn) crosswind, the engine was operated from ground idle to 2800 rpm physical fan speed. Steady-state speeds of 2500 rpm and 2800 rpm were set and held long enough to collect stress data. This sequence was repeated with 59.3 km/hr (32 kn) crosswind and repeated again with 79.7 km/hr (43 kn) crosswind.

6.4 TEST RESULTS AND DISCUSSION 6.4.1 Modes of Vibration Modes of vibration and levels of response of the improved fan blade were observed throughout the engine operating speed range for various levels of distortion. Blade characteristic vibratory modes and corresponding infinite - - ~ -- ---- """ t _ _ _ - - - ' life .tre •• limit. a.sociated with the strain aaae locations were determined and an analysis of the vibratory behavior of the tan rotor a •• emblie. a. a system wa. conducted.

Significant .ystem vibratory mode. Ob.erved in the fan are: firlt 3- diametral, second 6-diametral and second 5-diametral, given in the order of occurrence from around idle to maximum fan .peed a. shown in Fiaure 46.

In sy.tem mode., the blades respond in pha.e with each other. A .ys- tem mode of the N-diametral type for a rotating sy.tem i. u.ually character- ized by a vibratory wave which travels in the direction oPPolite to that of the system rotation. Thu., when the system speed and the vibratory .peed are equal, the vibratory motion appears standing still to a stationary oblerver.

This occurs when the freque~cy curve of the N-diametral mode cro.ses the N- cycle per revolution line. The speed at which this occurs is referred to as the N-diamelral critical speed. This is the only speed at which this mode can be excited to any .ignificant stress level. The first 2-diametral mode has also been observed at high engine operating speeds. Since the frequency curve for this mode does not cross the 2/rev line in the engine operating speed range, the observed levels have been low.

The other type of vibration observed on the fan blades is out-of-phase mode vibration. During this type of vibration, blades tend to behave inde- pendently of each other. Figure 47 is an out-of-phase Campbell diagram for the advanced fan blade. The only significant out-of-phase modes observed on the fan blade have been first flexural, first torsional, and second flex~ral modes. Peak response of these modes generally occur at the crOS80V"1" point of frequency curve and the N-cycle per revolution lines.

Vibratory characteristics of the improved fan blade are very similar to the current CF6 production fan blade.

6.4.2 Stress Limits Infinite life stress limits were calculated for each vibratory mode at each strain gage location where significant stresses occur in that particular mode.

The meaning of infinite life stress limit is that, as long as the stress at a particular gage location in a particular mode does not exceed this limit, the maximum stress experienced in the part is less than the endurance strength of the material, and the part will operate safely for an infinite time at this condition. The limits include factors to account for: (1) sharp corners, fillets and other discontinuities when necessary, (2) blade-to-blade varia- tions, and (3) electronics accuracy limit.

Infinite life stress limits for the principal vibratory modes of the im- proved fan blade are presented in Table I and Figures 48 through 51.

350~----~-- __ ~ ______ T- ____ -r ____ ~~ ____ ~ ____ ~ ______ ~ ____ ~

tt Significant Modes

,611/'1'''

300 r-----~------4_------~--- __ ~-- __ ~-- __ _

~~1a-

metral 250r- ___ -+ ____ +- __ - 6 Diametral 4/Rev 200~-----+ ______ ~ ____ _+-- __ CI) 3/Rev C.

<.I >.

I

5 150 r----+----4_--

"'ettal ---

;, D n... _-

i 8 __ .... __ C"

1 t - ----.------.1

QI ~ ... -- I

""

100 1-- ___ -+_ 501-- ___ o ~ ____ ~ ______ ~ ____ ~ ______ ~ ____ ~ ____ ~~ ____ ~ ______ ~ ____ ~ o Fan Speed, rpm Figure 46. CF6-50 Engine, Improved Fan Rotor Blade System Modes Campbell Diagram.

. -.

. -.

"...... --.- -

~0r-------~--------r--- ______________ T- __ ~

fiRST lORSIO~AL-"'t'-

I

400 SECOND F,LEXURAL ---+-+-+.....,""-"...,.._~ __ *_ ......

N 300

=

..

>.

u c Q) :s a' 3/aEV ...

flo 2/REV L-_-- I 1/1l£V

o

o 2000 1000 3000 Fan Speed, rpm Figure 47. Fan Campbell Diagr&~ Out-of-Phase Modes, Improved Fan Blade.

SO~--~--~--------~--------~-------- 40 ..... ----_+_=''-

-

~

-

-

~

30)-- ____ .......... _'- -

40 N.

c:: ~ .....

Str:.in .&J .....

20~--------+---------4-- -+---+----- lO~ _______ +- ________ ~~~ ____ __ o~--------~--------~--------~--------~o 1000 2000 3000 4000 5000 Fan Speed, rpm Figure 48. Improved Fan Blade Scope Limits, First Two-Diametral System Mode.

..

60 x 10

-

-

-

~

i!

5 ....,

-

-

N

Ne

.

~SD6

C ....

4 0

-

- "-

~ .c .-4 ..

.., '1'4 .., -;~ 3 0 ...

-

! SOA

,; ~ ....

..J GI C.

t-

-

U 2 0 o U) u Vl -1 0 o 2000 3000 4000 5000 Fan Speed, rpm Figure 49. Improved Fan Blade Scope Limits, First Flexural Mode.

L-.

'- .... ·~~tn.m J

t

.~ ~ ~~

J

j . ~ 'i 80 10 x l !

• . ~ ., '.~ 50 70 l

-

< Q

-

......

~ 40 ......

N 1 .

NIS 50 c:

1 'rI

u ., ........

........

.0 Z ~ ~ ...

...

'rI 'rI IS IS 'rI .~ ...:l ...:l SD6 <II <II Cl.

Cl.

SOA (.)

(.)

tf) tf) 10 SDl o ----~--------~----------~--------~o 1000 2000 3000 4000 5000 Fan Speed, rpm Figurp. 50. Improved Fan Blade Scope Limits, First Torsional Mode.

• __ .=.

60 x 10 40 t---- 50 "'"' < Cl '-" ''-I 30 ......... -----+-----+- .

!:: ..-1 .D

-

~ ~ .~ a .~ ~ Q) p.

10t----=--i-~ :: C)

VJ

I SD€

I

o 1000 2000 3000 4000 5000

I

~ , ~ , Fan Speed, rpm J , Figure 51. Improved Fan Blade Scope Limits, Second Flexural Mode.

~1 Table I. Improved Fan Blade System Infinite Life Stress Limits for Several Vibratory Modes.

2 2) Limits N/cm (lb/in. for Stress Frequency, Speed, Strain Gage rpm SRK SOl S06 Mode Hz SOA

l

First 2-Diametra1 See Figure 48 First 3-Diametra1 126 2520 25,230 12,061 35,164 37,302 (36,600) (17,500) (54,100) (51,000) 3400 9,374 29,780 22,415 Second 6-0iametra1 283

---

(13,600) (43,200) (32,600) Second 6-Diametra1 250 2500 11,031 31,987

---

2J.O~ (16,000) (46,400) (33,400)

•

R2 .. '- ~L ... , lri"~7nrr7· me 6.4.3 DC-IO-30 Inlet - Actual Stresses Baseline Test Stresses observed during this test with the DC-IO-30 inlet were well within infinite life stress limits. The maximum stress levels for each gage location were as follows: Stress Limit Gage( 1) Overall (2) at Gage Fan Speed, Stress (N/cm2) Point, % Location rpm

.Hi!!!

SHK 9,992 19 2500 Gusting Tailwind SHK 6,109 12 2500 Sligh t Tailwind Gusting Tailwind SDl 7,933 18 2500 12 2500 Sligh t Tailwind SDl 5,167 SD6 14,826 29 2500 Gusting Tailwind SD6 8,619 17 2500 Slight Tailwind SDA 7,933 17 2500 Gusting Tailwind SDA 5,167 11 2500 Sligh t Tailwind (1) See Figure 42 (DA) (2) Double Amplitude As shown in the above tabulation, the maximum stress for each gage oc- curred at 2500 rpm fan speed. The p~edominant vibratory modes are first 3- diametral and second 6-diametral system modes. The ambient wind was a slight tailwind 3.7-5.6 km/hr (2-3 kn) gusting to 9.3 km/hr (5 kn). The data reduc- tion clearly shows an increase in stress for the gusting condition. Since the lower stresses are more representative of baseline stresses, both conditions are listed here. For 2500 rpm, the infinite life stress limit at the critical point on the blade is 44% for the gusting tailwind and 25.6% for the lower tailwind (Table II).

90° Crosswind Distortion Testing Thp. engine was opera~eci tv 3900 rpm fan speed with up 0 60 km/hr (32.4 kn) crosswind, and to 3580 rpm with 65.8 km/hr (35.5 kn) crosswind. The test was terminated at this point because of the aforementioned facility failure.

The conditions tested exceed the aircraft operational limits in a (rosswind environemnt. ~ll fan blade stresses were well within infinite life limits (Table II).

Plots representative of the fan face total pressure distortion patterns observed in this test are shown in Figures 52 and 53. Distortion levels were generally low. Separation was encountered at 1180 rpm with 60 km/hr (32.4 kn) .. ~",,' t: , ..

-&tz '-ME [I' , , :,I!

w.

CXJ ~ ;~ Table 1I.

Improved Fan Blade Crosswind Test, Douglas DC-lO-30 Inlet, Measured Stresses J :~ and Life Limits.

I , Fan Sp .. ed (rpoo)

l l~oo (Flight Idle) 2500 2746 33~0 3670 3783

J I

I

>- '" ...

I

!

0 .. .. .. ., '" .... 'I]

.. • • to OJ .. •

VI " ..

.. ... • • .. .. •

>- .. OJ ..

... ..

> .. .. .. or, ..

.. ..

I :; :; ...

... ... " :l ::: • ... ... ...

:l .- ... .. ... ...

.... ." ", C ...

en en til ...

"'" ... ..~ E1 Ii " >-~ Ii "Ii -a S

........ '" S

s .... .... ... .... ... .... '"

... ... .... ... ... ... ...

...

,...,~ ,...,~ ,...,~ J

~ ... .c .. 0 .c ....... .... r. ......... ... ~ ....N ,...,~ .......

'" " ....'" ...~ ... ~ .., ... ","", .. S ... '" '" u_ .. S .. ... ~ E .. .. E ..

... .. S ..

,,~

WI 0 a .. .. .. " .. u .. ~ .. ~ .. ~

.. u .. u .. <; .. ~ .. .., ::11

0 ...... .. '" .. "

'" ... .. " .... ....... ow ....... ... .. .....

ow ow ....... ... OJ'>. ... ....... ...

. "

.... ~

0" ... >z ... >z ...

"'- ... >z ... ... ...

... "" ~z ,...,

u> <", ~z ~z ... gz

Z'" 0 0 ... ... ... ...

...

'"

I

,~ 2755 7.7 14,826 44.0 7,21b 21. 1 9,(.4'1 Baseline 0 28.2 1,580 21.0 6201 18.3 5,510 20.8

- 1

25.6 4,481 13.1 18.5 8,619 S,~10 4,U8 15.0 j 16,895 33.9 4 1000 4138 12.2 49.8 10,002 29.0 18,268 56.6 10,688 28.6 7,923 26.4 ~2.6 5 930 51\;.1 11.4 50.4 11,718 17,238 18,611 57.'1 30.5 9,l05 34.7

I

bO.O 7 1460 5863 21.5

I

bO.O 7 1460 4481 16.4 18,611 54.4 18,611 57.2 11,031 2J.4 ~ 12,061 36.5 ~i I 6~.8 9 1630 7933 29.4 .~ 6~.8 9 1630 :;510 17.6 16,199 47.0 5J.7 17,~82 13,444 41.2 .

NOTES: I) Attachment fan speed is the speed at which the flow attaches to the inlet lip.

~,'

Flow is s~parnted from the inlet lip below this speed.

!'

2' Highest measured ove'~~l dynamic _tress at given speed - double amplitude values COA).

r

:r J) Life lillit6 perce". of allOWable dynamic stres9 for infinite life at critical point.

I~ t' 4) Peak levels include effect of 9. J ltDi/hr rdlwind. Lower levels ignore peaks 'Iue to taJlwind and are more rerresentative of baseline stresses.

~

~.

~ , L ," • .JilIIl

Seven Facility Fans: Nl = 3887 rpm

Crosswind Velocity • 32.4 knots Top P - PAvg x 100% P crosswind~ Avg Direction

-2

I

!

, ,

I

IDC = M ax.

0.016 ID~ -~ax.

p - P Looking Upstream - ALF Max. rHn.

= 2.8% P J Avg.

i

'i CF6-50 Engine With Improved Fan and DC-lO-30 Inlet: Fan Face

l

Figure 52.

Total Pressure Distortion Patterns for 90° Crosswind at 32.4 l Knots, 388i RPt-1 (Attached Inlet Flow).

f :, t~ ') in mM';;iWW'"tr %t Nine Facility Fans: Nl = 1556 rpm 'Jelocity - 35.5 knots Top x 100% Crosswin~ ...

Direction IOC = M ax.

IDR

= 0.008

-}fax.

P - p Max. Min. = 4.9% Looking Upstream - ALF P Avg.

Figure 53. CFn-50 Engine With Improved Fan and DC-la-3D Inlet: Fan Face Total Pressure Distortion Patterns fur 90° Crosswind at 35.5 Knots, 1556 R~M (Separateu Inlet Flow).

crosswind and at 1400 rpm with 65.8 km/hr C35.5 kn) crosswind. Flow reattached at 1460 rpm and 1630 rpm, respectively. No separation wa. encountered at high- er fan .peed.. The maximum stre •• levels for each gage location were a.

follows: Streaa Limit O ) OVerall (2) Gage at Gage Fan Speed, Crosswind, Location Streas CN/cm ) Point, % rpm km/hr 42.6 SHK 14,483 28 2500 60.0 SDl 13,100 30 2500 3380 42.6 SD6 18,611 39 36 2500 60.0 18,611 39 3380 60.0 18,611 SDA 16,542 35 2500 60.0 (1) See Figure 42 (2) Double Amplitude (DA) The maximum stress observed was 18,611 N/cm DA on gage SD6 at 3350 rpm with 42.6 km/hr (23 kn) crosswind. The predominant mode was second 5-diam- etra1 system mode. At this condition, the infinite life stress limit at the critical point was 57.2%. ·.~his was the highest limit seen in the test.

6.4.4 B747 Inlet - Actual Stresses Baseline Test Stresses observed during this test with the B747 inlet were well within infinite life stress limits. The maximum stress levels measured at each gage location were as follows: Stress Limit C Overall (2) Gage !) at Gage Fan Speed, 2) Locat ion Stress (N/cm Poir;t, % rpm SUK 6207 12 2500 SDl 7580 17 SD6 21 3350 SD6 8619 17 2500 SDA 7236 15 2500 (1) See Figure 42 (2) Double Amplitude (DA)

- .. ---.--.-~--- ~--- - ... _.j

The predominant vibratory modes at 2500 rpm are first 3-diametral, and second 6-diametral system modea, and the infinite life stress limit at the critical point on the blade is 24.8%. The vibration mode at 3350 rpm is al- most entirely second 5-diametral system mode. and the infinite life stress limit at the critical point is 33.5%. The data reduction for gage SD6 shows the effect of gusting wind at the 9992 N/cm DA reading. A stress of 7236 N/cm DA is more representative of the baseline at 3350 rpm, making the infi- nite life limit 24%.

90· Crosswind Distortion Testing The engine was operated up to 3350 rpm with 29.6 km/hr (16 kn) crosswind, up to 3400 rpm with 19 knots crosswind, and up to 2800 rpm with up to 79.7 km/ hr (43 kn) crosswind. This test sequence is consistent with the operation of the aircra:~ :~ service. Rolling takeoffs and aircraft velocity during ma- neuvers prevent hi6h crosswind distortion without sufficient axial flow to clean up the inlet.

All fan blade stresses were well within infinite life limits as shown in Table Ill. Maximum stresses measured on this test are slightly less than those measured during similar testing on the original fan bla~e.

Plots representative of the fan face total pressure distortion patterns observed in this test are presented in Figures 54 and 55.

The maximum stress levels for each gage location were as follows: Stress Limit Gage(l) Overall (2) at Gage lo'an Speed, Crosswind, Location Stress (N/cm2) Point, % rpm km/hr SHK 7,923 15 2500 59.3 and 79.'1 SDI 9,649 22 2500 59.3 SD6 13,787 2500 79.7 SDA 11,718 25 2500 79.7 (1) See Figure 42 (2) Double Amplitude (DA) The maximum stress first obuerved in the test was 13,787 N/cm DA on gage SD6 at 2500 rpm with 79.7 km/hr (43 kn) crosswind. The predominant vibratory modes are first 3-diametral and second 6-diametral system modes. At this con- dition, the percent of infinite life stress limits at the critical point was 40.2%. This was the highest life limit s~en in the test.

f' .. '.i .....

.~ Improved Fan Blade Crosswind Test, Boeing 747 Inlet, Measured Stresses and Life Limits.

Table III.

hn Speed (rJIII) l~ (Flight Idle) ;:; ., ., ., c~ ..

., ., • • ..... .

;::; • • • .. .. .. .. ..

>. "- ., • ., .. .. .. .. .. .. ..

..

., • • ..

.. ~ ~ .. .. .. .. .. .. • ..

ge, .., <II .... OIl v, ... ... <II <II "' ...

Ii i1 c >. iI f! if '" .. ~~ .... ... .... .. .... • .. ... ... .... ... ... ...

- .. <J

.. -

... ,... ....... ....... ........

:. ..... .. .. ...t .~ -oN ...l~ ...l~ ... ~ ...t~ ...t~

-.~

., . .. .. .. .. .. .. B

.'.,J.e ... w .. Ii .. E .. .. II ..

.. ~ .. u .. ~ .. u .. ~ .. u . ..., .~ .. u

.,0' .. .... .. u

"'" .-

... .. " ...

... ... ....... ... ...... ... ......

... Of II'

.. '

.. - ,.z ..

e~.s o :i ..... :.z ... .... .. ... ..

>z ~z ,.'" .!;'" ...t

u:>- til III 0 .. 0 ...t .... 0 ...t 0 ...t

"''"- 8,619 S,16} 6201 18.0 l..seline 0 1725 4.8 24.8 9992 11.S 17.~ - !.

29.6 4 3200 2069 6.0 6,S!l0 }9.0 8619 211.8 8,961 28.6 I:.

li' n.2 5 2800 21SS 1,Sao 22.0 1l.2 9105 11.2 12,}.7 40.2 i·' 7.9 4,481

r

)451 10,688 40.7 7 Separated Up To 10.0 3LO 9,649 28.5 28JO rp1ll (HaJ<) Speed This Point ..

~9. ) 4118 11,374 10,345 10 12.1 n.l 10.6 ..

79.7 13,787 13 4824 13,9 40.2 11,374 J4.n NorES: 1) Sepa: aCion fan "peed 1& the hlah speed poInt at which the £low separates frOB the inlet lip.

2) Higheat 8eaRureJ overall dynaaic .tres. at given speed - double a~11tude velues (DA) J) Percent of allowable dynaaic stress for In(lnIt~ life at crItical point.

oc ID Top Steady-state Pattern - ALF IDC c 0.022 Ma x.

p - P Max. Min. a 6.6% P Avg.

Figure 54. CF6-50 Engine With Improved Fan and 747 Inlet: Fan Face Total Pressure Distortion Patterns for 90° Crusswind at 22 Knots, Nl - 3899 RPM (Attached Inlet Flow).

4. =-. £D, a ! l !'

I' .

i . , Top IDC - 0.107 Ma x.

PMax. - P • Min - 15.9% PAvg.

I

Steady-state Pattern, ALF Figure 55. CF6-50 Zngine With Impr-oved Fan and 747 l"let: Fan Face Total o Pressure Distortion Patterns for 9U Crosswind ~t 43 Knots, N1 • 2814 RPM (Sepal'ated Inlet Flow).

"¥ seE4$ 4 Q.~~·WT ..... ""*;s::;::;_ A ~.4L2Z$ZM - Engine crollwind talting demon.trated that the impt'oved fan blade hal .imilar cro •• wind/di.tortion characteri.tic. a. the original CF6 fan blade which ha. Q very .ati.factory record for operation under cro •• wind condition.

or high inlet di.tortion without any problem.. Re.ult. indicate that the new fan blade can operate lucce •• fully without exceeding vibratory .tre •• limit.

with both the DC-10-30 and 8747 inlet. at allowable takeoff cro •• wind. up to 64.8 km/hr (35 kn).

l-

I

i

j

t ~hZ" ...... .m .............. c ................. ~.;.·.; .. ne.;.-~7i2Fr •..•. i6.~~ .. ~: ••.• trn.-.t.77 .. 7- .... ___ ~# ____________ ~_. ______ • __ 7.0 ENGINE PERFORMANCE TEST The objective of the back-to-back engine performance te.ting wa. to mea- .ure the performance improvement that re.ult. fr~ rp.placing the original pro- duction CF6-50 fan with the improved fan .tage. Performance improvement was mealured at lea level and durin, limulated altitude operation. Modification to the improved fan package wa. required to achieve the predicted .fc perfor- mance improvement. Te.ting included evaluation of improved fan blad.I, a fan ca ••• tiffener for improved roundne •• , reduced fan tip clearance, fan caeing tip Ihroud configuration, and fan noz&le area variation.

In addition to the blck-to-back engine performance telt re.ult. dilcus.ed in thi ••• ction, production engine and aircraft flight telt re.ulte with the final improved far. pac kage are pruent ed in S(!C t ion 11.0.

7.1 ENGINE TEST FAC1LITY

7.1.1 Tut ee 11

Sack-to-back engine testing was conducted in indoor test cells which con- tain engine mounting facilities. overhead air inlet with turning vanes. and an exhaust stack. The engine is normally mounted in a test cowling Which forms the fan exhaust nozzle and ~.tes with Lhe primary exhaust nozzle. Figure 56 shows the engine test configurat ion for both sea l~vel and II ~mtJlated alt itude performance tests. Altitude cruise was simulated by testing with fan and core exhauet nozzle diffuse~s to create choked exhaust nozzle operating conditions on a se. l~vel test stand.

7.1.2 Special Test Equipment Due to the requirements of this test to measure performance at sea level and simulated altitude conditions. special test hardware was provided. Spe- cial test equipment includ~d an inlet bellmouth. a modified CF6-50 fan rever- ser, a fRn nozzle diffuser, and a primary n~zzle and diffuser extension for simulated altitude operation.

Bellmouth - Initial testing was conductecl with a bellmouth without an inlet screen. ntis bellmouth d Hfers from the convent ional development bell- mouth in that it mounts to thp. enghe front !tange. instead of being facil ity mounted. in order to simulate the weight of the flight inlet. Initial test- ing showed unacceptabl~ inlet temperature measurements with th~ bellmouth- mounted thermocouple rake.. Consequently. a convp.ntional development test bellmouth with inlet screen-~unted air thp-rmocouples was utilized for inlet temperature measurement.. Inasmuch as thiu bellmouth was facility-mour.ted.

inlet weiRht simulation was provided by attachment of a dummy weight to the fan casing.

• " ..... ,,'e, '",~""""" ... - ~."~'''''.1"·:' ...

:: ,,!

I

l co

J

~ :1 "", iffening Ring Engine ~ounted Bellmouth Fan Nozzle Diffuser

l

Case Stiffener

Iii

••

I:i

" !

.J -':""'::~ Primary Nozzle

II

Diffuser I'J I ....

J ! ' ALTITUDE CONFIGURATION

I'

i

SEA LEVEL CONFIGURATION

I

I ~ !

I

j

':1 ()

l .1 \

I

CF6-50 Fan Reverser

~

t~ ....

i

: '"

i§

i ."':";.- ,Ii

I

Figure 56. CF6-50 Eng5.ne Test Configurations for Sea Level and Simulated , : ~,...

I"' Altitude Performance Test.

'8~

I ,~I

lSi ") ,...", ~;;;

I

- - - ~:-: :-~.-::-" "~~::-. :-:-:-',-:- ~';;;1:':":"~:t:::--~"'" , .""",-,>.I ... ,""I,"".a .. .u.1'"""""""udb..i"~~~1I1, i&I~~d kl.

Fan Reverser - A CF6-50 fan reverser with standard fan nozzle was used for the back-to-back testing. The reverser had been modified for an ejector seal and by the addition of a stiffening ring near the aft end of the trans- lating cowl to take the extra load imposed by the fan exhaust diffuser for simulated altitude operation. The fan nozzle area was varied by trimming the fan reverser trailing edge.

Fan Nozzle Diffuser - The fan nozzle diffuser is a 48-inch cylindrical extension to the fan nozzle (Figure 56). The diffuser causes the fan noz- zle exit static pressure to be reduced below choking pressure provi~ing a simulation of the cruise operating line of the fan. The diffuser is hinged to the pylon to permit opening for engine access. The diffuser is sealed to the reverser by a flexible seal between the forward end of the diffuser and the stiffening ring on the reverser. Figure 57 shows the diffuser, fan rever- ser, and core cowl door attached to the test pylon.

Primary Nozzle and Diffuser - For these tests, a special primary nozzle has been construct~d with the inner flowpath contour identical to a CF6-50 turbine reverser and a mounting flange for a diffuser extension (Figure 56).

Installing the primary nozzle diffuser extension causes the primary nozzle to choke at a lower pressure ratio to simulate altitude operation. Figure 58 shows the primary nozzle with the diffuser extension attached. This special primary nozzle was repla .. ed with a convent ional nozzle without the diffuser mounting flange for sea level testing. Early testing has shown that the flange had a marked effect on primary nozzle flow coefficient without the diffuser installed.

7.1.3 Data Acquisition System The system basically co~sists of a cell system and a site system. TIle cell system performs steady-state and transient data acquisition, conversion to engineering units, quick-look performance calculations, and short term storage. Converted data are automatically transmitted to site system for fur- ther on-line processing, graphic display, and hard-copy output. The site sys- tem utilizes a data base concept for efficient storage, retrieval, and repro- cessing of current and historical data. In addition, data may be transmitted to the General Electric Evendale time-sharing computer center for further pro- cessing such as cycle deck analysis and comparison.

Data acquisition capability consists of 400 pressure channels, 400 tem- perature channels, 10 frequency channels, and 128 dc voltages such as: load cells, individual pressure transducers, position pot~ntiometers, etc. The pressure system consists of ten 40-port scan valuQs with available pressure ranges from % 0.7 kg/cm2 gage (% 10 psig) through ± 35.15 kg/cm gage (t 500 psig). The system incorporates autoranging and multiple sampling capability for all data channels to assure optimal resolution and precision in addition to variable averaging time for frequency measurements. Data may be achieved and processed in either a steady-state or transient mode. Typi- cal steady-state acquisition time is 30 seconds with each parameter sampled 49 times over the 30-second time period. Transient acquisition rates are I .

,

I ! . 1 '1 I,, ' , r' :, 1' 1 1

Figur' 5 7 . ran R'v r s' r with Fan ozz l Dif f ll r ln sta ll d .

~ Pl ' I' i , I !

!' ri m .!r \' Fi g llr ' 5H.

7.7. 1 , ; lll d ill 1I l l s l ' r Ill l- S i llll l l .!LI'! \It i l lldL ' variable from one sample per second per channel to 250 samples per second per channel. Redundant measurements are made of key parameters such as fuel flow, fan speed, and thrust. Automatic data reje~tion techniques, ratio of redundant measurements, and on-line .ystem verification analysis further enhance overall data quality.

7.2 ENGINE TEST CONFIGURATIONS Initial testing was performed using a CF6-50 development engine. The in- tent, in belecting hardware for this engine, wal to build an engine with per- formance equivalent to a new engine that had been subject to initial perfor- mance deterioration.

Significant features of the initial engine test configuration included: fan booster, high pres.ure turbine and low pressure turbine rotors were assem- bled primarily from parts used on a previous build; the high pressure compres- sor rotor had new blades which had been ground to provide clearances con.i8- tent with the engine new front mount configuration; the high pressure compres- sor stator had a refurbished forward casing and the aft casing used on the previous build; vanes were used parts; the high pressure turbine tip shrouds were new production configuration and were ground to obtain clearances slightly opt~n from nominal; and the low pressur~ turbine stator had been used on the previous build except that the tip shrouds and interstage seals were new.

For the baseline performance tests, the CF6-50 engine was run with a set of original production fan hlades. The fan casing tip shroud was open-cell aluminum honeycomb and was ground to the original production minimum tip clearance of 4.45 mm (0.175 in.).

Following the baseline tests, the original production fan blades were replaced with a full set of the improved fan blades. A new forward fan case and the fan case stiffener ring were installed. New fan casing tip shrouds of open-cell aluminum honeycomb were installed and initially ground to a mini- mum tip clearance of 2.92 mm (0.115 in.), a reduction of 1.52 mm (0.060 in.).

Refer to Figure 1 for the improved fan engine configuration and to Section 3.0 for a detailed description of the improved fan.

Improved fan blade sets were fabricated in-house by General Electric and by an outside vendor. In order to fin~ tune the fan engine match to obtain the predict~~ sfc improvement, the blade part-span shroud interlocks were subsequentlY1lodified (restaggered) to close the running blade stagger angle by about 1.5· at the part-span shroud. The fan casing tip shroud was also modified by installing microballoon epoxy in open cell aluminum honeycomb to provide a smooth casing tip shroud. The tip shrouds were then ground to pro- vide tighter fan tip clearances down to about 2.16 mm (0.085 in.), a reduc- tion of 2.29 mm (0.090 in.).

A second new CF6-50 engine was utilized in order to complete the improved fan performance test program in an expe~itious manner. A third (production) CF6-S0 E2 engin~ was also used to test the restaggered improved fan blades back-to-back against production-type improved fan blades.

~- !f4 =rt< ___ .' , .. ~'-~·'·~·"l 7.3 INSTRUMENTATION AND DATA REDUCTION t1 ~~ Enline pe~formanee inlt~umentation ineluded p,e •• ure and temperature mea- ~ .u~ement. at the inlet and exit of the fan, boo.te~ and hilh pre •• u~. comp~el .o~. and at the exit of the hilh pre •• ure and low pre •• u~e tu~bine.. Inlet bellmouth in.t~umentation wal p~ovided to determine fan inlet ai~flow. In addition, both fan and eo~e ~oto~ .peeds. fuel flow, and engine th~u.t we~e measu~ed.

Telt in.t~umentation uled to measure fan performance and to monitor en- line operation i. b~oken down into two l~ouPI: lene~al inlt~umentation and aerodynamic instrumentation (Figure 59).

7.3.1 General Instrumentation • Baromet~ic Pressure - The local barometric pressure measured uling a recordinl microbarograph.

• Humidity - The absolute humidity mea.ured in grains of moisture per pound of dry air using a humidity indicator.

• Cell Static Pressure - Test cell static pressure measured at four locations in the cell.

• Fan Speed - Low pressure rotor speed measured using two fan case- mounted, fan speed sensors.

• Core Speed - High pressure rotor speed measured using engine core speed sensor driven off the end of the lube and scavenge pump.

• Main Fuel Flow - Volumetric flowmeter, facility-mounted.

• Verification Fuel Flow - Second fuel flowmeter mounted in series with the main fuel flowmeter.

• Fuel Temperature - Temperature of fuel measured at the facility flowmeter using a single chromel/alumel probe in the fuel line.

• Fuel Sample Specific Gravity - Specific gravity of the fuel sample measured using a hydrometer.

• Fuel Sample Temperature - Fuel sample temperature measured during the specific gravity measurement.

• Fuel Lower Heating Value - Lower heating value of the fuel sample as determined by a bomb calorimeter.

• Thrust - Thrust frame, axial force measurement using three strain gage-type load cells for redundant measurements.

1IiiiIiio_ ..... ________ .....oiiiIi.-. ..... _.o....iiio-.";,,,.;, ... __ .- iilii£-~iiiil-iIli'·ii- iiI- .\IIi1L ___ ~-.-IIiI-IIi"'IiI~.--I--.· .- .. -~.-II! ..... _--.- ... _ .... -~ ______________ .~ o o Plane 49 - LPT Inlet 11 TT Probes - Averaged Plane 13 - Fan Discharge 5 P Probes T 4 Arc Rakes - 84 P , 84 TT T 2 Wall Static Pressure Plane 50 - LPT Discharge 4 Radial P Rakes - 5 Elements Each T 2 Radial TT Rakes - 5 Elements Each '--- Plane 1 - Fan

£,K" , '&1' '.' ' iI' *' - . i::i2S7'~ ~

4 Bellrnouth Rakes 24 PT' 24 PSt 8 TT

~)

J

.,. .

[

.-

1 L, I ' ~'~ " ,.

Plane 3 - Compressor Discharge

---

4 Wall Static Pressures IS-Element TT Probe

..... - -

Core Inlet General Radial Rakes - 25 P , 25 TT T Thrust I 5 Wall Static Pressures Fan Speed 2 Core Speed I Fuel Flow 2 Fuel Temp. I Throttle Angle 1 VSV Position I VBV Position !

Humidity I Barometer I Figure 59. CF6-50 Engine: Instrumentation for Fan Performance Test.

.,,~ :, ," ,,!

",,~ ....

'.~; o o :j , Plane 49 - LPT Inlet 11 TT Probes - Averaged Plane 13 - Fan Discharge ·~~i 5 P Probes Plane 23 - Booster Discharge T 4 Arc Rakes - 84 PT' 84 TT----~ 4 Arc Rakes - 24 PT' 24 TT 2 Wall Static Pressure 1 Wall Static Pressure Plane 50 - LPT Discharge -"t:::..~==- 4 Radial P Rakes - 5 Elements Each T Radial TT Rakes - 5 Elements Each

'---

Plane 1 - Fan 4 Bellmouth Rakes 24 PT' 24 PSt 8 TT

-'~ .~

~) • I •• ;* ,:

[

.-

. - -'~ ~; ~C'- , ; 3 - Compressor Discharge It:iII' Wall Static Pressures ~ ~~ .~ ..... _ - - I 5-Element Tr Probe 1 ,Ii:

rr Plane 25 - Core Inlet General

5 Radial Rakes - 25 Pr' 25 Tr Thrust 5 Wall Static Pressures Fan Speed Core Speed Fuel Flow 2 Fuel Temp.

I Throttle Angle 1 VSV Position 1 VBV Position Humidity 1 Barometer I Figure 59. CF6-50 Engine: Instrumentation for Fan Performance Test.

!

I

J

"'

J Variable Stator Vane Position - Read~ut of the linear variable dif-

•

ferential transfo~.r attached to the hiah pressure comprellor vari- able stator pump handle.

Variable Bleed Valve POlition - Readout of the linear variable dif-

•

ferential transformer attached to the variable bleed valve actuation aaechani ...

7.3.2 Aerodynamic Instrumentation The followina rakes, probes, and static pressure tapi were installed to measure airflow, temperature, and pressure as required to define component performance: Fan Inlet <Station 1) Bellmouth rakes were installed to measure static pressure, total pressure, and total temperature at the fan inlet. Four rakes each having six total pressure probes, six static pressure probes, and two total temperature probe.

were used.

Fan Discharge <Station 13) }o·o~r arc rakes with 21 probes/rake were installed in the aft fan case to meawure total pressure and total temperature. Seven different radial immer- sions with three elements per immersion were sampled for a total of 84 temper- ature and 84 pressure probes. Two pressure taps were installed in the aft fan case to measure fan discharge static pressure.

Booster Discharge (Station 23) Four arc rake& each having six temperature and six pre~sure probes were installed to measure booster discharge total temperature and t~tal pressure.

A single tap was installed to measure booster discha~ge static pre.~ure.

Compressor Inlet (Station 25) Core inlet pressure and temperature were measured uling five rakes and five elements per rake. Five flowpath wall static pressure taps were alsu installed.

1('1 1IIIIIII __________ • ____ ... __ ... iiiIiI-: .. ~lil~iI.-'-Ii-II .. I'L.-I·I-111-- -1""1- .-~a.tlrl·~ -.- 1~;'I111'.- .... ____ " .-- -_ .. __ Compreaaor Diach_rle (Station 3) Four of the boreacope port pluga in the compreaeor rear frame were modi- fied to permit compreaaor diacharge .tatic pre •• ure meaaurement. A .lngle 5-element thermocouple probe wa. u.ed to mea.ure compre •• or diach.rge temper- ature.

Low Pre •• ure Turbine Inlet (Station 49) Temperature in thi. plane waa meaaured by a production configuration KGT harne •• con.iating of 11 dual element thermocouple probea electrically aver- aged. Pre.aure waa meaAured uaing five probea each having five elementa all feeding a aingle fitting.

Low Preasur~ Turbine Diacharge (Station 5) Low pre.aure turbine diacharge pre.sure was mea.ured using four rakes having five elements each. Temperature was meaaured by two rakea having five element. each.

7.3.3 Data Reduction Preliminary performance calculat ions were done on-line by the "quick-look" data reduction program. This program ia part of the cell data acquisition and proceuing system descr:bed in an earlier sect ion. Test measurements, cal i- bration curves, configuration con.tant., and fixed data input. are combined in the.e calculations.

Most parameters of interest are calculated in the quick-look program.

However, detailed cycle analysis was augmented by design point .tuuies using the status cycle deck. This teChnique has the capability of providing a balanced cycle e~aluation of any test point. The user has the option of selecting between redundant mea.urements for input to the cycle match while letting the cycle calculation provide parameter estimates where measurementa are not available.

7.4 TEST PROCEDURE AND HISTORY Performance improvements associated with the improved fan package were determined by back-to-back performance calibrltions that were conducted by testing a CF6-50 engine with the original production CF6 fan replacing the fan with the improved fan anJ repeat ing the test. Performance cal ibrat ions consisted of two automatic data recordings at each of 12 power settings from the maximum takeoff rating to ground idle. This performance calibration was then repeated in each case. Additional data were occasionally required if any uncertainty in performance levels existed after completion of the two power calibrations. For simulated altitude operation, the fan and primary nozzle diffusers were installed.

"led on the previoul t.lt relultl. the ori,inal fan blade part-lpan Ihroud val reworked to ~prov. blade reliltance to foreiln obje~t d .... e (POD).

Initial back-to-back enaine perfor.ance t.ltl eitabilihed that the part-lpan rework would be co.aon to both the ori,inal ancl wproved fan bledel. Mdt- tlonal '.lIine te.t. vere conducted to check out the operational characteri.- tic. of the fan and pr~ry nOI&le diffueer •• A liet of the fan perfor.ance te.te on the firet enaine ie pre.ented in Table IV. The fli,ht condition and the enaine confiluration are indicated.

A .econd nev CP6-S0 enaine va. introduced into the u-proved fan parlor- .. nce te.t prolr .. while repair. vere beinl .. de to the lirlt inltallatiDn.

A back-to-back telt of the orilinal to the Oeneral Ilectric fabricated ~ proved fan bladee va. run to e.ubJ :.eh a ba .. line on the .econd en,ine. A lilt of to.tl coapleted on thi. @naine il .hovn in Table V. Telt procedure •• inltrument~tion. and data reduction vere liailar to thOle vith the orilina1 en,ine, and thil te.tina val completed in a limi1ar telt cell. The enaine wa. te.ted in I 11ave cov1inl with a ,lave primary nOllle. Mo.t .pecial in.trumentation va. reaoved from the ori,ina1 enline and inlta11ed in the .econd engine. except for the boo.ter dilchar,e rake •• Additional telting wa. done to evaluate the perfo~ance of improved per- for.-nce fan b1adel which had been vendor-fabricated. Teltin, va. allo ac~o.

plilhed define a further red~tion in tip clearance below the 1.5 .. (0.060 in.) reduction. already i~corporated with the fan cale .tiffener. A further reduction in clearance of 1.0 am (0.040 in.) with a .aooth-.urfaced .hroud material in place of the open-celled honeycomb wa. te.ted.

7.5 DISCUSSION OF RESULTS The predicted perfo~ance improvement of the improved fan packa,e, in- cludin, the 1.5 mm (0.060 in.) reduction in fan tip clearance, i. Ihovn in Table VI. Predicted cruile Ifc improvement. and equivalent value. for ,round te.ting are Ihown in this table. Sea level performance comp~rilon. in Table VI and lublequent curvel are made at a lea level .tatic thrult of 17.800 daN (40,000 Ib) which correlpondl to a typical altitude cruhe power leuina. Aa shown. the improved fan provide. a silnificant improvement in part-.peed effi- ciency at cruise power lettinll. This providel a lilnificant fuel lavinal, particularly at the intermediate power .ettinal typical of altitude cruile.

The early portion of the te.t leriel evaluated the improved fan blade with 1.50 am (0.060 in.) tip clearance reduction compared to the original fan confiauration. Relultl of Teltl 1 to 4 are Ihown on Figurel 60 through 67.

Initial telting with the original fan b1Adei in the lea level and the limu- lated altitude cruile confi,urationl with exhault diffuler. Ihowed expected engine operating characteri.tic.. The diffulerl provided choked exhault noa- a1el for an effective limu1ation of altitude crui.e engine operating linel.

Actual .fc performance i. plotted on an ab.olute .fc Icale, and the curvet .how th. incremental improvement in ablolute Ifc. The %4lfc improvement il 1,11.

:~

Table IV. CF6-50 Engine I - Fan Perfonaance Tests.

" .....

, :,.

.~ Fan Tip Fan Case , '."

Clearance Moz&1e easing Tip Flight Fan Blades! St iffener Test ~ ;r _ (in.)

ec-ent.

Area, Al8 ,,' Ring Shroud No. Condition Fabricator ~ 4.45(0.175) Mo..

Open Cell Orig./CE 110 I SLS '~ Al u. in ..

Prod.

~J Honeyca.b i ~ Noll. + Fan 110 l Sia. Orig./CE ,,~ Diffuser Prod.

Alt itud~

I

" ~ ,,'I 'ij Mo.. + Fan 2.92(0.lU) "' laproved!GE Yes Si •• ;i Diffuser Altitude " :~i ,~; ....

SLS Iaproved/CE II ~ Data H.G., '~ 4.45(0.175) Orig./CE VSY Opened 2· RiCh Flow 2.92(0.115) Iaproved!CE ';1 Ii Booster

j

~

,!,'

.... Std. Booster

:'1' : 1:11 1Ioa. + 1%

8 SLS 1

I

IJ .... + 1% ~ Iaproved/CE 9 si •• i1 + Fan Alt itude ~~l Diffuser

~

,'~i .... + 1% Oata II.C., Iaprovedl 10 s1..

"l~ Fa il ed Core + Fan Vendor Altitude Diffuser Cowl Doors .... + 1% I I.proved!

11 SLS Vendor SLS Restaggered Iap./Vendor 2.92(0.115) 1Ioa. + 1% Open Cell Orig./CE Yes 13 SLS A1 .. i01._ Prod.

Hone yc a.b

:~

:1 , J " I Table v. CF6-50 Engine 2 - Fan Perfor.ance Tests.

Fan Case Tip Fan Test Flight Fan Bladesl St iffener easing Tip Clearance Nozzle _ (in.)

Cond it ion Fabricator Ring Shroud No.

Area. A18 blproved/GE Open Cell 2 • 92 ( 0 • 115 ) 14 SLS Yes ....

AlUMin ...

Honeyco.b 15 Orig./CE Open Cell 4.45(0.175) Alwlinu.

HoneycOIIIb

i~

16 lap ./Vendor Open Cell 4.45(0.175) !:,,~,

I'

~,", Al ... inUM !i ",'" Honeyco.b I,' 11 R.!staggered Oval Grind 2.92(0.115) ....

I ~, :~: 2.16(0.085) IJIp./Vendor Open Cell Al ... inUM Min.

" HoneycOMb 18 Restaggered Oval Grind 2.92(0.115) .... + 1% I.p./Vendor Open Cell 2.16(0.085) AlUMinUM Min.

!

Honeycoab 19 SLS Re s t a,ge red Yes SMooth 2.29(0.090) .... + 1% lap./Vendor Micro-

balloon 1

, Epoxy in I Open Cell Al ... inUM HoneycOMb

- o

(II ,,<>, , "

J

Equivalent % Predicted , .fc at SLS.

Iaprov ... nt 17 .. 800 da ..

It .. % A Cruhe .fc ( .. O~ Ib)FN I.proved Fan Ilade -1.0 -1.9 Reduced Fan Cl.arar.ce -0.6 -0.8 1.5 am (0.060 in.)

1% Increased Fan Noul. -0.2 0 Area (AU)

-

-

TOTAL -1.8 -2.7 _______ ;,.. .. ,;. ... '.·~_z=iI· II" 'iI'Izll' ~iiII~ ...... ' .Sll ... r_· __ --~------- .- --- .......

Corrected Net Thrust, 'N' Ib 35 40 45 50 Tip Clearance _ (tn.)

Symbol Tip Shroud Fan Blade Open Cell Honeycomb Original 4.45 (0.175) Open Cell Honeycomb 4 Improved (OE) 2.92 (0.115) ,...

.c ...

....

~ ,c: ....

.c ...

.....

.-

0.01 "0 == .....

...L

,c:

'"

.....

co ~ .

CJ .....

III • 1.8% b sfe at Cruise Equivalent Power 230 2 170 180 190 Corrected Net Thrust, F , kN N Figure 60. CF6-50 Engine: SFC Performance Versus Net Thrust for Original and Improved Fans, Sea Level Static, Initial Testing.

- -"WE- Net Thrust, F , lb N 7 8 9 Simulated Sruiae 0.85 Ho ; ;; 35.000 ft

T

, Standard Day 0.005

--L

,...

,.CI

o

...

......

i

...

.c ......

.n ...

'-' ~ ......

...

.c ......

:f Test Tip Clearance u .... Symbol Fan Blade l1l1I (1n.) Tip Shroud. [tiJ ~ co 0 1 Original 4.45 (0.175) Open Cell ijoneycomb 4 Improved 2.92 (0.115) Open Cell Honeycomb (GE) 30 35 40 45 50 Net Thrust, F • kN N Figure 61. CF6-50 Engine Simulated Cruise SFC Performance Versus Net Thrust for Original and Improved Fans.

"_,.,,,,::i "", .".~,,~~'.~:~:--r-,,:~~~.~~~.;~~:,:,~,~::;;: .iii., -W,.·- £4'- H'¥,I't'!"*,,~-t~.~.~~~:,,'--=~·r~~,.~'~-1'~~·"· ."_ .. ~~'ii-""~,~=--'f ...... ~,"'''''~ _~o_''''jj( .. - ___ 5':~~_~~~ ~

t

t 1300 1500 1200 1400

ImprQV~-d--i--'-- I

~.

Fltn Pretest -··-·[]·n;--·-------·-~--"8Jf:::::_Ii)......ii-t-----1I------4 ~ Estimate ...

....

"':::::-"',:--t ...

',I ~ 4.2 Points A , u W <II Co

FA' "fldo.oy f' 1 0% 11 Ilt Cruise-Equi "

.;,

I

U

't~F-= lent Power ,,-

4.7 Points '.

~ ....

'"

u ....

...

...

.---- ------4-- .--~--!....I ~C0~t -~,.!"Ir~~(i°G-<0~fJ '~-~

1>.1 e OrlllolnAl 0" I l'II ~ "~dn rr~t(,Mt .----------------- -------- ........

Ellttr."'tl'! - Test Tip Clearance ~. ---1-

~ ~o. r.n..n~ladp' _ nun (in.) Tip Shroud ~.-.--- .- 0 O\"lgtllAl 4.4~ (O.llS) Open Cell HoneycOlib [] 4 Improved 2.92 (0.115) Open Cell Honeycomb

t

(GE) 580 600 620 Figure 62. CF6-50 Engine Fan Efficiency Improvement Versus Fan Airflow for Original nnd Improved Fans. Sen l.evel Static, Initial Testing.

Corrected Fan Airflow 1200 1300 ISoo 1.80 1. TS .

Test Sea Level Static No.

1.70 ~N .......

1 Original Blade (If) ~ ...

Blade

0 4 .blproved

..

.... Slllulated Blade +I l.6ti

,: Original Blade

~ 2 3 Improved Blade

0 5

•

...

lJ ::s ~~ .,

•

l~ II

•

' ~

t. .'

''': 1.60 II s:I "1 III ! ~ !lot

q

1.55 1.SO 560 580 600 620 640 660 680 700 Corrected Fan Airflow, kg/sec Figure 63. CF6-S0 Engine Fan Operating Lines for SLS and Simulated Altitude Operation With the Original and Improved Fans.

i -~ .-.;:, _ ..

. _~."''''<;* %; -; ~,.;O.-,:-"",,",- -~r:t"'E'~'-"·~ .. -:_~~ = -.,.". =- - - -~ -~ - .~---~.-- "'!' ..... ~~ ____ s ...... ""'~~h.~:~,~~~~~~~=::~.:~~'~:.,~:.~~2~: :=.:=-~:~.':::""~~ '44,--- ,- .--~?--. ~~=_~~~~~.=t=="'ee~~==_'=!'__r~='c= _",-,~==~,_ !

~------~------~--------~------~--------~--------54 x 10

Te.t~---------+---------+-------- ..l!2.a.

o 1 Original Blade

o 4 Improved B1 ade

.c f"'4 ..

~ ..

ra:.

=

46 .; : I) ..

~ e

I)

is

e

~

is ~

lI; ~ ~ Ii!: L- ______ ~ ________ ~ ________ ~ ________ ~ ______ ~ ________ ~38 3300 3400 3500 3600 3700 3800 3900 Corrected Fan Speed. N , rpm 1K Cf6-50 Engine SLS Thrust-Speed Characteristics With Figure 64.

I Original and Improved Fans.

I

~~~~~ __ a""t ... ~ 11"1*11" ~"~II~~ ~it.eln.;Itf1I'·IMIM __ ....... _ ... ~ _:::... .... ~'~~":~=::~:~==,~:~:":~a:J

=~- _._~ ....... _ ~=';: -~5 ~,~: ;£ -w-- --' __ _ ••• 1_._ -

~

•

......

.a ~ .

:.c COl

•

.

..

S

: ....

'104 ..

'fo4 < a ..

...

'i ..

u f 'i ...

..

f

...

Teat .JiSla..

() 1 Original Blade

CJ 4 Improved Blade

3300 3400 'J500 3700 3800 Corrected Fan Speed, N , rpm 1K Figure 65. CF6-50 Engine Fan Airflow Characteristics. Sea Level Static with Original and Improved Fans.

Corrected Booster Airflow, W K. lb/eee 220 240 260 2S0

i

u

Nominal Fan Nozzle Exit Area (A1 >

S "" 8.

l~ Test JW- () 1 Original Blade [] 4 Improved Blade 2.4 ..

o ....

+I (\, c= 2.2 CI

It

, "" =' III III CI

t. 2.1

"" III +I

I 2.010""0---1...J.0-5----'1l'-0---1-i15

120 125 130 Corrected Booster Airflow, W • kg/sec 23K Figure 66. CF6-50 Engine Booster Performance. Sea Level Static With Original and Improved Fans.

Predicted LPT"\ Map lope j.....-- •

o

...... 1% Teet

~Original

Blade

o 4 Improved Blade

o ::.:: 0.072 .....

I

..c .....

~ .....

~ :I .....

.., ~

CQ

I

..

..

eft eft t::.. ..

~

0.070 ~.

!

..

..

S 290 ~----1-- S

~ ....

....

+' +' U ~ II ...

...

~ ~ ~ DIll 285 ---+----+-----t------:::::::II 0.068 r.:I ....

....

.a .a ...

...

e. 280

e.

At ..:I

eJ

0.066 104 106 108 110 112 114

LP TUrbine Corrected Speed, N /.;r-

Ffgure 67. CF6-50 Engine Low Pressure Turbine Performance, • Sea Level Static With Original and Improved Fans.

indicated for the cruile power .etting for both •• a level and .imulated alti- tude condition •• Piaure. 60 and 61 Ibow .fc improvement. at both .ea level and .imulated cruile frOM the initial Anaine te.t. with the im~oved fan blad.. At a ••• level It.tic thru.t equivalent to altitude crui.e power, an improvement of 1.8% wa. obtained, Which repre.ented approximately two-third. of the improve- ment required to achieve a 1.8% improvement at altitude crui.e conditionl.

The curve. indicate that significant fuel con.umption improvement. had been demon.trated but were not con.i.tent with the oblerved fan efficiency increa.el Ihown on Figure 62. The mea.ured fan efficiency increale of 4.2 pointl was within 0.5 point of expectation.. On the ba.il of perfo~.nce .en.itivitiel, this would have been expected to produce a much larger improvement in enaine fuel conlumption. Detailed analy.is of the enaine cycle revealed that the fan efficiency improvement wal not reflected directly in improved sfc but wa. off- .et tq a .ianificant degree by 10. Ie. in the boo.ter and low pre.lure turbine effielencie •• Fan operating line. are .hown in Fiaure 63 for the sea level and .imula- ted altitude run. on both fan bladel. An unexpected rollover in the fan oper- ating line occurred at higher airfiowl with the improved fan blade. The oriainal fan bladel were reinltalled and te.ted to check out the in.trumenta- tion and met.lured performance. This testina indicated that a "preuure cor- relation" il required to be applied to the mealured pre.lures to determine true averaae .tation pres lure with the new fan blade. Thele adjustments do not affect sea level sfc results where performance i. baled on measured thrust, but they are needed for the performance asseslment for the simulated altitude te.tl with the exhaust nozzle diffusers where thrust mu.t be calculated from exhau.t nozzle pressure.

AI had been expected baled on prior teltina of the improved fan blade on an earlier engine, the thrult and airflow characteri.tics versus fan speed show. con.iderably lower fan Ipeed requirement for equivalent thrultl with the improved fan. These higher flow pumping characteristici at a given lpeed are shown on Figures 64 and 65 for the improved fan.

Data analYlil and delian point cycle studies Ihowed that the higher flow pumping characteristic of the improved fan resulted in performance penalties in the booster and low pressure turbine, detracting from the fuel laving.

projected for the fan efficiency improvement. Figurel 66 and 67 .how these deviations from the expected performance in the booster and low preslure (LP) turbine. The lower required fan lpeed to produce a aiven thru.t cauled the boo.ter and LP turbine to operate at lower lpeed and lower efficiencies com- pared to operation with the original fan. Prior engine telting had given an indication from limited inltrumentation that increased booster lupereharging with the increaled hub camber on the improved fan blade could ellentially recover the boolter operating line from the lOll that would be expected from the required low rotor lpeed. Figure 66 IhoWI that this did not occur on Test 4, which had complete booster di.charge in.trumentation. The lower boo.ter operating line relulted in a 10.1 of operating boolter effici .. ncy of 0.5 to 0.75 point. A1thouah ~ 10" in LP turbinl efficiency v •• eXPIcted due to off-d •• i.n oper.tion vith thl lover turbine .peed. thl d.t. iadic.ted th.t the LP turbine effici.ncy 10 •• v., .ore th.n expected by .bout 0.6 point.

Th ••• d.t •• bowed tht furth.r cycle mprov_ntl could be obtained by i.prov.d aatchina of the DeV f.n .ad other enaine coapon.nt ••

Th. 1.0% incr •••• d fan noaal. ar.a. A18. v •• then te.t.d vith the u.-

prov.d fan blade at •• a 1.v.1 .nd .mu1.t.d crui.e (Te.te 8 and 9). Specific fu.l con .... ption r •• ult. ar •• hova on riaure. 68 and 69. Th. u-proveaent.

were not con.id.r.d .ufficient to warrant the chana', .ince th. calculation • • ctu.lly .howed a .liaht 10 •• (0.2%) et cruil' with tb. open A18' Th •• e ... 11 frection. of a perc.nt are vithin the .bi1ity to calculate perforalace fro. noaal. pre.lure •• The major te.t ina needed to ident ify add itional perfora&ncl iaprov ... ntl for the iaproved fan perforaanc. enline w •• directed to the •• coDd Cr6-50 enline. The fan di.ch,rle arc r.ke. were reaoved froa the fir.t eDline and in.ta11ed in the .econ~ enline. The b.ck-to-back ILS te.t of the orilinal to General E1eetric-f.bricated improved f.n blad ••• howed a 2.3% .fc mproveaent .t the crui •• -equiva1ent power of 17.800 daN. Which w ••• oaewbat b.tt'r th.n the 1.8% .fc improveaent ob.erved on the fir.t enline but .ti11 1e •• th.n the objective 2.7% at SLS.

Te.tina Qf the fir.t enline had .hown that the key to a acre optia ... ea- aine cycle wa. to reduce the fan blade flow, ther.by inerea.in, fan rotor .peed to obtain the .ame thru.t. Thi. became even 80re imper.tive with the vendor-f.bricat.d improv.d f.n blade.. Th •• e blad •• h.d .hown • further in- erea.e in fan .irflow relative to the General Electric improved b1ad ••• nd • higher .fc by 0.6%. Althoulh the intent we. to produce identic.l blad.I, the procel./toolinl at the vendor had re.ulted in .liaht .taaler ch'na" in loae .ection. of the airfoil.

A reduction in blade flow w •• aceoapli.hed with a .et of prototype vendor- fabricated blade. by aodifyinl the part-'pan Ihroud interlockl. The aodifica- tion produced a reduced incidence .nale in the runnina condition, .inc. the blade anile during encine operation i •• et by the blade-to-blade •• atina of the interlock.. The objective w •• 1.5· clo.ure of the runninl .talc.r anile at the p.rt-.p.n .hroud.

The re.taClered (vendor-fabricated) improved f.n blade. were effective in reducinl the f.n flow at .pe.d or incre •• inl rotor lpeed at a liven thru.t.

Thi. effectively improved .fc at crui.e equiv.lent power of 17,800 daN SLS thrult by 0.5%. Filure. 70 through 72 lhow the effect. of the v.riou. f.n bl.de. on SLS Ifc improvement ver.u. thrult, thru.t verlu. fan 'peed, .nd f.n efficiency ver.u. fan airflow, re.pectively. The re.t'a,ered improved (vendor-f.bricated) fan blade. Ihow a con.iderable reduction in en,ine thrult above SLS takeoff thru.t level. due to the reduction in fan airflow. A 10 •• in fan efficiency with the reltallerina at higher fiowl i •• hown in Figure 72, although aerodynamic analy.i. would not predict a 10... Ielu1t. of the f.n blade evaluation are .u.mariaed on Filure 73, Which Ihow. the Ifc u-proveaent a. r.lated to fan airflow increa.e over the ori,inal fan blade at con.tant 'peed.

- ~- .~# _ .. - -

42 44 46 so

4.

Te.'

110.

0 laprcwed l1ade ,.) __ Ai ~ • laproved l1ade ,.) +11 A • 110 110 200 110 lao :.., "'.,.., • rill' kII Figure 68. CF6-S0 Engine: Effect of Fan Nozzle Area (AI8) on SFC Perfor1ll8nc:e, Sea Level Static With IlIIproved Fltn.

-d']~i+t 0 rtar -

--~

Net Thru.t. '0' lb .,

• •

r--,-----------T~------------~------------~'~----------~~

Simulated Cru1 •• 0.85 "0. 35.00~ It, Standard Day

t

i

..

..

'" Teet

•

No.

o 3 I.proved Blade (01) N_ All

o • I.proved Blado (01) +1' All

40 45 Net Thruet, 'NI. kN Figure 67. CF6-S0 En,ine: Effect of Fan Nozzle Area (Ale) on Si.ulated Altitude SFC Performance with Improved Fan.

11.

._ 2L;,~-.~~~::.n. --A· )lIll¥'W'~~~=-'''''''.''''f*-i''''~:".''''''. ·",,·iF ... ·-o ... _ .... _ ... ;.."""_.""'~_~ __ ... .zF, ."..,,, -;.-¥","~¥#~, .•. - ."' T-.- ·~e;,~~"'~~;.!T:~~~¥~~~·~~T:"'~.,,~-~,"'" .,...."=""§7?"~~,..'"'"'" -'".~- '·i;'·';··';··;~--··- Net Thrult, 'NIt lb 40 45 SO 55 x 10 ~--~~----~----~~--~~--~--~--~----~

-

~ ~------~----~ ......

..

.c ......

A ~

-

, ."

T12 Clearance Test N~ ......

...

.c 14 Improved Blade, (GE) 0.115 15 Original Blade 0.175

:4

16 IlIIproved Blade 0.115 ..

~

U (Vendor) \W III 17 0.084 Resta&&cred 1m- proved Blade, Closed 1. 5- (Vendor) 170 190 220 Net kN Figure 70. CF6-50 Engine No.2: Effect of Varioue Fan Blades on SFC Performance, SeA Level Static. Nominal Fan Nozzle Area A • l8 l_ I L tip C1U'.'

illS III. Ii _~4 aleft, 14 (01) 0.115

I

Or1l1ul II ...

15 0.175 16 l.,nHtI 11He 0.115

I

~

CY ...... ) I 17 ... t ... .,M Ie- 0 0.014 i ,rcwM 11_, (I ... 1.S· (VeMo") 54 • 10l l -j 23~----~------~------~----- 'U T .0. 'ower SO 2Z J

. n

•

-

.... .

. ....

..

.

..

!

t.

!

..

t.

..

•

•

It 17~ ____ ~ ______ ~ ____ ~~ ____ ~~ ____ ~~ __ ~~_~~ noo '400 )500 '600 3700 )900 Fi,ure 71. CF6-SO Enline No.2: Effect. of Variou. Fan 11ad •• Oft Thruat Ch4racteri.tic8, Sea Lev~l Static, ~1nal Fan Nozzle Area.

--~ :"J.~~:-. t:_,1';=:;:~~!':~~~""'!:~'" MJL.

Teat No. Ti2 Clearance (GE) 0.115 14 Improved Blade, Original Blad. 0.175 16 Improved Blade 0.11~,

.. ~

(Vendor) 0 17 R.atallered 1m- 0.084 proved Blade, Closed 1.5· (Vendor) Correcred Fan Airflow. W , 1b/aec 2K 1250 1300 1400 1450 1350 1500 1550 ;, t-----;.----+-- rr-::---9;-1~B-~~~~ :---~----I u C III ...

u ....

.....

...

III c: ~ 1-- __ --+-_ ~------+_~0--~--------4_------_4--------~------_+--------~ 580 600 620 640 660 680 700 Corrected Fan Airflow, W , kg/aec 2K Figure 72. CF6-50 Engine No.2: Effects of Various Fan Blades on Fan Efficiency, SLS, Nominal Fan Nozzle Area AlB' Eltimated Airflow Effect on Cycle at Constant Fan .

Engine 2 - Telt 19 Efficiency

! Restallered Vendor Bladel

l!! 3 1---- Reduced Clearance. Smooth U Shroud. +1% A18

'"'

.5 -"'-..,1

~ -I ....... """ Engine 2 - Test 14

IS -...

.. Engine 2 - Test 17 ~ __ OGE Blades III 0 --"-- I ~ Restaggered Vendor Blades -.-.

~ 2 r-----~~----T-------------~---------~~~------ Engine 2 - Test 16

o Vendor Blades

~ Engine~ - Tes~ 4"-

'~

! GE Blades

(j ...

III) ""'"

1 2 3 4 5 6 Fan Airflow Increase. percent Figure 73. CF6-50 Engine SFC Improvement Versus Fan Airflow for Various Improved Fan Blades. SLS Thrust of 17.800 daN (40.000 Pounds).

Nominal A • l8 Fan tip rub button te'tina on the .econd enaine va. initiated to deter- mine if a further reduction in tip clearance wa. po.aible (Fiaure 74). Thia teatina ,howed that the fan case 'ti~fener had provided a aianificant ia- proveaent in caaina rou"dn~.s compared to the unatiffened case. This perait' tiahtenina arind clearances such lh~t aver .. e runnina clearance. are corre- spondinaly reduced for improved performan~e vithout reduc:~a minimum runnina clearance.. Back-to-back testina with a further reductio • in tip clearance of 0.090 em (0.035 in.). along with a chanae in tip .hroud material from open- celled honeycomb to a .mooth surface "microbanoon" •• howed a significant im- provement in fan efficiency (1.2 points) and sfc (0.7%) at part power somewhat more than expected. Conversely, th~ performance improvements previou.ly .. a- .ured on an enaine with the restaaaered blade at the .... clearance were .ome- what less than expected, considering the airflow reduction achieved as .hown in Fiaure 75. Although the performance increment. measured for the further reduced tip clear.ance and for the blade restaager at the same clearance were inconsistent with expected improvements, these increments are small and with- in the capabilities of engine/data system repeatability. The net total ef- fect of these changes, however. was consi.tent with prediction •• The final improved fan package includes the restaggered improved fan blade, a total of 2.5 em (0.100 in.) reduction in tip clearance with the fan case stiffener and the smooth fan casing tip shroud.

7.6 APPLICATION OF RESULTS Results of this testing and plans for further development have been co- ordinated with the aircraft manufacturers currently using CF6-50 model en- gines. Flight test programs directed toward recertification of the respec- tive aircraft and a production engine progr .. for the CF6-50C2 and E2 model engines have evolved from the fan development program. The' first set of pro- duction-type fan blades having the desired stagger angle was tested back-to- back against a set of the modified restagsered blades in a production CF6-50E2 engine installed in a production test cell. This testing essentially showed the same performance for the two blade sets. A comparison of sfc performance for the third engine is shown on Figure 76.

The production CF6-50C2 and CF6-50E2 model engines are demonstrating th~ expected sfc improvements relative to the CF6-50C and CF6-50E models, and the performance results are presented in Section 11.0.

:""","iP;: '111'111:;;-"-:;'=:;-- ;;;:o- .. ;;:_-=,,:;;~: .. :.:~="'":. _:;- -':;:z::;:'; :-:;:.-:'"'¥':;:.,.......,.;:., ~:~.""'*="":--=.-=. _:'""'=. -::,:' f"'""""=,"": . . :71!;:.=-=-'fF"i':::,.'!'£_::''''' ':~:':"H,:.::::-;.,,,,:>-':!f~:}--::" :_!':. ;..::4:"::_.R:--,_,:,¥,,,:~.,o;~TW~~--'~~:--'-!'¥'~7" ~;;r;:::~,~~ljff"~"=""!~ .·:-->""~''"'';'~'''H,.-:· = I: :l so!I ___ iii '"'"_ .. __ _ _,_~._~-=--==~ :""'-- ... '..,..... ___ EF

r

corrected Fao Airflow, W , lb/ .. c 1200 1250 1300 1350 1400 1450 1500 .

>- u Test No.

e

... D 18 Open Celled Honeycomb Shrouds, 2.92 _ (0.115 In.) ___ +-

u ....

Tip Clearance ....

....

IaI

o 19 Smooth Microballoon Shroud" 0.9 II1II (0.035 in.)----+-----I

c: Reduced Clearance :.

560 580 600 620 640 660 680 Corrected Fan Airflow, W ' ka/aec 2K Figure 74. CF6-50 Engine No.2: Effects of Fan Tip Clearance on Fan Efficiency. Sea Level Static, Nominal +1% Fan Nozzle Area with Restaggered Improved Vendor Fan Blades.

Net Thrust, F , kN NK 180 190 200 240 210 220 230

D 18 Open Celled Honeycomb Shrouds, 2.92 mm (0.011

in.) Tip Clearance C 19 Smooth Microballoon Shrouds, 0.9 mm (0.035 in.) Reduced Clearance .0

-

,....j

--'--

......

...

.c 1% ......

.0 ,....j

-

:z: III

"C T

......

...

.c ......

~ a U \I-< en 0.7% 40 45 Net Thrust. F , lb NK Figure 75. CF6-50 Engine No.2: Effects of Shroud Tip Clearance on SFC Performance, Sea Level Static, Nominal +1% Fan Nozzle Area with Restaggered Improved Vendor Fan Blades.

)E~~~~v-.,-_ ,*-~., ,.,' - P-J

i

Net Thru.t. 'NK' lb 41 44 41 41 so , .

-

j

rJ

=lll

09J

r--

~e.FJ

OftP

.

u ...

5D

•

g

o ... hgpred laproved Blade. (Welded)

[J Production Restaleered I.proved Blad •• -

I I

180 190 :110 220 230 240

Net Thru.t. 'NJ(. kN

Figure 76. CF6-S0 Engine 3: Back-to-Back Improved Fan Blade Performance.

• 8.0 ENGINE ACOUSTIC TEST A leriel of static back-to-back atoultic testl was conducted on CF6-50 engine configurations with the original and the Uiproved fans. Objectives of the acoultic test program and subsequent data analYlel were to eltablilh the acoustical effects of the Unproved fan performance improvement package on CF6-50 engine noile, and aSless the impact of the subject engine modification on community noise levels for typical aircraft approach and takeoff flight conditions. This section describes the acoustic test facility, engine con- figurations, instrumentation, acoustic testing procedure, and data reduction methods utilized. Results are presented comparing noise levels of the origi- nal CF6 fan to those of the improved fan configuration.

Community noise exposure estimates are presented in terms of simulated effective perceived noise level (EPNL) determined from the static test data and typical aircraft operating characteristics for takeoff and approach flight conditions. These estimates were developed analytically assuming the static measurements to be flight levels.

8.1 ENGINE ACOUStIC TEST FACILITY The static back- to-back noise tests were performed in the outdoor perfor- mance/~coustic test facility. The site was paved with concrete exte.,ding a minimum of 6.1 m (20 ft) beyond the microphone positions. The acoustic field was free of obstructions for 45.7 m (150 ft) minimum distance beyond the far field microphone locations. The engine was mounted to a thrust frame supported by an open-trussed cantilever structure with the engine centerline located 3.96 m (13 ft) above the concrete as shown in Figure 77.

The engine was operator-controlled using a minicomputer-based engine and , facility data monitoring system located in the control room complex. In addi- i· tion to the real-time monitoring of engine performance parameters, the compu- ter system provided real-time data acquisition and analysis for a maximum of 878 channels of steady-state engine performance data and ambient conditions.

The data channels were continuously sampled and time-averaged for a 52-second interval for each engine power set point. Acoustic data acquisition was syn- chronized with the data monitoring system (OMS) data acquisition.

8.2 TEST CONFIGURATIONS One product ion CF6-50C engine (Douglas configurat ion) and one product ion CF6-50E engine (Boeing configuration) were used for the acoustic tests. Both engines were fitted with a referent c' dcoust ic inlet with bel1mo'lth 1 ip, a long reversing core nozzle (LRCN), and current production fan and core duct acoustic treatment. The acoustic treatment is described below: .. -~

'I ' . \

Fl g ur 77 . CF6 - 50 Eng in Mount d on Thrust Sta nd for Acoustic T sts .

OR ' GINAL PAGE IS Of POOR QUALITY Treatment Area Location Treatment Type 2 2 'an Inlet Sinlle Delree of Freedom (SOOF) 5.57 m (60 ft ) 2) F.n C.linl Multiple Delree of Freedom (MOOr) 5.85 m (61 ft 2 2) Fan Exhault Duct Single Delree of Freedom (SDOr) 4.65 m (50 ft 2) Multiple Delree of rreed~ (MOOr) 4.37 .2 (47 ft 2 2) Long Reveninl "Tophat" (SDOF) 1.94 m (21 ft Core Nozzle The engine. were modified in order to run the confilurations described in Table VII. All performance ra~es were excluded from the fan inlet, fan exhaust duct, and core exhault duct for the.e test. in order to prelerve the acoultic char- acteriitici of an aircraft installation.

For the baleline acoustic tests (No.3 and 5), the CF6-50 engine wa. run with a let of orilinal production fan blades. The fan calinl tip Ihroud wa.

open cell aluminum honeycomb and ground to the production minimum tip clear- ance of 4.45 mm (0.175 in.). For the fan blade tip clearance acoultic telt (No.1) with the original fan blades, the fan caling tip Ihroud was modified by installing microballoon epoxy in open cell aluminum honeycomb to provide a Imooth casing tip shroud. The tip Ihroud was ground to provide a fan tip clearance of 3.30 mm (0.130 in.), a reduction of of 1.14 mm (0.045 in.).

For the improved fan acoustic tests (No.2 and 4), the original produc- tion fan blades were l'eplaced with a full set of restaggered improved faD blades. The Imooth fan casing tip shroud. were ground to provide tighter fan tip clearances down to 1.90 mm (0.075 in.), a reduction of 2.54 mm (0.100 in.). Refer to Section 3.0 for a detailed delcription of the improved fan.

8.3 INSTRUMENTATION - ACOUSTIC 8.3.1 Far Field Microphones Acoustic data were obtained from a set of microphone Iystaas which con- sisted of the microphone, cathode follower, power supply, and pistonphone, no windscreens were used.

The far field microphone. were positioned 3.96 m (13 ft) above the con- crete surface and at 10· to 160· in 10· increments measured from the inlet direction. Additional microphonel were positioned at 85·,95·, 105·, 115·, and 125· at the same height. All microphones were located on a 45.73 m (150 ft) arc measured from the fan exhaust nozzle and oriented toward the source (normal incidence) at a sufficient offset distance from the microphone stands t9 minimize reflection effects. A sketch and photograph of the sound field are shown in Figurel 78 and 79.

... ~ --m-- ,

'~ .. '

, ,.

~

(1 1'1 ...

j

~ .: 'I ~ • i

!

Table VII. CF6-S0 Engine Test Configurations for Acoustic Tests.

I

l

Rominal Buildup !

I Tip Clearance Core Test Fan Casing Tip

-i

Shroud ..

Configurat ion Blades .. in. J!fozzle

No.

,~ '" Original Fan Original SIIooth 3.30 0.130 uCll* ,M 1 • Microballoon Reduced Tip

i

Clearance Restaggered S .;loth 2.29/1.78 0.090/0.070 LRCR 2 Improved Fan Improved Kicr'·.;,alloon

'1

Original Fan Original Open Cell 4.45 0.175 LRCR

I

Baseline Honeycomb 4 Improved Fan Restaggered Smooth 1.90 0.075 LRCM

II

Iaproved Microballoon I , , ~ ', j " Original 5 Original Fan Open Cell 4.45 0.175 LRCM Honeycomb Baseline , : *LRCN - Long Reversing Core J!fozzle

, !

I

,.: - .. __ ~_~~" ~'""'T'''r'''"'''''' -','

"" ... . ....... ~, ...... , .• ,." ....... ,.. .'

~

". ",,~ •.. "."," ..'I,'"""o:.",.il!llIiuillWlll,: ;"""81 .. 011II1II1.,,,, ~:ijj,",,",.I,",iI,'",, .. ~,,"" .. oII ...... "'"",J~"".IItl .... ",,'.

_____ +_-,0 __ cE:lt-

~ \

-

..

X

~160 PORTABLE MICROPHONE LOCATIONS ~ @ ENGINE G.. HEIGHT OF 3.96 III (13 ft -1-150 45.7 • (150 ftl ~ --.....,...

+-140

51.8 III

+-

TO COHTROL (110 rt) ROOtI +130

+

+

+'20 115

~ PORTABLE

~ -+ 105

ENVIROtU£HTAl

~70

DATA WEATHER

110~ ~

CONCRETE TEST PAD

1l 1e X

STATIONS

~ 80

Figure 78. CF6-50 Sound Field Layout.

w

-

-

Ml c rophnn l e r O I. hoo lt' --_.a.

~ I ,nJ F6- 0 Eng ln Scund Fl Id at T s t S!. t • FI A ur 79 .

,.".".--"""""' ...... -- ....... ---- ....... .,,-'~~~.------- -------

8.3.2 Acou.tic Oat. Recordin.

Acou.tic d.t. vere recorded on •• ,netic t.pe .t • tape .peed of 76.2 em/ .ec (30 in.l.ec). The recorder v ••• et up for 40% carrier deviation (t 40%) at full .cale record level. Si,nal .. plification va. provided by a Cener.l Electrie-de.i.ned ae/dc pre .. plifier .odule. Durin, t~.tin •• the tape record- er input .nd output vere monitored to a"ure that adequ~te .. plification va.

u.ed and to c"ure proper operation of the r8eorder. Oat. vere recorded for at le •• t 2 .inut •• at each .peed point.

8.3.3 Atmo'pheric T •• t Condition In.truaent.tion aarometric pre •• ure w •• recorded for each t •• t point. Wind .peed, dir.c- tion, air temper.ture .nd dew point w.re .11 .... ur.d u.in, • port.ble .nviron- ment.l data .cation (PEDS). Tva of the.e .t.tionl were loc.ted approximately 45- from the inlet on a 51.2 m (168 ft) arc. The .en.or. were po.itioned .t • 3.96 m (13.0 ft) heilht. Wind .peed .nd direction, •• me •• ured on one of the PEDS, were recorded continuou.ly on .trip ch.rt.. The .econd PEDS incorporated General Electric deatgned wind speed and V COle wind direction instrumentation.

These signals were also recorded continuously on strip charts. Ambient temper- ature was measured by e~oirated resi.tance temperature devices. The dew point measurement was made w1~.1 a hygrometer which sampled air from the J.96 m (13.0 ft) location. All tht ~ measurement. were also recorded on the OMS compur.er system.

8.4 TEST PROCEDURE ANr DATA REDUCTION 8.4.1 Atmospheric Te.t Condition Limit.

Atmolpheric condition Iimita were le't prior to th~ teat which include w.ter. Inow, or ice cover on lound field •• nd no viaible to, or precipit.tion.

Any data recorded outlide the iimita lilted below were dilcarded and no telt- ing wal permitted under conditionl of: Relative humidity 20% < RH < 95%

- -

Temperature Headwind < 4.1 m/lec (8 kn) includin, gUltl Crollwind < 2.6 m/lec (5 kn) includin, guat.

I

Tailwind o m/tec

GUlti < 1.5 m/lec (3 kn) ~ ~ 8.4.2 Enlin. Te.t Condition.

Variou. enaine confiluration. vere run to obtain data for ca.pari.on. at the .... corrected thru.t over a ranle of condition. thet enco.pa •• ed the ap- proach. cutback. and takeoff pover ranae. for aircraft powered by the CF6-50 enaine. For .ach te.t confiauration, the .equance .hovn in Table VItI wa. re- peated tvice in the .... order for a total of three r •• dinl' at each power .ettinl. A .hutdovn of at lea.t 30 ainute. occurred between each te.t •• ri ••• At each pover .ettina, the enline wa •• tabiliaed for at lea.t 2 ainute. prior to recordinr acou.tic data. All perform.nce p.r ... ter. were determined from the OMS comput.r .y.te. and corrected to .t.nd.rd temper.ture, .t.nd.rd .ea level pre •• ure, aero huaidity, and aer~ vind u.ina .... ur.d .abient data for t .. perature, pre •• ure, ab.ol~re hu.idity, and wind velocity and direction.

8.4.3 Calibr.tion of Far Fi.ld Microphone Sy.t ...

Prior to •• ch confi,ur.tion te.t .erie., each .y.t .. VI. c.librated to det.~ine frequency re.pon.e and .en.ttivity. tach .icrophone he.d va. re- .ov.d and a known volta •• lev.l of pink n~i.e va. input to the .icrophon.

pr.amplifier. The output .ianal from each .. pli£i~~ va. recorded on .alnetic t.pe. Sub.equent pl.yback .nd proc ••• ina throulh the data reduction .y.t ..

determined .y.t .. frequency r •• pon.e correction.. The .icrophone c.rtridge re.pon.e, a. determined from the individual microphone l.boratory c.1ibr.tion curve., va •• 1,ebraic.lly added to the foraer correction. to determine the overall .y.tea re.pon.e for inc1u.ion in the data reduction prolraa.

The .icrophone c.rtridae va. then rep1.ced and. 124 dl pi.tonphone va.

applied to e.ch .icrophone. The microphone .en.itivity v .. (~omp.red to the mo.t rer-ent laboratory calibration data to a •• ure compli.nce vithin t1.5 dB.

Any .y.t .. fallin, out.ide thil band v •• replaced. The microphone output.

were then normaliled u.ina variable .ttenuatorl in order to record the ftaae voltaae level vith the pi.tonphone .ource input. At the conclu.ion of e.ch te.t .erie., the pi.tonphone va. re.pplied. and the volt.ae level wa. record- ed a. a verification of microphone 'Yltem intearity.

nn .everal occa.ion. throuahout the te.t leriee, 2-.inute recorJina' of _bient noile vere made vith "facility on" and "facility off". The.e record- ina. were .ade at ,.in .ettin,1 u.ed durin, the lound .ea.urement. to aa.ure acceptable aisnal to noile ratio. for the acou.tic data.

8.4.4 Acou.tic nata Reduction Off-line acoultic data reduction va. performed uling an Avtomated One- third Octave Band nata Reduct ion Sylte.. The recorded data ve~'e played back on a 28-track .y.te.. In the automatic operatina mode, control of the 'YI- tem va. provided by meanl of a computer and operato~-provided information.

The data to be ... pled were located by .eans of a time code reader. indexing from the ti .. code .ianal recorded ~n the data tape. Thi. tape-Ihutt1inl va.

continued for each data channel vith .a.pling performed over the .ame time Table VIII. CF6-50. Engine Nominal Test Conditions for Acoustic Tests.

Corrected Thrust, Corrected Speed. Nll1.r~m~ Fn/4 N Original Fan 1b Im..R,.roved F.n.

52,700 234,420 3875 227,962 51.248 3850 216,810 48,741 3780 3710 204,488 45,971 3700 3600 168,921 37,975 3450 3350 158,477 35.727 3370 3275 33,382 148.491 3290 3210 126.592 28,459 3100 117,113 26,328 3010 2910 109,173 24,543 2930 2835 101,686 22,860 2850 2760 ql~ ,654 21,279 2770 2685 88,079 19,801 2690 2610 81,958 18,425 2610 76,291 l7,151 2530 72 ,341 16,263 2470 2405 68,053 15,299 2400 59,094 13.285 2230 2190 53,632 12.057 2100 !I"~'%*"",,\_' . _. ,··-,_"",,~,.,~T,r_~,-,,- _"" .. "" "'~'.,.,....,,-,~"". F,', ~""~,,, ...... _ • .4 .. .to"" .~,_,., .. --,~ __ .,~~"~~ ..,.. ., $ ""'--:>l""':, I • increment until all channels of a particular reading were processed. The system then advanced to the next data point. based on the operator-supplied time reference. and repeated the shuttling process. After the processing information (including reading identification. reading time. gain changes, etc.) was set up by the operator. the system ran without further operator assistance until a magnetic tape change was required.

All one-third octave blil\d analyses were performed using a one-third octave band analyzer. An integration time of 32 seconds was used to provide adequate sampling of the low frequency portion of the data signal. The fre- quency range of the data reduction process was 50 Hz through 10 kHz. Each datd channel output was passed through an interface to the minicomputer where data were corrected for both the frequency response of the acquisition and reduction system (as determined from the pink noise calibration) and for the microphone head response. The minicomputer was interfaced to a main frame computer to generate a file containing the one-third octave band data for fur- ther processing. The one-third octave band data were also punched on paper tape as a backup for the communication interface system.

Noise data at each point were processed using a digital computer program to normalize the data to a 298- K (77- F)/70% relative humidity. standard day and perform data extrapolations to various sideline distances. Overall sound pressure levels, PNL and PNLT, were computed for each angle at the sideline distances. The sound power level for each one-third octave band and overall sound power level were computed for each test point. These results were used for subsequent analysis and data comparisons.

8.4.5 Instrumentation Accuracy The accuracy of any sound measurement is dependent on the accuracy of the acoustic data recording and reduction system which is dependent on the toler- ance of each independent component in the system. A 1 ist of each component and the accuracy (3 a tolerances) is presented below: Component 3 ,T To lerance (± dB) ~ Microphone Cartridge Calibration 0.2 f < 10 kHz Cathode Follower Amplifier 0.2 Pistonphone 0.2 Noise Generator 0.5 Power Supply 0.09 Variable Gain Amplifier 0.2 Tape Recorder 0.5 Tape Deck 0.5 1/3 OB Analyzer 0.25 Since these variations are independent of each other, the estimated variance of the sound level can b~ computed as the rms of the variances due to each component separately. The accuracy of the acoustic data recording and reduc- tion system is then ± 1.0 dB.

The accuracy of the system does not define data reproducibility, which is dependent on many other factors. The intrinsic variation of acoustic data.

due to meteorological conditions, source variation, random instrument error, etc., defines data sample variance about the "true" absolute level of the noise source under evaluation. Hence, the instrumentation accuracy defines the tolerance (systematic error) on a "true" noise level determined from test sample statistics (random error). Noise level differences obtained from static back-to-back testing remove any data bias introduced as a result of instrumentation systematic error.

8.5 TEST HISTORY A production CF6-50E engine with the original CF6 production fan blades was modified to a reduced tip clearance and smooth shroud configuration and then tested. Upon completion of the clearance testing, the engine was con- verted to an improved fan configuration and various performance and acoustic tests were run. The engine was then converted to an original fan configura- tion with the original CF6 production fan blades, open cell aluminum honey- comb fan tip shroud, and production fan tip clearance. Checkout, power cali- bration, and acoustic tests were performed to provide an acoustic baseline for the original fan.

A production CF6-50C engine was converted to the final improved fan con- figuration and installed at the same test site. Checkout of the engine and facility was completed, followed by a power calibration of the engine and the acoustic testing. The engine was then modified to an original fan engine configuration and a baseline acoustic test series was completed.

8.6 TEST RESULTS AND DISCUSSION 8.6.1 Effect of Fan Tip Clearance on Acoustic Characteristics Static acoustic tests were conducted to evaluate possible differences in engine noise produced by reducing fan tip clearance. Results are presented for the original product ion fan configurat ion with the product ion open cell honeycomb tip shroud with a minimum tip clearance of 4.45 mm (0.175 in.) in Test 3 (Table VII) and the original fan blades with a smooth microballoon casing tip shroud and a reduced clearance of 3.30 mm (0.130 in.) in Test 1.

The 45.7 m (150 ft) arc data were extrapolated to a reference sideline of 122 m (400 ft). Spherical divergence was used to correct for distance, and SAE Aerospace Recommended Practice ARP866A (Reference 3) was used to correct for atmospheric absorption. A 122 m (400 ft) sideline distance was selected for data analysis, because it is representative of FAA certification altitude for approach engine pow~r settings.

The use of longer sideline distances representing takeoff flight paths woui~ tend to mask potential differences in high frequency fan noise that could be produced by the tip clearance/fan shroud modification. To facili- tate data comparisons between the engine configurations, all data were aver- aged, and comparative plots were made. Data are presented for the power settings summarized in Table IX. Averaged one-third octave spectra at peak forward and peak aft angles are presented in Figures 80 through 85 for low and high approach, and cutback thrust levels. PNL directivity behavior comparing the engine configurations at these thrust levels are exhibited in Figures 86 through 88. PNL data at peak forward and peak aft angles are shown versus thrust in Figures 89 and 90.

No significant acoustic differences between the engine configurations are apparent in the spectral comparisons at peak angles. lbe perceived noise rli- rectivity behavior and thrust behavior comparisons do not exhibit any system- atic differences in engine noise as a result of the tip clearance/fan shroud mod ificat ion.

8.6.2 Improved Fan Acoustic Characteristics Results comparing the original and improved fan configurations, obtained from two separate static back-to-back engine test series, are presented herein.

Data analyses to evaluate small differences between the engine configurations are discussed.

Corrected thrust versus corrected fan speed (NIK)' as determined during the test series, are compared in Figure 91. Good agreement was maintained be- tween repeat test runs for each engine configuration.

The 45.7 m (150 ft) arc data were extrapolated to reference sidelines of 122 m (400 ft) and 305 m (1000 ft). Spherical divergence was used to correct for distance, and ARP866A (Reference 3) was used to correct for atmosphere absorption. These reference distances were chosen because they are typical of FAA certification altitudes for approach and takeoff, respectively. To facil- itate acoustic data comparisons between engine configurations, all data were averaged and comparative plots were made. Results are presented in terms of one-third octave band spectra at peak forward and peak aft angles, PNL direc- tivity, and PNL as a function of thrust for peak forward and peak aft angles in Figures 92 through 107 at typical takeoff and approach power settings summarized in Table X. The small differences in acoustic characteristics be- tween engine configurations were not systematic and were considered to be within normal data scatter for static engine noise testing with the excep- tion of the 315-630 Hz band levels shown in Figure 96.

nle spectrum level differences shown in Figure 96 between the improved and original fan engine configurations at 50· and 158 kN corrected thrust (cutback) are the result of mult iple pure tones (HPT) , Le., "buzz saw" noise.

: ~.~ _..i.1..4.:'f: ..:..

'~*rtE'Z. r Table IX. Nominal Engine Power Settings Used for Data Presentation Comparing Reduced and Standard Tip Clearance Configurations.

Nominal Thrust Flight ~idel ine ~istance kN lb Condition meters feet 33,400 Cutback 122 400 148.5 88.0 19,900 High Approach 122 400 ..

Low Approach 122 400 59.0 13.300

I

Table X. Nominal Engine Power Settings Used for Data Presentation Comparing Improved and original Fan Engine Configurations.

Nominal Thrust Fl ight Sideline Distance kN lb Cond it ion meters feet 51,200 Takeoff (T/o) 305 228 1000 158 35,600 Cutback (C/B) 305 1000 88 19,800 High Approach 122 400 13,300 Low Approach 122 400

-

~~~ __ """"",- ___________ ",. _____ ?"··"'··"'--''''·§iiii5_ •• W_iiiIiiIliII_n_~_··;lii"ii1i .... ··~""""":" .... ,,,,;,,;;._~~;,,,,,;;;;;,,,;,,,.-. _____ ,,,,,,_~ r- - ,4f!i!!!f - ..

~

.. 100

Cnlnl Tip Tip Clearanea 5)'111001 Teat ran Blade

I

Shroud Cold _Un.)

.; 1 Orilinel Open CeU 4.4S(0.17S) II ..

lIcmeyeOlib 90 (Prod.)

•

...

Smooth 0 1 Orilinel 3.30(0.130) MierobaUoon

~

!

~

§

7u ~ ~ I Ii!

10000 FREQUl::NCY. Hz Figure 80. One-third Octave Spectrum Comparison of Original Fan With Original and Reduced Tip Clearance at 60 Peak Forward Angle (122 m Sideline and 59 KN Corrected Thrust - Low Approach).

N Tlp Charance

• Caslnl Tip

5)'111001 Teu Fan lllade ~ Cold _(tn.)

Shroud N § 3 Or1g1na1 Open Cell 4.45(0.175) w' Honeyeomb e (Prod. ) II Original 3.30(0.130) I Smooth

.. 0

M1crobaUoon • ...

.; ...

> !j

!

...

~ 70 i-< g Q ..

'"

i5 I ~ 100 1000 10000 FREQl1ENCY. Hz Figure 81. One-third Octave Spectrum Comparison of Original Fan With Original and Reduced Tip Clearance at 120 Peak Aft Angle (122 m Sideline and 59 KN Corrected Thrust - Low Approach) • callna Tip Tip elaarlnca Syabol T .. t Pan Ilada N Shroud Cold _Un.)

i

3 Orlalul Open Clll 4.U(0.175) HoU)'cOIIb (Prod.)

i

0 1 Ofia, ... l SIIOoth 3.30(0.130) Mlcroballoon ..

"

•

"D

~

'0

a ~

S

~ A • ..

~

I

10000 100 1000 PUQIl!IICY. HI Figure 82. One-third Octave Spectrum Com~arison of Original Fan With Original and Reduced Tip Clearance at 60 Peak Forward Angle (122 m Sideline and 88 KN Corrected Thrust - High Approach) • e .. lna Tip Tip eleerence $yabol Tut 'In Ilade Shroud Cold _(In.)

3 Orlalnal Open Cell 4.45(0.175)

i

HoneycOllb o (Prod. ) Smooth 1 Or1atul 3.30(0.130) MlcrobeUoon 10000 PUQUENCY. HI Figure 83. One-third Octave Spectrum Comparison of Original Fan With Original and Reduced Tip Clearance at 120 Peak Aft Angle (122 m Sideline and 88 KN Corrected Thrust - High Approach) • - --------------- -.-~-- ~ .. ~_,,'''''~ __ ~~'"*__.:-.~~'"'' __ ~~''''_. _'."0 __ o",....iiiiiI en

I

I

~ ..

~

'I

I

fA

Ii

'0 ..

II ..

i

•

..

iQ 6

•

Q

Qi 6 6a i

GQ lil QG

~ .. &~e&'~ i

I

o $&

<:)

!

Tip Cl .. ranea III Ca81nl Tip S,.bol T .. t ran Ilad. t' :- Shroud Cold _Un.)

g

~ ] Orilinel Opan C.,U 4.4S(0.17S) Hon.y~OIIb ~ ..

(Prod.)

/: t /:) I Orlainll !IIIIooth 3.30(0.130) Microballoon ~ I 1000 10000 rnQutlfCY • Ha Figure 84. One-third Octave Spectrum Comparison of Original Fan with Original and Reduced Tip Clearance at 60 Peak Forward Angle (122 m Sideline and 148.5 KN Corrected Thrust - Cutback).

N ~ 100 ...

I

ci II ..

• ...

~.

III .0 Tlp Cl .. ran~e C .. inl Tip

!

5)'Il1001 Te.t ran Blad" Shroud Cold _(In.)

:- '" inel Open Cell 4.4HO.175) 3 Orll

§

Ho"ayc_b III (Prod.)

• ..

1 (l r1 1 1n el S""",th 1.)0(0.130) j:; II t "!nob.lloon ~ 100 1000 10000 Fl!QU!IfCY. Ha Figure 85. One-third Octave Spectrum Comparison of Original Fan With Original and Reduced Tip Clearance at 120 Peak Aft Angle (122 m Sideline and 148.5 KN Corrected Thrust - Cutback).

-

I

~

l

e

& ~

e

a

S-I_~ ~.

C 100 .

...

~

I A_

...

.8 , ...

II $

Clal .. Tip TiP CI ... IIIC.

T.lt .... w.

',-01 III ....... Cold _(Sa.)

i

!:l

0, .. c.u 4."(0.175) -

1 Od,1 .. 1 IIoM1c ..

( ...... )

).lO(O.llO) _th 1 Or", .. 1 NlcroNU_ 120 140 160 20 40 60 80 100 Ie to Inlet, dearee.

Figure 86. Influence of Reduced Fan Blade Tip Clearance on PNL Directivity (122 m Sideline and 59 kN Corrected Thrust - Low Approach).

; 110 .

~

... ~

~ 6) ~

a

~ ~

j

~ 4> II tb ! 100

~ e

II :> ...

II Tl, CI .... ncl 101 Cliina Tip S,-.kl feU .. 'a. lied.

Shroud Cold .U •• ) II ... 90

_yc_

) 4.4HO.17S)

Oflltnol Op<rn Cell -e

~ ("od.)

3.30(0.110) I Oriliul StIooth t:.

"l"ON11_

I

I I .~ 120 140 160 o 20 40 60 80 100 Angle to Inlet. degrees Figure 87. Influence of Reduced Fan Blade Tip Clearance on PNL Directivity (122 rn Sideline and 88 kN Corrected Thrust - High Approach).

~t"" 1ii ...... • __ -'-_.-.l __ -'-_~ ...... _~'"'··"",..,.I!!Ii·L··::..' .:.,' "'z';;'·· _-:;;..i..' ·:';;;::;4:'·,;';-:::·.oi.''':~i:1io' ",",' .-=~oio:-:.=~.: i"!:3·~·.·Il:kit'.·iI:I:·:I··=7Iitr.·I!Ii'I'·IF1 1' =.Iii. __ .... ___ .. ~ ... &~,.~."" .. ~. '......,~~ .. ~~-, ,.,.-~ .. ~.~~ o """'""'":=-.----p.=- '-~-=-~~~;=c~ "~.-'''~--#'' ,¥¥= '" - *o~-, ¥"t='~ ,_ :""~~""~.§#:~- -_ >!~- _~ _- M ___ ~-_ ~=="'_""..,~~~~~-~ ;O,;~ ___ ~_~~~-._-, .-.""'E.~~~ ",,~~¥>""";"'f1 ~~

~

$

...

e

8_$_._8

~-.

" 110

!

•

•

&

J 100

H " CUilll fI, '" cu.r_ .,...a r.at raa.u caw _Ua.)

ArM

i

, o,..c.u 4.4t(0.1fS,

•

I 90 0 0riPMl

....,...

-

( ...... )

...... 1.)0(0.110)

1 OI'lalMl A IU.cHMlu.

I

I I

40 60 1:0 140 160 o 20 10 100 Filure 88. Influence of Reduced Fan Blade Tip Clearance on PNL Directivity (122 m Sideline and 148.5 kN Corrected Thrust - Cutback).

Conanad Thl'ue,. lit 111 20 30 40 , II s

I

0 11:.

o 0& Jit

110 •

0& Iifl~ 4» I', Cl.., ....

_"" fl' THt 5-~~~ ',..1 'MIlaM Cel. _C, •• ) ,

- 0 ClrltiMl a.oo c.11 '.')10.1lt)

11:.

*-,- ChM.)

Clrap.l 11:. I I.JOI".,O) "u"'U ...

--

Htall '- Approach Approacll C"tback 9 5 60 100 140 110 Corre.~.d Thruet. IrJI Influence of Reduced Fan B1ade·Tip Clearance on PNL as a Filure 89.

Function of Corrected Thrust at 60· Peak Forward Anile at 122 m Sidelin,.'

10 20 40 I

!

•

o

0'1:,

j 110

'If ClMt_.

'*tl .. TIJI ... alMo fo ..

.. , ....

I COW.l •• )

--

i 4>1:9 0 J .,.", .. 1 '.ueo.lIS) """ Coli -,.- (P, .... ) c{]0

! _.h

1 .,.", .. 1 J.JOCO.1101 ~ 105 IUlnNll..,.

:.

6%6 Lw IIllh Approach Cut""~k Approach f-._- 1- , 22b 140 1 0 Corrected Thrun. III Figure 90. Influence of Reduced Fan Blade Tip Clearance on PNL as a Function of Corrected Thrust at 120· Peak Aft Angle at 122 m Sideline.

- --~-~"~~"'-'f"---~"----~" ~1

I'

f!

I

I'

4Yeoo 2000 2200 2400 2600 2100 3000 3200 3400 3600 3800 4000 Corrected ran Speed. filii Filure 91. CF6-50 Engine Corrected Thruat/Corrected Speed aelation- .hip with Original and Improved Fana.

_e-~_ ~-- -~.~ - -7Ets- -~------.....~ 1ft * 14" ,.., ...

'1 __ Hr ...... L''''' ~ 'lit,," lilt , KC_r " . .,.

• II.

Q .M Ct· UPt .. "'ll(lI tl' ..... " IN ",.ac" IHtl U Utll U IlIU U 114" LI ,. ,.

_I "I" ""1

""I

" " --

I

tNt

..... .._- 1", .. -

... , .. "' .....

...

"''''. _ fl •• , e I ",-" .. _,.

.... 1 .. '

4." CI.""

-c~ ......

........ ....... .,., I

•

0 ............ ...... ~,-

• '.""." CI ... ".'"

.... ' .. _n_

'''- I." ... .,.,

•

•

@) @ IJ

•

.·El

••

-- E)~wi fJ ~

~

~I

4) ~

: e

•

~~

~ i ft":!.

I ~

•

I , I I I • • ~ .. I I •• + •••• I ........ 1 • I I II .. I I ItllM )1111 " In II' - ten tIM

f""'fufOht"Y. "t

Figure 92. One-third Octave Spectrum Comparison of Improved and Original Fan at 50' Peak Forward Angle (122 m on Sideline and 59 KN Corrected Thrust - Low Approach).

r-------~·~--·---- - - 11" "ft. ,., I ~ \OUtta ,.., "~t l t wn y~ ' .. OUt fit , II'.; """r Il ,,, (1\'1"" ,,, U -,', ... ,~. {l-utt .. II '"~ tI·\et ..

IJU' ,. Iltt' I' ... ., I' ",' I ..

! ,~~~-!'--~-' ~-----''''''' .. ~~.'_" ______ --~-----~

, .. CI_ ".,

.,... e..1 .. ,I,_

.e.,.,. _ C11 ••

....

... .. -

, .... 011 .. _po e .... II.'U'

"'f''''

1:[ I'

.-1- .... '

"'II .. , .... c •• ""

• -T

............. 11- 0 .... .-.- .

I ... "'," II."".".

•

,-

.... ,,, ..... ,,-

0 '.11 II.""

•

, . I ~ o

I

I

I +--+--4-+--+-+ ~ i -+-"+--i~ -t -T-' ~~ +- I -i - -t~-.-+ -;--+-+--+ --+---+ --+- +--+--+-+J In 1" _ ,_ !>NIt ,,~ 1'''' fr .. "ue'H''f. tb Fip,ure 93. One-third Octave Spectrum Comparison of Improved and Original Fan at 115· Peak Aft Angle (l22 m Sideline and 59 KN Corrected Thru8t - Low Approach).

).' ,

" ,,'n f r_

I lie IIIoQI

•

II •• ea'LIIt" G •• U'LIIt" ••• U'LIIt" • .M "'Lilt" • .... u ." .. LI

I

''''IU

I"" L'

.....

I I'"'

M -" " "

"

'"" "

",et_

..............

.....

...., _ca ...

.... - ...

'"'

g ...-n .. ..,.

• .•• " ... a, ..." ..

•

-,..-. ....

....... . ..........

.a !

•

'.l1li." ...........

• ...... --1-

.... -- .

.......

...... Itt)

0 ....... --,-

•

•

i

, • ia

tI Il

. ~

I

,

•• I Ii

!

!

~_~i81i~

I 1 ;1.

•

I.

I

. . .. .. 1

I I I I • -+ • + ..

+ + -+--+- +- -+

• .+

• •

•

.n

•

•••

Figure 94. One-third Octave Spectrum Comparison of Improved and Orig~nal Fan at 50' Peak Forward Angle (122 m Sideline and 88 KN Corrected Thrust - High Approach).

I '" ~ I •• n, .. ~I ~ft~""". 1.1Yf" v, '''Outl!ttt i 'I' ICC ... "-1 I [I,.ee ..

I) 'M t" .... 0 ,II U ...... ~. fI-.ee.

" ."

. ......

I"" •• "", .. "," ..

...,.

..

,,."

!

.'", J

.- " " "

-"

...... YI, __

YI."_

...... ... ......

....

« ... ,. _ u •• '

-

a ... , •• 1 .... " .... ,

• "'-11"-"

_C_II., .a •••• c ..... , ... ".,

•

........ _11-

~."I\." ..........

•

.... -- .

........ _n_

0 ... c .... .,

•

'--

II

I

ii

I

• i

•

lale. I

I ,

• •

~8~~~AIIAee~M

i

I

•

II W.

i , --~I

.. I-+-j

+ • ~+-+-+ I

+

• • • • • •

• • •

In .,.,..

•

- - -

One-third Octave Spectrum Comparison of Improved and Original Fkn at 115' Filure 95.

Peak Aft Angle (122 m Sideline and 88 KH Corrected Thru8t - H1gh Approach).

, - ... -_ii'

!

'hot Tlp C1 • .,." •• • ,.-.1 re. IIlde. CU'.I T', ..... 1Id (Cold), _ U •• ) .0.

,

C o, .. -con .. HO",- 4." (o.la)

Or'I'''' ._~ ( .. _.\l0ft) Or'I, .. 1 4." (o.nn II

•

_ .. II,onuU ...

I 1.2111.,. (0.0110.01) .. • .. r •• '", _,~ lI'onull ...

4 I~"\f" 1.10 lO.on) <> ~ ):( a El

•

~ €I I , €I El

Q e ~

~ ~

.,llI

iI •

!g

~ €I El !

•

i El .

. i

:

e

; I I

I

• ..

5.

ll' 811 2011 ,eel 1251.

l""

Fri:'qu~ncy, Hz Figure 96. One-third Octave Spectrum Comparison of Improved and Original Fan at 50 Peak Forward Angle (305 m Sideline and 158 KN Corrected Thrust - Cutback).

M fl011 rTl 5L S.OUND P RE.5SU~F lE"iEL v~ 'JPfOUf~CY JI' 11' DEC "NCL.E ,e C!',LflCN • 15" Cl'lR.CN [1/lR[N El [~'I R[N () I El 1'1 1'.

] ,&~. LB ]5&86 II ]5~1Z II ]5B~6 L e 1~1I:I4'J7 N I 5BJ7c N 15874111 N 15B&17 II ~ 'RIGINAL PAGE IS f POoR QUALITY • ..

• I!! • @ .~ • i CD

~ ~

il

... I!!I

."

..

,.j t- o.

III , 'hot Tt, ct •• ntlee .,.-.1 eool.1 TI, ... rolld .. r ••• 'ad ..

(CoU), _ U •• ) No.

, 0,.., oCon.d _,- 4.4& (0.1'" 0 Or'I, .. 1 rrcttNcU •• ) CMb, 0"11 .. 1 4.41 (0."" Xl

•

0 I 1.2111." (0.011D.07) -,.

., • _ .. ro'led "".""", 4 "oUt <> , I -+--1-+-+ -t--+-: f-+-+-+-+- ! ~""I ~-4~~~1 -+ t I I --+-+-+-+- • ..

]IS eee e... 5... 125.. ]15RI

J."'-,"equency, Hz Figure 97. One-third Octave Spectrum Comparison of Advanced and Original Fan at 115 Peak Aft Angle (305 m Sideline and 158 KN Corrected Thrust - Cutback).

IllUND fflilUIE .(YEl n '.(QUEII!:.

lIS " 11_ FU IL 51 DU IIIG.I .,. () .,. [tiLIleN I,.

El [I,UKN Cl'llleN II IS. CI 'llleN

•

117ft L.

~Il" II SI'" 1I "". U ,nvi II U"U N

,,7'" N

"'17' II !

,....

TI.ca .... _ Cao' .. TI. __ ...-.a ... Il_ ...

eN ••• _ ca •• ' I 1Ir111 ...

a •• 41 C •• nl' ._ e_tl.., ...-" .. _,.

tl lIr'II ...

• • ••• CD •• ".

..... _ ....

I 0 ....... I_n_ .... " ... CD ..... .,.

a .• (1.1,.,

• .....~.I-I\-

'.-

l( I (> ~ .11 )(

lIillla~I'il i

...

~ D

!

El

•

W " ..

0.1 ..

~ r.

e; ~I W I ~

•

,

I I I I I I-f-++ I I I I I I I I I I I I I I I I

• , ...

.1 ll~

·

Frt-quency, Hz One-third Actave Spectrum Comparison of Improved and Original Fan at 50° Figure 98.

Peak FOJ'"ard Angle (305 m Sideline and 228 KN Corrected Thrust - "alteoff).

-- ----._-----_.- -- -----------------------~

ll5 ~ II'.' FTI SL 'iDIJNQ PIlESSUIU l£ vEL vs. rRroU£"fC' II~ D[~ RN~lE El .,. CI'l~N f) .,. C~'""CN 0 I" (2,'IP(N W "II CI'lRtN ,,119 LI 51t .. l' ~'Zl& L" 51878 LI I U'.'I N U714J N U71Z1 .. U11l1 ..

• ..

I ...

" ..

... ..

'" '" ,.. ..

In .. Cl •• r ....

I_a ro.lI_

Ceo' .. T""_

...

CCol.,. _ II •• ' Ii

I ... ..c:.U .. _po

, IIrlt, .. 1 4." C'. ItS' __ C_tl ... ) • ...

II •• ' .... 1 4.45 (0.175) • I 0 ............. _., I''''',,,,,.

1.1.11 ... ( ....... Of' .........

_ .. Ilc .. kll ...

a, .. C._OtS'

•

• ..

Fr"qtJenc,! t Hz Figure 99. One-third Octave Spectrum Comparison of Improved and Original Fan at 115 Peak Aft Angle (305 m Sideline and 228 KN Corrected Thrust - TAkeoff).

~ ""*<.... I - =- ~ ;; -' -::- .. ..., '-~- - • • - ,,- - - ,- -' - - ~ . . - ~" ~." - .- ~-~~~~~ .. - - %~ - ~---- ~ -~--~~ .--=-.- .-~~ ~----~.-- '-'--~~-"------ .. tlt(fl'[~ .. O.~[ lC¥tl III " •• 1. r"!tl "'Ii "NGU elf -llfCIN (I CI-IIKH 0 I~' II EI lSi [I·-Lilt" 0 ". lift" 1~' , LI IJr,] II IUH! II Ilt& I HI] II s"el H ,_ N 1 !I'lU1 II II ".'1 --- ,.." fttCI .. N_ .... 1 .........

CU, .. ft ........

(C.I.), _ caa.)

110.

o,.a-c.U'" _po

I a .... c'.nl) __ C_IIM' Dr'I'.!!!!..- 4.4' (o.nll

.n. 0.11 '

•

0 "',a.c., ...

1.'111.71 10 ....... " _" M'4""'U ...

•

_"Mle_ll_ I.to ('.on)

'''''-

• ..

I

I

J

•

•

t ~ II iI

I

iI

I

~ J----+-----+-----+-----+------+----- ---+_._--+-_ ... - --- +-------+---J

I el •• be e. 'al ..... I.' 1&1 1111

An~l,' tQ tnlrt. dr~r~~s Figure 100. Perceived Noise Level Directivity Comparison of Improved and Original Fan (122 m Sideline and 59 KN Corrected Thrust - Low Approach) • ., ,.f '''It l''It II 1I,j111~f If 'fl t I

I ,l' "IoU!. , t I 'Jl

I

I 0 I~. [t ·lJf[N ('\ ~~D (t- ·l~(H 4) 1 ~1.1 (i" tift... .. 1"'11 r 1 lPCN

\ ll1nl; II l't"ttH La 1'.,,40 t" 8''11. N 1l88/~ IN 6'1-.' N _:'.-. , -- ~.- ~- - --

"1

1'1, ca • .,..,," Tw •• . ,.., r •• I"tI •• C •• I .. ", A .....

(Cel." _ II •• , ....

I

Or.,t .. , 3 o.e.~.n ....... ,- • • ., (O.IU' D e: .... CJ'1'ettwc:u .. , I 4 • ., (0.1>" n 0"'1'''' I 0 ~t ... le .... ".' ... I.nll." (0 ....... " ".'.1,.'''.

' ..... .,..

..,." .lo,..,.n.,..

4 I.to cO.on)

•

.0 " .

• II !it ..

...

ii • ..

e

•

..

..

_ ....

,-- -- ------+- ---+ -I f --- ---t-

J

o. ...

1'1 n • •• a Jr.

• I., Au.: I., lul,-r.

dl·~rt·.·~ Figure 101.

Perceived Noise Level Directivity Comparison of lmprovl"'d and Original Fan (122 m Sideline and 88 KN Corrected Thrust - High Approach).

- . ~~ - ~ - ' :- - '- ~ -~- .,s 2 .;. gO !t±,

-~; -;::t'

~'CUYfD NOI5£ llyn .• ~ RN~l( CI'Llth f) I!I ct"lltN lSi C~"LICN I( lSi CI.'LIICN EI lSi

•

3St7t LI lSiSI LI )'101. LI »6111 LI ISI)1! II 1_77 N 1".'17 N 15111.1 N ~- - !

!

M

! .@lles.il'IW

i ~

I

I

•

I .'

~

•

I

•

i

tI, CI ___

•

..... a '", .... 1- CuI .. TI,_

... (Cea.,. _ (I •• '

I J 0.._11 .. _,.

0 OrI.I .. 1 4.41 CO. I'" -_C_tl .. , t1 OrI" .. 1 4.41 CO.I'"

•

0 I ........ , .. , _~.I .... I1_

1.1'".71 (0 ....... ,

. ., .. - _ •• 10_1\_

• 0 • I." (0 ••• ' ...

•

I I I I I I I I I I

• ..

II a. 1~. I •• IU

I H

••

1- 1-

Angle to Inl .. t. degrees Figure 102. Perceived Noise Level Directivity Comparison of Improved and Original Fan (305 m Sideline and 158 KN Corrected Thrust - Cutback).

, , .5 ANGLF.

PlRCllVfU NOISE LEVEL " I .... Fll Sl J'

I

C' .. l.CN E) 15. ce'lRCN $ "~II ce'LatH ..

EI I'. C.'LIft .. 1'.

512'44 II 5ll!'£o LB ",7 •. 1 5118'1 L' ll7sel N ll71711 II ZUUI N U7'1U N ~ Ta, Cl •• r ....

ft., ... 11_1 COl'" ,.., ........

..... a CCol.,. _ lIa.'

...

t.n CO. In) O ... -C.U'" _,.

I OrI,' .. 1 -_I-II .. , 4.41 CO.175) • 0.1,1 .. 1 l!

~

•

... ,_ ....

I I.ltll." CO.OI .... " _" ." .... U ...

_,,1 ..... 1\_

I." C, .•• ) 1 .... - El !

il

I

18!l lll

I ..

~.~ ."

..

I !

I rl ~ ..

I

• • .

•

I ""- •

= t-

,el III IU &I 81 II.

I 1I 14'

••

Angle to Inlet, degrees Figure 103. Perceived Noise Level Directivity Comparison of Improved and Original Fan (305 m Sideline and 228 KN Corrected Thrust - Takeoff).

1S2 ~ ~ - ~ ~ . - -' ~ --- -. .

-l- ... ....-_ _____ ____ • __ _,,~ -_ T • .--- = _ .,.. ... _ .... "' ..... _ ~ _ ~ * ,.- •• _~,. _, - -._,--,-,_ - - - - - - -- - ~~--~-.

III rll !II pr_C£IVED NDI~£ I (VEl VS lHIfU~T

" It"

,.

D(~ "~Ll CI'L~H • ,~.

(I'L~H El I" C"L.CH IS. (VL~H fl I'.

•

•

:- _t

.,... ft, ca .......

PI •• 1 •••• CU' .. fl, ......

... (Cal.', _ II •• ,

Opa.-c.u ... _po

a J OrI., .. 1 4." ('.171' --I_U ... , I 0.'.1 .. 1 I:l 4.41 (1.171'

•

..... 111._11_

...t .........

•• nll.tI (I."".'"

•

I .... "' 4 ... "MI._ll_ 0 1 ... (0,"1' • ~ /---- ~ ..

."

I

~ ~.-~ ~

t;-a,-£*-

~ .-

.. __ •• f i ~

",I

1--- ~

J---

+-

--- ---- f---.

---- • •

I

I -LJ

I la • za. ,U I" 1(0' lB' II' •• IlL '31 118 IJI ''1 'I t I"~ 91

I'. 'J)

Thrust. kN. (1000 Ih) Perceived Noise Level Directivity Comparison of Improved and Original Figure 104.

Fan as a Function of Thrust for 50° at 122 m Sideline.

PUil(I"[D NOJ~E. llvtl V~ 1HIluST -·-·-------l

I . ~-.- o 1 ........ , .... _ ......... u ... 1."11.71 (O."IO.OJ)

o • I........ _ot_ .' ..... Ilo.. 1.10 (o.on'

• • ..

..

---4---.--~---+ :

.. 6. a. Il' lb •

I.'

'" Il7 ~I ilL q'

I. S I 118 'I) I'. I)J

Thrust, kN, (1000 Ib) Perceived Noise Level Comparison of Improved and Original Fan as n Figure 105.

Function of Thrust for 115 at 122 m Sideline.

.~ ... .,...,.~ ~~-. ~ ,., c _, .. _ "" * __ ..... ~~~ .... _;a~~ ...- .... ..,..,.._-'"' ~_.,.,...~, .. _~_-oo-....., ___ ~~."",..-~:..-.-,:,"",,f' ....... __ -. ...... ==.""". ___ ",,_:i;- __ r'~_"'~~ 5-~ ~ .... ~ ">-_ '-"''-'", .. ~'- ",- ... ..c.~. ,- ~ - • s PlKU~ED IIOISE LEVEL Vi THNI' " II .. rTI ....

~ DIll IINJ:I.I I) I~ ct,UCN • I~ Cl!'LKN 101 I~ I!l I. C\'~.CN CI'LIICN '- n'c' ...._

.,.., fttIl CMI .. n, __

... " .... CeN.', _ caa.'

...

o,oao(eU" _,.

a I Drill.' 4 .• ' (I. In' ._c_" .. , 4.41 CO."., II Drl .... '

•

... ,_ ....

.......... _H_

0 '.ltll.7I c ...... It,

•

1 .... - 4 1." (' .• n, ""~'I_'\-

•

."

, !iii I

,-I

~-,-,~- ...

~

•

~",.,. f i

~fJt : • ...

,~. ,44 B' ,II'

•• ••

'"

I~ '! I Il7 ~l 13& '1) '18 '!I ThruRt. kN. (1000 Ihl Perceived Noise Level Comparison of Improved and Original Fan as a Figure 106.

Function of Thrust for 50° at 305 m Sideline.

----, ]., PERCEIVED NOISE LEVEL V~ TH"UST

" It'" rTl 5L

'" DlC ANGLE tl/l.CH _ ["LatH Cel'L.KN o 1511 I.

El 1511 E> I. CI'LIfCN

~ I I I I n. el .. ' ....

.,.., -, .... ta ... ea .... TI ........

... CCol.', .. CIa., I .... -c.n •• _,- 4." co.,n, a Drl" .. l • _I_tl ... , ~ 4.4' CD.nl) DrI,I'" II

•

..... Ile .. _.II ...

0 I I.UIl.7I CO.n".It' "".".,'' .

. _ .. ,,"

.... _ IUNHn_ 4 1." CO.07I' • -- I ~~I- fA ..

..,

I

*> •

• -----

!

" ----

ri