UNCLASSIFIED

AD NUMBERAD378480

CLASSIFICATION CHANGES

TO: unclassified

FROM: confidential

LIMITATION CHANGES

TO:Approved for public release, distributionunlimited

FROM:

Distribution authorized to U.S. Gov't.agencies and their contractors;Administrative/Operational Use; 10 DEC1966. Other requests shall be referred toFederal Aviation Administration,Washington, DC.

AUTHORITYFAA ltr, 10 Oct 1972; FAA ltr, 10 Oct 1972

THIS PAGE IS UNCLASSIFIED

N " 1IM IA ~F-22131 C' DEC 1966~

A:Z- (UNCLASc-IFIED TITLE)

<ONTHLY PIROGRSS REPOTN.L

DEVELOPMENT OF

A SUPERSONIC TR,%ANSPORT

AIRCRAFT ENGINE

PHASE Il-C

i NOVEMBER THROUGH 30' NOVEMBER 1.966

CONTRACT NO. FA-SS-66-8

(Competit~ive Data)

T tot OOcomENr CONTAINS INFORMA4.TION AFFECCIING THENAT ONAL DEVLNSK 0OF THE UNITED STA-FS WITHIN THfFKIPANING OF THE EPIO1NAGE LAWS. TITLE III U S. C.-SECTIONS 7III ANDj 7WA. ITS TRANgMIPSION OP THCRE V EL-T.014 CF rT CON ENIS IN ANY MANNER TOAN UNAUTHORIZED PEMIIYON is raoi0npiDEP Ov LAW.

N' ,, iA5I -~ DIVISION OF UN-TEIM 4`lRCRAr-T P-0Rr0,YATlfl?.

4'-SA..C AMD DEVELPPME~oT CENTER IlGAF-TER IZ YLFp'I 0oo ,1

[1Pratt &Whitney Airi

[I ~~CONTENTS -

SECTION PAGE

ILLUSTRATIONS ...... . .. .. .. .. ... . .. .. . ....

I SUMMARY OV PROGRESS .. .. ........................ . . -

1I PROBLEM REPORT............... . .. .. .. .. . . . ......

II DESCRIPTION OF TECHNICAL PROGRESS . . . . . .. . . . . I-A-1

A. Engine Design. .. .. ......................... . . . I-A-!B. Engine Test. .. .. ............................. . . II-B-1C. Compressor....... .. . . .. . ... . .... .... Il-C-1D. Primary Combustor .. .. .. ................ ......III-D-lE. Turbine. .. .. .......... ......... ......... ......-E-lLIF. Augwentor. .. .. ...... ...... .. .. . . . .III-F-lG. Exhaust System........ . .. .. .. .. .. .. .... . III-G-lH. Controls s..... .. .. .. .. .. .. . . . . I-H-1I. Bearings and Seals . . .... .. .. .. . . . . II-i-

J. Fuels and Lubricants. .. .. . . ..... ....... .III-J-1IK. Inlet System Cotmpatibility.......... . .. .. . . .. I-K-1L. Noise. .. .. ........ ........... . *.. .. . .III-L-i 1

M. Mockups .... .. .. .............. . . . . . . . . I-N-

N. Coordination. .. .. .................. ......... Ill-N-i -

0. Maintainability . . . . . .. .. .. .. *.. . .111-0-1i'P. Value Engineering . . . . ... .. .. .. . . . . I-P-1Q. Configuration Management. .. .. ................. III-Q-lR. Quality Assurance .. . ..... ........ .. ..... III-R-iS. Reliability .. .. .. ................ ......... . II1151

IV AIRLINE CCMMENTS............ .. .. .. .. .. .. .*. IV-l

V STATE-OF-THE-ART. . ... .. .. .......................V-i

r

Pratt & Whitney Aircraft

Pa& PWA FR-2213

ILLUSTRATIONS

HFIGURE PAGE

IiI-A-I Alternate Ist-Stage Turbine Blade CoolingScheme ....... ...................... .... III-A-4

III-B-I JTFI7A-21 Prototype High Pressure CompressorRig Performance ..... ................. .... III-B-14

H III-B-2 Comparison of JTF17A-21 Prototype and Demonstrator

High Pressure Compressor Rig Performance . .#. III-B-15

III-B-3 Inlet Distortion Screen and Strain GageInstrumentation Duct ..... ............... .... III-B-16

III-B-4 Change in Pressur Profiles Because ofSimulated Inlet L-ocortion ............. . . . III-B-17

III-B-5 JTF17 Effect of Inlet Distortion on Fan SurgeLine for Two-Stage Fan Compressor Rig ...... III-B-18

SIII-B-6 Schematic of Fan Inlet Distortion Test. . . . .. . III-B-19

III-B-7 Distortion Screen for 0.6-Scale Fan Rig ..... III-B-20

SIII-B-8 ist-Stage Turbine Blade Damper Modification . . . III-B-21

III-B-9 Increased Strength Outer Fan Shroud .... ....... III-B-22

III-B-10 Reworked Outer Fan Shroud Incorporating Bands forAdditional Stiffness and Dampirg. ........ III-B-23

III-B-lI Revised Turbine Exhaust Section atia Gas GeneratorExhaust Nozzle Contour. ................ III-B-24

III-B-12 Effect of Inlet Distortion on Turbine AverageRadial Temperature Profile. . . .......... III-B-25

III-B-13 Waspaloy Sheet (PWA 1030) Stress Rupture Test vsDesign Curves ........ ............... .... III-B-26

II!-B-14 Waspaloy Sheet (PWA 1030) 0.5% Creep vsDesign Curves (Revised) .... ............. . . III-B-27

III-B-15 IN-100 (PWA 658) Stress Rupture Test vsDesign Curves ..... .................. . III-B-28

III-B-16 TD Nickel (PWA 1035) Stress Rupture Test vsDesign Curves .......... .................. III-B-29

1 III-B-17 TD Nick4l (PWA 1035) 0.5% Creep vs Design Curves. III-B-30

III-B-18 PWA 654 Stress Rupture Test vs Design Curves. III-B-31

III-B-19 Summary of Sulfidation Testing on CandidateMaterials and Coatings ....... ............. . II-B-32

III-B-20 Condition of SST Alloys Following SulfidationTesting . . ................... III-B-33 I

I1n-B-21 Condition of SST Alloys Following Oxidation -

Erosion Testing . .. ............... III-B-34

_- iii

* 1L1 Pratt &Wh~tney Pirtrnrfti'WA FR-2213

-1I1 LLUSVTRJAT1.1NS Cniud

U FIGUJRE -- PAGE

III-B-22 Oxilation -Erosion Test Data Obtained at1800'1" Metal Temperature. .. .. .. ......... . ..... III-B-35

Ill-C-1 Fan Stream Performance, Two-Stage Fan CompressorRig, Build No. 11. .. ... . ........ . .............. I-C-4

III-C-2 Engine Stream Performir-ce, Two-Stage Fan'I'Compressor Rig, Build No. 11. .. .. ............... III-C-5

III-C-3 JTF17 Prototype Compressor R3.g. .. .. ............. III-C-6

III-C-4 JTF17 5ch-Stage Compressor Blade Comparison-Trailing Edge View. .. .. ... . ........ ........... III-C-7

III-C-5 Rotor and Stator D Factor vs Hub-Tip Ratio. ., Ill-C-8

I II-C-6 Rotor and Stator Chord vs Hub-Tip Ratio .. .. ..... III-C-9

III-D-l Orifice Cross Sections of 1st-Stage TurbineJlVane Pressure Instrumentation .. .. ............... III-D-4

III-D-2 1st-Stage Turbine Vane With PressureInstrumentation..................III-D-5

III-D-3 Two Dimension Cold Airflow Diffuser Rig on

D-32 Stand. .. .. ......... . .......... ........... III-D-6

IIIII-D-4 Two Dimensional Experimental Diffuser Test Rig. III-D-7

III-D-5 Two Dimensional Prototype Diffuser Test Rig.. III-D-8

5III-D-6 Engine EX-161-5 Primary Combustor Nozzle FlowTest Results .. .. .. ... ... .. ........ ........... III-D-9

III-D-7 Pri-mary Combustor Fuel Nozzle and S Lpport

Details from Engine FX-161-5. .. .. .............. III-D-101:III-D-8 Primary Combustor Fuel Nozzles and Seals fromEngine EX-161-5 .. . . .. .. .. . . .... I-D-1l

III-E-l JT=L7 Is t-Stage Turbine Blade - 2 Piece,P&WA Thermal Skin~m. .. ... ...... ................ III-E-2IIIII-E-2 JTFl7A-20 1st-Stage 4-Wall Turbine B.'sde-I Piece, P&WA Thermal SkiaM . .. . . . ............. E-3

III-E-3 1st-Stage Turbine Blade - Impingement CooledLeading Edge, Trailing Edge Exit, Tip D~ump. . . . III-E-4

I1I-G-1 Boeing Wing/Nacelle Model. .. .. ... ..... . . . . III-G-2

III-G-2 Engine FX-163 Instrumentation (Pressure Taps)IIfor Reverser-Suppressor. .......... 3.11-G-3

III-G-3 Engine FX-163 Instrumentation (Thermoco.,ple) fox,

Reverser-Suppressor . .. . . . . ... .. .. 1--4

iv

7-7 7777__ ___ 7

LA

Ihi[ jjPratt &Whitney Aircraft

PWA FR-2213

[ ILLUSTRATIONS (Continued)

SFIGURE F'AGE

Ill-H-1 JTFl7A-2' Response to PLA Modulation fromMinimum Noise Abatement to Maximum Augmentation

Sat Sea Level .. . . . ..........................

III-H-2 JTFl7A-21 Response to PLA Modulation fromMaximum Augmented to Minimum Noise Abatement. . . III-H-6

III-H-3 JTFl7A-21 Response to Noise Abatement ModeSelector Being Moved Into and Out of NoiseAbatement ............... ................... III-H-7

III-K-I JTF17 Fan Distortion Test - Cruise Distortion

Simulation ....... .................... .... III-K-4

III-K-2 JTF17 Fan Attenuation of Circumferential TotalPressure Distortion ......... ............... III-K-5

III-K-3 JTF17 Fan Attenuation of Circumferential TotalPressure Distortion .... ............... .... III-K-6

IIi-K-4 JTF17 Fan Attenuation of Radial. Total PressureDistortion ....... .................... .... III-K-7

I IIl-K-5 JTF17 Estimated Dynamic Performance, DuctHeater Light. .... ................... .... III-K-8

III-K-6 JTF17 Estimated Dynamic Performance, DuctHeater Fuel Cutoff at Climb Thrust ..... ....... III-K-9

III-K-7 JTFI7 Estimated Dynamic Performance, PrimaryCombustor Fuel Cutoff at Cruise Thrust .... ...... III-K-10

III-L-i Nozzle Noise Suppressor Models ............ .... III-L-3

III-L-2 Nozzle Noise Suppressor Models ............ .... III-L-4

III-L-3 Acoustical Performance of Turbojet Mixing NozzleModel - 4-Lobe-50% Penetration - Short ..... ...... III-L-5

III-L-4 Acoustical Performance of Turbojet Mixing NozzleModel - 4-Lobe-50. Penetration - Long .... ...... III-L-6

III-L-5 Fan Noise Absorption Linert - Existing Design . III-L-7

"III-L-6 Fan Noise Absorption Liners - Growth Version. . . III-L-8III-L-7 Spectral Analysis of Jet Noise With "Pure

Tone" Impressed ...... ................... ...-. L-9

III-R-I JTFI7 2nd-Stage Turbine Blade Wail ThicknessMeasurements. ................... III-R-2

-

•] <'; Pratt WVhitney Aircraft -

NWA FR-2213

SECTION ISUMMARY OF PROGRESS

1 Testing bf tile prototype configuration high compressor in the high

Scompressor rig has resulted in performance which either meets or exceedsthe airflow,-efficiency, and pressure ratio requirements for the JTFl7A-21 7production rating. This configuration is being assembled into engine FX-163and is ischeduled for sea level maximum thrust calibration testing early inDecember.

0 Engine FX-161 has completed a program at simulated cruise conditionsw -th-inlet distortion representative of the maximum level for the' intended

Sinstallations. The engine did not encounter stall or surge and no measurableeffect on engine performance was noted. Engine disassembly to date hasshown no parts discrepancies and the engine is scheduled to be rassembled_

for sea level testing in December.

I * Testing :of the 0.6-scale fan rig with inlet distortion in excess of

both Boeing and Lockheed cruise distortion has shown no effect on theengine-side surge line aiA only a 3% loss in surge margin on the ductside. The gas generator aid duct sides attenuated the distortion morethan predicted, This rig has been reassembled with redesigned Ist -nd2nd-stage bladee ,%nd is scheduled for testing early in December.

The Source Selection Council vP-ited FRDOC on 10 November for JTF17Iengiae discussions. The FAA Supplementary Engine Evaluation Task Force [visited FRDC on 16 and 17 Novembe to review the SST engine and rigtest results since Lhe SST evaluation team visit the week of 19 September.Numerous meetings were held with Boeing and lockheed personnel in a con--

tinuing effort to keep engine/airframe coordination current.

-- ~ .--.

E u ] CONFIDENTAL Praft&Whitnev Aircraft

,4A FR-2213

jiti The following significant achievements have been made on the JTFI7A-20

if Bexperimental engines:

__---7 i•Total engine time 137.89 hours1 Total duct heating time 26.20 hours*

Time at 2000OF and above 52.15 hours

Time at 2200OF and above 14.27 hoursat cruise (M 2.7, 65,000 ft)I Heated inlet time 32.34 hours

Time at cruise conditions 28.59 hours(Mach 2.7, 65,000 ft)

*This time was erroneously reported as 35.00 hours in the summaryof last month's report.

i1-2

!] ~COHFIDEM'IAL

IT-5 2 E-I-

=aPratt &Whitney AircraftPWA FR-2213

SECTION 11PROBLEM REPORT

Engine FX-161-5 completed 7.52 hours of testing in the altitude test

stand, and subsequent disassembly has shown no further problem in the

areas found after build No. 4 reported in last month's report (NWA FR-2156).

Pevisions incorporated and tested during this build of engine FX-161 wereoo

[ as follows:

i. Harder insert bushing- at the duct heater support bosses

and wear-resistant coating on the support pins

2. Inner duct heater quarter paael supports with heavier tie

strips and rivets replacing spot welds

3. Addition of sharp leading edge extensions to the main

- . !diffusei case struts to reduce individual peak temperatures

171 f!4. Reoperation of the aft end of the film-cooled transition Iduct to provide additional cooling air over the turbine

F ivane ID platform

5. Reoperation of the 1st-stage turbine vane cooling air

I-. tube ID by installing a bleed slot to purge the vane 1ID platform support cavity.

A revision to seal the riveted rear flange of the No. i seal support

was incorporated in engine FX-161-5. This change eliminated the coking

r •problem experienced on engine FX-161-4, resulting from leakage past the

rivetE into the labyrinth seal pressure supply area of the intermediate

case. Disassembly of the engine after 7.52 hours of running indicated iFno leaks at this flange. This revision has been incorporated in both

No. I and No. 2 seal supports on engine FX-163-2.S~ITablocks to replace indented lockwashers for the duct heater Zone II

fuel nozzle covers will be evaluated on the next build of engine FX-161.

•-$II-I

P 1 Pratt& Whitnav •__aKA FR-2213

SECTION IIIDESCRIPTION OF TECHNICAL PROGRESS

A. ENGINE DESIGN

I. Fan

Layouts of a new 0.6-scale fan rig for Phase I1 were completed.

I £ 2. Compressor

The prototype compressor design layouts were completed. This com-

F pressor has the same aerodynamic design as the compressor rig that demon-

strated performance capability above specification requirements in early

November.

Design layout work continues to define the incorporation of this proto-

jjj type compressor into a third compressor rig for use during Phase III

development.

3. Primary Combustor

SThe design of the primary combustor is 907° complete. The design

concept is the same as described in the Phase III proposal.

VI anA convectively cooled transition duct design has been completed and

an alternative film-cooled transition duct design has been initiated. n

The engine diffuser case design is also 907. comnlo.te.

4. Duct Heater

Design layouts of the combuster, duct heater support case, and rear

mount case are complete.

i Design is continuing on the duct diffuser case, inner and outer liners,

and Zones I and II fuel injection systems.

S5. Turbine

An alternative hat-stage blade illustrated in figure III-A-l has

been designed.

The design of the turbine high spool rig for Phase III is continuing.

I ~ ~I II-A- l" '

•;•-_-- _ - • • •__..-.•-S.-_-----;.•- • • •• •-i• -t ---•---.-

ni Pratt& WhItey Pircraft

,II'WA '• 421. .

6. Shafts, Bearings, and Seals

SThe design layouts for the No. 1 and 2 bearing compartments are com-

plete. The No. 3 compartment design is approximately 90% complete. The

No. 4 compartment is approximately 70% complete.

7. Accessory Drives

Design layouts of the Boeing power takeoff hydraulic pump drive I

gearbox and of the Lockheed engine-driven compressor gearbox were com- .

pleted. Design layout work for the engine accessory gearbox continues.

1 This gearbox housing design is being coordinated with the engiae hydraulic

and fuel pumps and the quick-disconnect feature for the engine unitized fuelS~control.

Accessory and starting drive bevel gear sets located within the No. 1

and 2 bearing compartments have been completely designed.

8. Fuel System

j The design layout for plumbing through the struts, utilizing strut

covers with integral fittings, was completed during this report period.

Revised plumbing layouts were provided in this period for the experi-

mental engine gas generator to accommodate the incorporation of the proto-

type compressor in the experimental engine.

9. Control System

Coordination with vendors on the unitized fuel control, main fuel [pump, and hydraulic pump was continued. A mockup of the revised con- rnection between the fuel control base plate and the accessory drive

gearbox was constructed to evaluate the accessibility of the connection

for easier gearbox replacement. A preliminary layout of the unitized

fuel control-gearbox mounting system, and a stress analysis of this system

were completed in this period.

The design layouts of the combined drain valves and the fuel-oil cooler

were completed in this period. Design work on the aerodynamic brake

_I actuator and the duct nozzle feedback system was continued. Design of the

[A reverser interlock mechanism and the coupling to the unitized fuel control

is progressing.

III-A-2

_77Fli -MfIa0ý

tLl

1A

10. Lubrication System

The design layout of the secondary (oil pump-cabin air c•mprt,U-1 drive gearbox for the JTFITA-21L engine was completed in this period.

11. Reverser-Suppressor

The design layouts of the JTF17A-21 reverser-suppressor and duct

heater variable nozzle are continuing.

I|

I

I -:1•

ROii2 n

UN

IM

itt

AL Dp" &%Afl.*&~- -

%M Wv I l PyALr~r-PWA YR-2213

II+

+

3V

4 4

J58 Blade withI Same Cooling Scheme

S•Figure III-A-I, Alternate Ist-Stage Turbine ED 17094I Blade Cooling Scheme

111 -4

S~~CONFIDNTAL

*.c c . - . .. • = .

'I

aI Prat &Wbltrey Arrf

I B. ENGINE TEST

November Phase I1-CTimey hours Time, hours

Engine FX-161 FX-162 FX-163 PA-161 FX-162 FX-163 Total

Total 7.52 10 87.41 47.55 2.93 137.89Heated Inlet 7.52 :430.27 2.07 0 32.34

JCruise Gondition 5.90 26.94 1.65 0 28.59

Duct Heater

Testal 1. 20 .0 20.43 5.77 0 26.20

IICruise Codto 1.0> 617 0 0 61

if2300*F and above 4.0 2.0.159 6 0.38 040.97

2200rV and fulsaloe rig8 has4 5eslte in1 2efr4nehteihr8et

totyp ilutaF byd thev dat presnte in3 fiur I0-.9

Testir g ilusrae the potyecnimprovedn surge chapracessristic e reaige

tor txedshe airlow, hfiighcy compressore ai requirements anfuid o.r

the and 7-2 Nrouc.o 37A0cpressog. This ecollntefigrmatinc is being

assebledintoengieP1-163 and is scheduled for sea level maximum

Uthrust clbaintesting early in December.

WIP ~IWhIWmey AircrartPWA FR-2213

LI Engine FX-161 has completed a program at simulated cruise conditions

with inlet distortion representative of the maximum leve.j for the intended

installations. The inlet distortion was simulated by mounting a gradated

density screen, figure I11-B-3, eight feet ahead of the engine. The over-

U all inlet distortion, (Max PT2 Min PT2 )/Average PT2' was 8.97. with

Kd2 =374, as shown in figure I11-B-4. The engine did not encounter stallflor surge and no measurable effect on engine performance was noted. Engine

disassembly to date has shown no parts discrepancies and the engine is

scheduled to be reassembled for sea level testing in December.

51 AL Testing of the 0.6-scale fan rig with inlet distortion in excess ofI D both Boeing and Lockheed cruise distortion has shown no effect on the

ngn-side surge line and only a 37. loss in surge margin on the duct£1side. See figure I11-B-5. The gas generator and duct sides attenuatedthe distortion more than predicted. Aac~mtc ftets cnfgrto

IUin shown in figure 111-B-6, and the distortion screen used is shown in

figure I11-B-7. This rig has been reassembled with redesigned 1st- and

2nd-stage blades and is scheduled for testing early in December,

1. Engine FX-16

a. Rabuild

fi The fifth build of the engine was completed and it was delivered to the

altitude test facility on 14 November for evaluation of inlet effects and

altitude performanze. The following is a sunmary of features included in

U this build:FAN -Same as build No.* 5 configuration; strain gage instru-

mentation on both stages.

1IMRMHEDIATE CASE -Drooped fan discharge splitter as part of

U the build No. 5 fan; pressure instrumentation added to strutleading edges on main engine side.

I ~H1G1t COHPASSOR -Same as build No.. 5 configuration.

KAIN DIPIUSER -Sharp leading edge extensions on struts;Iinstrunmetation bosses for compre-sor discharge traverses;

heAtvy e4gh% splitter from previous builds.

111-B-2

IWA FR-2213

PRW COMBUSTOR - Original primary combustor with film-cooled

transition duct with the trailing edge of the inner duct modifiedSto provide additional cooling air for the TD platform of the

lst-stage turbine vanes.

STURBINE - Additional cooling airflow to the Ist-stage turbine

vane ID platform support cavity as used on engine FX-163-1. TheSdamper weights were reoperated to assure uniform loading of the

dampef weight on the 1st-stage turbine blade platforms. The

[j round hole in the damper weights was changed to a triangular shape

hole to reduce frictional contact area. The width of the yokes,

which retained the damper weights, was also decreased to reduce

the possibility of binding on the tips of the adjacent dampers.

(See figure III-B- 8.)

T[RBINE EXHAUST SECTION - Primary nozzle resized for speedmatching with the build No. 5 fan.

FAN DIFFUSER - Modified to accommodate the compressor dia-

. charge traverse probes.

DUCT HEATER - Hard sleeves of L-605 material installed in the

support pin holes and tungsten carbide hard faced pins;

octagonas nozzle from previous builds.

SC(PONENTS - Manual control of main engine fuel flow, ductheater fuel flow, duct heater nozzle, and vane angles.

GEARBOXES - Same as build No. 4.

b. Test Program

3 The engine was installed in the altitude test stand with sa inlet

distortion screen to simulate the Boeing distortion pattern. The screen

and the duct for fan strain gage instrumentation are shown in reference

figure 11I-B-3. Two runs were made for evaluation of airflow patterns.

The distortion screen was rotated between the runs to effectively double

the instrumentation. The overall inlet di-stortion, (Max PT2 -

Miin PT2)/Average PT2, was 8.9% with Kd2 = 374, reference figure Ill-B-4.

The inlet distortion screen was removed and a baseline calibration was

run without distortion. For further evaluation of the data, see para-

graph 3-B, Performance. Fan blade strain gage readings were less than

10,000 psi for: all testing.IU-1-3

PWA FR-2213

H ~The altitude program included testing ait cruise conditions with

heated fuel and lubricant. A fuel temperature of 3600 F was measured

at the inlet to the nozzle support. This would result in a maximum fuel

temperature of 370 F at the fuel nozzle. Refer to section TII-iD for

additional information on the fuel nozzles after test and sec~tion III-J

U for the fuel quality. During the engine testing at cruisc conditions,

ri a maximum lubricant temperature of 360*F out of the No. 4 compnrtment

was measured.

The engine was removed from the altitude test stand and delivered to

the assembly floor for inspection and to incorporate changes for a seaIsvel performance program.

c. Disassembly laspection

ii The engine has been disassembled and inspection of all parts is in

progress. No obvious parts discrepancies were seen during disassembly.

II2. Engine FX-163

Engine disassembly was completed on 19 October. All parts wereif inspected and reviewed in preparation for rebuild. Reference October

Progress Report No. 16, NWA FR..2156. Assembly was initiated, replacing

~j.damaged parts-- onlyo but this plan was revised to incorporate the high

U pressure compressor to the configuration that demonstrated performance

capability above sped.fir~ation requirements in the compressor rig. (See

paragraph C, Compressor.'.

A general description of the major features being incorporated in

engine FX-163-2 follows:

FAN -Same as rig build No. 5 configuration with a new, iL~creased

strength outer fan shroud. The increased strength outer shroud,

figure T11-S-9, eliminates the need for dampeaing bands required

II on the original design as shown in figure III-B-10.

1N1TJNEMIA CASE -Same as on engine FX-163-1 which >.,cluded

the drooped splitter as part of the protutype fan rxe!&~

HIGM CC!0ESSOft - Same as prototype configuration featuring

increased chord length vanes and-~blades incorporating new vane

and case assemblies, new inlet guide vanes, steel blades from

ii. _

111-W4

IWA FR-2213

the compressor rig, (AMS 5616 Greek Ascoloy), No. 3 through 6

I disks from the compressor rig, and No. 7 and 8 disks from

' hI engine FX-163-l.

This configuration is three inches longer in overall length

Sbecause of the increased chord length ot the vanes and blades

and includes a 5th-stage starter bleed system.

I The inlet guide vanes are variable and the 3rd- and 7th-stage

vanes are fixed. See section I11-C for a detailed description.

II MAIN DIFFUSER - No changes from engine FX-163-1 that had sharp

j rLE extensions on the diffuser struts. For this build, four

IFI starter bleed valves were mounted on the diffuer outer wall to

provide supplemental bleed air capacity for starting only.iI,-IMARY COMBUSTOR - Same as engine FX-163-1 which included

heavyweight splitter, standard combustor unit, and film-cooledP !transition ducts. The outer transition duct is two-piece,

1 and the inner transition duct provides a flow of cooling air

over the ID platform of the 1st-stage turbine vanes.

TURBINE - Same configuration as on engine FX-163-1. First-stage

blades are of the nongated airfoil type from engine FX-16:-5.

TURBINE EXHAUST SECTION - The OD wall contour on the exhaust

Scase is revised to reduce the back pressure on the turbine,

and the exhaust nozzle contour is revised to reduce base area

drag losses. See figure III-B-ll.

FAN DIVFUSER SECTION - Same as engine FX-163-1, original con-

.0 figuration.

DUCT HEATER - Same as engine FX-163-1, original combustorU section wich round, balanced flap variable nozzle.

CONONENTS - Same &as engine FX-163-1 which included the automaticduct heater control.

GEARBOXES - Same as engine FX-163-1.

III-B-51L~~7111 -~

k __

i F Pratt&Whitney Aircraft

nIMf *OLLL.Jt~

PLUMBING - Same as engine FX-163-1 except for changes necessary

to accommodate the prototype compressor, the starter bleed system,

and the revised turbine exhaust case.

REVERSER-SUPPRESSOR - Unit No. 2, identical to unit No. 1, will

be installed at test after completion of the initial performance

test program.

i Completion of engine assembly and delivery to A-4 sea level test

stand is schrluied for early December. The primary objective of this

tEit is to demonstrate increased sea level static engine performance.

3. Performance

During November, engint FX-161-5 was tested at supersonic cruise

conditions with and without simulated inlet distortion to demonstrate

engine-inlet compatibility. The inlet distortion was simulated by

mounting a g--adated density screen, reference figure III-B-3, eight

feet ahead of the engine. The pressure distortion generated by this

screen,(Max PT2 - Min PT2)/Average PT2 was 3.9% with a distortion factor,

Kd2, of 374, The differences in measured engine parameters, with and

without distortion, are compared in table 11-B-I and show no effect on

performance resulting from the distortion screen. The turbine inlet

temperature radial profile with distortion was uL~chcnged from the profile

without distorticn as shown in figure III-B-12. Peak 1st-stage turbine

vane temperature with distortion was within 150 ol the peak temperature

without distortion.

Table Ill-B-1. Effect of DistortionMach 2.7, 65,000 ft

1 Parameter Change Due toDistortion

N1/JJ + 0.05%

N2<1 3 + 0.4%.

P'T4/PT2 + 1.07.

(PT4 - PT5)/PT4 No change

TT5 . 12 F

SPTT/PT2 + 1.2.W 0T2/6T2 + 0. 1%1

-7:!

* • Pratt &Whitney AircraftMA rR-2213

S4 , iacer~ais and Fabrication

SLong-time stress rupture and creep rz, !ure testing is nearing coin-

pletion on candidate SST materials. Stress rupture testing has been

completed on Hastelloy X Sheet (AMS 5536). This curve was submitted in

prior reports and will not be repeated in this or subsequent reports.

A list of candidate materials, the proposed applications in the SST

design, and the limiting creep and stress design criteria are tabulated

as follows:

Material Application Limiting Design Criteria High TimeCreep, Creep and Stress Specimen,

R upture Temp Range, hr0

Astroloy (PWA 1013) Disks 0.1 1200-1400 4745

Waspaloy Sheet Cases 0.5 1200-1600 5113(NWA 1030)

Waspaloy Forgings Disks, Shafts, 0.1 1100-1300 3587(PWA 1016) Hubs

Inco 718 Sheet Ducts, Cases 0.5 1000-i200 3220

_ I (PWA 1033)

L-605 Sheet Ducts, Liners 0.5 1400-1800 5203(AMS 5537)

IHastelloy X Sheet Burners, Ducts 0.5 1400-1800 4761(AMS 5536)

IN-100 (TWA 658) Blades, Vanes 1.0 1400-1800 4950

Inco 625 Duct, Liners 1200-1500 998

TD Nickel Vanes 1700-2100 2023I (PWA 1035)

Titanium Blades 0.1 700-900 7541

(PWA 1202) 1 3

PWA 664 Blades 1.0 1400-1800 3060

A-110 (AMS 4910) Cases 0.1 700-900 ',52

The results of lon•g-time stress rupture and creep testine of the

j following current alloys are plotted in figures IIi-B- 13 through III-B-18.

"III-B-7

A Pratt& Whitney Aircraft 1PWA PR4-2213

Material Stress Rupture Creep FigureFigure No. No.j

i4 Astroloy (NWA 1013)**A O Waspaloy Sheet (NWA 1030) III-B-13 III-B-14

Waspaloy Forgings (NWA 1016)**

1-718 Sheet (NWA 1033)L-605 Sheet (AMS 5537)1I ~Hastelloy X Shee~t (AMS 5536)**

Zý-IIN-100 (NWA 658) I11-B-iS10 ~Inco 625**TD Nickel (NWA 1035) III-B-16 III-B-17N PA 664 111-B- 18*

~~ti Titanium (1NA 1202)**[

*Tcsting completed. Curves in previous reports.

.4 5. Sulfidation and Oxidation - Erosion Testing

Uidation testing is being continued on the most promising candidate

SST materials and coatings. The following is a summary of sulfidation

testing conducted at accelerated test conditions of 1.0 ppm NaC1 contentI;in air, maximum sulfur content allowed (0.3%.) by NWA fuel specifications

4 522 and specimen metal temperature of 1800*F.

SIMaterial CoAting Protect-on,hr

NWA 1035 (TD Nickel) NWA 62 1250 F

A rA 6 64 NWA 47 1350KrPWA 664 IWA 64 1850NWA 658 (IN-100) NWA 64 1150I I A 658 (IN-100) *2500-

NWA 1035 (TD Nickel) * 100**

*NWA number has not been assignedI ~**Testing of specim~ens still in process,

F A graphic representatlon of these data is shown Ln figure 111-B~-19

4 ~The sulfidaticn test.--g and retesting results preaently in process-onthe following -iaterials and coatings are:(!) NWA 1035 (TD Nickel) coatedwith the newly Ceveloped coating showed encellent sulfidation protection

after 1100 hourr& of tc 1 testing, (2) NWA 658 (YN-IQO) coated with thenel~developed coating anowed excellent sulfidutionpoeto fe

1100 hours of total testing, (2) WA 658 (N-100) and NWA 664 coated with

U1 EWA 64 ithowed excellenr sulfidation-vrotection af~ter 550 hours of 'retesting.~

The conditions of tI~ese specimens are shown in figure III-B-20.

-B-3L

Pratt &Whkmny Aircraft_Rýý7W Egfl-2213

U'Y- in Long-time oxidation-erosion testing of candidate SST materials andi

coatings (also other materials-and coatings.for comparigon) is con~tinuinv at 1800OF specimen-metal temperature- The resul-ts for the following

materkal and coatings being tested a-re: (1) NWA 658 (IN-l_00)- coated withPWA 58, And PWA 664ý coated with NWA 47 showed excellent oxidation-eresion

protection after 1300 hours of -testing; (2) P~A 657 (SH-302) coated with] NA 45 showed excellent protection after 500- hours -of testing; (3) NWA 1035( TD Nickel) coated with NWA 62 showed exceilent-protection after 350 hoursof testing; (4) FWA--664 and NWA 658 boated wdith PWA 64 showed excellentI]protection after- 450 hours of testing. The conditions o~f these specimensafter the- testing time indicated are shown in figure flI-B-21. Graphicpresentation of these data are shown in f'I~gure III-B-22.

J6 -AvneiMtra Deeomn (Rlae Aehooy

Thrie- N'co-Atingor -66iio h8as ubderong of ficiall o-ersionate stN g at8

Cf o a f dCoatedt-es&t bars .look excellent after 504O hours of rig testing.

II-b. Astrolpo Sheet- Program -

In-addition to the production quantity-of-Allvac sheet, which is

being--evaluated, a siaL414r quniyo eaten stainless sheet is being

-evaluated, The weld stu-dies, outlined in last month's pro~gress report,

are con~tinuing. A seconidafterburner duct wilIl be fabricated from the-Eastern 4ta nlss imaterial.--jpc. UX-1500 (Turbine Disk-Alloy)

The L_ý-1500 inigot' has -beden given a se±cond homogenization cycle toclean-up an undesir~able- segre_8#te found-in the -subacale forgings. This

_paeor sgrgate: cbud- not- be reTmovedtrihea treatmdent and negated-Ithe obtaining of ipt -imum- inechdtiical propertieit.. -Additional subscale

51' -forgings Are to -be. made -por t0 or .na 11 -41 e disk.

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I LI d. Extrusion Forgings of INi-100 and Modified Mar M-200 Alloys-(Turbine Blade Application)

-A Evaluation of the effect of prior work on the mechanical properties

of iwrought INI-.lO is being made. Further work is being performed to

Sanalyze grain boundary condition and the effect on ductility.

e. Advancee Titanium

b Testing is continuing on forged pancake specimens of Ti 6-21-4-2,

Ti 6-2-4-2- + 1.5 Si, IMI 679 (NWA 1205), and Hylite 60 for comparisonr! with NWA 1202 properties.

7. Engine Accessories

a. Compressor Bleed Valves

The JTFl7A-20 compressor bleed valves are air actuated, springloaded poppet valves that bleed compressor discharge air to the~ cavity

surrounding the gas generator. These valves were used during engineII testing to vary the percent of bleed airflow during starting and operation

of the engines. Operation of the valves during sea level and altitude

running of the engines was flawless.

j ~Calibration and environmental bench tests were performed. The comn-

ponent calibration test consisted of measurements for actuation pressure,

piston ring leakage, and valve seat leakage. The environmental test

consistec1 off cycling the vaive for 5000 cycles in a 1050*F environment.II Two valves were subjected to the environmental test. Pre- and post-test

calibration dat6 of the valv*us were in good agreement, and Within the

limits of the component calibratior schedule as illustrated in table III-B-2.

The detail parts of each valve* were in excellent condition.

'I The total test timie accumulated is as follows:

Engine - 218.19 hours

Calibration - 28.33 hours

Environmental 9 7.32 ho~ura

Total - 333.84 bours

II~~I _ _ _ -_ B-_ 1__

1111Pratt &Wthtney Aircraft

The features of these bleed valves that have demonstrated successful

operation during this phase have been employed in the Phase III design

of the JTFl7 engi"e valves.

Table III-B-2. Xleed Valves-

Parameter Pretest Post-Test

Pressure-to-Actuate':

Open to close 1.5 psi 1.5 psi

Close to open 0.5 psi 0,5 psi

Pistoa Ring Leakage 0 pph 0 pph

Valve Seat Leakage 0 pph 0 pph -

b. Actuators

(~~1 Linear hydraulic actuators wa_.e employed on the JTFi7A-20 engine

for operation of the high ,ompressor variable vanes, the variable area

e duct heater exhatist nozzle, and the reverser-suppressor clamshells.

These actuators employ enaIne fuel as the working fluid and incorporate

design featurus that were developed for actuators used on the high

performance JS8 engine. Operation of the a-etuators durinn engine

testing at sea level and altitude conditions was very satisftctory.

Calibration and environmental bench tests were performed. The corm-

Sponent calibration test consisted of measu.ements for actuation prtssure,

cooling flow rate, length of stroke, and dynamic seal leakage

A The environmental test was a 75-hour decign verification test that

was conducted under the following conditLins:

1. Ambient tei .-rature - 900PF

2. Fuee rnlet tempera, .re - 350°F for 4 hoursr50OF for 1 hour

- cycle repeated every 5 hours:

3. Actuator stroke - 1 in. (nominal full stroke - 4.2 in.)

4. Actuator cycle rate - 90 cycles per ciinute

5. Open to close pressure - 1500 to 2000 psi V

6. Overboard drain leakage - 5 cc/rein (max).

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Mal a Pre- and post-environmental test calibration data were in excellentagreement as illustrated in table III-3-3, and the actuator parts showed

little or no distress.

The total test time accumulated is as follows:

Engine - 1097.14 hoursSCalibration - 89.09 hours

Environmental hours-_ 62 hours

Total - 1282.47 hours

Table III-B-3. Actuators

Parameter CalibrationPretest os-Teut

fl Pressure to Actuate:4Open to close 60 psi- 20 psi*A Close to open 3 5 psi 10 psi*

Overboard Drain Leakage 0 cc/min 1.5 cc/minStroke 4.369 in. 4.352 in.

- Cooling Flow Rate 215 pph 216 pph' **onrmal range after environmental cycling -as determinedSfrom 358 experience.

4c. Lines and Fittings

The JTWI7A-20 lines and fittings that have been used in the Phase 11-C-engine program are fabricated ftrm stainless steel tubing (AIWSIt_ 347)

and Inconel materials. The fittings are either bra.-d and integral tubeAN type that use soft nickel conical gaskets or the threaded or bolted

iflange type that use metal K-Meats.

These lines and fittings, which were designed using the criteriaSand experience from the development of the J58 engine, have provided

excellent service during- engine and rig operation. Only one tubefailure, induced by preload as a result of- impropert component installation,

has occurred during all engine testing.

An investigation using titanium tubing-has been conducted withcomzercially pure tubing, grade A-40, ard the ally tubing1 3&-2.5V.These materials were successfully upst to the standard integral ferrule

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___ configuration for AISI type 347 material (PWA 770). Vibratory fatigue

V tests under simulated engine environmental conditions were conductedL i on the integral ferrule and on the welded tube'-connector configurations.In addition, salt water corrosion tests, bend tests, and brazing tests

(for brazing tubing standoff supports to the tubing) have been con-ducted.

From the testing of titanium tubing it is concluded that the tubing

alloy 3Al-2.1 5V is comparable to or better than the AISI type 347 tubing3 material and will provide a suitable tubing material. Because of low

-vibratory fatigue strength, the coummercially pure titanium tubing is

unsatisfactory. The investigation of the titanium alloy is being

continued.

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j 1.1C. COMPRESSOR

f 1. 0.6-Scale Pan Rig

November Phase II-C TotalTest Time 103.00 hours 562 hours

j Build No. 11 of the 0,6-scale fan rig completed its overall per-

formance calibration and was tested with inlet distortion during November.

Build No. 11 incorporated the build No. 5 configuration fan with a

7-degree ove-rcamber of the 1st-stage vane root, It was intended thatl~t overcamber of the 1st-stage vane rooc would improve the 1st-stage rootmatch and increase the root stall margin. Although the engine side surge

U line was slig~ntly improved, the overall performance of build No. 11 wasessentially that of build No. 5 with no change resulting from the 1st-stage

- wie-overcamber. The performance of build ýNo. 11 -relative to the build No. 5- -- s urge line is illustrated by, ý.e compressor mapt, of figures Ill-C-I and

7, -A4

A Fan inlet distortion testing was accomplished through the use of

r screerss instaiied in a straight cylindrical-test-section approximately

one duct diamet.er upstream of the 1st-stage blades. A shematic of the

test- conliguration is shown iii reference figure 111-11-6. Twonty-four totallo prsuerksusram -of the fan- i-alet vere -used- to define the inlet pattern.

A similar set of rakes at the fan-discharge wore used to determine thet

ta -atnuation tharacteristles, of the fan.

-The distortior. screen used is shown -in reference figure 111- -7, The dis-I torvriou factor (K ) of the tests was 15%above the maximum anticipated for 4

d.2IMach 2-.7 cruise operation. As shotun in reference figure IIl-'3-5, the surge

H margin on the-engine side was unaffected by the simulated distortion andlowered only 3% on the duct aide. The fan attenuated the distortion in pass-

-Fin& through the two fan atages. No indication of flow inatebility or stallwas encou~ntered. Details cot the distortion test analysis are containedj

Sin Inlet System Compatibility, section I1I-K, >~f this month's report.

'The redesigned 1st- and 2nd-stoige blades scheduled for resting in

December are now assembled in the 0.6'.acale fan rig. The purpose of the

build No. 12 crifiguration is to romatch the compressor stages for peak

92 performance. The 1st-stage blades are a revised version of the buildNo. 10 rotor that failed bec~ause of cracks resultino from rework to

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overc-amber the build No. 7 blades. The 2nd-stage blades inc~orpor-ate a

closed leading edge root and an opened lead-.ng edge from 3"/ span,

Ii relative to the build No. 7. This rotor w'11 match the stages for high

duct pressure ratio and have the capabilicy to operate to lower flows.

Test results from build No. 12 are expccted the 1st week in December.

2. Full-Scale High Compressor Rig

November Phase Il-C Total

Test Time 12.92 hours 114.50 hours

1L1 Rig testing of the prototype JTFl7A-21 compressor was comnpleted early

J in November. The compressor rig, ready for test, is shown in figure III-C-3.

The performance met. or exceeded the airflow, efficiency, and pressure ratio

requirements for the JTFI7A-21 production rating. The excellent performanceH of the prototype compressor is illustrated by the data presented in ref-

erence figure III-B-i. Reference fl'gure III-B-2 illustrates the improved

surge characte5ristica relative to the current Phase 11-C high compressor

requirements and builde No. 3 and No. 5 of the JTFl7A-20 compressor.

if The prototy~pe compressaor is a redesign of the JTF17 high compressor

to improve the -surge margin of the build No. 5 compressor and to meet

- the reauirmaelts of thso JTF17A-.21 engiil The rig tested is aerodynamically

Iidentical to tb-a coinpre~sair p-,zerntfd in cn, 6 September, Phase III proposal

and was tested with :interstiaga bleed and wit!. only the inlet guiee vane

m or chantger, from the eorlier com~pressors. -First,-the camber of blades

and vanes has been modified to flatten the~ star- flow and pressure profiles'ithin -the comxpressor, An exml f h aber change in blades is shown'Cby a comparisou c f the 5th-stage blades of the prototype; compressor and

-of build N4o. 5, as shown in figure, 11l-C" 4 The compressor work has been

rddistributed in the-design to reducee~the loading of rear stages. This

design change is illu-stratod- by plots oi rocor and stator D factor,

j figuire 111-C'-5. -Compressor chords have been selectively increased,

- pimarily in vones. These changes; are showna in figurer III-C- 6 for

rotor and stator..

No further compressor -rig-testing is planned for Phase 11-C The

In major effort will-be directed at engine testing of the compressor.

Although the prototype rig parts vere desigried for rig teating only, the

11i-C-2M_______5i______I

I I ~.Pratt& Whitney Aircraftj NWA FR-2213

disks and blades are adequate for sea level running. As soon as possible,

new vanes and cases will be manufactured to replace the rig 347 stainlessI steel cases with PWA 1010 cases, structurally adequate for engine testing.Engine testing with the improved compressor is expected early in December.

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j flD *PRI1M1ARY COMBUSTCOR

• VFull-Annular Rig Test Time

November Phase II-C

t JTF17 0 5.36 hrSRelated Technology 0 65.8 hr

I {! 1. Engine Combustor Testing

Testing of the primary combustor continued on engine FX-161-5 at

altitude conditions.

Engine FX-161-5 acc-,mulated 3.68 hours of Mach 2.7 and 65,000 feet

£1 Htesting at cruise turbine inlet temperature during November for a, total

of 14.27 hours. Refer to section 1114- for additional significant test• ftimes. The general condition of the primary combustor and the transition

duct was good. This combustor with 87.41 hours is suitable for continued

testing. The primary combustors in engines: FX-162 and FX-163 did not

accumulate any testing during this report period.

1 2. Full-Scale 120-Degree Segment Rig

Testing on the primary combustor was continued to investigate diffuser

splitter configurations for possiblt incorporation in the JTF17 engine.

SThe configuratians were ter7ed ,or a total of 17.53 hours in November.

Th Th• tests included blockage of the ID and OD diffuser passages to reduce

the expansion area ratios, iemoval of the center wedge of the origi..al

I•- splitter, and forward extension of the orginal splitter.

l •3. PT Instrumentation Caiibration

A program has continued on the vane (-scade rig to develop improve-

mente in the combustor discharge pressure in4trument-ation. The original

pressure instrumentation in the leading edge of the first turbine vanes

.:Jii Uwas a sharp edge crifice as shown in view A of figure III-D-l, TheA October Report, FA FE-2156, reviewed cascade rig tests showing that -

this configuration. was 1.37 to 4.0% low in total pressure readings.

VI-D :

[1 Pratt & Whitney Aircraft) PWA FR-2213

By adding a chamfer to the e.dge of the orifice as shown in view B

of figure IIA-D-1, the readings improved on one of the vanes but not

on the other. Yaw probe traverses showed air angles fro. i n es on one

j ~ side of the center to 2 degrees on the other side. Interchanging vane

[J positions showed that the air angle was the determining factor for a

reduction in measurement error. The chamfer was an improvement, but it

(ti is s't.11. sensitive to the air approach angle.

A tube extension on the leading edge, as shown in view C of figure .II-D-l,

L -increased the air acceptance angle. Instrumentation laboratory tests indi-

cated that the acceptance angle for 1% error was approximately doubled.

Ill Cascade rig tests at 23000F and 107 psia showed pressure readings with

zero to 11. error. A set of pressure instrumented vanes has been reoperated

as shown in figure III-D-2 for engine testing.

4. Two-Dimensional Diffuser Rig

A two-dimensional diffuser rig was assembled and used for JTF17

primary combustor diffuser airflow studies. i"he rig, F-33449, is shown

J- on D-32 stand in figure III-D-3. Tests were conducted on the experimental

A-20 and the prototype A-21 diffusers as shown in figures III-n-4 and

III-D-5, respectively. A summary of the test results is shown in

table III-D-l.

Table III-D-l. Two-Dimensional Diffuser Tests

Diffuser Configuration AP/P, %

L Experimental - No struts 2.73

j Prototype - No struts 2.49

Experimental - Round nose struts 2.25Prototype - Struts 3.42

- Experimental - Sharp leading edge struts 2.16

Experimental - Sharp struts - forward splitter 2.49extension

U |jTesting is continuing on this diffuser rig to establish changes to

Sthe prototype diffuser strut and splitter design.

III-D-2

f [j! Pratt &Whitney AircraftPWA FR-2213

5. Fuel Nozzles

-I U Post-test flow checks are in process for the Zone I duct heater

nozzles and they are completed for the primary combustor nozzles from

fri rengine FX-161-5. The pretest and post-test flow checks do not indicate

any flow shifts as shown by the comparison at three flow levels in fig-

ure III-D-6. These nozzles have 87.41 hours of JTF17 engine time and over

3 hours of operation with 370 0F fuel temper. ture at the nozzle. Refer to

"engine section III-B (FX-161 test program). Figure III-D-7 shows the

{j condition, of details from a nozzle and support assembly after test.

j Nozzle tips and seals, from two assemblies that were instrumented to

measure the fuel temperature at the inlet fittings, are shown in fig-

ure III-D-8.

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I E. TURBINE

i. Thermodynamic Cascade RigNovember Phase lI-C Total

Test Time, hours 23.91 445.95

The test results from the P&WA two-piece Thermal SkinTH lst-stage

blade (figure Ill-E-l) equaled the predicted values of the integral cast

P&WA Thermal SkinTM blade (figure I11-E-2). Modification of the removable

internal airfoil to reduce thermal gradients across the blade is now in

process and testing will continue next month. The P&WA two-piece Thermal

SkinTH 1st-stage blade is considered to be a significant improvement over

the integral cast P&WA Thermal SkinTM blade with respect to casting, fab-

rication, and inspection.

2. Related Technology

Testing of the impingement leading edge lst-stage blade continued

this month. The test specimen, which was fabricated using production

tooling, incorporated 0.020-inch diameter leading edge impingement holes,

trailing edge exit holes, and tip dump holes (figure III-E-3). Fig-

ure III-E-3 includes a metal temperature profile recorded at 2200OF

TIT and 1100°F cooling air temperature. A change in the design has been

initiate4 to reduce the high leading edge metal temperature and slightly

increase the trailing edge metal temperature, increasing thermal fatigue

life.

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t aNovember Phase I1-C Total

SFull-Scala Rig Test Time: 0 hours 44.97 hours

The duct heater has now accumulated 26.20 hovrs of engine testingU (with a maximum fuel/air ratio of 0.057) in addition to the 44.97 hours

previously accumulated on the full-annular rig resulting in a total test

time for Phase II-C of 71.17 hours.

Engine test results obtained to date compare closely to those of

the full-annular rig. The duct heater was successfully operated at

ci-ise conditions in engine FX-161 with the inlet distortion simulating

the maximum level representative of the intended installation. The

operating characteristics and performance of the duct heater were unaltered

• Iby the inlet distortion. All 67 attempts, at sea level and cruise con-

ditions, to light the duct heater in JTP17 engines were successful. Smooth

SI lights have been made with duct diffuser inlet Mach numbers as high as

0.9 (design maximum M - 0.56) and fuel/air ratios as low as 0.00157.

AiII

III--

Pratt &Whitney AircraftPWA FR-2213

SG. EXHAUST SYSTEM

The stat" pressure tap cruise test program has been essentially

completed. The objectives of the tests were to further investigate the

supersonic performance potential of the JTF17 exhaust system and to corm-

I pare measured pressure distributions with theoretical predictions. Com-

plete data for these tests will be available by early December,

The scale model reverser teat program to investigate reverser per-

S* formance, targeting, and flow characteristics was initiated. Checkout

of the flow visualization technique is complete, and photographs of the

reverser flow patterns are being taken.

The Boeing wing model has been received frotu the vendor and is under-

going instrnmentation prior to test (figure III-G-l). Approximately 50

I pressure taps will be installed to establish the wing local flow field

in the vicinity of the exhaust system. These tes-s will be conducted in

!I the United Aircraft Research Laboratory facilities in early Decemberto gain additional exhaust system compatibility data.

Instrumentation of the first reverser-suppressor unit and fabrication

of t'he second unit was completed during November. Engine testing is

scheduled to be resumed during December. Figures III-G-2 anJi III-G-3

show the pressures and temperatures which will be recorded during these

I tests.

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"1 1. Initial Experimental JTl7A-20 Control System

a. Duct Airflow Computer (Breadboard)

The Hamilton Standard S!N 1 breadboard computer unit was bench tested

at FRDC to investigate the overall accuiacy of the unit. The data are

[ b.ing evaluated.

b. Duct Fuel Controls

V! (1) Modified JFC-51 Duct Fuel Controls

The modified JFC-51 control continued to perform satisfactorily on

engine FX-161. The second control is engine ready.

[ (2) AA-MI Duct Heater Fuel and Nozzie Area Control

A T bias cam, a revised fuel flow schedule cam, and a reduced gainT2

Sintegrator cam have been installed in AA-Ml Control S/N D07COO1. The

T bias cam will allow the fuel schedric rate to be varied by mechanicalT2

- •trimming of the TT 2 servo. This bias feature will be used to match the

fuel flow a:,d exhaust nozzle area schedules to provide optimum control

system operation as evidenced by minim,im integrator piston motion. A

[r linear potentiometer has been installed inside the control to produce a

remote indication cf integrator piston travel. This control is under-

going final bench calibration at P&WA.

AA-Ml Control (S/N D07C002) was partially disassembled at P&WA and

the integrator cam was sent to Be-idix for rework to reduce the integrator

gain. Bendix is manufacturing a T bias cam and a revised fuel flow camT2

i I for installation into this control. A linear potentiometer as described

in the preceding paragraph, is also being incorporated during this con-

trol modification.

AA-Ml Control (SIN D07C003) was shipped from Bendix on 17 November.

The schedules in this control are the same as in SiN D07CO01, described

above.

III-H-i

tInPratt & Whitney Aircraft

PWA FR-2213

i c. Ignition

X The JTI2, 4 -joule, low tension, ignition system with the JTF17 type

shunted gap igniters has ignited the gas generator 132 times and the duct

heater 67 times on JTF17 experimental engines. Nineteen of the duct

lights were altitude lights.

ed. Quick-Fill

Bendix has completed a layout and is producing detail drawings forSthe prototype quick-fill kit to be lased with the AA-MI duct heater con-

trrols.

e. Duct Fuel Pump

A bench test was conducted that demonstrated the maximum flow cap-

ability of 86,000 pph using the duct fuel pump that is assigned to

engine FX-163-2.

R • 2. Prototype JTFi7 Engine Control

a. Conceptual Design

Both control vendors have supplied schematics which satisfy the

requirements of the noise abatement mode. Noise abatement is achieved

by allowing the duct heater to be operated below maximum gas generator

power. The noise abatement mode is selected by placing the shutoff

lever in the noise abatement position.

Operation of the system would be as follows:

1. Just prior to or during the takeoff roll the shutoff lever

would be placed in the noise abatement position.

2. Takeoff would be accomplished at part or maximum augmented

power setting. After the takeoff has been completed, the

engine power would be decreased to the part gas generator

power setting necessary to maintain the desired rate of

climb or loiter altitude. The duct heater will remain on

and will supply some augmentation which is matched for

minimum noise level for the thrust level selected.

iI-H-2

Ljj Pratt & Whiney AiroraftPWA FR-2213

3. Acceleration out of noise abatement into normal augmentation.10 is accomplished by power lever motion the same as before

addition of the noise abatement mode.

S4. Nvise abatement can be shut off by moving the shutoff lever

into the normal run position. Noise abatement is also shut

i • offlhen the power lev.r is moved below the minimum noise

abatement power lever position.

b. Digital Electronic Airflow Computer

Both vendors being considered for the unitized fuel and area control

have submitted proposals for an electronic module that would control air-

flow and certain other functions as an alternative approach. The vendors

S are continuing design studies to ascertain that when the module is mounted

[ on the unitized fuel and area control, the allowable engine dimensions will

not be exceeded. Evaluation of proposals is conLinuing.

Sc. EPR Control

Evaluation of the proposal for the optional EPR controls is continuing.

d. FRC Computer Studies of the JTF17 Control System -Dynamic simulations of the proposed JTF17 noise abatement syctem

were conducted on the digital computer. Results of these studies are

presented in figures III-H-2 through III-H-4.

The schedule used in these simulations is a 0.005 duct heater fuel/air

i ratio at minimum thrust setting with a linear decrease in fuel/air ratio

to 0.002 at maximum gas generator power.

Duct heater fuel/air ratios are limited to the noise abatement level

until the high rotor speed exceeds 80/. of maximum during a rapid powerI [1lever transient.

Figure I11-H-i illustrates a transient into minimum noise abatement

and from minimum noise abatement to maximum augmentation.

The operating point was set Just below 80% N2 when noise abatement was

selected. The duct heater lit at a 0.002 fuel/air ratio and increased to

0.005, Three seconds after noise abatement was selected, the PIA was

III -H-3

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f rPratt &Whitney AircraftPWA FR-2213

advanced to maximum; three seconds after the PIA was moved, the engine

was producing maximum thrust.

ji UFigure III-H-2 shows engine response tu a reduction in power from

maximum augmentation to minimum noise abatement. The transient required

j I U 1.75 seconds to reach a steady-state speed Lrom the time of the PLA

change.

Figure IUI-H-3 illustrates engine response to noise abatement when it

).s turned on and off. The cutoff of noise abatement is shown as a step

- j !| function. During the development program, the cutoff may be revised to

provide a time rate decrease of thrust.IF A dynamic siviulation of the latest Boeing inlet wats recei-ed and isbeing integrated into the system. it will be used to check engine/inlet

compatibility and to ensure optimuo' operation throughout the operationalpi range.

f- If •e. Prototype Duct Heater Fuel Pump

Hamilton Standard has incorporated the oil cooler bypass valve as

an integral part of the duct heater fuel pump. Revised installation

drawings showing this feature have been received at P&WA.

3. Failure Mode and Effect Analysis (FMEA)

The control system FMEA was completed for Phase iU-C with the

issuance of the reliability block diagram as reported in last month's

progress report. Therefore, this item will no longer be included in

f this report.

4. Advance Control System Program (Related Technology) - Exhaust Gasill Temperature (EGT) Control for the J58 Engine

Thirteen production models of the electronic (analog) units have been

-ldelivered, and flight suitability tests have been completed. Preparations

for flight testing of these systems is underway. Engine tests of the

experimental digital breadboard version of this equianent are continuing.

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Pratt &Whitney Aircraft[1PWA R21

I.BEARINGS AND SEALS

No rig development testing has been accomplished on bearings and

seals during this report period.

I Pratt &Whftey AircraftPWA FR-2213

SJ. FUELS AND LUBRICANTS

-I 1. Fuels

Fuel coker tests on aviation keroaene have continued to confir that

the fuel used in the experimental ITFl7 engines is meeting the purchase

specification requirements. This continued monitoring of the fuel

deliveries disclosed a problem with one of the fuel suppliers storage

taks and prevented the delivery of unacceptable quality fuel.

Testing has been accomplished in the JTFI7 engine aC cruise con-

ditions with fuel temperatures up to 370 0 F. Refer to engine section III-B

(FX-161 Test Program). Refer to section III-D for additional information

on the condition of the fuel nozzles.

2. Lubricants

Laboratory tests were continued on candidate lubricants to ensure

conformance to specification requirements. Testing has been accomplished

in the JTFI7 engine at cruise conditions with the lubricaut temperatures

up to 36G6F. Refer to engine section III-B (FX-161 Test Program).

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PWA FR-2213

K. INLET SYSTEM COMPATIBILITY

1. Inlet Distortion

Distortion tests were conducted on the 0.6-scale fan rig and the

i JTFI7A-20 engine. In both cases the inlet total pressure distortion was

generated with screens located at least one duct diameter ahead of the

fan in order to minimize the influence of the screen on the fan inlet

static pressure profile. V

a. 0.6-Scale Fan Rig

These tests were performed to determine the effect of distortion

en the fan surge line at cruise and to document the capability of the

[ fan to attenuate the distortion.

The surge line on the engine side (reference figure III-B-5) of the

fan was substantially unaffected by the inlet distortion. On the duct

side a slight loss (3%) in surge margin was measured. The distortion

introduced to the fan exceeded both the Boeing and Lockheed cruise dis-

tortion. The values of the distortion index, Kd2, are tabulated below:'=tCruise Distortion Simulation

Test Boeing Lockheed

Kd2 466 153 kCritical) 405

306 (2% Super-critical)

The pressure gradients created at the fan discharge by the inlet

distortion are presented in figure III-K-I together with the inlet

distortion pattern. The excellent capacity of the fan to attenuate

distortion before entering the duct heater is ev; at since the fan

discharge pressure is little affected by the distortion. Small dif-

ferences (- 1%, are noted in only scattered regions comprising a minor

portion of the fan duct discharge area. On the engine side (entering

the compressor) the attenuation is also excellent although not as com-

:1plete as that on the duct side. As shown, the regions in which the

preasure is more than ± 2% different from the average are significantly

reduced in size upon leaving the fan. These fan attenuation data show

that the high pressure compressor will be required to tolerate only a

Ifraction of the inlet distortion.

Ill-K- 1

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Pratt &Whitney Aircraft114 A th{-4'

The attenuation of the circumferential component of the distortion

is depictea in figures III-K-2 and III-K-3. These data show the pres-

suie distributions entering and leaving the fan on the mean striamline

for the engine and duct streams, r_-spectivelv. The difference be:ween

the maximum and minimum pressures at the fan dischar,k- is approxi .'atelyA 3% of average on the engine side and 2% on the ducL side., Entering

PK the fan, this difference is approximately 87- Good attenuation of the

average radial component of distortion is also attained (f7i._re III-K-4).

The change in the average radial discharge profile cue to the inl.t

qdistortion reaches a maximum of about 1/27., or 1/3 of the radial com-

ponent of the inlet distortion.

b. JTF17 Engine FX-161-5

[ The engine was run at simulated cruise conditions in an altitude

test cell with distortion generated by inlet duct screens. Normal

engine operation was attained at a distortion level representative of

tr'e intended installations. (Reference fi.gure III-B-3.) As anticipated

on he besis of the fan rig tests, the engine did not encounter stall

or surge ard no measurable effect on engine performance was noted. These

data are detailed in engine performance section MII-B-3.-i

The attenuation of the distortion on the duct side of the fan was

found to be substantially cot.nplete as in the case of the fan rig tests.

The distortion did not alter any discharge pressure by more than 2%.

The influence of the high pressure compressor on distortion attenuation

was measureu for the first time during these tests. At the compressor

discharge the largest incremental pressure attributable to the inlet

distortion slightly exceeded 2% over a very small region. The effect of

the distortion on turbine inlet te.,nerature was found to be negligible.

2. Engine/Inlet Compatibility

A digital dynamic simulation of the Boeing inlet was received and

Sin,! -porated into the JTFI7 engine simulhtion. Figures III-K-3 through

III-K-7 demonstrate the stability of .he system at cruise for the

following failure transients: duct heater light, duct heater blowout,

and primary combustor blowout. A failure analysis is proceeding with

the Boeing inletiJTFI7 simulation.

II -K-2S°I

Pratt & Whitney AircraftPWA FR-2213

Lockheed is preparing an updated inlet simulation for the engine/inlet

compatibility study. The-estimated completion date is 30 November.

IBM system 360 JTF17 simulations are being prepared for both-airframe

manufacturers. Boeing han requested a formulation of the JTFI7 engine

Fimulation to enable them to begin a hybrid simulation of the propulsion

system. A tape of the Fortran language statements is being prepared for

transmittal to Boeing.

A digital dynamic simulation of the JTF17 engine and control for

transmittal tc. the Air Force Aero-Propulsion Laboratory is being pre-

pared. £

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L. NCISE

The single jet noise model test program discussed in NWA FR-2098 and

NWA FR-2156 was completed in the East Hartford anechoic chamber on

22 November 1966. During this test series nine primary nozzle configu-

rations were tested both with and without a blow-in-door ejector. Atotal of 290 runs were made over a range of jet velocities from 1510 ft/sec

to 2270 ftisec to ensure a complete defin-tion of the nozzle performanceof each noise model. Figures III-L-l and IfI-L-2 are photographs of severalfof the mixing nozzles tested. rreliminary results from the initial nozzles

tested indicate from 5 to 6 PNdb attenuation from the base case conical

nozzle for the 4-lobe-50M penetration-short and for the 4-lobe-507,penetration-long, respectively. See figures III-L-3 and III-L-4. The

acoustical performance of the remaining nozzles will be reported next

month.

Construction has begun on a resonant liner test section to be run

in the East Hartford dual reverberation facility. Delivery of the base-

line, nuntreated section is scheduled for I December, w.th the treated

section delivery scheduled for 8 December. Fabrication and assembly

of the associated equipment has begun in East Hartford. This unit was

described in the last progress report, NWA FR-2156.

A review of the acoustic performance of eight study engines wasif initiated. These studies included the evaluatiun of predicted engine

noise level at takeoff, community takeoff, and community approach.

i None of the engines studied offered any appreciable advantage in noise

abatement over the basic JTFl7A-21 cycle engine.

Studies were made of recently available data to reevaluate the

predicted effect of fan duct resonant liners on rearward propagated

fan noise, Data used included absorption coefficient studies measured

with an impedance tube at FRDC and Douglas Aircraft, together with

results of simulated JTF17 fan duct tests in the dual reverberation

chamber at East Hartford. These data substantiate a 15 PWdb attenuation

of rearward propagated fan noise as a result of the current diffuser

design (see figure III-L-5). Increasing the treated area as shown in

figure III-L-6 will provide 24 PNdb attenuation of fan noise.

III-L-1

A

Pratt & Whitney AircraftPWA FR,-2213

71 A study program has been initiated to determine psycho-acoustic

reactions to impressed pure tones on a jet noise background as illustrated

in figure III-L-7. Previous studies by other investigators suggest that

imposed pure tones may be more annoying than the current PNdb scale would

indicate. To date, individual reaction to the various levels of pure

tone noise indicates that a better method of analysis is required. A

suggested approach to - wr,e the validity 'of the psycho-acoustic studyis to evaluate indjv*-:, ceactions through Lhe use of a polygraph which

will monitor the sub -ts reaction to the intensity level of the imposed

pure tone. This approach will be evaluated during the initial studies

which are scheduled to begin in December.

The rig for determining the reflection of reartard propagated fan

noise because of a density gradient in the fan discharge duct has been

constructed. (Reference EWA FR-2156.) Au analytical evaluation program

is being run concurrently with Dr. Ingard as consultant. Testing is

I scheduled to begin in early December.

The outdoor noise test rig, D-33, is approximately 907. complete.

Further delay in the delivery of major parts requires rescheduling of

the 1/8-scale JTFI7 engine erhaust system to December. The first series

i of tests on this rig will evaluate the effect of various primary plug

configurations run in a coannular exhaust system with blow-in-door ejector

I to confirm results obtained on the 0.07-scale model run in the East

Hartford anechoic chamber in August.

I The full-scale, 4-lobe mixing nozzle and blow-in-door ejector

(NWA FR-2098) has been constructed and will be tested on a sea level

static test stand early in December. This particular configuration of

a single jet mixing nozzle has been tested in the anechoic chamber in

East Hartford (4-lobe - 50% penetration - long length). Preliminary

data from this test indicated attenuation of 6 PNdb.

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M. MOCKUPS

Revisions to improve the main fuel pump and hydraulic pump

configurations for accessibility and maintenance were accomplished

on the full-scale engineering mockup.

A scheme for quick disconnect of the accessory gearbox from the

unitized fuel control base plate was mocked-up in full scale during

this period.

I!I

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S• i Pratt&Whitney AircraftPWA FR-2213

N. COORDINATION

I I. General

Messrs. H. T. Luskin, Assistant General Manager, SST Propulsion

I and J. Stroud and E. Bragden, Propulsion Group, LCC, visited FRDC on

4 November, for briefing on JTF17 hardware development progress sinceI 6 September Phase III Proposal submittal.

The following Source Selection Council members visited FRDC on

j 'b' November for SST discussions:

Mr. D. D. Thomas, Associate Administrator, FAAII Mr. A. Dean, Associate Administrator for Administration, FAA

Mr. J. Blatt, Associate Administrator for Development, FAA

Mr. 0. Bakke, Eastern Regional Director, FAA

Mr. N. Goodrich, General Counsel, FAA

Mr. J. H. Hoover, Special Assistant to Associate Administrator, FAA

Mr. E. W. Stimpson, Assistant Administrator, Office of CongressionalLiaison, FAA

Dr. R. Bisplinghoff

The FAA SST Supplementary Engine Evaluation Task Force visited FRDC

on 16 and 17 November to review the SST engine and rig test results since

the SST evaluation team visit during the week of 19 September.

Mr. F. A. Maxam, Chief Engineer, SST Program, The Boeing Company,

visited FRDC on 16 November for SST discussions with Mr. W. L. Gorton,

FRDC General Manager. Other Boeing SST visitors during this report

period included: Messrs. R. S. Neuhart, Chief, Power Plant Design, and

J. J. Hamry, Power Plant Design, on 4 November and Messrs. K. Chun,

D. Swanson, and D. Smith, SST Propulsion Staff Group, on 7 and 8 November.

Dr. N. Golovan, Office of Science and Technology, visited FRDC on

18 November for JTF17 SST engine discussions.

Messrs. W. L. Gorton, FtDC General Manager, B. N. Torell, Chief

Engineer, W. H. Brown, Assistant Chief Engineer, and H. N. Cotter and

P. L. Bosse, Performance Engineering Group, visited Lockheed on

21 November and The Boeing Company on 22 November for engine and airplane

growth discussion.

III-N-1

Wresti K vvrn1Uu my 141t ;F cauitPWA FR-2213

ii FRDC performance engineers visited LCC on 14 November and Boeing

on 15 November for thrust measurement systen discussiuns.

JTF17A-21 noise abatemient study decks were transmitted to both

Lockheed and Boeing. Both IBM 7090 and IBM 360 system decks were

provided.

Data on the JTF17 engine comb'istor relight characteristics were

sen to both Lockheed and Boeing.

2. JTFl7A-21L Engine

A P&Wk-Lockheed installation coordination meeting was held at FRDC

on 27 and 28 October. The meeting covered general, review of both air-plane and engine programs and open items on instnllation design.

As requested by Lockheed, an IBM 360 study performance deck, com-

parable in performance to the specification IBM 7090 deck, was provided

to aid Lockheed in setting up their IBM 360 system.

Studies are in progress concerning relocation of front mount pads,

a revised rear mount lug configuration, and possible engine movement with

respect to airframe in the event of a failed front or rear mount system.

Studies are continuing on control drive motor locations and reverser/

fuel control interlock configurations. A Lockheed drive motor specification

has been requested to aid these studies.

In response to Lockheed's request, a reverser-suppressor outer skin

F temperature profile, from the nacelle/reverser-suppressor mating face to

the tail feather exit plane,was generated and transm'itted to them.

Lockheed was provided with data and drawings to initiate coordination

on the design of their 0.6-scalc . .fuser to the run with the 0.6-scale

fan rig at the Willgoos Laboratory in July 1967.

3. JTF17A-21B Engine

FRDC performance and installation engineering personnel visited

Boeing on 7, 8, and 9 November for discussions on performance, mission

analysis, economics, engine and airplane growth, and engine cycles.

III-N-2

weI~B Pratt &Whitney Aircraft

PWA FR-2213

Boeing SST staff performance personnel visited FRDC on 7 and 8 November

to deliver and review the Boeing dynamic inlet simulation. The study will

be used to investigate engine/inlet compatibility in the flight Mach

number range of 2.2 to 2.7. Boeing and FRDC will periodically meet to

Slcoordinate study results.

During t•,e Boeing visit of 4 November, progress was made in improving

areas of installation compatibility relative to engine mount system,

nacelle/reverser-suppressor mate-up, inlet/engine mate-up, power takeoff

gearbox, power lever and fuel shutoff lever attachments, air bleed and

power extraction requirements, windmill brake a-tuation requirements, fuel

: tank drain installation, ground handling, and engine operating envelope.

The schedule for updating the Boeing JTFI7A-21B mockup engine was reviewed.

II An FRDC acoustic engineer visited Boeing on 21 November to review

F&WA fan rig noise test data and noise calculation methods.

U Data on the JTF17 reverser-suppressor were transmitted to Boeing in

support of their reverser-suppressor design study for the SST.

I

i ITI-N-3

!

jI Pratt & Whitney Aircraft

FWA FR-4`21

0. MAINTAINABILITY

The split primary burner case configuration has been redesigned to

an annular case ,, is translated to gain access to the primary burner

modules. The elimination of the split case reduces the number of bolts

required, thus reducing maintenance man hours. Also the problem of

alignment of distorted case halves at reassembly has been eliminated.

F The translating of the annular burner case can be readily accomplished

as presently done on the P&WA JT3D engine.

The access panels to the gas generator have been redesigned to reduce

the number of bolts required for disassembly by approximately 20%. The

previous design required a large number of bolts in each panel since the

panels were structural members of the outer engine cases. Present access

panels are nonstructural members and therefore require a minimum amount

of fasteners to attach to the engine. This reduce: the elapsed time

_ maintenance man hours required to gain access to the gas generator fuel

nozzles, fuel manifolds, and start bleed valves.

A mockup of the access panel area will be built to permit demon-

stration of maximum accessibility of the gas generator area through the

panel openings, and to determine the size and strategic location of

the panels.

Further studies of the Customer Maintenance or Overhaul Errors,

documented by our Service Records department for the perioa of January 1961to April 1966 have been made. Analysis of the data has assisted the DesignMainLainability group in preventing error-inducing designs from being incor-

porated into the JTF17 engine.II The following design layouts and studies are in process which will

improve the maintainability characteristics of the JTF17 engine.

1. A study and redesign for the attachment of the fan stator

assembly versus the location of the controls. The ultimate

objective is to have the bolts that attach the f:n stator

assembly accessible without having to remove any controls.

2. Incorporate p~ovisions whereby individual compai lent

breather flow checks may be made.

!~11--IIIO

_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ -o

'I ~Pratt &Whltny AircraftPWA FR-2213

1 3. A design for a simplified attachment between the variable

inlet guide vane and the connecting lever.

1 4. A study as to the possibility of incorporating borescope

provisions for inspectivn of compressor disk rims andI weas outside of compressor disk spacers.

!Iaifttainability presentations were made at the ATA conference inI Los Angeles, California by Service and Design personnel. Brochuresealttled l~aintainability and Reliability Trends for the New Generation

I ~of Aircraft Engines" were distributed at this meeting.

A visit was made to American Airlines in Tulsa, Oklahoma, by Service

Department personnel to present films of maintainability and to provide

verbal desci-iption of JTF17 engine configu.ration and maintainability

featares.

Documentation of procedures for the repair of turbine exhaust case I-vanes has. been prepared.j

41

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I _ I Pratt &Whitney Pircraft�wA h-22i.s

9�.I P. VALUE ENGINEERING

Value Engineering proposals completed during the month of November

Uinclude:

1. Value Engineering proposal No. 3.18 revises the anti-Ivortex tubes. Potential savings of $198 per engine

3 2. Value Engineering proposal No. 4.07 proposes a change in

the material of the duct heater case support from a forging

3 to c� casting. Potential savings of $1308 per engine

3. Value Engineeri.ng proposal No. 3.20 proposes a change in

the material of the 6th, 7th, and exit compressor stageIshrouds from INCO 718 to INC0NEL. Potential savings of

$45 per engine.IValue Engineering cc�st studies c.ompleted during this report period

include:

[ 1. Forged versus cast fuel nozzle supports

2. Rear mount case as a �eldment vs single forging

3, Estimated cost of titanium honeycomb noise suppression

liners in the fan exit duct and duct diffuser

4. Cost of cast module tracks for primary and duct burner

1 5. Aerobrake

if 6. Anti-vortex tube

7, "Z11 spacer in high compressor

1 8. Integral versus welded borescope pad on rear mount case9. Proposed material change to Waspaloy on inne�: and outer

- wI transition duct cases.

During this month, Value Engineering adso:

I 1. Revised the contour of the primary combustor case and the

diffuser case to straight c.onical sections, whiuh permits a

I gear shaping process instead of contour milling

Ill-P-I

Pratt & Whitney AircraftPWA FR-22113

1. . nvestiigated possi-bil~itv of casting linkc am~ or-, Akr'hrake

in place of m~achined forging

1 3 Conti-nued liaison with East llartforu Production Enginleeriný

co ensure that the engine design is adaptabi;e to produc-

IF

Pratt & Whitney AircraftPWA FR-2213

.. CONFIGUP!rION MANAGEMENT

Design of the prototype engine is continuing WiLh incorporation of

I coordinated interfaces. Detail changes required on some irteL'face items

are being coordinated with the airframe manufacturers as design of the

engine and airframe progresses. All pioposed changes are being trans-

mitted by Field Survey layouts with a log of dates of transmittal and

acceptance or rejection by the airframe manufacturer. The basic con.-

figuration of the engine is unaffected by these changes.

[ • On 4 November Boeing propulsion system engineers visited FRDC for

installation coordination. The major topics of discussion included

inlet/engine mate-up, reverser-suppressor mate-up, windmill brake

requirements, drain system, and horsepower extraction. Problems in each

category were dAscussed, and the required action by both parties was

determined. Boeing provided a revised presentation of horsepower extraction

( •which is being reviewed.

M~.

Nii

r -

Pratt& Whitney Aircraft

Li PWA FR-2_2i3

R, QJALI•I.Y ASSURANCE

Exptrimentation with the Defectometer, an eddy current defect detector,

manufactured by Automation Foerster, has denonstrated the capability of

indicating tightly closed cracKs 0.010 inch deep by 0.060 inch long in cast

turbine blades. Work with this equipment is continuing, and purchase of

the equipment for use during Phase III is under consideration.

rf Work with turbine blade airfoil wall thickness measurv.,ient techniques

and equipment has greatly increased confidence in the ability of the

inspection procedures now in use to control turbine blade airfoil thick-

ness in critical areas. Points of measurement for the 2nd-stage blades

I"

N are shown in figure III-R-I,

i&

LIE

I i- R- i--

AR

Pratt &Whitney AilrcraftFWA FR-2213

£ 1 IFUl Length of Convex Side

gI~ 2Places

0 0

M

Figure 111-A-1. JTF17 2nd-Stage Tu~rbine FD 19064

Blade Wall Thickneas

|4

'I Pratt &Whitney AircraftWA FR-2213

S S. RELIABILITY

1. Design Reviews

Approximately 24 prototype engine layouts are in process of review.

S2. Failure Mode and Effect Analysis

The rough draft of the third edition of the Failure Mode and EffectAnalysis is now being reviewed by Design and Project Engineering.

3. Special Reliability Studies

A computerized data retrieval program has been devised to permit

rapid compilation and presentation of statistical information pertining

to commercial engine inflight shutdowns and premature engine removals.

Coding of all entries and preliminary tabulations have been completed.

A study of all P&WA turbine engine disk failures resulting in engine

case penetrations on commercial sircraft since the beginning of service

has bccn completed.

The pertinent results of this study were as follows:

II 1. Fifteen disk failures in 34,000,000 engine operating hours.

2. Of those failures which were primary disk failures, two

resulted in fuselage penetration.

3. One instance of a disk failure resulted in fuselage

penetration at altitude. This was a secondary disk failure.

4. The JTF17 engine design incorporates features to preclude

the primary failures found in this study that can escalate

into disk failures.

5. The engine design also incorporates configurations to reduce

the possibility of primary disk failures from fatigue, £such as the elimination of bolt holes in the disk portion

*1 of tne titanium fan rotors, elliptical broach slots,

integral spacers and disks, and long snap diameters for

the high compressor rotors.

A study of the JT8D No. 4 breather tube failures has been completed

and all pertinent information transmitted to the engineers designing

the JTF17 breather system.

7 O0-S-I

A

Prait & Whlttney PircraftWA FR-2213

A study is in progress to determine the feasibility of retaining the

inner platform of the ist-Rtage turbine vanes to prevent failure escalation

in the event that both airfoils of a paired vane have been severed.

A study of the JT8D 6th-stage compressor blade root problem is in

progress to ensure that a similar condition will not exist on the JTF17

engine.

4. Parts History Survey

The parts history data retrieval system was utilized to determine

the test exposure time of the 3rd-stage compressor air seal on the high

compressor component test stand.

i 5. Reliability Training

Three reliability training films listed below were obtained from the

U.S. Air Force and viewed by the Design and Development Reliability

I, Groups.

1. "No Second Chance"

2. '"Maintainability - Design/Living"

3. "AGREE in Action."1 6. Statistical Engineering

To determine jet fuel storage requirements, an analytical study is

Ii being made of the jet test fuel demands and fuei storage supply capa-

bilities. The study has produced a curve showing :he probability

distribution of fuel consumption for each test area. 0

4MI .

Q I1, -7

MuffiS 2

PWA FR-2213

-i SECTION IV

AIRLINE COM{MENTS

< IPreliminary test stand requirements for the JTF17 engine have beenestablished qn transmitted, in answer to specific requests, to Pan

I ~American World Airways, Am~erican Airlines, United Airlines, and Trans-World Airlines for use in planning their engine test stands. These

:1 requireml~nts are also being made available to all the other domesticand world-wide airlines who have expressed an interest.

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KU IPratt &Whitney RArcraftNWA FR-2213

SECTION VSTATE-OF-THE-ART

I Development testing of main shaft seals for advanced air-breathing pro-

pulsion systems under Contract NAS3-.7609 is continuing with completion

scheduled for June 1967. This program was discussed in Section V of the

July Progress Report NWA FR-1953. Progress to date follows:

1. A total of 70 hours of calibration time has been accumu-

lated on rubbing face seals. Testing is continuing in

order to establish the operational limits of this type

seal configuration.

2. Approximately 18 hours have been accumulated on the orifice

compensated concept. Difficulty has been encountered in

achieving leakage rates comparable to those of rubbing face

seals. Testing is continuing.

More specific information may be obtained from Progress Re-

ports PWA-2879 and 2958.

A similar Contract (NAS3-7605) that covers development of compressorend seals, stator interstage seals, and stator pivot seals has been initiated.

I The object is to provide development data for the design of the above type

seals for advanccl air-breathing propulsion systems. The program will

entail concept feasibility studies leading to the design and testing of

hardware. The ultimate goal is to achieve increased compressor efficiency

through the use of seals that have much lower air leakage rates than seals[ ~carrently in use while retaining or improving the reliability anid weight

factors.

I. Preliminary designs have been compl'kted and shmitted to NASA for

approval.

V-1

37