F/6 HIGH BYPASS TURBOFAN COMPONENT DEVELOPMENT. … · inni/illlllluu Illllllllll. AFAPL-TR-79-2103...

AD-AOG8 621 SDNERAL MOTORS CORP INDIANAPOLIS IND DETROIT DIESEL --'ETC F/6 91/5HIGH BYPASS TURBOFAN COMPONENT DEVELOPMENT. PHASE 11. FAN DETAI--ETC(UlDEC 79 0 C CHAPMAN F33615-78-C-2014

UNCLASSIFIED DA-EDR210026 AFAPL-TR-79-2103 ML

EilEliEliEE-EllEllllEllEEinni/illlllluuIllllllllll

AFAPL-TR-79-2103

HIGH BYPASS TURBOFAN COMPONENT DEVELOPMENT

Phase Il-Pan Detail Design

0 D. C. Chapman

Detroit Diesel AllisonDivision of General Motors CorporationP. 0. Box 894Indianapolis, Indiana 46206

December 1979

TECHNICAL REPORT AFAPL-TR-2103

Final Report for December 1978 to October 1979

IThis documuent has b~approved

Air Force Aer@ Propulsion Laboratory for public relcae and sale.- Us

SAir Force Wright Aeronautical Laboratories distribution is unlimited.

Air Force Systems CommandCD Wright- Pattersonl Air Force Base, Ohio 45433

_80 4 4 052

/I

NOTIC.5

r.hen Government drawings, specifications, or other data are used for any pur-pose other than in connection with a definitely related Government procurementoperatiol, the United States Government thereby incurs no responsibilit.y nor anyobligation whatsoever; and the fact that the government may have formulated,furnished, or in any way supplied the said drawings, specifications, or otherdata, is not to be regarded by implication or otherwise as in any manner licen-sing the holder or any other person or corporation, or conveying any rights orpermission to manufacture, use, or sell any patented invention that may in anyway be related thereto.

This report has been reviewed by the Information Office (O) and is releasableto the National Technical Information Service (NTIS). At NTIS, it will be avail-able to the general public, including foreign nations.

This technical report has been reviewed and is approved for publication.

ERIK W. LINDNERpcoject Engineer Tech Area Manager

.# 4 FOR THE CO :f:.ANDER

aSAK J. GERSHONActing Branch ChiefPropulsion BranchTurbine Engine Division

"If your address has changed, if you wish to be removed from our mailing list,or if the addressee is no longer employed by your organization please notifyA.J.T.I .._,W-PAFB, OH 45433 to help us maintain a current mailing list".

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() REPORT DOCUMENTATION PAGE SE2ADDICOMPLTN ORM- r aV? AC MCI~ i. PCI Ab'I AA.248 -4UMVn 1

EV PI~I~ ERIOD COVEREID

6 IGR JYAS S TURBOF N% DEVEL0it, pAna' '2TAIL DESVA, _91 '4.£P 4 mean

~~~Pev D..AnRm~- ;

D. C 1 Xha/al F33615-78-(-1f

9. PERFORMING ORGANIZATION 4A!7 AMC ADDRESS 00 POQOPAW VMCNT. VR0JIMC7 7ASl

Detroit Diesel Allison, Division of General MotorsCorporation, P.O. Box 894, Indianapolis, Indiana 3 446206 AeoPouso aoaoy i oci

I I. CON TROLLING OFFICE M AME AMC ADDRESS SP~.&

Ai oc eoPouso aoratory, AirFre Decembe 179/W'right Aeronautical Laboratories, Air Force Sytk. 91Coamand. Wizht-?atterson Air Force Saaa Ohio 141 68

G4 4~7~N AGENCY SiAME & AODRICSS(f different tram Caftlaod1114 Office) i S. SZC'IIY CL.ASS. (ot :hia *art)

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16. DISTI41SUTION STATEMENT (at thgs Report)

Approved for Public Release; Distribution Unlimited

17. DI57N13UION STATEMENT 'at th abstract entered In Block 20, if iffrone from Report)

18. SUPPLEE)ANTARY NOTES

APAPL-TR-79-2034, High Bypass Turbofan ComponentReference report - Development, Phase I-Preliminaryi Design and Life

CPhase I Cycle Cost Analysis of Candidate Engines, 0. C.Chapman and W. A. Redmond

tl. KEY WORDS (Carrno~ an reoaea side it nocoaerec aIdentifr y black number)

-~ Turbofan Engine, Trainer Enginer, Fan, Small Fan

ZO. ABShTRACT (Continuean onwer**. side ft necooa~p dod idenitity bw block num~ber)

The objective of this program is to develop an advanced, small, hiigh laypass tur-bofan component applicable to an advanced, fuel-efficient engine *for a 'Iew AirForce primary trainer aircraft. The program consists of two phases. Piaseconsisted of the preliminary design and life cycle cost analysis of candidateengines. This report documents Phase 11 which wias the detail design of theselected fan configuration. Primary design point of the fan IS 1.8 pressureratio at 55.9 ibm/sec flow. At a lower speed, the fan will produce 1.63 (cont

DD I iNI 1473 t0 o~mo OFINOV 63 ISOSLT~- 2~.z? ~'~''~SECURITY CLASsIWICAtluN1 OF -MIS 4GC t$%oft Des Sneered)

SCURITY' CZLAMPlICATION OF ?T4eS PA*OE(Wh@ mama tn#@p.EJ

pressure ratio at a flow compatible with the GMA500 advanced technology coreengine. The resultant engine satisfies all requirements of the contract.The fan stage meets all requirements of >1il-E-5007D, including bird ingestionand predicted noise levels are below PAR Part 36 requirements. <--

SCCIRIT : .ISSRIC TIOHIF is MG91ho"Oarsentred

Preface

This report was prepared by D. C. Chapman of Detroit Diesel Allison, Division

of General Motors Corporation, Indianapolis, Indiana.

Design details of the fan stage designed in Phase 11 of Contract No. F33615-

78-C-2014, High Bypass Turbofan Component Development, sponsored by the A.F.

Aero Propulsion Laboratory are presented. The Air Force contract monitor was

Capt. Larry Gill.

rrL.-~on

Eyi

Ave i.

i- ,iI- '

12s specia

TABLE OF CONTENTS

SECTION TITLE PAGE

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . ix

I INTRODUCTION .......... ........................ I

II DESIGN REQUIREMENTS .......... ................... 2

III AERODYNAMIC DESIGN ........... .................... 4

Flow-Path and Vector Diagrams ...... ............. 4

Blade Design ........... ...................... 9

Vane Design ......... ...................... .. 16

IV STRUCTURAL ANALYSIS ........ .................... .. 20

Steady-State Stresses ...... ................. ... 20

Vibration Analysis ....... ................... ... 30

Flutter Analysis ........ .................... ... 32

Bird Ingestion Analysis ...... ................ ... 36

V NOISE PREDICTION ....... ........................ 42

Appendix A-Axial Compressor Design System ..... ........ 44

Appendix B-Design Point Vector Diagrams ..... .......... 46

Appendix C-Rotor and Stator Blade Coordinates ... ....... 49

~33SD3BPAZ xAumrjo num

7

LIST OF ILLUSTRATIONS

FIGURE TITLE PAGE

1 Schematic of fan flow path ................. 5

2 Blade Mach numbers .......... ..................... 5

3 Vane Mach numbers .......... ..................... 6

4 Radial distribution of total pressure loss coefficient. 7

5 Blade and vane loading distributions ...... ............ 7

6 Single stage surge margin correlation ..... ........... 8

7 Blade air angles ........... ...................... 8

8 Vane turning angles .......... .................... 9

9 MCA airfoil definitions ......... .................. 10

10 Blade chord .......... ....................... ... 12

11 Blade solidity ......... ....................... ... 12

12 Blade maximum thickness/chord ratio .... ............ . 13

13 Blade incidence and deviation angles .... ............ ... 14

14 Blade metal angles ......... .................... ... 14

15 Blade maximum thickness and inflection locations ........ .. 15

16 Blade conical airfoil sections ..... ............. *.. . 15

17 Vane solidity ......... ....................... ... 16

18 Vane chord .......... ......................... ... 17

19 Vane passage throat minimum critical area ratio ....... .. 17

20 Vane incidence and deviation angles ...... ........... 18

21 Vane metal angles ........ ..................... ... 18

22 Vane conical airfoil sections ..... ............... ... 19

23 Principal blade stresses at 20,223 rpm ... ........... ... 22

24 Principal blade stresses at 18,950 rpm ... ........... ... 23

25 Fan blade and wheel S-N diagram ..... .............. ... 25

26 Fan blade Goodman diagram ...... ................. ... 26

27 Wheel equivalent stresses ...... ................. ... 28

28 Frequency-speed interference diagram (first 3 modes). . . . 31

29 Frequency-speed interference diagram (first 32 modes) . . . 32

30 Relative dynamic stress distribution--LB mode at 20,223 rpm 33

31 Relative dynamic stress distribution-IT mode at 20,223 rpm 34

vi

FIGURE TITLE PAGE

32 Stall torsional flutter analysis .. .. ............ 35

33 Bird impact area and blade thickness illustration .. .. ... 39

34 Bird ingestion damage index-titanium blade. .. ....... 40

35 Bird ingestion damage index-steel blade .. .. ........ 41

vii

LIST OF TABLES

TABLE TITLE PAGE

I Structural design criteria . 21

2 Steel bl.ade stress summary .. .. .............. 24

3 Steel blade stress sumary at maximum dynamic response .. 26

4 Steel wheel stress summary .. .. .............. 27

j5 Titanium blade stress sumary. .. ............. 29

6 Titanium blade stress summary at maximum dynamic response .30

7 Titanium wheel stress sumary . ............... 30

8 MIL-E-5007D bird ingestion requirements .. ..........37

9 Fan and aircraft speeds for MIL-E-5007D conditions. .. .... 38

10 Undergraduate trainer noise levels .. .. .......... 42

viii

SUMMARY

An advanced technology fan stage adaptable to a small, high bypass turbofan

engine for a future Air Force primary trainer has been designed. Primary de-

sign point is at 1.8 pressure ratio and 55.9 lb/sec flow. At a lower speed,

this fan stage will produce a 1.65 pressure ratio at a flow compatible with

the GMA500 core engine to form a high bypass engine meeting all requirements

for the trainer application.

At the 1.8 pressure ratio design condition, the fan rotor operates at 1606

ft/sec tip speed with an inlet annulus specific flow of 42.3 ibm/sec/ft2.

The inlet hub/tip radius ratio is 0.40. The rotor has 20 blades of multiple

circular arc airfoil sections with an aspect ratio of 1.64. Maximum thick-

ness-to-chord ratio varies nonlinearly from 8.5% at hub to 2.5% at tip.

Forty-wo vanes of multiple circular arc cross section are tilted rearward at

the tip to increase blade to vane spacing for noise considerations. The vane

aspect ratio is 2.32 and the maximum thickness-to-chord ratio varies from 6%

at the hub to 8% at the tip.

The design meets all structural design requirements including steady and vi-

bratory stresses, blade flutter, and bird ingestion. Substantial margins

exist for blade and wheel permanent set at 122% speed and for burst at 130%

speed. Adequate margins also exist for low cycle fatigue, considering 12,000

cycles to design speed with Kt = 3 at blade leading and trailing edges, Kt

1.4 at blade crown, Kt 0 1.4 in the wheel rim, and Kt = 2.0 in the wheel

web. Allowable blade vibratory stress exceeds the required +15 ksi at reson-

ance points and the required +5 ksi at nonresonance points.

The blading was also checked for torsional stall flutter and found to be sat-

isfactory.

Bird ingestion requirements of Mil-E-5007D forced a slight thickening of the

blade leading edge region in conjunction with a material change from titanium

to stainless steel.

The predicted noise levels are substantially below FAR Part 36 levels at take-

off and approach. Furthermore, the ground idle noise levels are greatly re-

duced from those of the current trainer configuration.

ix

' -7 i-7 ..... -1

SECTION I

INTRODUCTION

The current Air Force primary trainer (T37) fleet is approaching the end of

its useful life, and a replacement aircraft will be needed. A fuel-efficient

engine for the replacement aircraft must be developed. Advanced technology

engines, such as the GMA500, suitable for the core of such a high bypass en-

gine are being developed. Advanced technology fan stages in this size class

are not available. To fill this void in technology, Detroit Diesel Allison(DDA), Division of General Motors Corporation, has conducted the High Bypass

Turbofan Component Development Program for the United States Air Force Aero

Propulsion Labratory, Wright-Patterson AFB, Ohio. The program consisted of

two phases:

9 Phase I-Preliminary Design and Life Cycle Cost Analysis of Candidate

Engines

0 Phase II--Detailed Design of the Fan stage chosen from Phase I

The Phase I studies were reported in Report No. AFAPL-TR-79-2034, High Bypass

Turbofan Component Development, Phase I--Preliminary Design and Life Cycle

Cost Analysis of Candidate Engines by D. C. Chapman and W. A. Redmond. This

report documents the design of the fan stage completed in Phase I of the

program.

Now= ........... .

Section II

DESIGN REQUIREMENTS

The core engine of choice in the Phase I selection process was the QASO0, an

advanced technlogy turboshaft engine which was a winner in the United States

Army Advanced Technology Demonstrator Engine (ATDE) Competition. DDA is cur-

rently under contract to complete 500 hours of running on the GMA500 engine

starting early in CY 1979. The engine consists of a two-stage centrifugal

compressor, foldback annular combustor, two-stage gasifier turbine, and two-

stage power turbine.

U In Phase I of this program, fans of 1.5, 1.65, 1.8, and 2.0 pressure ratio

were matched with the GMA500 core engine to form candidate high bypass ratio

engines for system life cycle cost analysis. The performance of all engines

met or exceeded the requirements of this contract. Using representative air-

craft characteristics, these engines, designated PD418, were applied to the

mission requirements established by the Air Force. Both aircraft gross weight

tj !and system life cycle cost were minimized with a fan pressure ratio of 1.65,

although the advantage of that pressure ratio over 1.5 and 1.8 pressure ratios

was not great. DDA, therefore, recomended to the Air Force that the 1.65

pressure ratio fan be selected for detail design in Phase II of the program.

The Air Force identified a higher technology level with the 1.8 fan pressure

ratio and because the Life Cycle Cost (LCC) penalty was nmall, selected that

pressure ratio for Phase II. DDA preferred the 1.65 pressure ratio not only

because of the LCC analysis but because the engine sea level static thrust

level was approximately 7.3% greater than the engine with 1.8 fan pressure

ratio. A mutually agreeable set of design conditions were established wherein

the fan would be designed to achieve 1.63 pressure ratio at a flow compatible

with the GMA500 core engine and also to achieve 1.8 pressure ratio at a higher

flow and speed. The primary design point thus established is:

* Pressure ratio 1.8:1

* Corrected flow 55.86 Ibm/sec

e Efficiency 85%

2

NOW

The secondary design point, which matches the GON core engine requirement

at Z5,000 ft, 0.5 Mach number is:

* Pressure ratio 1.65:1

* Corrected flow 52.8 Ibm/sec

, Efficiency 87I

Structurally, the fan should meet the requirements of Mil-E-5007D, including

bird ingestion capability. Furthermore, the fan noise levels should be withinthe limits of FAR Part 36 at both takeoff and approach. An unofficial goal

was to achieve a substantial reduction in noise at ground idle compared to the

existing primary trainer.

.. 3

SECTION III

AERODYNAMIC DESIGN

The aerodynamic design is presented at the 1.8 fan pressure ratio operating

4 condition. Flow-path and vector diagram details are followed by rotor and

stator blading information.

FLOW-PATH AND VECTOR DIAGRAMS

The design parameters for the small high bypass fan are:

e Stage pressure ratio 1.8:1

* Corrected flow rate, ibm/sec 55.86

* Adiabatic efficiency, % 85.2

* Rotor inlet hub/tip radius ratio 0.40

. Corrected tip speed, ft/sec 1606

* Corrected speed, rpm 21685

• Mechanical speed, rpm 20223

o Corrected specific flow rate,

ibm/sec/ft 42.28

The velocity diagrams of the fan were obtained using the DDA Axial Compressor

Design System. A description of the design system is given in Appendix A.

The fan was designed at an altitude cruise condition of 25,000 ft at 0.5 Mach

number. This point represents the maximum mechanical speed achieved by the

fan in a representative trainer mission.

A schematic of the fan flow-path is shown in Figure 1. The fan has a constant

tip diameter of 16.974 in. and a rotor inlet hub-to-tip radius ratio of 0.40.

The rotor hub ramp angle is 31.25 deg. The number of rotor airfoils is 20

while the stator has 42 vanes. The number of vanes and the vane-blade spacing

were consequences of acoustical considerations.

The average value of the blade inlet absolute Mach number is 0.617. The blade

inlet relative Mach numbers are supersonic for the outer 75% of the span. The

exit relative Mach numbers are all subsonic (Figure 2). The average vane exit

Mach number is 0.465. The inlet and exit Mach number profiles for the vane

are shown in Figure 3.4

9

7

Rotor 120) ttr(2

5

4

3 L0 1 2 3 4 5 6

Axial dimension- in. TE-MM3

Figure 1. Schematic of fan flow path.

1.8- ne

1.6 -Exit

~1.4

~1.2

S1.0

~0. 8

0. 6--

0 20 40 60 8010Percent span from hub

TE-8037Figure 2. Blade Mach numbers.

5

" 1.0- ne, ----- Inlet

0.9 - -- Exit

0.8

~0.7

.=0.6

.~0.5 -

i0o.4

0.3-

0.2 -.

0 20 40 60 80 100

Percent span from hubTE-8038

Figure 3. Vane Mach numbers.

The predicted blade and vane total pressure loss coefficients are illustrated

in Figure 4. The resulting average efficiencies are 88.4% for the blade and

85.2% for the stage. The spanwise distribution of the design point loadings

(diffusion factors) are shown in Figure 5. They are moderately high but the

estimated surge margin for the fan is 18.6%. This surge margin estimate is

* based on a correlation of blade aspect ratio, relative Mach number, and tip

loading at surge for various single stage compressors (Figure 6).

Figure 7 shova the blade inlet and exit relative air angles while Figure 8 is

a plot of vane turning angles. The exit air angle from the vane is designed

to be 0.0 degrees (axial).

6

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

0.22S lade total

0.20 B lade ShockVane total

0.18

*~0.14

v0.12 -

~0.10

~0.08

0.04 7

0 20 40 60 30 100fPercent span from hub

tE-3039

Figure 4. Radial distribution of total pressure loss coefficient.

0. 6-Blade

S0.5

0.4

S0.43

0 20 40 60 80 100Percent span from hub T-0

Figure 5. Blade and vane loading distributions.

0 9 ret a 5a Sin ge-sta e data

0.8 - 1.2

1.6X ~0.7M0.

~0.6-

0.5 -T

0.4

0.4 0.8 1.2 1.6 Z. 0 2.4 2.8 3.2Rotor aspect ratio

AW' TE -3041

I Figure 6. Single stage surge margin correlation.

70

*1 60

~50

~40

30- 1nMet

----Exit

10

020 40 60 80 100Percent span from hub

TE-8042Figure 7. Blade air angles.

8

50

_I 45

i40

!351 0 1 1

0 20 40 60 80 100Percent span from hub TE-8043

Figure 8. Vane turnling angles.

The design point vector diagrams, calculated along streamlines, are tabulated

for the blade and vane leading and trailing edge stations in Appendix B.

BLADE DESIGN

There are 20 blades with an aspect ratio of 1.64 (based on average span and

true mean chord). The blade consists of multiple circular arc (MCA) airfoil

sections designed on conical surfaces approximating streamlines of revolution.

An MCA airfoil is shown schematically in Figure 9. It is made up of two cir-

cular arcs which define three metal angles: inlet (01*), exit ('2"), and

inflection (0 *). A metal angle is the angle between the axial direction

and the mean camber line at a specified location. A blade section is designed

by adjusting the metal angles to satisfy incidence, deviation, and starting

9

Throat area

Passage inlet areaFirst Icaptu red4Mach 0waveAxa

Max thickness locationAxiaW

2 TE -8044

Figure 9. MCA airfoil definitions.

margin criteria. In the outer portion of the fan blade, where the inlet rela-

tive Mach number is eupersonic, the airfoils were shaped to minimize shock

loss. In the subsonic region of the blade, the airfoil shape transitions from

the first supersonic section down to a near double-circular arc airfoil sec-

tion at the hub.

10

A low aspect ratio, and, therefore, a long average chord, was selected to

meet flutter criteria without the use of part-span shrouds. The spanwise

chord taper was selected to satisfy the solidity requirements and also be

viable from a weight and stress standpoint. The radial distributions of chord

and solidity for the blade are shown in Figures 10 and 11, respectively. The

maximum thickness to chord ratio (Figure 12) was set to avoid responsive res-

onant conditions and to maintain radial uniformity of blade mechanical prop-

erties. One of the mechanical considerations in the design was blade integri-

ty with bird ingestion. The leading edge radius of the blade was set at

0.0125 in. from the hub to 60Z span and then tapered to 0.010 in. at the tip.

This maximized blade strength in the primary impact area while at the same

time minimizing the efficiency penalty in the high inlet Mach number area at

the blade tip from increased shock loss.

For the portion of the blade which has supersonic relative inlet Mach numbers,

incidence was set on the suction surface at a point halfway between the lead-

ing edge and the emanation point of the first captured Mach wave (point A' of

Figure 9). This incidence is the offset of the suction surface from a "free"

streamline, which would exist if there were no blade forces, and it establishes

the maximum flow the cascade can pass when the throat is not the limiting fac-

tor. The incidence value was set at 1.5 deg and is intended to account for

leading edge blockage, suction surface boundary layer, and the bow shock wave.

In the subsonic portion of the blade, the meanline incidence for each airfoil

section was selected to locate the throat near the passage inlet.

Deviation angles were calculated using a modified form of the NACA 2-D rule

for circular arc meanlines and then adding an empirical adjustment. The modi-

fication is a circulation correction based on the radius change of the stream-

line across the blade airfoil section.

The radial distributions of meanline incidence angle and deviation angle are

shown in Figure 13. The resulting meanline blade angles are shown in Figure 14.

11

1.5

14 t igure 10. Blade chord.

.0

1.4

1.2

1.0 I0 20 40 60 so 100


Figure 11. Blade solidity.

12

0.10

0.08

I0.06

00 20 40 60 s0 100


Figure 12. Blade maximum thickness/chord ratio.

The chordwise location of maximum thickness and circular arc inflection are

shown in Figure 15. These were selected for two reasons: (1) to avoid accel-

erating suction surface curvatures ahead of the anticipated passage inlet

shock wave location and (2) to set the passage inlet area and contour the

airfoil passage between the passage inlet and the throat to minimize the

velocity change through the passage. The design throat minimum critical area

ratio (A/A* min) distribution for the supersonic airfoil sections is set to

1.03 for a normal shock total pressure loss applied at the passage entrance

with a linear distribution of profile loss from the leading to trailing edge

of the airfoil section. Streamtube contraction and the effect of radius

change are accounted for.

Figure 16 shows the blade hub, mean, and tip conical airfoil sections in

engine orientation. For manufacturing purposes, the airfoil sections were

redefined on planes normal to the stacking line. The stack line is a radial

line passing through the center of gravity of the hub conical section. The

blade manufacturing coordinates are listed in Appendix C with definitions

given in Figure C-I. 13

16

14- - DeviationI nc ide nce

~12-

10

6-

-2120 0 0 80 100

-Percent span from hubTE-8048

Figure 13. Blade incidence and deviation angles.

70

60 -

50

140

-.- Inflection~20 - Ex it

0 20 40 60 so 100Percent span from hubTE84

Figure 14. Blade metal angles.

14

Sg -0 Max Thickness

0

C

.50

0 20 40 60 80 100Percent span from hub T 85

Figure 15. Blade maximum thickness and inflection locations.

Tip

Mean

Axial Hub

TE-8051

Figure 16. Blade conical airfoil sections.

VANE DESIGN

The vane is also made up of MCA airfoil sections. The vane axis is tilted

rearward from radial at an angle of 13.5 deg. This gives a more desirable

acoustic spacing between the blade and vane at the tip while minimizing hub

length for bypass engine applications. There are 42 vanes with an aspecttratio of 2.32 and a solidity of 1.78 at the I.D. and 1.30 at the O.D. (Figure

17). The radial distribution of chord is shown in Figure 18. The maximum

thickness to chord ratio varies linearly from 6% at the hub to 8% at the tip.

4 The incidence angles were selected to position the throat location at the vane

passage inlet. The passage throat margins were based on minimum loss cascade

data (Figure 19). Deviation angles were determined from the NACA 2-D rule

with an empirical adjustment. The incidence and deviation angles for the vane

are presented in Figure 20. Vane metal angles are shown in Figure 21. 'lane

hub, mean, and tip conical airfoil sections are illustrated in Figure 22.

2.0

~1.8

1.6

1.4

1.2

1.0 I

0 20 40 60 80 100Percent san from hub E8052

Figure 17. Vane solidity.

16

2.5

~2.0-4

1.0-

0 20 40 60 80 100Percent span from hub

TE -8053

Figure 18. Vane chord.

1.20

~1. 15

1.0

0 20 40 60 so 100Percent span from hub T 35

Figure 19. Vane passage throat minimum critical area ratio.

17

14-

OF 10 -------- - - -

8-

- Incidence6-Deiaio

-20 20 40 60 s0 100

f Percent span from hubTE -8055

Figure 20. Vane incidence and deviation angles.

60-

50

30 - -. - - - - - - - -

~230

20 InMet.......I nf! ection

10-Ei

0

-10

-20o 20 40 60 80 100

Percent span from hub TE-8056Figure 21. Vane metal angles.

18

TipHub Mean

TE-8057

4 Figure 22. Vane comical airfoil sections.

The manufacturing coordinates for the vane are given in Appendix C with per-

tinent airfoil section definitions given on Figure C-2. The section coordin-

ates were defined on planes normal to a stacking line. The stack line for the

vane is on a radial line passing through the vane hub section c.g.

19

...... ..19 . .. Ill'li~li~ . . ..

SECTION IV

STRUCTURAL ANALYSIS

Structural analysis of the fan rotor consisted of calculating airfoil and

wheel steady-state stresses, airfoil vibrational characteristics, and bird

ingestion capabilities. Satisfying all these requirements while retaining an

aerodynamically acceptable configuration required numerous iterations on the

design with changes in flow-path shape, spanwise chord and thickness distribu-

tion, and eventually a material change. The material change resulted from the

inability to analytically satisfy bird ingestion requirements in an aerodynam-

ically viable blade with titanium material. A stainless steel material (17-4

PH) was, therefore, substituted with consequent weight penalties. The final

design meets or exceeds the requirements of Mil-E-5007D.

A titanium rotor (Ti-6AI-4V) design was near completion when the material

change was deemed necessary. A satisfactory match-up of blade to wheel had

not been obtained, and a valid hot-to-cold run was yet to be completed when

the steel rotor analysis was started. Airfoil stresses reported here for the

titanium rotor were determined with internal program boundary conditions for

clamped hub, free tip. These automated constraints are not accurate for an

integral blade-wheel rotor; preliminary blade/wheel match-up attempts

indicated that airfoil crown stresses would be increased approximately 10 KSI

over the reported results while leading edge and trailing edge stresses would

be reduced.

The results of the analysis of the steel rotor are reported first. Titanium

rotor results follow in a skeletonized format.

STEADY-STATE STRESSES

Design criteria for the rotor are given in Table 1. No steady-state blade

stress will exceed 95% of the 0.2% yield strength of the material at 122% of

design mechanical speed. High cycle fatigue requirements for the blade are a

15,000-psi allowable vibratory stress at resonance with a Kt = 3.0 at lead-

ing and trailing edges to allow for foreign object damage. Low cycle fatigue

requirements for both wheel and blade are for greater than 12,000 start-stop

cycles (zero-to-maximun stress) with a reliability of 0.9999. Finally, the

wheel burst speed must exceed 130% of design speed.

20

TABLE I

Structural design criteria.

Blade

Permanent set 95% 0.2% yield @ 122% speed

Low cycle fatigue 12,000 start-stop cycles

High cycle fatigue 15 ksi allowable vibratory

stress at resonance. FOD (Kt

a 3.0) at leading and trailing

edge

Wheel

Wheel burst 130% speed

Low cycle fatigue 12,000 start-stop cycles

Design point stress analyses were performed for the fan rotor at mechanical

speeds of 20,223 and 18,950 rpm corresponding to 1.8 and 1.65 pressure ratios,

respectively. The analysis is accomplished with finite element computer

models which account for centrifugal loads, air loads, temperature effects,

airfoil tilts, airfoil untwist, and wheel deflection. Airfoil bending stress-

es were minimized by tilting the airfoil in the direction of air loads.

Steel Rotor

Airfoil principal stress levels on both suction and pressure surfaces are

shown for 1.8 and 1.65 pressure ratio design points in Figures 23 and 24, re-

spectively. The maximum level of 110 ksi occurs near the airfoil hub on both

suction and pressure surfaces at the higher pressure ratio and 100 ksi at the

lower pressure ratio. The 110 ksi local principal stress on the blade surface

compares with an average section stress of 61.5 ksi at design speed. To check

the requirement for no damaging permanent set at 122% of design speed, it is

21

;CIS

i' 9

P..

CY

I W

ai 9e

o an a v a a ~ i a i a vi a '22

0&

--

0 0 f a ar V m V

CaOU U.=,n

(N~ I l'sot

f 0!0 VI 0k uI w APC a a a

0 U I- 6 a an p 0 023

TABLE 2

Blade stress summary.

Allowable Stress CalculatedType of Failure Criteria and Location, ksi Section Average Stress

1.8 R., ksi 1.65 Re, ksi

Permanent set 95% FT @ 122Z speed 137 crown 91.5 80.4

Burst 95Z FTu 3 130% speed 151 crown 103.9 91.3

CalculatedSection Max Stress

1.8 Rc, ksi 1.65 Rc, ksi

Low cycle fatigue 12,000 cycles Kt = 3 66 lead edge 46.3 39.2@ 100% speed Kt a 3 66 trial edge 22.4 19.7

Kt - 1.4 125 crown 110.0 100.3

High cycle fatigue *15 ksi vibratory? resonance (Refer to Table 3)+5 ksi vibratory I nonresonance -5 ksi crown 412.0 -14.4

(required) (a'lowable) Tallowable)

necessary to scale the 61.5 ksi by the square of the speed ratio which gives

an average stress level of 91.5 ksi at the overspeed condition. As shown in

Table 2, this compares with an allowable stress of 137 ksi. Similarly, for a

check of failure at 130% speed, the average stress scales to 104 ksi which

compares with an allowable stress of 151 ksi.

Referring to the S-N diagram of Figure 25 at the 12,000-cycle requirement for

low cycle fatigue, the airfoil leading and trailing edge allowable stresses

are found to be 66 ksi based on a K- 3.0. The crown fillet allowable

stress is 125 ksi based on a Kr M 1.4. Again, referring to Table 2, the

calculated maximum principal stresses are well below these allowables.

The Goodman diagram of Figure 26 indicates that the 110 ksi maximum steady

stress at the hub fillet, in conjunction with a fillet radius concentration

factor of Kt - 1.4, provides a vibratory allowable stress at that location

of *12.0 ksi. Since this is not a potential resonance condition, a level of

*5 ksi would be considered satisfactory. The vibratory allowable stress at

the hub fillet at the lower pressure ratio condition is *14.4 ksi.

24

18 \ Low Cycle FatigueO1.-4 PH steel, 0 at 70'F

[o _

60 -

40

Cycles Izero max)TE -8060

Figure 25. Fan blade and wheel S-N diagram.

Blade vibration analysis, to be discussed more fully later, indicates two po-

tentially troublesome resonances in the operating envelope of the fan. These

resonances are a second engine order-first bend mode coincidence at 11,800 rpm

and a fourth engine order-first torsional mode coincidence at 18,800 rpm.

Allowing a reasonable scatter of individual blade frequencies, the maximum

static stresses at the maximum reasonance speed and at the critical vibratory

stress points are given in Table 3. Entering the Goodman diagram of Figure 26

with those static stresses and the appropriate Kt values, the allowable vi-

bratory stresses show.n in Table 3 are defined. All these allowables exceedthe goal of ti15 ksi vibratory allowable. It should also be noted that in the

torsional mode, the dynamic stresses at location A and D are 50%. and 85%, re-

spectively, of the maximum dynamic stress which occurs at location B.

5 4

so" J' . . . . " ' . . . . . . . . l . .. . . . . ~ elII lll I[ : +. .. . '

100Fan Slade

17-4 PH steel, H950 at 70OF(-3. 7 sigma Goodman diagram)

AA

I-I

20dngeg

Mea stes ns

Mean tress ksiTE-3061

Figure 26. Fan blade Goodman diagram.

TABLE 3

Steel blade stress summary at maximum dynamic response.

Hligh cycle fatigue allovables

17-4 PH steel SuctionPressure

Cast properties

Max Static Allowable

IResonance Stress at Vibratory

Mode Location Speed, rpm Resonance, ksi KStress, ksi

First bend A 11,800 35.0 1.4 +30.4

First torsion A 18,800 95 1.4 +15.7

D 23 3.0 +15.6

3 8 3.0 +17.3

26

The airfoil, therefore, meets all static and dynamic stress criteria with ap-

propriate Kt factors in cast 17-4 PH material.

Wheel equivalent stresses are shown in Figure 27. Referring to Table 4, a web

equivalent stress of 128 ksi at 122% of design speed compared with an allow-

able yield stress of 137 ksi assures no detrimental permanent set at that con-

dition. Checking wheel burst at 130Z of design speed finds a calculated mean

hoop stress of 102 ksi and a maximum web radial stress of 127 ksi at that con-

dition which compares with an allowable stress level of 137 ksi. Actual wheel

burst is assumed to occur when the mean hoop stress of the wheel reaches 95%

of the ultimate strength of the material. The burst speed thus calculated is

163% of design speed or 33,000 rpm. In terms of low cycle fatigue, the cal-

culated values of stress at rim, web, and bore are all well under the allow-

able stresses taken from the S-N diagram of Figure 25 at 12,000 cycles and the

appropriate K factors.t

TABLE 4

Steel wheel stress summary.

Allowable Stress Calculated

Tpe of Failure Criteria and Location, ksi scress, ksi

1.8 R 1.65 RC C

Permanent set 95Z FT @ 122% speed 137 web equiv. 128 112.4y

3urst 862 FT @ 130% speed 137 mean hoop :02 89.6u 137 web radial 127 111.5

Low cycle fatigue 12,000 cycles @ 100% speed

KT - 1.4 125 rim hoop 56 49

KTr a 2.0 (bolt hole) 92.5 web equiv. 86 75.5KT a 1.0 174 bore hoop 92.8 81.5

KT - 3.0 67 balance holes 40 35

Therefore, the wheel also meets all design stress criteria in cast 17-4 PH

material. The weight of the wheel and blades is approximately 16.3 lb.

27

/T

AI

7 , Stress (ksi) at design Doint:

\~ .1.830Altitude - 25000 ft

.7 I M 50

3_ Speed - Z2023 torm- aterial: 17-4PH steel

'tot Legend

40000 A

30000 F88000 GI Max stress 931080

TE -9

Figure 27. Wheel equivalent stresses.

Titanium Rotor

Blade stresses for the titanium rotor are summarized in Table 5 where allow-

able stresses for permanent set and burst are comfortably above calculated

stresses. In the area of low cycle fatigue however, calculated stresses ex-

ceed the allowable at the leading edge and equal the allowable stress at the

trailing edge for the 1.8 pressure ratio condition. These stresses are calcu-

lated for a rigidly clamped airfoil; including wheel rim flexibility would

lower the edge stresses significantly.

28

TAB LE 5

Blade stress snary.Allowable Stress Calculated

Tve of Failure Criteria and Location. ksi Section Average Stress

' .8 R_, ksi 1.65 Rc, ksi

Permanent set 95% F, 3 122% speed 92 crown 57.5 50.5

Burst 95% FT T' 130% speed 99 crown 65.3 57.4u

Calculated

Section Max Stress

1.8 Rc, ksj 1.65 1_, ksi

Low cycle fatigue 12,000 cycles K . 3 39 lead edge '0

3 100% speed K - 3 39 trail edge 29 34.5

K - 1.4 77 crown 53 !.7

High cycle fatigue -15 ksi vibratory @ resonance (Refer to 7able 3)

-5 ksi vibratory ? nonresonance -5 ksi crown 10.0 1I.1

(required) (alawable) (allowable)

Referring to Table 6, there are two resonances in the operating envelope of

the titanim fan. The second engine order-first bend mode coincidence at

11,500 rpm produces a maximum vibratory response at location A where the sta-

tic stress is 11.6 KSI. With a fillet radius concentration factor of 1.4 the

allowable stress iq cast titanium is -17.5 KSI which exceeds the requirement

of +15 KSI. Similarly, the fourth engine order-first torsional mode coinci-

dence at 18,800 rpm prr.duces an acceptable allowable dynamic stress of +18.4

KSI.

The titanium wheel stresses are s,-gmarized in Table 7 where the allowable

stresses are seen to exceed calculated stresses by comfortable margins for all

conditions. The calculated burst speed of the titanium wheel is 34,500 rpm or

171% of design speed.

The weight of the titanium wheel and blades is approximately 9.3 lbs, or ap-

proximately 7 lbs less than the steel rotor.

29

TA1 LE 6

Titanium blade stress sumary at maximum dynamic response.

High cycle fatigue allowables

Titanium-6-4Cast

Max Static Allowable

Resonance Stress at Vibratory

Mode Location Speed, rpm Resonance, ksi K Stress, ksi

First bend A 11,500 11.6 1.4 +17.5

First torsion B 18,800 7.1 1.4 +18.4

TABLE 7

Titanium wheel stress sutmary.

Allowable Stress Calculated

Type of Failure Criteria and Location, ksi stress, ksi

I.8R 1.65 R4c C

Permanent set 95% FT @ 122% speed 92 web equiv. 79 69

Burst 86% FTY A 130% speed 89 mean hoop 56.7 49.8U 89 web radial 79 69

Low cycle fatigue 12,000 cycles @ 100% speed

T 1.4 77 rim hoop 32 28

KT = 2.0 56 web equiv. 53 47

KT - 1.0 102 bore hoop 52 46

VIBRATION ANALYSIS

Dynamic analyses of the airfoil, both vibration and flutter, are unaffected by

the material change.

30

Frequencies, mode shapes, and relative dynamic stress distributions for allmodes up through the vane passage frequency (42 EO) were calculated, using

finite element techniques. The frequency versus speed interference diagram

showing the first three modes is shown in Figure 28. The overall interference

diagram is presented in Figure 29. Note that the first bending mode (1B)

crosses 2 EO at relatively low speed (60%) such that the excitation levels due

to inlet distortion will be low.

The relative dynamic stress distributions are determined to locate the maximum

dynamic stress location for each mode to assess, in combination with the

steady-state stress calculation, the allowable vibratory stress levels. Fig-

ures 30 and 31 show the relative radial dynamic stress distributions for the

first two modes. These first two modes are given particular emphasis since

the excitation force levels produced by the lover four engine orders are us-

100%

1200

1400-

130

0 2 4 6 8 10 12 14 16 18 20 22 24Rotor speed-trm - 1000

TE-8063

Figure 28. Frequency-speed interference diagram (first 3 modes).i

311

1004,

16000 142__--"---'--" vanes14000

12000 ________0_2________20__

S10000 -

R60004000

I Lyre2000 2

0 2 4 6 8 10 12 14 16 18 20 22Rotor speed -rpm 1000

TE-8064

Figure 29. Frequency-speed interference diagram (first 32 modes).

ually higher than those generated by higher harmonics of rotor rotation. As

previously discussed, the results of the allowable vibratory stress determina-

tion using the Goodman diagram, predicted steady-state stress at coincidence

speed, and relative dynamic stress distributions satisfied the *15 ksi vibra-

tory stress criteria.

Blade response because of a coincidence of high modes with vane passage (42

EO) are expected to be very low because of the large axial distance (1.5 chord

lengths) that the vane row is located aft of the blade row. This large

spacing is a result of the noise design criteria.

FLUTTER ANALYSIS

The results of the torsional stall flutter analysis are shown in Figure 32.

The predicted margin of safety is an adequate 4 deg of incidence angle above

the estimated operating line. The calculated bending stall flutter reduced

frequency parameter is well above the 0.25 criterion at 0.302. Similarily,

the supersonic unstalled reduced torsional frequency parameter at 0.68 satis-

fies the 0.60 requirement.

32

4-

43

4'

N 00 -4

U- ~- vi 43

4343 4' o0 2 4-. 4343

-~ '-'a ~ 0t'aw~ - N -

I'- .~ ~4

N 4N

4' 43~ -

- -.

-4

0-4

~1' -a-4

-000

-~$ a.'-a0

43,-0~0 U

-4

43

00

i43 -43 '0

0a.'0

43-4

A

~-' '- 4,'- -

43 4.14

43~

43,-

33

C1WlU

-,v

00-

'..61

030

6.0

66 -w -4

6-4

vt2212

it - 34

Flutter regime16

14 7Stable operation

12 Stall flutterbou ndary

a 10 50 4deg incidence margin

6,0

6 2 1 2 1 160 18-Jn i t e l t v e v e o c t - t / e cx7o0T18

6Figure 32. St1 0 to si nl flut e m , a yIis

35mtdoertn n

BIRD INGESTION ANALYSIS

The USAF requirements for bird ingestion are defined in Mil-E-5007D, paragraph

3.2.5.6.1. A sumary of the specified bird sizes and engine conditions is

given in Table 8. Tb small high bypass fan and typical trainer aircraft

speeds corresponding to the Hil-E-5007D requirements are given in Table 9.

In Table 9, the aircraft liftoff, climb, cruise, and descent speeds for the

1.80 pressure ratio fan are assumed to be the same as for the 1.65 pressure

ratio fan application.

The annular inlet area of this fan is less than 200 in. 2 which sets the max-

imum bird size at two pounds (Ref. Mil-E-5007D par 3.2.5.6.1e). For bird im-

pacts up to 2 ib, no failure shall result which will cause shutdown of the

engine although some damage to engine parts may occur.

The failure mode considered here for bird ingestion is local impact damage in

the leading edge region of the blade. The calculation of a local damage index,

based on a Lycoming criterion approach (Ref. FAA-RD-77-55), has been incorpo-

rated into the DDA bird ingestion analysis. This approach relates significant

bird slice, impact area, and airfoil parameters to a damage index value that

corresponds to critical blade damage. An acceptable damage index level is

determined by correlation with actual bird ingestion test data. Engine sur-

vivability for new blade designs at the critical ingestion conditions is then

predicted with some confidence by use of the damage index calculation.

The shear-penetration damage index is expressed as

D. *0.273 VN

36

where:

V the normal component of impact velocity (ft/sec)

3e the bird density 0.045 lb/in.

DB the bird diameter (in.)

h the target mid-thickness (in.)

= the target material shear yield (psi)y

KB the bird fragmentation parameter

TABLE 8

Mil-E-5007D bird ingestion requirements.

A. Birds weighing 2 to 4 ounces (a macimum of sixteen at a time) and birds

weighing 2 pounds (one at a Cime) ingested at a bird velocity equal to the

take-off flight speed, with the engine at maximum rated speed.

B. Birds weighing 2 to 4 ounces (a maximum of sixteen at a time) and birds

weighing 2 pounds (one at a time) ingested at a bird velocity equal to the

cruise flight speed with the engine at maximum continuous speed.

C. Birds weighing 2 to 4 ounces (a maximum of sixteen at a time) and birds

weighing 2 pounds (one at a time) ingested at a bird velocity equal to the

descent flight speed with the engine at an associated engine speed.

D. Birds weighing 4 pounds ingested at a bird velocity based on the most

critical flight speed with the engine at maximum rated speed.

Note: Condition D does not apply since the trainer fan inlet is less than 200

in 2 . A maximum of four birds weighing 2 to 4 ounces must be consid-

ered for the trainer fan.

37

TABLE 9

Fan and aircraft speeds for Mil-E-5007D conditions.

Pressure ratio Pressure ratio

1.65 1.80

Condition A

Engine take-off fan speed (rpm) 17,303 18,465

Lift-off speed of typical aircraft (kts) 90 90

Climb speed of typical aircraft (kts) 198 198

Condition B

Engine maximum continuous fan speed (rpm) 17,292 18,454

Cruise speed of typical aircraft (kts) 194 194

Condition C

" Engine descent fan speed (rpm) 10,377 11,074

Descent speed of typical aircraft (kts) 198 198

Condition D

Engine cruise fan speed (rpm) 10,496 11,201

Cruise speed of typical aircraft (kts) 194 194

The values of DB and KB are defined by:

D 3.48 (bird weight)0 33

DB

K 3 = 1.0 VU 260ft/sec

S2.40-0.0OS4VA for 260 < VN4400

2.5 400 < VN

Figure 33 illustrates the region of leading edge impact.

38

Airfoil thickness

/12

Bird diameter

Impact Impact radiusforce F

= bird slice length

TE-8068

Figure 33. Bird impact area and blade thickness illustration.

The ballistic-limit velocity for total shear penetration is reached as the

damage index approaches 1.0 and theoretically should cause maximum structural

damage. In practice, the maximum allowable damage index must be determined by

correlation with test.

The calculated damage index along the airfold span for a base-line titanium

design at a 1.80 pressure ratio is shown in Figure 34. The maximum damage is

predicted to occur at the outermost radial position permitted by the outer

case and the 2-lb bird diameter. The maximum allowable damage index for

shroudless airfoils is set at 0.40 which limits the expected damage to span-

wise tears. A damage index greater than 0.40 could produce loss of airfoil

section (for several blades at the 2-lb size) and would cause rotor unbalance

that would require engine shut-down. On this basis, the base-line titanium

design is judged unacceptable.

39

Tip R 8. 47

8.0

Take off.0 escent 2-1b 3-oz size bird

7.0 size bird -Climb an

I Z cruise at max continuous2 -lb size bird

d 6.0 "

I2-ib size bird

L• "b R - 3. 41 I . Acceptab~le damage Iia it3.0,

0.1 0.2 0.3 0.4 0.5 0.6 0.l 0.8Damage index

TE-3071

Figure 34. Bird ingestion damage index--titanium blade.

A design study was made to establish the damage index sensitivity to leading

edge radius, thickness/chord, number of blades, etc, to identify an acceptable

damage index domain for the titanium fan. A titanium airfoil with a 0.025-in.

leading edge radius was found to have an acceptable index. However, the per-

formance penalties associated with the required design changes were excessive.

Thus, based on the bird ingestion requirements (which are particularly impor-

tant for trainer engines), a switch to 17-4 PH stainless material was made.

The advantage of the material is an increase in shear strength from 66.5 to

106 kIi.

The calculated damage index along the airfoil span for the steel design is

shown in Figure 35. Again, the maximum index value of 0.39 is found to occur

at the outermost radial position permitted by the outer case for the 2-lb size

bird. The steel design (with a 0.0125 leading edge radius) thus satisfies the

acceptable damage index limit of 0.40 for the this type of blading.

40

Tip R 8.47 (NOTE: Reduced rotor speeds at the 1. 65pressure ratio would give reduced

damage index values).8.0

Descent7.0 . 2-4b bird I

Take off and climb and6.cruise at max continuous6.0 2-lb size bird

5.0 Take off

'I 3-oz size bird

... - cceptable damage limit

4 HubR - 3.41 i3.0 I { i0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.3

Damage indexTE-90T2

Figure 35. Bird ingestion damage index-steel blade.

Bird ingestion requirements have forced a change in leading edge thickness and

material for this fan blade. The penalties are approximately one-half per-

centage point in stage efficiency and an increase in rotor weight from 9.25 to

16.25 lb. Any degradation in engine response during power transients have not

been quantified.

41

SECTION V

NOISE PREDICTION

Noise goals for the high bypass turbofan powered undergraduate trainer are to

be in compliance with Federal Air Regulation Part 36 requirements and maintain

ground idle noise substantially below current levels. Engine cycle and fan

design data were combined to estimate trainer noise levels for Part 36 and

idle conditions. A brief description of the noise prediction methods used and

the noise levels estimated for the undergraduate trainer are presented in this

section.

Estimated noise levels for a PD 418 powered trainer are presented in Table 10

and show that the above noise goals are met.

TABLE 10

Undergraduate trainer noise levels.

FAR Part 36 Levels (EPNdB)

Takeoff Approach

Part 36 Requirements 89 98

PD418 Powered Trainer 78 93

Ground Idle Tone Corrected Perceived Noise Levels

PNdBt at 250-ft Radius (Single Engine)

Front Rear

T37 with J69-T-25 118.3 99

PD418 Powered Trainer 79 84

Noise estimates were made through a DDA computer program developed for turbo-

fan noise prediction. A noise generation model for each source, fan, jet,

turbine, and combustor, is contained in the program so that the noise output

42

from each source is dependent on its individual operating conditions and en-

gine total noise reflects the contribution from each noise source.

High bypass ratio turbofan engines are usually fan noise dominated with the

jet and combustor being secondary sources. The PD418 engine incorporates the

following design features to reduce fan and jet noise generation:

o Ample space between the fan and the outlet guide vanes (1.5 fan chords)

o Ratio of outlet guide vanes to fan blades >2 to cut off blade passing tone

o Internal mixer to reduce nozzle exit velocity

The noise estimates for the PD418 include the noise reductions provided by

these features. The PD418 incorporates a 1.8:1 pressure ratio fan (design

point) matched to the GMA300 core and operates at part speed for takeoff,

climb, and approach. The engine is fan noise dominated at these conditions so

that fan duct acoustic treatment could be used to achieve levels lower than

the 78 and 93 EPNdB predicted for takeoff and approach. These levels are 11

and 5 EPNdB below the FAR Part 36 requirements.

At the ground idle condition, the PD418 peak levels at 250-ft radius are ex-

pected to be 79 and 84 PNdBt (tone corrected PNdB) in front and rear of the

aircraft. In dBA units, front and rear levels are 60 and 67. These levels

translate into noise reductions of 34 PNdBt or 35 dBA when the PD418 trainer

is compared with the T37B with the J69-T-25 engine.* In predicting the ground

idle noise levels, it was assumed that ingested ground level turbulence would

prevent cutoff and so the blade passing tone is included in the above levels.

*Speakman, 1. D., Power R. G., and Lee, R. A., Community Noise Exposure

Resulting from Aircraft Operations, AMRL-TR-73-110, Vol. 4, February 1978.

43

6 6 - U .. ... .

APPENDIX A

AXIAL COMPRESSOR DESIGN SYSTEM

The vector diagram calculation used for axial compressor design assumes an

axisymmetric flow field and obtains a solution of the continuity, energy, and

radial equilibrium equations. The design analysis is identified as the Axial

Compressor Design System (ACDS) Program BD76. Viscous terms are omitted; how-

ever, the equations do account for streamline curvature, radial gradients of

total enthalpy and entropy, and blade force terms arising from non-radial

blade surfaces. Calculations may be performed at the leading or trailing

edges of the airfoils by slanting the calculation stations.

Enthalpy rise across a rotor is given by Euler's turbine equation, and the

continuity equation is adjusted for local as well as endwall blockage.

tUsed as a design tool, the calculation provides detailed examination of the

aerothermodynamic solution of the floov process through the compressor. The

solution is iterative and must rely on profile loss estimates which are cor-

related as a function of aerodynamic loading (diffusion factor). This data

has been obtained from test data for a wide range of compressor designs and is

continually updated.

The equilibrium equation is in the form of:

dV 2 d(V82) - d (Vr) [(d2 ds 1,- - _i _ + -T

dr c dr dr L dr iIdrr..dVr VdV dz

- - ZV -+ 2V- 2[ z dz r z dz d,

d(rV*) d

2V - -IZdz dr ;

44

where:

r radial distance

z axial distance

5,0 tangential distance

V radial velocityrV axial velocityzV0 tangential veolicty

T total temperature

8 entropy

HO total enthalpy

c projection of the calculating station on relative stress surface

0 relative to stream surface

The continuity equation is:

W a 2 7rfK, p V sin (k-e) rdy

:" { Yh

where:

W airflowaV meridional velocity

mK7 blockage factor

P density

Y length along the calculating station

angle between tangent to the streamline projected on the

meridional plan and axial direction

angle between calculation station and axial

45

APP Ml IX B

IDESIGN POINT VECTOR DIAGRAMS

46

~~~~~~~ .~~*.V D~ . .-j.p-... .A,.. . .

x,446

14. . . . ..............

44 4J.~ P4 4vccW zw a 444440 WW

Saw. ...........

-~ ~: 34N4NO

NM IceOOOOC OO4_ *13w

'M~4, ...~ ,' .....

QR -".4 'A 115 4sE..... .. .w A.. ... ~

- 4 'M4 0.0N01A.D4 .410 a11 2.- ---- -Sa. -A - 2 4 A'S ... E

4 19*4s1. 54.J. . . . ..#- 0..

Q. Qcw,.'Q.

.~* - oomemooo

..... ~......-t Q C=_1 4 4f.C.'~

2w Uwa~v~sAE=W.,.

.11~~l -roam MEOO~O .

.Ao O -,S OQLe O~

'MA ~ .. . . . .Ml~ 4 . . . . .~O-~~I-.1e47

n.

Z.C .44-4. .jU ~ ~ z '. 73a~A, ~ ~ 4A0

Q.-W . i

zo. A~ . :..

4.. u.

-WXI 4..-W-C 00 0 0

-W-J WOi. J

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

..........

-0 46O 0 .- 9a c c c

ZmM ..- 4% -A . . . .

-z 4=00 .--.. 4- '&00-co coe 2

4"d

.d -.

......... ......... -

4W !u P#.U .4460-

W 49-~~A.l

48-N N-~.

I

APPENDIX C

ROTOR AND STATOR BLADE COORDINATES

49

Leading edge Trailing edge

!.__Engine cente r "Sakpit_ (

*+Setting angle

TReference airfoil thickness DoealangleD - Distance to leading edgeW - Setting angle minus dovetail angle T 86

Rotation is counterclockwise from the rear

Figure C-1. Blade manufacturing dimension definitions.

so

Leading edge Trailing edge

R S,-

(+Y)

StackpointEngine center I ine

T -Setting angle

(-x) (+x)

T -Reference airfoil thickness0 -Distance to leading edge

XI -Y)Rotation is counterclockwise from the rear

TE-8070

Figure C-2. Vane manufacturing dimensioni definition$.

A 51

mil

23 AUG 7q PAGE

FAN COMPRESSOR SLACE CaC 12120

REFERENCE DISTANCERAOIAL SETTING AIRFOIL TO LEAOING RADII

OISTANCE ANGLE THICKNESS EDGE L.E. T.E.

3.2500 -9.175 0.2982 1.300C 0.014 0.015-9C 1ON 26S

LEADIG EDGE AXIAL TANGENT POINT -0.9448

Q OIMENjICN 0.0368 U DIMENSION 0.0322R DIMEN ICN C.0600 S CIMENSION 0.0600

STACK PCINT CCOROINATES 0.0 , 0.3CENTER OF GRAVITY CCOROINATES 0.1476 -Ca!87qCCNPlESSOR ROTATION IS COUNTER CLOCKWISE FR P THE REAR

REFERENCE COORCINATE POINTS

STATICN

No. X Y X Y x X

1 -0.8491* -C.6670 21 0.413 0.1936 4 0.4470 -0.11492 -4.8472 -c.6594 12 0.564 0.1567 42 0.3554 :3.oQN63 -4.8133 -0.5831 23 0.5919 0. 073 43 0.2831 .0802_4 -0.7769 -0.5081 4 0.6987 0.c2e6 44 C.1943 -G.97435 -0.7276 -(.4167 25 0.736 -0.0540 45 0.1243 -0.07398 -4.680 -C. 3 57 26 0.8675 -0.1494 46 0.0556 -0.u828

-0.6396 c.2 69 27 0.9703 -. 2901 47 0.C286 -C.09828 -0.5789 -C.1943 28 1.0501 -0.4210 48 -0.0940 -4.1160q -0.5271 -C.1314 29 1.1461 -0.6077 49 -0.1579 -0.1386

tO -0.4724 -C.o717 30 1.1643 -0.6482 54 -0.J360 -0.172911 -4.3q98 -C.0023 3 .1706* -0.6664 51 0. 969 - .2C4812 -. 3383 C.0486 32 .1469 -0.6659 52 -0.3713 -0.2499

13 -0.2574 C.1056 33 1.1267 -(.6388 53 -0.42q4 -0.289814 -0.18q3 C.1453 34 1.0250 -0.5158 54 -0.4864 -0.3329

15 -Q.1181 C.179 35 0.9448 -0.4311 55 -0.5560 -0.3907

16 -0.0250 C.2112 36 0.8*6L -0.3410 56 -0.6105 -0.440017 0.0521 C.J282 37 0.7682 -0.ZT7q 57 -0.6639 -0.491818 0.1314 .32 38 u.6903 -4.2277 58 -0.7292 -0.559819 0.2329 C.2359 39 0.5962 -. 1745 59 -0.7803 -C.6166

20 0.3160 C.2237 4C 3.5212 -0.1410 6C -0.8304 -C.6754

a INDICATES EXTREME POINTS

23 AUG 79 PAGE 3

FAN COMPRESSOR @LACE C8C 12120

REFERENCE DISTANCERACIAL SETTING AIRFOIL TO LEADING RACII

OISTANCE ANGLE THICKNESS E3GE L.E. T.E.

3.9500 11.933 0.2241 1.300. C.013 0.014i0 SSN SOS

LEADING EDGE AXIAL TANGENT POINT -0.9108

C .f337 U C'MINSISN 0.831tRI 8 1 CMH I C.QeOO0 5 0 IM NSI N 0

STACX PCINT CCOROINATES 0.0 1 C.3E ENTER CF GRAVITY CCOROINATES C.0754, -0.3674COPPRESSOR ROTATION IS COUNTER CLOCKWISE FRCM THE REAR


STAT I N

N4. X Y X y x y

1 -10362 -. 5001 21 1.4402 0.1430 41 0.4084 -0.07312 :101 -C.4a6 12 0.5133 k.1149 42 C.J4 :1:2-0.5561 -. 9711 -C.4315 3 0.5640 0. 0776 4 6 0.04754 -1.91 -. 3758 24 3 0.0172 44 81712S -8321 -. 3083 25 0.7317 0.0429 4 102 -0.04686-.768 -. 2,6 1 6 0.7 910 0. 3 46 0.0336

-o.70o3, -B o,' .8585 -0.IMI 1: -. 56"-0.619q -C.144 8 3.9464 -0.3L484 -0.221 4 14 4S - . -0.046

- .39 -0.00a 31 0.96qi1 -0.4940 St -0.3494 -0.15712 -0.3205 008 32 0.9520 -0.49 9 5 -0.4364 1313 -0 924 0.0760 33 0. 9 -0.4741 53 -. 5099~ 414 - 536 .1073 34 0.82 -0.3e9 54 - .1T 5 7615 -. 0780 C.L324 35 0.81 -0.3182 55 -0.6:1, :.3i1616 11 0.569 36 0.7423 -0.2481 56 -. 7316 339117 0:0 14 C. 172 37 0.6816 -0.200?6 5 - .7987 37710.6114 :0:216 5

17S9 0.5397 1 5 -0.915 421'

a 3 61 i.L66 40 0.446 .q31 60 -1.192 .5093

0 INOICATES EXTREME POINTS

52

23 AUG 79 PAGE S

FAN COMPRESSOR 8LACE Cac 12120

REFERENCE DISTANCERACIAL SETTING AiRFliL TO LSACING RADIIDISTANCE ANGLE THICKNESS EOGE L.E. T.E.

4.6500 29.957 C.1476 1.300C O.J13 C.314290 57" 25S


Q 0IMENSICN C.0323 U CIMENSICN 0.0375A DIMENSCN 0.06ou S 0INENSION 0.0600

STACK POINT CCOROINATES 0.0 , C.0CENTER CF GRAVITY CCCRDINATES .0286 -C.3344CC"PAESSOR POrArroN is cauNrER C LcCWSE FR6 THE REEAR


STATICN

NC. X y X X y

S:1:171* :C:2951 21 0.4744 0.1136 41 0:4503 106-0.2827 22 0.5525 a.oql 42 0.3575 -at

3 -1.0928 -0.2472 23 0.6287 0.0747 43 0.2824 -0.01614 -. 0167 -C.J4 0.7205 0.03e2 44 C.1877 -0.0L815 0:9210 -'.M1 5 0.7906 0.0017 45 0.1115 -4.0236& -0.8441 -1395 26 0.8554 -0.0425 46 _.0351 -0.03237 -0.1667 -0.1089 27 0.9283 -0.I086 47 -0.0619 -0.0473a -0.6695 -C.0726 28 0.9802 -0.1C3 48 -00p

0 3 4 0.8610

9 -0.59t3 -q.0452 Z9 1.0359 -C.2586 4 -0.186 -0.275710 -a.S12? -t.0191 30 L.0457 -0.2778 50 L.166 -4.095211 -0.4L40 CAW 31 1.0462 -. 2,e79 51 -C.3950 -0.111812 -0.3347 C 033 32 1.0318 -0.2996 52 -0.4931 -0.13353 -0.2352 C.0594 33 1.0191 -0.2658 53 -0.57t6 -. 15174 -0.15S2 C.0780 34 0.9514 -0.2228 54 -U.6501 -C.1705

is -0.0751 0.0947 35 U.8936 -0.17S3 55 -0.7484 -0.191916 0.026U C.1129 36 0.8171 -0.1328 56 -0.8270 -0.215117 0.1085 0.1236 37 0.7529 -0.1015 57 -0.9057 -. 23588 0.19tu %,.1299 38 0.6843 -0.0764 S8 -1.0042 -0.2623

19 0.2934 C.1311 39 0.5959 -0.1518 59 -1831 -0.2e4L20 0.3746 3.L266 40 0.5236 -0.370 60 -1.1821 -0.3C62

*INDICATES EXTREME POINTS

23 AUG 79 PAGE 7

FAN COMPRESSOR SLACE C8C 12120

REFERENCE DISTANCERACIAL SETTING AIRFOIL TO LEACING RADII

DISTANCE ANGLE THICKNESS EDGE L.E. T.E.

5.4000 43.936 0.L180 1.3000 0.013 0.01343C 56N 9S

LEADING ECGE AXIAL TANGENT POINT -0.8401

C G OINSICN C.0305 U DIMENSION 0.0358R a MENSION 0.0600 S CIMENSION 0.0600

STACK POINT CCOROINATES 0.0 , c.4GNTER QF GRAVITY CCCROINArE5 :60196, -C.0104

opNES50R RaTATION IS COUNTER CLOCKbWISE FC THE REAR


STATICNNC. X Y X Y X y

1 -1.28620 -0.1293 21 0.4983 0.0712 41 0.4873 -0.03032 -12752 -5.162 0.532 3Q~1 42 0.3832 - *07

f-.L9o -2 094s0.6675 C. 0574 43 3.2998-. 1049 -00776 24 0 .7720 0.386 44 0.1953 -C.0278

S -0.9984 1:C 0552 f5 0:8547 0:0203 45 G..1116 -4.4302.1 9 :6:U2.J23 J7 o0369 -U.0327 4? -0.77906 -0.213 -. 01 8g 111 -010620 4a -0.1827 -4.4453

9 -:0.6360 . 2 29 1.2121 -0.1041 4 0.2475 -0.450810 -0.5507 0.J231 30 1.2310 -Q.1133 51 -0.3536 -0.058Z

1- .44 7213 - .2518 0.0605 33 1.2024 -. 1335 53 -0.630L -. 079514 :2.13 1.09 34 1.24 -. 81 A- -. 066

8 C0763 35 1.0 17 -0.il - .82 -0.095916 162 9 . 639 36 0: 1 C6 -0.-0.107 -0.1038

19 3C59 C.0893 39 0.6531 -;.j4C4 5 - 1.1865 -0.131620 2*.3q6 C.JR60 4C 0.5703 -0.0345 6c -1.2725 -Q.1408

* INOICArES EXTREME POINTS

53

23 AUG 79 PAGE 9FAN COMPRESSOR SLACE C8C 12120

REFERENCE DISTANCEOIJNC SETTING AIRF~j TO MSSN ADII

NG5 LE~ THIC4s ~IG L.E. I.E.640 5467 0.0933 1.3000 0.013 C.013


0 OIRlENS[CN C.2287 U 01;INSION 0.0318A DIMENSION C.0600 SC NNSION 0.0600

SrACK POINT CORIATS4.0 1 1..0CE 4OFGRVIY CODIATS 0.u127 C.S2167

ST~ Ar1 iE FROM THE REAR aiei[ -t4tl - 539 21 0.51 34 4 .18 4062 -1*04 .C0071 H22 01 CJ?6 4 .97 0073 :1.319 C a01 23 0.6990 C.0654 43 0.3066 -0.0174 -1.1172 C.0098 24 0.8138 0.0571 44 0.1924 -10.01925 -t.1 04 C.0194 25 0.9054 0.045C 45 0.1007? 400

1 005 26 .9969 30397 46 0.0090 0217 :1:9101 :032 27 1.11L2 C.3264 47 -0.1057 QC

a -4.7979 C.04111111 8 1.2024 0.4143 48 -01 6 G zo15 -0.41 5 .3 0 71 31 1.1060* -0.0182 51 -0.9369 -C0i250

FAN0.89 CO.0ES 55 33AC C.3L C 00110 1.04 06

75 .05 S 044 0.07 8s 1.1 .0C 2 5 0.9L36 0.011

LEAI6 EDGE9 CAIA 4 rANEN PON0. 81609 5 56 -. 67 -0.#217 OI .101C .058 37 0.E 999O 0 .12 7 -. ~9-.02zc 3.4NSCN C.0773 S 4IMNSO 0.0 60. 0 6 144 001

SCKIN C RIATES 0.TEM POINTS

FNE CFMPESRAT BLCDNAE 008 1C10

REEEC ORCINAT E DISAINTSCI

8:8A2 j *54? 0. O4S 41G L.546 0

7..000 6,1.23 0.756 1.0.^ 4 0.12 .0 68

200 1.0 4430

- .~vSc '0 020 U IESO 0552.96

REFERENCETCSCEXTRETE POINTS

STATIC4

21 AUG ?9 PAGE 13FAN COMPRESSOR ILACE CDC 12L20

REFERENCE DISTANCERIACIA6 SETTING AIRFOIkS TC LEACING RADIXsrh.4c ANGLE TH CKNES EDGE L.E. T.E.

8.2500 64.569 0.1)724 1.30c.^ 1.010 C.010

LE401IAG EDGE AXIAL TANGENT POINT -0.8253

SCK. ON R I NACS3. 0.6

JE.I 2 -1.73674 23 0 083

EJR:VTaSC U W CP 1 TH2

REFERC CER EPONCETSStC

THITNSSTOE CX

:I~~yANGEN PON -.?3 0. 8L2 CCC7 0.ATES . 19, 2 0.40 0.313 ~ ~ ~ ~ ~ ~ ONE CLC1S 6FRH 1.HE89 23 0825 .19 43 0390R094 ~ ~ ~ ~ CODNT POINTS00 4 .44 .l3 44 OLO3 008

9 : 360 .031 13 1.54 0116 5 .I94N.28

6 1:561 094 6 1166 0.117 4 :0 5 0558

10 AUG 7V PAGI I

*E)~Ca ~2ISTANCRADIAL SE I tro. ^I,(PtJIL T4S L.EAU1NG RADII

01 s AN~i A fo IA TmLC~ftSkS E0(k L.E0 T.5.

LghADLNI. EUa AAIA#. TAfthf.tT #"UitT -. *;.27

Q~ jIMnt4la3N 0.32'.Z (1 4t~N.%Z(* 0.021--R DIQMSIONO 0.0*00 4 40ZPEfhsbuN 0.000STACK POIMT COORDOINATES 040 1 0.0

0.o(f i( GAAV11.V CJURU1#,ArES -0.1099 -,.05L6%U4(t,%2u.Q ADTATION IS 4 uTo QCL.AWLSEE t;,QM Tme REAR

*Eewtmcx C008004Tqa POINTS

STAT! CM

1. -"4s52 zi 0.1,320 -0.0459 1L 0. 1#4 -. 1060 09 02692 22 .104 .0396 1#2 006 -

10 -0.6358, * *0-9 .2s5* -0.0 at" -0.0291 -0 .o26Cz

5~ 01.0 0.32 2 0.2973 --0:006 1- -0 . 04f -0.1269

-9 0.0L9 3?03 002 4 0 "9

9 -).3994 S0.03 VA 9 0C450 0312 129 .28 017

11-.37 -Z.1-0.0 7534L O-ot 0.0450 0.0:08f,5 0469Z3 - fS9-0 .374& 31 ,.r S4 S 049 005

0w 0:1 9EDGE AIM7 34NG0N 494INT -0.5719 -. *0 00215 OML3SIO .48 ftdp 0INENS t 0?10200.06 -005ft OI0.0SIOI -0.0*0 36 0.3453Z 0.32000542 -000

1 1 -. .fts 0tIRATES 0.0 ,004d 5 -.5as - 00

19, 0043 0 0.1Apd 3-P 00469 -0.3 5 711:1

IN-- ICATkS &CXIAPITS ONS-

Z -. 05* 0105 2 00520 0.74032 4 0.035 (Y.06*4

R o- "EQ17rrT0.0600 Z4 01MENION0 ~00600 ~g w0iS [email protected]* 0.0*6 25 379 0 .027 45 0.011 -00

9 -0.3392* 0.0309 21 0.241 -0.0093 41 5526 -0:07"00 -0.2909t 226 30 0.37 -0. 00 50 9673 __o 064

21024 .18 3 0.544 0.04 9*3 5 0.13Z -0083 --9ok :4 - 3£ 0.3410- %0061 - _ MZ . Q7" y a

0.00 3353 QM3 Z 0.537 0.065* 53 00.3199 -0.031-0. tO5 -00CSZ 34 0018 0059 4 - 34 0952

5-.05 -0.009 35 0- .4 0.30 55 -0.8683 - .pa,*M* -0z1o-3 0.452 0.00r3? 4r -0.87 0.0133

91 03392 -0.03 37 00.979 -00095 579 -00.55 -002750

3 :," 0001 3 0.593 0:08O 32 35 -0.53 0.01

0 M -. 07 7 0:30979 -0.054 59 -0.5299 0.0299 0.1101 ::018?TZ 092o389 -0.014 59-)629 00

*INDIC ATES IA ThEN POIN4TS

56

IC AUG 79 pAct& 54Ats .MPRC.,U~A 'VAft CSC 12122

RIEERNCE 0 IS TA4C ESF.[TkG £AL-qt..jl ri LEAOING RAOZI

5.1500 -22.56T C.3aCC 0.7453 0.308 0.005-Z20 :31* 'S

6tAUINrv .Cj(c AAAL rAh.cfvr W%41NE -..377

(Q JIlft4410f4 0.0254 j .4t4451 0 ae.25A O1MtIt4~uf r..0600 j.imem~izu: 0.0.00

L-r.Ttp4 cjI 4-44VOCT C3C. UJ1NArcS 1.1169 0.209J'PktS~o% mJT&rt1% i .~c~u.43m-- C#..LA.ll 146;'i E qE~

4EtepqNCt COGKUIEAIe POINIS

L -0.62,430 0 .12*2 4 0 -dIsl7 0 .0 1 48 41 0.2963 -0.049L2 0.6128 a3.1311 e. 0.3233 03257 42 C.Z3o -0*059Z

. -0:9231 0 .10'0 f). 0.4159 0.0446 44 0.1395 -0.0464a 0:4,178 0.0800 &b C.,*965 0.06?9 46 0.0450 :0.0669-0.343 0.0963 27 -..5*5 0.0555 -V? -0.0086 0o0e0

19 :.33 0.666 46 .586 0101 48-0.0534 -0.05629 -0.2136 0 .-0589 2SP .. a3 2 0.1237 '.9 -0.0983 -0.049613:1. Z4jo 11.05 1> jo 10.6-21 0.121!4 50 :0:.154-L -0.0397

It -017 0 ."29 A 1 13099-9 0.1235 5 1 -. 986 -0.030612 -G115 0.0366 iz 0.a0*78 1.1193 !Z -0.254-0 -0.017303 :-C0910 0 .02 1" 33 0.-z 31v5 o.1ia 53 -0.29dL -0.0054r1. -0.0406 a .C d.4 3.5406 C.30* 14 -0.3.1L9 0.00691.7 -0.QJ&-2 0 .0 1 v 3~ 0. bas 0.2567 ti5 -0.39*2 I.Q2'.5,a 00*, 0.,, . 0 1.13 li 56 -47.4392 0.039917 C.cv.a J.Ca. 07C7,* ',.a03 17 0.49 0.0563

19 J- 2 , .A3 -J.O z 59 59 -0.5763 0.097120 :.ZZ93 0.3i45 6C 3j9 -C.C366 60 -O.a1 75 0.1109

a INOICA1ca t.ATAEMft PQ1 P rS

to AUG 79 MAGE 7

I AN 0JM9Pr;&Uft VANEt CZC 12122

A0IAL SEITING alxvrUtL TU L.6AfL'4G ftAOIIJISTANCE A W.L c r..CNiSS E L.b. T.E.

5. 5000 -21.134 0705 0. 7,w51 0.008 0.006-tlo 3:0p is

a ~j1potumf O.~t u .194 ~ 0.0230it 0UZ4WISIGN 0.0600 i ;31Ml1.ft 0.0600

ASVA CI00Nf C3CROINATES 13.3 p C.0jc~j~K .~?AT~l4 i~b~de " l&1i-id1Sc FAOM TtHe REAR

gel-C~ CW,401IAFE 0011MI:

STATIOW

g 0503 0166 . .3aw#7 C:30 41. %1.3792 -0.02162 -0.55dd lb13zp 22 ;6.0?q 0.0576 i3 .3 2 18 -0.0319A -Q.SLJS O.LtO26 4z 0 .. 513V 1.0631 .3 1.2752 -60.0375

*-0.40~ Of 0.1.021 24 0.5043 v.0760 44 0.2165 -0.0413,-O.ftL27 as s. 0.425 d 5. C 66 1.0676 45 O.1e92 -0.0416

6 -0.3679 0.42r,; eb 0.5866 0.100 40 0.1218 -0.03957 -Q.322V 0 .41 &;, 7 3-".4k G:1J62 4? 0.0634 :0.0340

: 0.46.8 0.104 20 1.613 0.5 48 0171 -0.0292%O-.Zazo 0 .09. 1. 0 717 1 1 2 49 -1.0291 -0.0212

10 -0.1772 0.09-' 40 0:7417 Q*LeO 50 -0.0866 -0.0109a .il- 0.371.*1 iL '.gas6 :.158 51 -0#1324 -0.0014

1.0.9 O*.7.7S0156 52 -0.1894 0.01210~ 02 0.v&23 33 0.7367 0.1"ao 53 -0..23.*7 0.02.1I". 0.0a". C,3 OM 3. -7.*933 3.Le 5^ -0.2790 0.037'b15 3.6692 0.0 213 3, 154SS 0373 55 -0.3356 0.0554bto CaI,?#. it*a C.4060 03.0595 56 -0.3799 0.071017 3..' 3.0'9^1 37 0.5050 2.13J96 57 -0..*23S 0.08791 - ;.~ 0..0 11 J! .i~e9 .421 7 St -0.47111 0.110219 .2O06 3.0'.* Q : 0.-*d7 .)"g4 5* O.S210 g.12 9 3

20 0.si 0 .0*eZ ;0 11O..4. -0. a10l7 .0 -0.5634 0 .Ilb9&#* NCI~a~c -:lkEMie 0POIS

57

.C AUG 79 P AGE 9

rAN 4 . JM tSS; VAME CSC 12122

A&,A aTP 14eEFRENCE DISTANCEAA(.IL Sa"ING AIMP(J1L TU LtA.J1MGRAI

-AJISTA4CE a00 Tn1CA.tSS LOGS L.E. T.1I.

6.Z:C -Z1.Zg6 '..C933 0.7450 .0 0.0-210) 17M 1-6

L. AiG Z0Gk; £114,. r~ht4:Nr PULpMr -0.3321

aj U~IPINS1ON 0.0264 U AmCNSlaft 0.0086K a1Dhms10ft a 0.w !S MINQZSL* 3.0600

s1Atr Pmmr co~vo NA res 0.0 9 0.0(.Errik OF GiAAvITY C£uRWAu.IS 0.2050e F0.0919,,JMPSSiA .WTtATIJ4 IS CJU94rL k C4.wI RO~M THE RtEAR

R8p.FtrnNCk COAD~IA8 POINTSt~ STArTIOr

~ o.7* 0.1998 9 0.105? 0.0945 41 0.5000 0.0151~ 0.4779 0...070 Z. 0.5 LO 3.1002 42 0. 397 0.0047

: -. '.309 0.1967 43 0.5761 0.1075 43 0.3908 -0.0008

-0.3040- 0 1867, 24 0.632Z 001199 44f 0.0292 -0.004z2-0.32514 0.17-09 25 1.6766 0.1299 .5 0.2797 -0.00t2

6-004?SS 0.loss 26 0.7210 001425 46 0.2302 -0.0016

I. -*.I?3Z- 0 .14.6a Z3 0.J192 0.1768 '8 0.1211 0.01109 -0.1kft 0.1390 .It 0.37?7 0.19911 '09 0.0731 0.0185

10 :9.0798 O.L310 30 0.8833 0.204-3 50 0.0133 0.029311 %0..02110 0.125 .3 0.89L?* 0.1993 51 -0.0343 0.039312 0*0ZW -01163 32 0.5892 0.1:45 52 -00O935 '0"1 0036 0.1064b 33 0.4799 0.1872 53 -001406 025

0.45 0.4029 34& 0.4318 0.1528 54 -0 18Th 0.379415 007 0: 0970 35 0.7915 0.127st 55 -0.24f54 0.0975to 0*.Z31a 0.31 36 0.1393 0.0966 56 -0.2914 0.113't 7 Q*1SQ6 3 .3886f .7 0.0960 0.0779 57 -0.3371 0. 130d#18 0.3262 0.0871 39 0.0515 0.0594 58 -0.393S 0.153019 0.3033 0.0875 4(# .it9054 0.339.* 59 a-0'382 0.17B3to 0.4-eft *.Cove '0 0.5476 0.0261 60 -0.44Z4 0.1927

* ZCAT&S EX TRE ME POINTS

I1 1 AUG 79 PAGE 11

PA N~3~4VANt isc 12122

~AL4RE.1Zi FERENCE DISTANCERAIO1SANCIA; ScTN A^ImF-0L 10 LEAiJING T.O1

6.0w V.911 '-i012 3.7'50 0.308 1.00-00 54-m 37S

LjA(jLNG CI,.C AAIAL rAN%.1,T 01Nt *-0.Zz%9@

0 *141N310F 0 027^ ki .1'4tN 1Cf4 3.3246a 31ENSLW 0.0600 4 J11M SION 50600O

STACK P0OINT C01,AOINAT£S 0.3 r 0.06tT KS 104AW171 C~j6.AU1NAi': 3.j.&0, 1i30

wzi.tc .JK ,oi3TAj IS .,U'c ;Li 41 qMp THE .(IAA

RtjtNCt: C£E3NOINA(I h'OINTS

NO x y A y 4

I-0.40236 0.2436 21 3.b079 0.1372 .1 3.6225 0.05162 -0.3932 0.2008 22 0.h651 0.14za '4 0.5595 0.04JA -.. *-j I .209 23 1.748Zl 0.1501 (* 3 0.5085 0.03 *4 -v.ZS 0.2313 Z* 0.7607 0.1615 '.4 (1.4-63 0:0321

S-0.230#5 0.-l9e 4!P 0.8072 0.1726 45 a.392? 0.0324

3 OLl 0.1110 ib 0.6535 0.453 *6 0.3444 0:8353

S-0.0704 3 192'. Zs 1.9564 3.ZZO1 43 0.2250 0.24899-0.027' 1 7.if7 Zv L.1126 M.2.31 .9 0.178?2 0.0508

£0 0.0ic7 0.1173 10 .*J38 0.2401 50 0.162 0.0682

11 a.0014 3.1685 11 1.33ZS* 0.24P30 51 0:0669 0007364 0t19_* 0.1616 3Z 1.0298 0279 52 0.0055 0.0931

13 0.1906 0.1539 33 1.GZ01 C . 304 5# 3 -0.0431 0.1059I'* 4.239L C..7 34. 0.9696 0.194.5 54 00 9 1 0.1i"9~5 0.2476 0.1-023 J1 et.927!1 1: 0..51 55 -0.1519 0.1308Is 6:367b 0 t~3 q 36 3727 0 13:1 56 -0.1996 0.1551

20 .5463 0.13213 40 3.6722 c.0,29 60 -0.3977 0:1 36 4

58

13 AUG 7q PAGE 13

%.JmqtS~b.A d$eCSC 12122

4 EFERENCE JISTANCEKAO1a s u SIT NG A I Pt I IL 10 LiAOLNG RADII

01S7AN c AW*Lc rhLC#(NtSS £CW; I.E. 7.1.

?.0O.~ 2O~1L 4~9 0..500.008 0.006-200 36M 374

L.=AOPNG tj*C AXIAL TiAiWCNT 0'UINV -0.1''3

.~UIMENtSI0N * U07b 'J ulaIS,1JM 0.0255q Emwe.slum 0.0400 i .iIMtNSlCfe, 0.0600

STACK Paw~ CO;UROINATES 0.0 , C.3,cG- Do; CAAVITY CI.URIohIAt, 3.-30 3.1076JMP~cS40g W~TATLah IS L.~UIcft C.CP~dli~ FROM Tpofi 4AA

9PEFetC - COORDINATE POINTS

* ~-J.30. 0.4.908 t 0.7300 0. se58 0..0 0.07 93 -0.2557 0...a7 Z3 0.6i.3 M.932 -.3 0.62" 0.0710

-.-0.2090- 0.2Z770 4 1 0.8$98 0.2050 44 0.5409 0.067542-O.L19 0.2*53 25 0.9381 0.21.63 45 0.5073 0.06796 -0.0411v 0.5603 26 0.9802 13.2294 46 O.4b531 0.0709

7-0.0410 0 .247e V7 1 .3"o 0:2482 4*7 0.3800 0*07611a .0220 0.2373 21 1.09'34 a.?652 48 0.3362 0.08529 .3744. 0.22," 49 L.1521 0.M90 109 0. 2947 0.0935

1 143 .17 3L 1.174tio 3.2889 51 0.1693 0.11*s

1.* 0.34d9 0.11M 3. 1 1075 0.1381 54 0:.05 Q19p 03992 0.195. 35 1.0637 0 Z Of# 55 -0.0 S 790S6~~~ 89*1 0.70 o 10081 -0.1064 0. 970

- .7 0. o 03 3.175'. 37 0 .9!1,? 6 .1563 57 -0.1553 0.2153Im. 3.559. Q.7 i ~i 0.1.361 55 -.10.2158 0.23971 v 0.*06 .13,* J9 :390 .114. 59 -0.2636 0.260520 .0497, 0 .713 .0 0.961 0.399 60 -0*3110 0.2824

0 INGICArts rXT IEME P0114TS

10 AUG 79 PAGE 15

R ADIAL se 17ING AIRFOIL rTl L-:AOING RADIIDISTANCE AkGLc 7nLcxfttss EKtE L. 1. T.c.

7.53c: -20.695. 3.11*3 3.7450 0.008 0.006-200 -.LI4 .4

LrAULN4. EDGc- AXIAL TAN%*ET P'ULt1 -0.094-o

3 aOIMSN1IY 0.3283 J DIAEW~al A 0.0263A almNSI5UN 0.3bO00 JVMtNb1lN4 0.0600

STACK POINT COLAo14ArES 0.0CQNTCA JF LsAAVITY C3j~QL.'4ATcS 0.5441 t 0.2072

* -.JP~SSJR 4AT141 IS W~ifttiR CLuCAw~i FROM THE 4.EAK

IUPeftINCa COORDIF4ATE POINTS

S TA rtm~4U* x y 9 xL -0.22910 0.3406 ZI 0.d5*3 0.2251, 40L 0:3707 0#1265i. -0.Z±93 0.3.*79 2z 0.9050 0.2311 4z 0.8025 0.1147

*3 -04.68 -0.3375 23 3*9550 0.2389 4.3 0.7472 0.10854 -0.3273 Z4 1.0187 Cr0.7 CF. 1.47

0..12 25 L so ft9 0.013H 45s 0.6220 0. 036 0.0033 0.4056 20 t.l18 0*4769 ##0 0.5662 0.1 67

7 .085 98 '7d? *6 0 j 0 47' 0.4981 o "t go01203? Mj~O 28 t1300 0. 148 6* I 8 0.4445 0.1 49r9 0 .1721 1.2777 ZIP X.2909 13.3399 4b9 0.3910 0.1329

10 0.22Z44 9.2099 30 i. 3 0 3 0 0.3453 50 0:344.5 0.j'SIL. 029 0.2603 JA 1.312U" 0.1396 51 0.216 0. 574La - *-.2S30 3 Z 1.3090 0.3346 s 0.2059 0.IT34V

0.3 0 0 .2##44 33 1.299Z 0.3262 53 0.1A537 0.1600I 049 0376 3-* 1.453 04J62 J4 0: 3J9 1.1136aS 0.11 .Z 5 sz 1.2000 0.2507 5 0 647

10 0.57194 0.24 34 1.I11 0.2232 So -0:0134 0.2f*3017 0.6 5', 0.2.0s 37 1.0 93 0.991 tP7 -0.0637 .202#40 0.oief 0 3.2± 3b .~ 0.17 77 S8 -0.:j6l N998319 ..7601, 028 39 0.j:974 0.15'.4 59 -0. 5 0.310440, 0.7909 0.2202 '.0 0.92#*5 0.1391 60 -0.2241 0.3334

0 11.01CATiLS lX1.st~fi P081lot S

59

In AUG 79 PAGE 17

i-AM k.M0~tSLA VANt CSC 12122

R!EECE DISTANCER ADIAL Sk1TTA.G aA.FUIL TO LEAOING RADIX

31STANCt AW*L% ralICKNtSS E(,c L.E. To.8.~.oooc -Z1.~ .T-50 .06 0.004-21(j 1714 ti

6.:Au1I46 E~tb AXIAL TAPN.ENT PI~uNI 1.0118

0 Df?4tMSI0N 0.0259 J ulI9EtSI3N 0.3271it 60ImMN~l O.46uO ! u1peems1ON 0.0600stACx Pa!Nr C0URGINA~c.. 0. a , 0.3

1.01C UF GAV~ITY cw.0c1'ArES 17 at9b . .2576Cjptcbu 4ur~rjuN 15 6JJtE CLUCKm1H2 FRJM TE REAA

RhFeREMCE COORDINATE W01?4S

STATION'40. x 'V A X

1 -0.143*e 0..9029 21 0.9772 0.2762 41 0.9951 0.17082 -0. i37 0.,*098 z. 1.0296 0:4827 44 0.9240 0.16

3 0 94[39 4 a L.0420~i 0~jg ;4fj La It I(1 6b 010 AUoir r384 24 L13 0304 4 0.7939 01

vADIG.e 0.3336 /I10 L.270 4012IN "90.95 01

i 0 03i.5 0 .4ft- 3 a L.-oOL40 0:13 s 0.400.9

&Z .hAI44 81 0AIA1. 3z'~ft POINT3 0 .420.04 22ii 0NGNS O .0Z9 33 L.4377ZI 0.723OZ3 02is DIMbeNiO 0.0600 1.3E36IO 0..50105 027.6CPtC~IA! 0.0a' .15 36017 .TI O (PRAVIT 37OORDI2 3.AT7ES 0.7204. 0.31

19 0.0640. 0.40V3 51 1.0961 0.2313 59 -0.087 0.3209

20091 .7a 10 L1051.1 0.416* 40 -0.138 0.39057479 1.63 0.4799~c 49 0.5962 0.240

3.r w -Z .94 0.1329 0.46560 0.00 4 yDT84516GEOG 3XIA TA .N POIN 0.25 02108

36 1.,091O 0..;94 56 0.36to 0.9709R DIMESIO 0.0300 7 DIM3NSIO 0.11 7 0.10 .021.AC P.09014 3 3OVRM0 1 0.04 5 .43 04j

C .fif F jAVT CORIAE .794 39S007 0.3233

* INDICATES CORDPINTS ONT

STATI60

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