N 8 4 - NASA CR-174688

PWA/GPD FR-17826

V/STOL MODEL FAN STAGE RIG

DESIGN REPORT

by J. G. Cheatham and T. L. Creason

United Technologies Corporation

Pratt & Whitney Aircraft

Government Products Division

P.O. Box 2691

West Palm Beach, FL 33402

Prepared for

National Aeronautics and Space Administration

NASA Lewis Research Center

Contract NAS3-22779
1. Report Number

NASA CR!174688

2. Govt Accession No. 3. Recipient's Catalog Number

4.Title (and Subtitle)

V/STOL MODEL FAN STAGE RIG,

DESIGN REPORT

5. Report Date

August 1983

в. Performing Organization Code

7. Authors)

J. G. Cheatham and T. L. Creason

8. Performing Organization Report No.

PWA/GPD FR17826

9. Performing Organization Name and Address

United Technologies Corporation

Pratt & Whitney Aircraft Group

Government Products Division

P.O. Box 2691, West Palm Beach, FL 33402

10. Work Unit No.

11. Contract or Grant No.

NAS3"22779

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546

13. Type of Report and Period Covered

Contract Report

31 JULY 1981 to 30 SEPT 1983

14. Sponsoring Agency Code

15. Supplementary Notes

Project Manager: Anthony J. Strazisar

NASA"Lewis Research Center, Cleveland OH 44135

16. Abstract

A model single!stage fan with variable inlet guide vanes (VIGV) has been designed to

demonstrate efficient design point operation while providing flow and pressure ratio modulation

capability required for a V/STOL propulsion system. The fan stage incorporates a split!flap VIGV

with an independently actuated ID flap to permit independent modulation of fan and core engine

airstreams, a flow splitter integrally designed into the blade and vanes to completely segregate fan

and core airstreams in order to maximize core stream supercharging for V/STOL operation, and an

EGV with a variable leading edge fan flap for rig performance optimization.

The stage was designed for a maximum flow size of 37.4 kg/s (82.3 Ib/s) for compatibility with

LeRC test facility requirements. Design values at maximum flow for blade tip velocity and stage

pressure ratio are 472 m/s (1550 ft/s) and 1.68, respectively.

17. Keywords

V/STOL Fan Stage

18. Distribution Statement

Unclassified — Unlimited

19. Security Classif. (of this report)

UNCLASSIFIED

20. Security Classif. (of this page)

UNCLASSIFIED

21. No. of Pages

258

22. Price*

*For sale by the National Technical Information Service, Springfield, Virginia 22151
TABLE OF CONTENTS

Section Page

I SUMMARY 1/2

II INTRODUCTION 3/4

III DESIGN STUDY 5

Definition of Overall Design Parameters 5

Preliminary Design Results 11

IV AERODYNAMIC DESIGN 17

Radial Aerodynamic Design Results 17

Definition of Metal Geometry 25

VIGV Geometry 28

Blade Geometry 30

Stator Geometry 30

Off-Design Predictions 31

2-D Transonic Time-Marching Analysis 32

V STRUCTURAL ANALYSIS 61

Recommended Strain Gage Locations 75

Conclusions and Recommendations 76

VI MECHANICAL DESIGN 87

Inlet Cases 87

Slip Ring Mount Tube 99

Inlet Guide Vane (IGV) 99

Rotor 101

Discharge Ducts 110

Diffuser Ill

Actuation System Components Ill

Position Read Out Components Ill

VII AERODYNAMIC INSTRUMENTATION 113

VIII FABRICATION 139

Specifications 139

Estimated Cost 141

Recommended Vendors 142

IX ASSEMBLY 155

Inlet Guide Vane (IGV) 155

Exit Guide Vane (EGV)/Discharge Duct Assembly 157

Rig Assembly/Facility Installation 159

in
TABLE OF CONTENTS (Continued)

Section Page

APPENDIX A AIRFOIL AERODYNAMIC DESIGN SUMMARIES .. 167

APPENDIX В AIRFOIL GEOMETRY 181

APPENDIX С AIRFOIL SECTION COORDINATES 195

APPENDIX D DEFINITION OF SYMBOLS 239

APPENDIX E DEFINITION OF VARIABLES 245

REFERENCES 249

IV

4104C
LIST OF ILLUSTRATIONS

Figure Page

1 Typical V/STOL Fan Map 6

2 V/STOL Fan Operating Line 7

3 V/STOL Bypass Ratio Estimate 8

4 Fan Design Point Efficiency Comparisons 9

5 V/STOL Fan Flowpath 11

6 Fan Work/Loading Experience 13

7 V/STOL Fan VIGV — Tip Section 14

8 Preswirl Effects at Nominal Takeoff 18

9 Effect of PT Profile at Nominal Takeoff 19

10 Total Pressure Profile Supercharging Effect at Max Control T.0 20

11 V/STOL Fan Blade Loading ID Stream at Max Control T.0 21

12 V/STOL Fan Blade Loading OD Stream at Max Control T.0 22

13 V/STOL Fan Blade Loading OD Stream at Max Control T.0 22

14 Rotor Loading Experience at Surge 23

15 Rotor Loading Experience at Surge 24

16 Fan Loading Experience 25

17 V/STOL Fan Stator Loading ID Stream at Max Control T.0 26

18 V/STOL Fan Stator Loading OD Stream at Max Control T.0 27

19 V/STOL Fan VIGV Choke Margin 29

20 V/STOL Fan VIGV Incidence Tip Section 33

21 V/ST01 Fan Blade Loading — ID Stream Incidence at Max

Control T.0 34

22 V/STOL Fan Blade Loading — ID Stream Deviation at Max

Control T.0 35

23 V/STOL Fan Blade — ID Stream Choke Margin at Max Control Т.О. 36

24 V/STOL Fan Blade — OD Stream Incidence at Max Control T.0 37
LIST OF ILLUSTRATIONS (Continued)

Figure

25

26

27

28

29

30

31

32

33

34

35

36

37

38

39

40

41

42

43

44

45

46

47

V/STOL Fan Blade — OD Stream Deviation at Max Control T.O

Page

38

V/STOL Fan Blade — OD Stream Choke Margin at Max

Control T.0 39

V/STOL Fan — ID Stream Stator Inlet Air Angle Relative to Cascade

Loss Bucket Range at Nominal T.0 40

V/STOL — ID Stream Stator Inlet Air Angle Relative to Cascade

Loss Bucket Range at Max Control T.0 40

V/STOL Fan — Stator ID Capture/Throat Area vs M2A 41

ID Stator Incidence 42

ID Stator Deviation 43

V/STOL — OD Stream Stator Inlet Air Angle Relative to Cascade

Loss Bucket Range at Nominal T.0 44

V/STOL — OD Stream Stator Inlet Air Angle Relative to Cascade

Loss Bucket Range at Maximum

Control T.0 44

Stator LE Flap Schedule 45

OD Stator Deviation 46

V/STOL Fan Map — ID Stream 49.

V/STOL Fan Map — OD Stream 51

V/STOL Fan Map — Overall 53

Speed vs Flow 55

V/STOL Fan Static Pressure Distribution 57

V/STOL Fan Static Pressure Distribution 58

V/STOL Fan Static Pressure Distribution 59

V/STOL Fan Static Pressure Distribution 60

Blade Steady Stress Finite Element Model 63

Blade Radial Stress Contour Plot — Convex Surface 64

Blade Radial Stress Contour Plot — Concave Surface 65

NASA V/STOL Fan Maximum Concentrated Stress at the

Flowpath Splitter 66

VI
LIST OF ILLUSTRATIONS (Continued)

Figure

 Pa

ee

48 Blade Free Vibration Finite Element Model 67

49 Blade Alone Vibration Analysis 68

50 Blade Modal Deformation at 17,750 rpm, Mode 1, 1010.988 Hz 69

51 Blade Modal Deformation at 17,750 rpm, Mode 2, 1637.072 Hz 69

52 Blade Modal Deformation at 17,750 rpm, Mode 3, 2472.997 Hz 70

53 Blade Modal Deformation at 17,750 rpm, Mode 4, 2690.768 Hz 70

54 Blade Modal Deformation at 17,750 rpm, Mode 5, 2863.475 Hz 71

55 Blade Modal Deformation at 17,750 rpm, Mode 6, 3497.499 Hz 72

56 Blade Modal Deformation at 17,750 rpm, Mode 7, 4360.125 Hz 72

57 Blade Modal Deformation at 17,750 rpm, Mode 8, 4950.768 Hz 73

58 Blade Modal Deformation at 17,750 rpm, Mode 9, 5377.659 Hz 73

59 Blade Modal Deformation 17,750 rpm, Mode 10, 6052.523 Hz 74

60 NASA V/STOL Fan Supersonic Flutter Analysis 77

61 NASA V/STOL Fan Blade Goodman Diagram 78

62 NASA V/STOL Fan Rig!IGV LE Strut Vibration Analysis 79

63 NASA V/STOL Fan Rig!IGV ТЕ Core Flap Vibration Analysis 80

64 NASA V/STOL Fan Rig!IGV ТЕ Fan Flap Vibration Analysis 81

65 NASA V/STOL Fan Rig!EGV ТЕ Strut Vibration Analysis 82

66 NASA V/STOL Fan Rig!EGV LE Duct Flap Vibration Analysis 83

67 NASA V/STOL Fan Rig!Vane Bending Flutter Parameter 84

68 NASA V/STOL Fan Rig!Vane Torsional Flutter Parameter 85

69 Recommended Strain Gage Locations 86

70 V/STOL Fan Rig!Vanes and Cases 89

71 V/STOL Fan Rig!Vanes and Cases 91

72 V/STOL Fan Rig!Vanes and Cases

 93

vn
LIST OF ILLUSTRATIONS (Continued)

Figure Page

73 V/STOL Fan Rig-Vanes and Cases 95/96

74 NASA V/STOL Fan Blisk 97/98

75 IGV Strut Case (Computer Graphics) 102

76 NASA V/STOL Fan Blisk (Computer Graphic) 103

77 NASA V/STOL Fan Blisk Airfoil/Section (Computer Graphics) 104

78 LCF Type I Life Cycle Summary 106

79 V/STOL Fan Rig Rotor/Case Thermal Map 107

80 Critical Rotor Assembly Clearances (Minimum) 108

81 NASA V/STOL Model Fan Stage Rig — Instrumentation 135

82 NASA V/STOL Model Fan Stage Rig — Assembly 161

83 NASA V/STOL Model Fan Stage Rig — Assembly Inspection Stations 165

84 ' V/STOL Fan, IGV, 1 of 2 183

85 V/STOL Fan, IGV, 2 of 2 185

86 V/STOL Fan Blade, 1 of 2 187

87 V/STOL Fan Blade, 2 of 2 189

88 V/STOL Fan Stator, 1 of 2 191

89 V/STOL Fan Stator, 2 of 2 193

90 Typical Airfoil Sections 196

91 Multiple Circular Arc Airfoil Definitions 241

92 Multiple Circular Arc Incidence Definitions ; 242

93 Cascade Loss Bucket Definitions 242

Vlll
LIST OF TABLES

Table Page

1 V/STOL Overall Fan Design Parameters 10

2 V/STOL Fan Meanline Design Parameters (from Preliminary Design) . 12

3 VIGV Comparison 15

4 V/STOL Fan — Predicted Performance

5 V/STOL Fan Airfoil Geometry '47

6 V/STOL Fan Summary of 2-D Transonic Analysis 56

7 Blade Shroud/Splitter Configuration 61

8 V/STOL Blade Frequency Summary 74

9 V/STOL Blisk Stress and LCF Summary 105

10 V/STOL Blisk Integral Arm/Balance Flange Stress Summary 105

11 Critical Speed Information 109

12 V/STOL Model Fan Gage Instrumentation Summary 114

13 V/STOL Model Fan Stage Instrumentation Listing 115

14 Parts 139

15 Materials 140

16 Processes 140

17 Drawing Interpretations 141

18 NASA LeRC V/STOL Model Fan Stage Rig Inlet Cases and Slip

Ring Support 144

19 NASA LeRC V/STOL Model Fan Stage Rig Inlet Guide

Vane Assembly 145

20 NASA LeRC V/STOL Model Fan Stage Rig Rotor Case and Tip

Shroud Insert 146

21 NASA LeRC V/STOL Model Fan Stage Rig Exit Guide Vane Assembly 147

22 NASA LeRC V/STOL Model Fan Stage Rig Discharge and Diffuser

Ducts 148

IX
LIST OF TABLES (Continued)

Table Page

23 NASA LeRC V/STOL Model Fan Stage Rig Core Actuation/Position

Readout Components 150

24 NASA LeRC V/STOL Model Fan Stage Rig IGV Flap Actuation/Position Readout Components 151

25 NASA LeRC V/STOL Model Fan Stage Rig EGV Flap Actuation/Position Readout Components 152

26 Delivered Hardware 153

27 Rotor Forging Mechanical Properties AMS 4973 153

28 Axial Location of V/STOL Fan Stage Rig Component Interfaces 163

29 Inlet Guide Vane — Nominal Takeoff ID Stream 168

30 Rotor One — Nominal Takeoff ID Stream 169

31 Stator One — Nominal Takeoff ID Stream 170

32 Inlet Guide Vane — Nominal Takeoff OD Stream 171

33 Rotor One — Nominal Takeoff OD Stream 172

34 Stator One — Nominal Takeoff OD Stream 173

35 Inlet Guide Vane — Maximum Control Takeoff ID Stream 174

36 Rotor One — Maximum Control Takeoff ID Stream 175

37 Stator One — Maximum Control Takeoff ID Stream 176

38 Inlet Guide Vane — Maximum Control Takeoff OD Stream 177

39 Rotor One — Maximum Control Takeoff OD Stream 178

40 Stator One — Maximum Control Takeoff OD Stream 179/180

41 Airfoil Section Coordinate Printout Explanation 195

4104C
SECTION I

SUMMARY

A program has been conducted under Contract NAS3-22779 to design a model single-stage

fan rig with variable inlet guide vanes capable of providing the flow and pressure ratio

modulation required for a V/STOL propulsion system while maintaining constant rotor speed.

The rig was designed for testing within the NASA LeRC W-8 facility.

The fan stage aerodynamic design goals were to provide maximum airflow, pressure ratio

and engine thrust at takeoff for maximum attitude control which was 25% above the thrust

level at the nominal takeoff condition. The takeoff maximum control operating point specific

flow (1.82 kg/sec m2 — 43 Ibm/sec ft2), pressure ratio (1.68), tip speed (472 m/sec — 1550 ft/sec)

and bypass ratio (6.35) were determined from a canvas of airframers. Stall margin (15%) and

thrust modulation (40-125%) were established by the statement of work. Maximum rig flow size

(37.4 kg/sec — 82.3 Ibm/sec) was determined by the NASA LeRC W-8 facility.

The resulting fan stage incorporates the following features into the aeromechanical design:

• A flow splitter which extended from the VIGV leading edge through rotor

and EGV to prevent mixing of the core and fan airstreams,

• A split flap VIGV with independently activated ID and OD flaps to permit

independent modulation of the core and fan airstreams to maximize core

stream supercharging for V/STOL operation, and

• An EGV with a variable leading edge fan flap for rig performance

optimization. Predicted adiabatic efficiencies at nominal takeoff for the

core, fan, and overall fan stage were 85.0, 82.7 and 83.0%, respectively.

The fabrication and assembly of the V/STOL model fan stage rig was cancelled due to

NASA funding priorities and the contract effort terminated with the completion of the

hardware drawings.
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SECTION II

INTRODUCTION

One of the basic requirements of a V/STOL aircraft propulsion system is the capability for

rapid thrust modulation for attitude control during takeoff and landing. This capability has

been demonstrated in previous V/STOL concepts through the use of variable inlet guide vanes

(VIGV) to vary the flow and pressure ratio of the fan at constant rotor speed and thus modulate

the propulsive thrust of the engine. Other V/STOL requirements include high thrust to weight

ratio which is necessary for vertical flight, and low specific fuel consumption which is desired for

maximum aircraft range. Of equal importance for operational considerations in multi engine

aircraft is the capacity to maintain sufficient attitude control to perform an emergency vertical

landing with one inoperative engine. To meet these stringent V/STOL requirements, a VIGV

fan stage has been designed employing unique design features to provide a highly responsive,

flexible means of thrust modulation while optimizing aerodynamic performance to give

maximum thrust with minimum core engine size.

The fan stage employs a split-flap VIGV with an independently actuated ID flap to allow

separate modulation of fan and core airstreams. The stage also employs a flow splitter which is

extended forward by incorporating it into the blade design to separate the fan and core

airstreams and prevent mixing between the higher pressure core stream and the fan stream. The

fan blade has been designed to deliver maximum pressure ratio in the root region in order to

provide maximum core supercharging and thereby reduce core engine size and weight.

The design features incorporated into the model V/STOL fan stage will enable a

comprehensive rig test to demonstrate their feasibility and observe in detail the effects of VIGV

variation on a typical V/STOL fan stage. The rig test will evaluate the various methods of

maximizing core engine supercharging to minimize weight which will be useful in future cycle

studies to define an optimum V/STOL engine.
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SECTION III

DESIGN STUDY

DEFINITION OF OVERALL DESIGN PARAMETERS

The overall design parameters for the model V/STOL fan stage were derived from proposal

objectives expressed in the statement of work, a canvass of airframers for typical V/STOL

mission requirements, and parametric studies from past and present Pratt & Whitney V/STOL

designs. The statement of work specifies that the model fan stage include a variable inlet guide

vane (VIGV) that will allow identification and evaluation of critical V/STOL design requirements. The model fan stage should demonstrate adequate thrust modulation capability during

takeoff and hover operation while minimizing required engine size due to multi-engine cross

shaft drive operation during one engine inoperative situations. For responsive attitude control,

rapid thrust modulation is to be accomplished through VIGV variation with no change in rotor

speed.

Figure 1 presents a typical fan map for multi-engine V/STOL operation. Maximum

airflow, pressure ratio, and engine thrust occur at takeoff for maximum attitude control which is

approximately 25% above the thrust level at the nominal takeoff condition. Individual fan

operating points are indicated for a one engine inoperative condition where the imbalance in

thrust between engines is corrected through the appropriate combination of VIGV variation on

individual fan stages. The statement of work specifies a range for certain fan parameters at the

nominal takeoff point as follows:

• Pressure Ratio 1.5 — 1.6

• Specific Flow (W/A) 1.6 kg/sec2 (38 Ibm/sec ft 2)

• Tip Speed 450 — 500 m/s (1450 — 1600 ft/sec)

• Inlet Hub/Tip Ratio 0.35 — 0.40

• Stall Margin 15%

• Thrust Modulation 40 — 125%

In addition, the model V/STOL fan stage is required to meet LeRC test facility dimensions

which set the maximum outer diameter at 50.8 cm (20 in.) and the minimum exit inner diameter

at 20.32 cm (8.0 in.).

A canvass of airframers was conducted to further define typical V/STOL design

parameters. As a result, the following parameters were selected for the takeoff maximum control

fan operating point:

• Specific Flow 1.82 kg/sec m2 (43 Ibm/sec ft2 )

• Pressure Ratio 1.68

• Tip Speed 472 m/sec (1550 ft/sec)

To determine airflow and pressure ratio at the nominal takeoff point, a normalized

operating line was calculated from a PW2037 engine cycle simulation and forced through the

takeoff maximum control design point. The PW2037 engine operating line was normalized for

this estimate due to its similarity to the V/STOL fan stage pressure ratio and bypass ratio at

maximum flow. Figure 2 displays the resulting operating line for the model V/STOL fan stage

and the estimated nominal takeoff point. The relationship between fan airflow and pressure

ratio to engine thrust was established from NASA tests conducted on a YTF34 engine as

reported in References 1 and 2. The nominal takeoff point was estimated to be on the

normalized operating line at a point corresponding to a 25 % reduction in engine thrust.
ОС

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2

a.

One Engine Out

Landing,

Maximum Control

One Engine Out

Landing,

Takeoff

Maximum

Control

VIGV = Y

VIGV = X

Q] Design Speed

О 95% Speed

Corrected Airflow

FD 265647

Figure 1. Typical V/STOL Fan Map

The fan bypass ratio for a representative V/STOL application at nominal takeoff was

estimated from a previous V/STOL "A" parametric study in which engine size was optimized at

varying fan pressure ratio and bypass ratio to satisfy thrust and one engine out requirements.

Figure 3 summarizes the trade study for a tandem fan V/STOL "A" application from which high

compressor flow requirements were established to permit extrapolation from a single fan

V/STOL engine (STF 521). As a result of these studies, the model V/STOL fan stage bypass

ratio at nominal takeoff was defined to be 6.0. Bypass ratio at the maximum control takeoff

point was calculated by assuming that high compressor corrected flow did not change from

nominal takeoff. The resulting bypass ratio at the maximum flow condition for the fan is

calculated to be 6.35.

Efficiency for the model fan stage at the nominal takeoff condition was predicted from the

Pratt & Whitney streamline design system which incorporates empirical fan blade element loss

data as a function of Mach number and aerodynamic loading. This efficiency prediction is

compared to state of the art efficiency versus tip speed in Figure 4.

The state of the art efficiency curve has been established by adjusting test results as

follows:

• Part span shroud effects have been eliminated

• Efficiencies have been adjusted to reflect the same flow size to eliminate

Reynolds number and relative roughness effects

Efficiency is measured at 15% stall margin at design speed.
t

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3

3
BPR vs P/P(an at Nominal Т.О.

Fan Average Pressure Ratio

1.7

FD 268954

Figure 3. V/STOL Bypass Ratio Estimate

For comparison purposes, model V/STOL fan efficiency has been increased by 1.0% for

part span shroud and splitter effects and 1.1% for size effects in Figure 4 and is predicted to be

within 1 % of state of the art technology.
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Table 1 summarizes the maximum control and nominal takeoff design points established

from statement of work specifications, rig flowpath constraints, a canvass of airframers for

representative parameters, and parametric studies and calculations for typical V/STOL fan

applications.

TABLE 1. — V/STOL OVERALL FAN DESIGN PARAMETER

Corrected Speed (N/V«) ~ RPM

Corrected Flow (W\/0/S) ~ kg/sec

Pressure Ratio

Stall Margin ~ %

Adiabatic Efficiency (ij) ~ %

Bypass Ratio

Specific Flow (Wc /A) ~ kg/sec m2

Corrected Tip Speed ~ m/sec

Maximum

Control Takeoff

17762

37.42 (82.32 Ibm/sec)

1.68

15

79.5

6.35

1.82 (43 Ibm/sec ft2 )

472 (1550 ft/sec)

Nominal Takeoff

17762

33.50 (73.70 Ibm/sec)

1.53

15

83

6.0

1.6 (38 Ibm/sec ft2 )

472 (1550 ft/sec)

Once overall fan design parameters had been established, consideration was given to

identifying features unique to V/STOL applications which would improve engine performance

by decreasing weight, maximizing thrust, and providing more flexible thrust modulation.

V/STOL aircraft are limited in range due to the overall engine weight which includes such

items as large vane actuators, variable nozzles, additional ductwork for attitude control, and

gearboxes and cross-shafts unique to V/STOL operation. Additional complexity occurs for

multi-engine V/STOL aircraft with cross-shaft designs due to the one engine inoperative

condition which requires that the gas generator of one engine drive the fans of both engines to

provide thrust and attitude control for emergency landing. During emergency landing, the fan of

the inoperative engine must run with its VIGV open to offset the loss of engine thrust. The fan

of the operative engine must run with its VIGV closed to equalize engine thrust and provide

attitude control while its gas generator provides the power to drive both fans. The power

requirement during emergency landing determines the size of the core engine and has a large

impact on engine weight. Consequently, it is desirable that the fan VIGV in the operable engine

not desupercharge the core flow while operating with closed vanes to achieve balanced thrust

with attitude control margin.

A part-span VIGV has been successfully tested at NASA LeRC on a YTF34 engine to

modulate thrust and improve the desupercharging condition during one engine inoperative

conditions. The results of these tests are described in References 1 and 2. The part-span VIGV

does not modulate the core stream coming into the fan blade; consequently, core pressure ratio

does not decrease as much as fan OD pressure ratio during VIGV closure. The tests indicate that

the core pressure ratio drop may be further reduced by extending a flow splitter upstream

toward the blade trailing edge to reduce mixing between the higher pressure core stream and the

lower pressure fan stream. A reduction in core size of 12% is reported to be achievable for a

part-span VIGV relative to a full-span VIGV configuration. An additional 2% core size

reduction is estimated when a flow splitter extension is included. The reduced core size,

improved cruise SFC, and decreased engine weight will provide substantial benefits to a

V/STOL aircraft

Even with a part-span VIGV fan, a significant reduction in core pressure ratio occurs as

the fan stream VIGV closes and fan stream pressure ratio is reduced due to mixing and

communication between fan and core streams. In the configuration proposed for the model

V/STOL fan stage shown in Figure 5, a flow splitter has been incorporated into the blade design

to allow extension upstream through the VIGV. An independently actuated ID flap on the

VIGV is provided for core stream airflow. The flow splitter completely isolates the core stream

10
from the fan stream to eliminate any desupercharging effect. In addition, the ID flap is available

to provide additional supercharging to the core stream regardless of the setting of the fan OD

flap. Isolation of the core stream alone should permit a reduction in core size of 8% according to

the referenced reports. By supercharging with the ID flap, additional reduction in core size

should be achievable depending on stall margin and turbine inlet temperature limits. The

benefits over competing designs of an isolated core stream, split-flap VIGV fan design for multiengine V/STOL application which include reduced engine weight, improved SFC, and more

flexible thrust modulation capability may be assessed relative to its added cost and complexity

in the model V/STOL fan stage rig test program.

Flow

FD 265646

Figure 5. V/STOL Fan Flowpath

Also included in the model V/STOL fan stage configuration is a variable leading edge flap

for the fan exit guide vane (EGV). The purpose of the variable flap is to align incidence into the

EGV and, thereby reduce loss through the OD vane. This flap is desirable in the rig test for

optimization and trade studies of fan performance over the wide range of incidence encountered

during VIGV modulation. This configuration will enable the rig test to assess the feasibility of

including a leading edge EGV flap in a full-scale V/STOL engine.

PRELIMINARY DESIGN RESULTS

Once overall design parameters were determined and the basic configuration was selected,

a preliminary design analysis was performed using the Pratt & Whitney Equivalent Cone Angle

(EGA) Meanline Design Program. The meanline analysis was used to define the flowpath and

meanline aerodynamics at the maximum flow condition to balance the aerodynamic loading

between the blade and vane. Meanline studies resulted in the definition of aspect ratio, solidity,

number of airfoils, and average chord which provide sufficient loading capability to ensure 15%

stall margin at maximum flow. Table 2 presents meanline design parameters for the overall fan

stage as well as core and fan streams. Figure 6 illustrates that the work and loading levels of the

11
V/STOL fan stage as derived from the meanline studies are within the experience level

demonstrated by previous fan designs.

TABLE 2. — V/STOL FAN MEANLINE DESIGN PARAMETERS (FROM PRELIMI!

NARY DESIGN)

N,/\/0 ~ rpm

Wx/0/г ~ kg/s

Pressure Ratio

Efficiency

Bypass Ratio

Stall Margin

VIGV

UTIP — m/s

We/A ~ kg/s • m2

A

1N

AOUT

Exit Mn

DF

ДР/Ро!Р

E'

(Cx/U)2

AR

a

(NoA)w/0 IGV

Length (R1LE!S1TE) ~ cm

Rl OD ~ cm

Flowpath

Overall

17762

37.42 (82.32 Ibm/s)

1.68

76.6

6.35

15.0

+18°

472.0 (1550 ft/s)

1.82 (43 Ibm/s • ft2 )

0.35

0.52

0.46

0.40

0.36

0.66

0.48

2.1

1.58

62

13.72 (5.40 in.)

50.8 (20.0 in.)

COD

ID Stream

17762

5.09 (11.2 Ibm/s)

1.74

81.9

—

15.0

+18°

236.5 (776 ft/s)

1.70 (40.3 Ibm/s ! ft2)

0.70

0.87

0.68

0.35

0.38

1.51

0.93

0.40

2.26

62

13.72 (5.40 in.)

25.4 (10.0 in.)

MIXED

OD Stream

17762

32.33 (71.12 Ibm/s)

1.66

75.8

—

15.0

+18°

472.0 (1550 ft/sec)

1.87 (44.4 Ibm/s ! ft2)

0.52

0.61

0.48

0.41

0.33 ,.

0.60

0.48

1.63

1.54

62

13.72 (5.40 in.)

50.8 (20.0 in.)

COD

The fundamental premise of a V/STOL fan design with a variable inlet guide vane is that

fan airflow and pressure ratio can be modulated by large amounts without a change in rotor

speed through manipulation of the VIGV. This manipulation necessitates a VIGV with more

turning capability than is found in conventional fan designs. Figure 7 illustrates the typical

range of flap positions that a V/STOL fan stage encounters from maximum to minimum control

at a given speed. The VIGV solidity level required to achieve the turning necessary for V/STOL

performance was verified through calculation of the Zweiffel loading coefficient for accelerating

cascades. Table 3 presents a comparison of basic geometry between the part!span YTF34 VIGV

of References 1 and 2 and the model V/STOL fan stage. The similar gap!to!chord and aspect

ratio of the two designs further verify the turning capability of the V/STOL VIGV.

12
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Figure 7. V/STOL Fan VIGV — Tip Section

14
TABLE 3. — VIGV COMPARISON

YTF34 P.E.T.

(Part Span VIGV)

ID

% Span 100

Average Diameter ~ cm 34.29

(in.) (13.5)

Chord ~ cm —

(in.) -

r/b -

Chord — Flap ~ cm —

(in.) -

r/b — Flap —

tma, ~ cm —

(in.) —

t/b —

LER/b —

TER/b —

Number of Airfoils

Airfoil Series

Flap Camber

BPR

Aspect Ratio

Hub

66.1

60.50

(23.82)

8.573

(3.375)

0.739

4.257

(1.676)

1.488

0.754

(0.297)

0.088

0.021

0.015

30

63

0°

6.2

1.84

Tip

0

111.66

(43.96)

17.734

(6.982)

0.659

12.842

(5.056)

0.911

1.270

(0.500)

0.072

0.014

0.010

ID

100

13.72

(5.4)

4.788

(1.885)

0.53

2.248

(0.885)

1.128

0.635

(0.25)

0.128

0.014

0.114

P&WA V/STOL

(Split Flap IGV)

Split

73.3

23.62

(9.30)

6.894

(2.714)

0.631

4.013

(1.580)

1.088

0.635/0.794

(0.25/0.3125)

(0.091/0.113)

0.008

0.010

17

63

-6°

6.0

1.90

Tip

0

20.0

(7.874)

4.987

(1.963)

0.74

3.487

(1.373)

1.060

0.3125

'• (0.123)

0.063

0.003

0.006

15
Page intentionally left blank

Page intentionally left blank
SECTION IV

AERODYNAMIC DESIGN

RADIAL AERODYNAMIC DESIGN RESULTS

After establishing overall design parameters and completing the preliminary design to

define average row geometry and performance, the resulting rig configuration shown in Figure 5

was analyzed using the Pratt & Whitney streamline analysis program. This program performs

an axisymmetric, radial equilibrium analysis of the flow field using a streamline curvature

solution technique to radially define aerodynamic parameters at airfoil leading and trailing edge

stations.

Preliminary studies were performed using the streamline analysis to optimize preswirl and

work distribution radial effects on fan stage aerodynamics at the nominal takeoff point. Figure 8

illustrates the effects of three possible swirl distributions on fan blade loss, aerodynamic

loading, and Mach number. By placing maximum preswirl at the tip and counterswirl at the

hub, blade Mach number and loss were reduced at the tip while aerodynamic loading was

reduced at the hub. As a result of this study, a nominal VIGV exit air angle distribution was

defined featuring 12 degrees of preswirl at the OD varying to no preswirl for the ID stream.

The effect of supercharging the core airstream to minimize core engine size by varying the

design stage radial work distribution was evaluated in a streamline analysis study as shown in

Figure 9. The figure compares the exit Mach number and turning of a design with a constant

spanwise pressure profile versus a design with a total pressure profile increased at the hub to

provide 10% supercharging to the core airstream. For the same average fan pressure ratio, the

supercharging work distribution requires considerably more turning at the hub and small

negative turning near the tip as contrasted to the more moderate turning for a flat pressure

profile. Blade loading limits for 15% stall margin determine the amount of pressure rise and

turning which can be achieved at the blade hub. From evaluation of blade aerodynamic loading

from streamline studies with varying amounts of core supercharging, the total pressure profile

displayed in Figure 10 was defined for the maximum flow point. The core supercharging

achieved by this profile was approximately half the 10% level initially attempted. The 10%

level appears unrealistic due to hub loading and turning requirements. In addition, the absence

of work in the blade tip sections where work capability is greatest raises the question as to

whether such a design is providing the most effective use of fan work capability. The rig test of

an isolated core stream, split!flap VIGV fan stage where supercharging can be adjusted through

ID and OD flap variation may provide a basis for additional engine cycle studies to define the

fan/core work levels for optimum V/STOL engine performance.

Fan blade loading for the ID and OD airstreams at the maximum flow point are shown in

Figures 11 through 13 in the form of diffusion factor and ДР/(Ро!Р) versus span. Predicted

blade loading levels at 15% and 20% stall margin for the OD and ID air streams are compared

to demonstrated rotor loading levels at surge for previous designs in Figures 14 and 15. The

demonstrated loading levels at surge substantiate V/STOL blade loading which will allow the

fan stage to demonstrate adequate stall margin for a V/STOL engine application. Figure 16

substantiates overall blade loading for the V/STOL stage relative to peak diffusion factor

demonstrated as a function of blade aspect ratio.

17
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Figure 8. Preswirl Effects at Nominal Takeoff

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Figure 9. Effect of PT Profile at Nominal Takeoff

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Figure 10. Total Pressure Profile Supercharging Effect at Max Control Т.О.

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Figure 11. V/STOL Fan Blade Loading ID Stream at Max Control Т.О.

21
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Figure 12. V/STOL Fan Blade Loading OD Stream at Max Control Т.О.

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Figure 13. V/STOL Fan Blade Loading OD Stream at Max Control Т.О.

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Figure 16. Comparative Fan Loading

Aerodynamic loading for the V/STOL fan stage exit vanes is displayed in Figures 17 and 18

at the maximum flow point. Both ID and OD stators are more lightly loaded than the blade and

are not expected to limit fan stall line at the maximum flow VIGV setting.

Table 4 summarizes V/STOL fan stage performance as predicted by the final streamline

analysis for the maximum flow and nominal takeoff design points which correspond to VIGV

settings of +18 and zero degrees, respectively. Stage performance as well as blade and vane

loadings are given for ID, OD, and combined overall airstreams.

A detailed tabulation of velocity triangle components, Mach numbers, loading parameters,

and overall airfoil row performance is given in both metric (SI) and English notation for each

airfoil row in Appendix A. The information is independently tabulated for ID (core) and OD

(fan) airstreams at both nominal and maximum control takeoff points.

DEFINITION OF METAL GEOMETRY

Once aerodynamic parameters had been defined for maximum control and nominal takeoff

design points to satisfy the flow, efficiency, and stall margin requirements of the V/STOL fan

stage, metal geometry was defined for the blade and vanes by employing empirical incidence,

deviation, and choke margin criteria in the Multiple Circular arc (MCA) blading program to

radially define airfoil section metal angles. The MCA blading program is based on a simplified

blade channel flow model in which an airfoil is constructed around a multiple circular arc mean

camber line. The program lends itself well to using empirical criteria to optimize camber, chord,

and thickness distributions along the mean camber line from which airfoil sections are defined.

4.5

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TO 265583

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Figure 18. V/STOL Fan Stator Loading OD Stream at Max Control Т.О.

FD 265582

27
TABLE 4. — V/STOL FAN — PREDICTED PERFORMANCE

Overall

N,/\/e ~ rpm

WvWe ~ kg/s

P/P

IAD ! %

BPR

VIGV

o3

Biade

Max Flow

17,762

37.42

1.680

79.54

6.35

+18°

—

Norn SLTO

17,762

33.50

1.531

82.98

6.0

0°

—

ID

Max Flow

17,762

5.09

1.751

84.97

—

+18°

20°

Norn SLTO

17,762

4.79

1.591

84.87

—

0°

20°

OD

Max Flow

17,762

32.33

1.669

78.68

—

+18°

0°

Norn SLTO

17,762

28.71

1.521

82.66

—

0°

0°

Dp . 0.452 0.360 0.468 0.307 0.449 0.369 "т

DP)TM,

ДР/Ро!Р

ДР/Ро!РС,

Vane

DF

DP)TM,

ДР/Ро!Р

ДР/Ро!Рти

0.605

0.432

0.686

0.323

0.596

0.220

0.464

0.470

0.379

0.557

0.410

0.530

0.312

0.407

0.576

0.586

0.686

0.382

0.596

0.384

0.464

0.465

0.431

0.440

0.385

0.530

0.391

0.407

0.605

0.408

0.638

0.314

0.397

0.195

0.227

0.470

0.370

0.557

0.414

0.488

0.298

0.322

VIGV GEOMETRY

The spanwise chord distribution and number of vanes for the VIGV of the model V/STOL

fan stage were established in the preliminary design such that the vane provided sufficient

solidity to achieve the flow turning capability required for V/STOL operation. Vane loss was

estimated from NASA Task II variable camber inlet guide vane blade element data as a function

of flap position and Mach number as shown in Reference 3. Vane angles were defined from

Reference 4 which details design procedures for selection of airfoil sections for a 63 series airfoil

developed specifically for inlet guide vane designs. The VIGV was designed to be in its true

airfoil position (nominal flap position) at the nominal takeoff point.

Due to the high specific flow of the fan stage, the VIGV has little choke margin at the

maximum flow point. This condition is aggravated by two factors: the presence of the flow

splitter through the vane which blocks part of the annulus, and a maximum thickness

requirement for the vane in order to pass a connecting rod through the OD flap to

independently actuate the ID flap. The choke margin of the final VIGV design is shown in

Figure 19 and was achieved by reducing the number of vanes from 18 to 17 and by modifying the

ID flowpath to increase annulus area.

28
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Variable inlet guide vane minimum loss incidence as a function of flap position was

evaluated for a typical section using the Pratt & Whitney Compressor Cascade Prediction

system. This prediction system is based on a correlation of standard airfoil series twodimensional loss and turning data versus typical geometric design parameters such as camber.

The cascade system prediction shown in Figure 20 illustrates that the VIGV will be within its

loss bucket at any camber corresponding to a flap angle setting to be used in the V/STOL VIGV

schedule.

BLADE GEOMETRY

The V/STOL fan blade was designed using the MCA blading program at the maximum

control takeoff point where airflow, pressure ratio, and engine thrust are at their maximum

values as defined in the design study. Once the maximum airflow point is demonstrated, the

nominal takeoff point can be achieved by closing the VIGV to its nominal flap position. The

maximum airflow and pressure ratio attained by opening the VIGV flap is achieved at the

expense of fan efficiency due to an increase in Mach number and associated losses. However, the

maximum control point is only run during transient V/STOL operation where fuel consumption

is not a prime concern. Design efficiency will be attained at the nominal takeoff point where the

VIGV flap is closed and Mach number losses are reduced.

Figures 21 through 23 display the incidence, deviation, and choke margin studies for the ID

section of the blade which controls the core air stream. At the hub where the airfoil sections

approximate a circular arc, the cascade prediction system was employed to verify that blade

incidence was within the loss bucket. Additional camber was required in the front arc of the

blade relative to a circular arc airfoil section to achieve the choke margin displayed in Figure 23.

Figures 24 through 26 display incidence, deviation, and choke margin studies for the OD

section of the blade which controls the fan air stream. Incidence and choke margins were set at

levels representative of typical fan designs. Deviation values became negative near the OD due

to the negative camber employed in the blade as a result of supercharging the core airstream.

STATOR GEOMETRY

Basic stator geometry such as aspect ratio, number of airfoils, and chord was determined in

the preliminary design to satisfy the turning requirements of a V/STOL fan without limiting

performance. The fan exit guide vane was configured with a variable leading edge flap at midchord to control loss and was designed to turn the airflow to the axial direction. The fan ID

stator which serves the core air stream was designed as a conventional MCA stator and leaves 20

degrees of swirl in the air stream as representative of preswirl to a high compressor. Incidence

into the stators was biased inside the predicted loss bucket at the nominal takeoff point to

provide the best incidence compromise into the vanes at the maximum control takeoff point.

Figures 27 and 28 display predicted inlet air angle relative to cascade loss bucket range for

the ID stator at the nominal takeoff and maximum control takeoff points (see Figure 93 in

Appendix D). Due to the relatively high Mach number at the hub of the ID stator, the loss

bucket is insufficiently wide to contain both air angle profiles. Consequently, the ID stator is

predicted to be slightly stalled at the hub at the maximum control takeoff point which will

result in additional loss and a small efficiency penalty. Since maximum control takeoff is a

transient point, the impact of stalled incidence is not expected to affect normal V/STOL

operation. Another measure of ID stator incidence, capture-to-throat area ratio, is shown in

Figure 29. This ratio is calculated from the MCA blading program and confirms the cascade

system prediction that ID stator incidence is within the loss bucket except at the hub at

maximum control takeoff. Figure 30 displays suction surface incidence and choke margin for the

ID stator at the maximum control and nominal takeoff points. The figure confirms that

30
sufficient choke margin exists at both points in the stator channel to pass the required flow.

Deviation of the ID stator is illustrated in Figure 31 for the two takeoff points.

The stator for the OD air stream, the fan exit guide vane, was designed as a conventional

circular arc airfoil due to lower (subsonic) Mach numbers encountered in the fan OD air stream.

The fan EGV was designed to turn the airflow to axial and was equipped with a variable leading

edge flap to help align vane incidence within its loss bucket. Figure 32 shows EGV inlet air angle

relative to cascade loss bucket range for the nominal takeoff point where the vane was designed

in its true airfoil position for minimum loss. Figure 33 shows EGV inlet air angle relative to

cascade loss bucket range for the maximum control takeoff point if there is no variation of the

leading edge flap position. From this figure, it was estimated that the leading edge flap must be

opened approximately three degrees to shift the loss bucket boundaries such that the maximum

control incidence profile would be contained within the loss bucket. As a result of this figure and

other part speed analyses, an OD stator leading edge flap schedule was developed as shown in

Figure 34 as a function of VIGV position. The schedule proved to be independent of fan speed

and may be ganged with the VIGV schedule for simplified rig operation. Figure 35 shows EGV

deviation for maximum control and nominal takeoff points.

Table 5 summarizes V/STOL fan airfoil geometry for the overall or average airfoil as well

as for the ID and OD air stream airfoils. Plots displaying airfoil parameter technical drawing

(TD) information are contained in Appendix B.

Upon completion of preliminary airfoil definition, a structural evaluation of the blade

revealed a first bending mode 2E intersection in the operating speed range that was judged

unacceptable for rig operation. The structural evaluation, which is detailed in Section IV,

necessitated the inclusion of a part span shroud into the blade design at approximately 60

percent span from the ID. To offset the blockage of the shroud, blade choke margin was

increased by increasing the leading edge angle and front camber over a length of 12 shroud

thicknesses such that the increase in choke margin offset the channel area blockage of the

shroud. The presence of the part span shroud is readily seen in Figures 24, 24a, and 26 which

display OD blade incidence and choke margin.

OFF-DESIGN PREDICTIONS

Once maximum control and nominal takeoff design points were established, V/STOL fan

stage performance was predicted for off-design speed and VIGV settings by executing the

streamline analysis program in its off-design prediction mode. In this mode, rotor and stator loss

and exit air angles have been correlated versus incidence such that the streamline program can

model fan stage performance as airflow is varied for constant speed and VIGV position. This

procedure was used iteratively in the design phase to obtain stall loading estimates for the fan

configurations under consideration.

Using this procedure, fan maps were generated for ID (core) and OD (fan) air streams for

VIGV settings at 0 and —20 degrees as shown in Figures 36 and 37. The stall line was predicted

for these maps by assuming a maximum blade diffusion factor limit of 0.72 based on fan

experience at V/STOL blade aspect ratio levels. Included on these maps is a speedline

prediction at 100% speed with VIGV equal to +18 degrees in order to model the maximum

control takeoff speedline. The predictions from the ID and OD fan maps were mass averaged

and combined to obtain an overall fan stage map which is given in Figure 38. Speed versus flow

characteristics were obtained from these maps for a typical operating line and are displayed in

Figure 39.

31
2-D TRANSONIC TIME-MARCHING ANALYSIS

To verify the empirical blading criteria used in the MCA blade and vane designs, a twodimensional time marching analysis was performed on representative blade and vane sections.

The time-marching analysis is a potential flow solution to the equations of motion employing

artificial viscosity to model the shock system in the blade passage. The analysis requires a

definition of the airfoil section, the static pressure ratio across the passage, and the streamtube

convergence through the passage. By matching predicted design values for leading and trailing

edge air angles, the analysis confirms that the blade passage has sufficient choke margin with

the calculated shock system to pass design airflow with the required turning.

Table 6 presents a summary of the calculated time-marching results as compared to design

values for selected blade and vane sections. The close agreement between the calculated and

design values confirms the blading criteria used for MCA airfoil sections. Figures 40 through 43

display plots of static pressure distribution along a section chord for selected cases.

32
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8.5 9.0 9.5 10.0 10.5 11.0 11.5 12.0 12.5 13.0 13.5

Leading Edge Radius, cm

(a)

8.0 8.5 9.0 9.5 10.0 10.5 11.0 11.5 12.0 12.5 13.0

Leading Edge Radius, cm

(b)

FD 265584

Figure 21. V/STOl Fan Blade Loading — ID Stream Incidence at Max Control Т.О.

34
20.0

5.04

12.0 12.5 13.0 13.5 14.0

Trailing Edge Radius, cm

(a)

14.5 15.0

15.0

12.5

о

10.0

7.5

5.0

2.5

0.0

12.0 12.5 13.0 13.5 14.0 14.5 15.0

Trailing Edge Radius, cm

(b) FD 285585

Figure 22. V/STOL Fan Blade Loading — ID Stream Deviation at Max Control Т.О.

35
1.150

1.125^

1.100

1.075

1.050

1.025

1.000

8.5 9.0 9.5 10.0 10.5 11.0 11.5 12.0 12.5 13.0 13.5

Leading Edge Radius, Cm

(a)

1.4

1.3

1.2

о

б

1.0

0.9

0.8

8.5 9.0 9.5 10.0 10.5 11.0 11.5 12.0 12.5 13.0 13.5

Leading Edge Radius, cm

(b)

FD 265586

Figure 23. V/STOL Fan Blade — ID Stream Choke Margin at Max Control Т.О.

36
со

со

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"2

12 14 16 18 20 22

Leading Edge Radius, cm

(a)

24 26

О

12 14 16 18 20 22

Leading Edge Radius, cm

(b)

24 26

FD 265587

Figure 24. V/STOL Fan Blade — OD Stream Incidence at Max Control Т.О.

37
о

12.5 т

10.0

7.5

I 5.0

2.5

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14 16 18 20 22

Trailing Edge Radius, cm

(a)

24 26

10.0

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14 16 18 20 22 24

Trailing Edge Radius, cm

(b)

26

FD 265588

Figure 25. V/STOL Fan Blade — OD Stream Deviation at Max Control Т.О.

38
1.150

1.125

1.100

1

!075

1.050

1.025

1.000

12

2.0

1.5

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15

16 18 20 22

Leading Edge Radius, Cm

(a)

17 19 21 23

Leading Edge Radius, cm

(b)

24

25

26

27

FD 265589

Figure 26. V/STOL Fan Blade — OD Stream Choke Margin at Max Control Т.О.

39
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Figure 27. V/STOL Fan — ID Stream Stator Inlet Air Angle Relative to Cascade Loss

Bucket Range at Nominal Т.О.

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FD 265597

Figure 28. V/STOL — ID Stream Stator Inlet Air Angle Relative to Cascade Loss

Bucket Range at Max Control Т.О.

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Figure 33. V/STOL — OD Stream Stator Inlet Air Angle Relative to Cascade Loss

Bucket Range at Maximum Control Т.О.

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47
Page intentionally left blank

Page intentionally left blank
22% SM :

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VIGV = +18 deg

Surge at DF э ) = 0.72 " Rotor Peak

4 5 6 7 8

Corrected Airflow (W\/e/S) ~ Ibm/sec

82 84 86

Efficiency, % FD 263970

Figure 36. V/STOL Fan Map — ID Stream
Page intentionally left blank

Page intentionally left blank
10 20 30 40 50

Corrected Flow (W\/0/6) ~ Ibm/sec

60

Figure 37. V/STOL Fan Map — OD Stream FD 268972

51
Page Intentionally Left Blank
0 deg VIGV

-20 deg VIGV

;100% N^Vflf

VIGV = 0 degr

20 50 55 60

Corrected Flow (W\/076) ~ Ibm/sec

65 70 75 80 85

FD 268971

Figure 38. V/STOL Fan Map — Overall

53
Page intentionally left blank

Page intentionally left blank
о

20

50 60 70 80

% Corrected Speed

90 100 110

FD 268952

Figure 39. Speed us Flow

55
TABLE 6. — V/STOL FAN SUMMARY OF 2-D TRANSONIC

ANALYSIS

Blade Analysis — Max Flow

Design Values Time Marching Results

SL

ID

No.

4

9

% Span

35.4

83.2

0,

50.1

54.3

MIR

0.95

1.07

ft

11.3

20.0

«

0.047

0.109

ft,

48.6

53.9

MIR

0.98

1.07

ft

11.3

20.6

w

0.049

0.104

OD 2 14.2 52.6 1.30 37.3 0.100 53.7 1.30 37.3 0.100

4 38.0 56.5 1.43 50.9 0.107 56.5 1.43 51.2 0.107

6 58.1 59.0 1.53 59.0 0.123 58.8 1.54 59.0 0.128

10 92.3 63.0 1.69 66.3 0.164 62.9 1.69 65.4 0.164

Vane Analysis — Norn SLTO

ID 2 11.9 40.3 0.97 20.0 0.138 48.8 0.95 19.2 0.137

6 52.6 36.5 0.89 20.0 0.037 53.5 0.89 19.9 0.036

10 90.2 39.0 0.83 20.0 0.136 50.1 0.81 19.5 0.134

56
0.80

ID Blade Streamline Number 9

0.20

0.20 0.40 0.60

Percent Axial Chord

0.80 •1.00

FD 265575

Figure 40. V/STOL Fan Static Pressure Distribution

57
0.80

OD Blade Streamline Number 2

0.20

0.20 0.40 0.60

Percent Axial Chord

0.80 1.00

FD 265576

Figure 41. V/STOL Fan Static Pressure Distribution

58
0.70

:L~: OD Blade Streamline No. 10 ~:

0.10 Ш.

0.20 0.40 0.60

Percent Axial Chord

0.80 1.00

FD 265577

Figure 42. V/STOL Fan Static Pressure Distribution

59
0.80

0.70

ID Stator Streamline No. 2 XT:

0.20 0.40 0.60

Percent Axial Chord

0.80 1.00

FD 265578

Figure 43. V/STOL Fan Static Pressure Distribution

60
SECTION V

STRUCTURAL ANALYSIS

The VIBRA/NASTRAN structural analysis conducted to determine the steady stress

levels and free vibration characteristics of the preliminary V/STOL fan blade revealed a 1st

bending/2E resonance condition at 80% design speed. This condition, involving a known

excitation source at a fundamental vibration mode, was considered unacceptable even at part

speed. High steady stress loads at the blade root precluded a decrease in thickness to reduce

blade frequency and thus drop the 2E resonance below idle. Consequently, an iterative process

was initiated to design a part-span shroud that would increase the blade 1st bending frequency

sufficiently to obtain a 10 % margin above 3E resonance at design speed.

The first iteration series determined the optimum radial location for the shroud. The

second series sized the shroud to minimize airflow blockage while maintaining acceptable stress

levels in the shroud and airfoil. As a part of this study, the flow path splitter thickness was

reduced from 0.457 cm (0.180 in.) to 0.203 cm (0.080 in.) and the blade root thickness increased

from 7% to 7.5% to reduce root stress. A third series of iterations was required to locate the

shroud as far aft axially on the airfoil as possible. The second and third iteration series were

repeated several times since stress levels in the shroud due to blade untwist increased

significantly as the shroud was moved aft of the airfoil center of gravity.

As a result of these studies, the optimum shroud and flowpath splitter configurations

presented in Table 7 were determined.

TABLE 7. — BLADE SHROUD/SPLITTER CONFIGURATION

Shroud:

Splitter:

Radial Location

Radial Thickness

Axial Length

Axial Location

First Bending Frequency

Radial Thickness

19.304 cm (7.600 inches) about 50% of fan

side airfoil span

0.318 cm (0.125 inch)

2.540 cm (1.000 inch)

65% airfoil chord

1011 Hz (14% above 3E)

0.203 cm (0.080 inch — reduced from 0.180

inch)

The resulting structurally acceptable blade as defined in Table 7 necessitated the

aerodynamic redesign of both ID and OD flowstream to transform the strucutural requirements

into an acceptable aerodynamic design. The addition of the part-span shroud increased blade

pull and steady stress at the blade root and added flow blockage to an already high specific flow

design. Consequently, adjustments to the flowpath around the splitter were made to provide

sufficient annulus area to pass maximum flow in both flowstreams while controlling aerodynamic loading. The redesigned blade featured revised metal angle and thickness definition to

accommodate the part-span shroud as well as flow splitter modifications necessary to achieve

the aerodynamic goals at the maximum flow condition.

The structural analysis and the flutter/stability margins of this final blade design and the

IGV/EGV struts and flaps are presented in the following sections. The inlet guide vane case

(AMS 5616) has. 17 integrally machined struts with a full ring splitter separating core and fan

flows and independently variable core and fan flaps. The rotor (AMS 4973) is an integrally

machined bladed disk (blisk) which also incorporates a full ring splitter separating core and duct

61
flows and a full ring part span shroud. The exit guide vane case (AMS 5616) has 38 integrally

machined vane struts with variable leading edge fan flaps and a full ring splitter separating core

and fan flows through the discharge.

Blade Analysis Steady Stress

Figure 44 depicts the finite element breakup used to conduct the steady stress analysis for

the fan blade. Plate elements were used to model the airfoil, and beam elements were employed

to model the integral splitter and part!span shroud. Considerable effort was expended to obtain

a geometrically accurate finite element model of the airfoil. The high rate of blade twist and

steep flowpath angle for this airfoil caused numerical incompatibilities between the aero

geometry preprocessor and VIBRA.

Radial stress contour plots for the airfoil due to inertia loads at maximum rpm are

presented in Figures 45 and 46. The plots show that maximum steady stress levels occur at the

junction of the airfoil and the flowpath splitter. These high stresses are quite localized and are

partially induced by the use of rigid body elements to attach the splitter to the airfoil. Figure 47

depicts the airfoil steady state stress distribution just above the fillet radius runout for the

splitter. This plot superimposes the stress levels due to inertial loads from the NASTRAN

analysis with the gas bending stresses obtained from the airfoil stress deck and is representative

of the predicted maximum stress levels which will be seen in the airfoil. The maximum

concentrated airfoil stress (a nom X Kt) is 51K Newtons/cm2 (74K psi) and occurs on the

concave surface at approximately 55% chord. This is within the design allowable of 62K

Newtons/cm2 (90K psi) or 75% of 0.2% Y.S.

Stress levels for the flowpath splitter and part span shroud were also obtained from the

NASTRAN analysis. Both the splitter and shroud were conservatively modeled as single span

beam elements attached to the airfoil with rigid body elements. The splitter and shroud were

sized to minimum thickness levels in order to reduce airflow blockage while maintaining

combined hoop and bending stresses to acceptable levels. Maximum steady stress levels are

predicted to be 58K Newtons/cm2 (85K psi = 30K psi dh + 55K psi o!bend) for the part span

shroud and 41К Newtons/cm2 (60K psi = 20K psi <rh + 40K psi ffbend) for the flowpath splitter.

Although analytical predictions indicate that the blade satisfies the basic design criteria,

this unique airfoil configuration requires that the blades be instrumented to monitor and verify

these steady stress levels during rig test.

Vibration

The finite element geometry used for the VIBRA/NASTRAN free vibration analysis of the

V/STOL fan blade is shown in Figure 48. The original airfoil design was predicted to have a 1st

bending/2E resonance in the high power running range. To eliminate this potentially severe low

order resonance, a full ring part span shroud was added to the airfoil at approximately 50% span

outboard of the splitter R = 19.431 cm (7.65 in.). With the addition of the shroud, the 1st

bending blade alone frequency was increased to 1011 Hz, providing a 14% frequency margin

above 3E at maximum operating speed as shown in Figure 49. Figures 50 through 54 are modal

displacement plots at 17,762 rpm for the rotor vibratory modes below 17E (17 IGV's).

62
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! !!*"!'x's?

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Figure 44. Blade Steady Stress Finite Element Model .

FD 269051

63
SYMBOL VALUE (PS*) SYMBOL VALUE

IT !1.16366&E+04

2 !5.140484Е+03

3 1.355707Е+03

4 7.85189SE+03

5

6

7

в

2.734047Е+04

3.333666Е+04

13 6.631756Е+0*»

7.281375E»0* "77930 99"4

16

FD 269052

Figure 45. Blade Radial Stress Contour Plot — Convex Surface

64
SYMBOL VALUE SYMBOL VALUE

I !37t

г !г.<

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~5—=Я7329437Г+ОГ

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12 4.140277Е+0*

14 5.S6119GE+04

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FD 269053

Figure 46. Blade Radial Stress Contour Plot — Concave Surface

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Figure 48. Blade Free Vibration Finite Element Model

FD 265643

67
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FD 265633

Figure 50. Blade Modal Deformation at 17,750 rpm, Mode 1, 1010.988 Hz

FD 265634

Figure 51. Blade Modal Deformation at 17,750 rpm, Mode 2, 1637.072 Hz

69
FD 265635

Figure 52. Blade Modal Deformation at 17,750 rpm, Mode 3, 2472:997 Hz

FD 265636

Figure 53. Blade Modal Deformation at 17,750 rpm, Mode 4, 2690.768 Hz

70
FD 265637

Figure 54. Blade Modal Deformation at 17,750 rpm, Mode 5, 2863.475 Hz

Tip mode analysis indicates there are 3 higher order complex modes which are within the

17E excitation in the 50% flow (12,000 rpm) to maximum power (17,800 rpm) running range.

These consist primarily of a complex mode at 3497 Hz (Figure 55), a LE tip flap mode at

4360 Hz (Figure 56) and a LE bending mode at 4951 Hz (Figure 57). They occur well above the

typical chordwise bending modes of concern in fan designs and, based on experience, are

normally moderate to low risk. However, due to the wide incidence range within which the

blades will be required to operate, it is recommended that extensive testing be done to position a

tip strain gage and establish stress ratios that will permit accurate monitoring of tip modes as

well as the fundamental vibratory modes during rig tests. Figures 58 and 59 are modal

displacement plots for the blade at frequencies intersecting 17E and 21E (# EGV-# IGV) near

max power, i.e. 5378 and 6053 Hz. Table 8 summarizes the rotor vibratory frequencies for 0 rpm,

12,000 rpm (50% flow) and 17762 rpm (maximum power).

71
FD 265636

Figure 55. Blade Modal Deformation at 17,750 rpm, Mode 6, 3497.499 Hz

FD 265639

Figure 56. Blade Modal Deformation at 17,750 rpm, Mode 7, 4360.125 Hz

72
FD 265640

Figure 57. Blade Modal Deformation at 17,750 rpm, Mode 8, 4950.768 Hz

FO 265641

Figure 58. Blade Modal Deformation at 17,750 rpm, Mode 9, 5377.659 Hz

73
FD 265642

Figure 59. Blade Modal Deformation 17,750 rpm, Mode 10, 6052.523 Hz

TABLE 8. — V/STOL BLADE FREQUENCY SUMMARY

Mode Shape 0 rpm 12,000 rpm 17,776 rpm

1st Bending

1st Torsion

2nd Torsion

2 Nodal Tip

1st Chordwise Bending

Complex

LE Tip Flap

Complex

Complex

Complex

716

1554

1834

2358

2674

3256

4200

4650

5298

5916

864

1593

2357

2416

2681

3371

4275

4790

5335

5979

1011

1637

2473

2691

2863

3497

4360

4951

5378

6053

74
Flutter

A supersonic unstalled flutter analysis was conducted for the fan blade at maximum flow

conditions. The results of this analysis are shown in Figure 60. The blade exhibits positive

aerodynamic damping for the first four modes of vibration, and no flutter/instability problems

are predicted.

Durability

A Goodman diagram for the V/STOL fan blade is presented in Figure 61. Combined steady

stress and vibratory stress levels were computed at several locations on the blade for 1st bending

and 1st torsion modes. Vibratory stresses were normalized by assuming a maximum vibratory

stress of 21K Newtons/cm2(30K psi) peak to peak for each mode of vibration. The plot shows

that the blade has in excess of twice the endurance limit capability of the material and will

provide adequate HCF durability during test.

Vane Analysis Steady Stress

Steady stress levels for the inlet guide vane and the exit guide vane were computed using

the airfoil stress deck. Very low stress levels (< 10% Y.S.) were predicted for all points

analyzed. Stress levels for the variable vane actuation rods were also predicted to be low even

during surge (5 X airload) conditions.

Vibration

Natural frequencies and mode shapes for the inlet guide vane and exit guide vane were

determined using VIBRA/NASTRAN and treating the fixed and variable segments of the vanes

as independent systems. The analyses predict that all vane fundamental modes will be below the

24E excitation source of the rotor and no resonance problems are anticipated. Campbell plots

summarizing these results are presented in Figures 62 through 66.

Flutter

Pressure ratio vs V parametric correlations were conducted in accordance with P&WA

guidelines for both the inlet guide vane and the exit guide vane systems. These results are

summarized in Figures 67 and 68 and indicate that the vanes meet the established flutter design

experience curves. However, we have no practical experience or calibrated analytical technique

for more accurately verifying the degree of stability of a LE flap configuration such as exists for

the exit guide vane. Consequently, we recommend that the exit guide vane flap be instrumented

to monitor vibratory stress levels for this airfoil.

RECOMMENDED STRAIN GAGE LOCATIONS

The airfoils of the proposed V/STOL rig satisfy the general P&WA fan design criteria but

in light of the unique blade design and inherently low mechanical damping of the blisk, it is

recommended that both steady and vibratory stress levels be monitored during testing as a rig

health precaution and to verify the analytical predictions. The recommended location and

orientation for four blade and one EGV strain gage are presented in Figure 69. Due to the

limited capacity of the W-8 facility slip ring, and the recommended use of two steady stress

gages (locations No. 1 and No. 3), only five gages can be monitored. These gages should be

distributed on two blades 180 deg apart. No routing provisions for this installation have been

provided by P&WA other than a 2.54 cm (1 in.) diameter hole through the blisk bore (L237441)

75
to permit rig testing in the 3.28m X 1.82m (9 ft X 5 ft) wind tunnel and the slip ring mounting

scheme (L237442-sheet 4) for the W-8 test facility. A drawing (HKJ 3397) was provided

detailing the EGV flap spindle hole required for strain gage routing to the suction surface of

flaps 12 and 30.

CONCLUSIONS AND RECOMMENDATIONS

1. Airfoil steady stress levels are within the design allowable of 75 % of 0.2 % Y.S.

2. The airfoils are free of any expected integral order excitations for all fundamental

vibration modes in the high power running range.

3. Blades and vanes satisfy the flutter design criteria.

4. Although the blade is predicted to have a 17E (17 IGV's) excitation crossing for 2 complex

tip modes in the high speed running range, both modes are well above the typical

chordwise bending tip modes normally of concern in fan designs.

5. The blades should be instrumented to monitor both steady state and vibratory stress levels

during rig test.

6. The exit guide vane LE duct flap airfoil should also be instrumented to monitor vibratory

stress levels for this unique airfoil configuration.

76
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Figure 61. NASA V/STOL Fan Blade Goodman Diagram

FD 265566

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86
SECTION VI

MECHANICAL DESIGN

The mechanical design of the dual flowpath V/STOL fan stage rig is presented in this

section. The rig was designed to fit within the existing NASA LeRC W!8 facility, incorporate

integrally machined rotor and stator cases to minimize cost and assembly time, and incorporate

existing NASA actuation/position system components.The basic rig cross section is shown in

Figure 70. The radial and axial flowpath rig station locations, dimensions, and actuation

component placement are shown in Figure 71. Vane position read!out component design and

placement are shown in Figure 72 and the slip ring support tube assembly is shown in Figure 73.

The integrally bladed rotor is presented in Figure 74.

Material selection was consistent with reference parts (usually SAE 1010!1030 mild steel

with an AMS 2485 black oxide finish) unless there were additional considerations. Inconel 718

(AMS 5665) material was used for the sync rings because of strength requirements. Greek

ascoloy (AMS 5616) was used for the IGV and EGV strut cases, airfoil flaps and flowpath

splitters because of raw material availability and instrumentation considerations. Aluminum

was used for the slipring support tube and permitted as an optional material for the exhaust

splitter because of weight. The rotor is to be machined from an 8.1.1 Titanium (AMS 4973)

pancake forging.

Rig components other than airfoils were designed consistent with reference parts.

Drawings of airfoils are done to engine specifications (PWA 360); all other components are of a

format similar to reference drawings provided by NASA. An explanation of drawing symbols,

material specification, and processes is provided in a separate volume (reference 5). The

information available in this volume is further defined in Section VII. Airfoil coordinates are

provided on tape in a 9 track, 1600 BPI, 80 block size format consistent with the requirements of

numerically controlled milling machines and are tabulated in Appendix С in P&WA 390 format.

The design considerations and philosophy for the major rig subassemblies are presented

from rig inlet to exit in the following paragraphs.

INLET CASES

The ID and OD inlet cases provide a transition between the existing W!8 plenum

configuration as defined by NASA drawing No. CF 851937!B and the V/STOL fan stage rig IGV

strut case. The cases are welded assemblies made of SAE 1010!1030 steel with a black oxide

finish. The 2.032 cm (0.800 inch) thick OD case is designed to accept the required NASA station

0 and 1 instrumentation as described in Section VII. Similarly the ID case contains the station 0

instrumentation and is designed to support the slipring mount tube as shown in Figure 73. Both

OD and ID cases have a 16 micro inch finish requirement for the flowpath surfaces and 6 jack

screw holes at each flange end to facilitate disassembly. The ID case flanges each have four

notches for routing the 56 scheduled static pressure lines into the plenum. NASA must grind 4

slots into the plenum strut assembly (NASA drawing No. CR 846186) matching those on the ID

case and provide any umbilical system needed to route the instrumentation through the plenum

wall.

87

S850C
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SLIP RING MOUNT TUBE

The slip ring mount tube is designed to permit installation and coupling of the NASA slip

ring to the rotor while the plenum is in the forward position. (See Section VIII for assembly

procedures). Once the slip ring/mount tube subassembly is coupled to the rotor, the plenum can

be moved aft into position as the slip ring mount tube slides forward into the plenum on the ID

flowpath forward and aft flanges. The split tube support (HKJ 3380) is then installed to rigidly

position the mount tube during testing. A spacer, item 115 (HKJ 3381) of Figure 73, can be

machined to properly locate the mount tube axially and prevent compression or extension of the

fragile slip ring to rotor flex coupling. The mount tube is designed of 6061!T6 aluminum to

minimize weight. The tube support is made of SAE 1010!1030 steel with a black anodize finish.

No aerodynamic fairing or nose cone was designed to cover the split tube support; NASA must

design this hardware after devising a routing scheme for the instrumentation through the

plenum.

INLET GUIDE VANE (IGV)

The inlet guide vane assembly shown in Figure 70 includes an integrally machined LE case

with 17 struts and a full ring splitter which extends from 6.6 cm (2.6 in.) upstream of the IGV

LE aft to the rotor LE (separating the core flow from the fan duct flow), independently variable

core and fan flaps, aft OD and ID cases and miscellaneous flap actuation components.

The IGV LE strut case was designed to be integrally machined as opposed to welded or

bolted to save tooling, machining, and assembly costs and to improve concentricity, tolerances,

and flowpath continuity. Greek ascoloy (AMS 5616) material was selected because raw material

was available, it will not rust when highly polished, and is easy to machine and instrument.

. The aft ID and OD cases bolt to the LE strut case along the plane of the IGV core and fan

flap spindle centerlines. The 8 deg tilt of the spindle centerlines decreases the fan flap ТЕ OD

endwall clearance in the axial position by approximately 0.25 cm (0.100 in.) and permits

installation of the core flap through the integral splitter. A computer graphics generated picture

of the strut case with the aft ID and OD cases but without core and fan flaps or LE and ТЕ

splitters is shown in Figure 75.

The splitter LE and ТЕ ring segments are located on snap diameters and riveted to the

strut case during assembly. It is important to note that the IGV core flap cannot be removed

from the IGV assembly without first removing the splitter ТЕ segment. The LE splitter/strut

case junction was located upstream of the instrumentation Station 2 to simplify installation.

The ТЕ splitter/strut case junction was located close to the core flap spindle to minimize the

raw material envelope and to maximize machining access. Station 3 splitter instrumentation

must cross this junction; making IGV core flap removal after this instrumentation has been

installed, very costly.

The aft ID case traps the core flap ID platform against the strut case ID to position the

core flap and limit its radial movement. The fan flap platform is positioned radially by a flanged

steel bushing with inexpensive Teflon inserts at the ends to provide low friction thrust and

bearing surfaces for the flap spindle. The bushing was required following the line boring

operation for the fan flap spindle platform. The bushing flange bolts to the strut case OD

permitting removal of the aft OD case without disassembling the fan flaps. The fan flap bushing

and the sync arms for both the core and fan flap are serialized to permit component matching

with strut location. Matching is necessary because the flaps are dimensioned to nominal and will

require individual hand fitting of components to obtain design endwall gaps during initial

99
assembly (see Section VIII). This dimensioning approach was taken to minimize performance

losses due to greater than nominal endwall gaps. Serialization ensures correct reassembly.

The sync arm to spindle interface design insures positive engagement and the axial

orientation of the tapered flats eliminate assembly problems due to tolerance stack!up. This

design provides accurate chord angle control and leaves the spindle end clear for routing

instrumentation. The sync arms incorporate available monoball bearing assemblies to reduce

cost and procurement time.

The IGV core flap (T4060682) spindle/sync arm locating flats were oriented so that the

X!axis of the airfoil as defined by the Engineering Master Drawings (EMD's) is parallel to the

sync arm centerline. The core flap actuation system and endwalls were designed to permit + 20

deg rotation from the SLTO nominal or 0 deg position. The ADP position is !(!18 deg.

To minimize the axial movement of the IGV fan flap sync ring and the length of the sync

arms, the fan flap (T4060681) spindle/sync arm locating flats were oriented such that the X!axis

of the airfoil as defined by the EMD's is 20 deg closed (—) when the sync arm is axial (parallel to

the rig centerline). To rotate the airfoil X!axis from +20 deg to !60 deg, the vane sync arm must

be moved from +40 deg to !40 deg. SLTO nominal or 0 deg orientation of the airfoil X!axis

therefore requires a +20 deg sync arm position. The ADP or +18 deg airfoil X!axis orientation

requires a +38 deg sync arm position.

Full 360 deg sync rings were incorporated to reduce machining costs and eliminate vane

angle variations inherent in split ring designs due to ovalization. The fan and core flap sync

rings are positioned by 8 equally spaced 1.27 cm (0.500 in.) diameter Teflon plugs that slide

along the IGV OD case as shown in Figure 70. The sync rings are each actuated at 2 points

approximately 180 deg apart by existing NASA actuation systems. This sync ring support and

actuation approach minimizes the number of components and maximizes sync ring concentrici!

ty. No mechanical stops to prevent over rotation of these sync rings and the damage resulting

from vane end contact with the flowpath walls have been provided. NASA is responsible for

setting actuator travel limits to prevent rig damage.

All IGV ID and splitter components were designed to accept the required station 1 and 2

static pressure instrumentation described in Section VII. The fan OD static pressures at the

IGV LE and ТЕ were installed in the inlet and rotor cases, respectively. The core ТЕ ID case

static pressure taps were designed as intersecting holes in a special "flange" to eliminate the

potential interference of routed hypo tubing and the installation of balance weights on the

forward flange of the rotor. Provision was also made for the routing of hypo tubing from the

inlet duct ID along the surface of the IGV core strut to the fan side of the splitter to

instrumentation Stations 1 and 2. When possible, instrumentation channels were made similar

to NASA design practice and large enough to accept 0.062 in. diameter hypo tubing.

A cursory vibratory analysis of the IGV strut case, core and fan flaps was performed. No

integral order resonances, torsional flutter response or other flow induced instabilities in the rig

running range were present. No strut case stress analysis was performed. The core flap 0.51 cm

(0.20 in.) diameter, 20.32 cm (8 in.) long spindle was calculated to deflect 0.5 degrees when a

conservative 5 x normal surge airload of 4 Newtons (18 Ibf) was applied; spindle shear stress for

this condition was 5.5K Newtons/cm2(8K psi). The allowable shear stress for AMS 5616 at

149°C(300°F) is 57% of the 0:2% yield strength or 41K Newtons/cm2 (60K psi). The fan flap,

although much larger, has a 1.27 cm (0.500 in.) diameter, 7.62 cm (3.00 in.) long spindle and no

spindle stress or deflection problems are anticipated. However, the surge loads transmitted to

the sync ring actuator pin were adequate to require a larger diameter pin 0.48 cm (0.190 in.) to

reduce the 22.5 Newtons (100 Ibf) airfoil surge load bending stress within the pin to an

100
acceptable 48K Newtons/cm2

 (7OK psi). The allowable bending stress for AMS 5663 at 38°C

(100°F)isl50% of the 0.2% yield strength or 152K Newtons/cm2 (220K psi).

Rotor Case and Tip Shroud Inserts

The rotor case (Figure 70) is designed to accept 2.54 cm (1.000 in.) deep tip shroud inserts

and is split on the vertical plane per NASA request to permit on!stand tip shroud configuration

changes. Alignment of the rotor case halves is maintained with 4 split line dowel pins;

concentricity is insured with fore and aft snap diameters which also minimize split case

ovalization. The IGV and EGV sync rings are located fore and aft of the rotor case flanges to

permit case removal and ready access to rotor tip shroud instrumentation. The solid rotor tip

shroud is coated with METCO 601 (a plasma sprayed mix of blended polyester and aluminum

alloy powder) to permit shroud abrasion without blade tip damage. The shroud design provides

15 static pressure taps across the blade tip as defined in Section VII. NASA is to provide a

scheme for routing this instrumentation to the rig exterior. Split line case bolt tensile stresses

were calculated to be less than 2.1K Newton/cm2(3K psi). No cold to hot rotor case flowpath

dimensional correction was made for the predicted 0.019 cm (0.0075 in.) radial growth at the

ADP thermal condition.

ROTOR

The rotor was designed as an integrally machined bladed disk (blisk) for cost savings, and

incorporates a full ring splitter extending from LE to ТЕ to separate core flow from fan duct

flow and a full ring!part span shroud to eliminate low order fundamental mode resonances in

the running range. Per NASA request, front and rear balance flanges were included for detail

and assembly balancing as was a 1.27 cm (0.500 in.) dia hole through the disk bore for strain

gage lead routing should the rig be tested in the NASA LeRC wind tunnel facility. These

features are shown in Figures 76 and 77 which were generated with computer graphics using

actual disk dimensions and airfoil coordinates, i.e., the "blisk" shown has yet to be fabricated.

The disk configuration was copied from NASA designs to insure compatibility with the

existing W!8 facility shaft and slip ring coupling. The blisk meets the general disk design criteria

and has good LCP life as shown in Tables 9 and 10. Burst margin and yield margin are 2.362 and

2.213 respectively. All fillet radii were made as large as possible to avoid life problems (Figure

78) and still maintain functionality of the parts. The rear balance flange design was checked

using a procedure based on P&WA experience which is conservative and calculates life using

worst case tolerances with the largest balance weight permissible. The front balance flange was

scalloped so that balance weights could be added to the rotor after assembly. The clearance

holes for the bolts were based on what was used on reference NASA prints (CF 851447 and

CC848383). The hole spacing as well as the single hole offset were selected to be compatible with

NASA rig hardware. The radial dimensions of the blisk were converted from "hot ADP" to cold

static conditions, i.e., the dimensions have thermal as well as dynamic growth factored into them

to give the desired running clearances. The thermal conditions for which calculations were made

are presented in Figure 79. The ADP running clearances are 0.051 cm (0.020 in.) for the blade

tip and 0.0741/0.053 cm (0.029/0.021 in.) at the LE/TE of the flow splitter. The cold gaps at the

flow splitter LE/TE are 0.056/0.910 cm (0.022/0.036 in.). Figure 80 presents a hot/cold running

clearance summary.

101
FC 83395-H

Figure 75. IGV Strut Case (Computer Graphics)

102
FC 83382

Figure 76. NASA V/STOL Fan Blisk (Computer Graphic)

103
FE 221508

Figure 77. NASA V/STOL Fan Blisk Airfoil/Section (Computer Graphics)

No tiebolt analysis was conducted since NASA is to provide the tiebolts.

A critical speed analysis is to be conducted by NASA because LeRC has all pertinent

facility definition and an analytical model of the drive system. The rotor/blisk information

necessary for performing a critical speed analysis is presented in Table 11.

104
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105
Max

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Typical

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Life Greater

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Life Greater

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Radius

Life 30,000 Cycles

Fillet Radius

Life Greater

than 50,000 Cycles

Fillet Radius

Life Greater

than 50,000

Cycles

Balance Weight

Attachment Hole

View A

• Life Greater than 50,000 Cycles

Balance Flange Scallop

Figure 78. LCF Type I Life Cycle Summary

106
Temperature - °F

Confidence Level ±50

FO 268924

Figure 79. V/STOL Fan Rig Rotor/Case Thermal Map

107
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All Dimensions Cold Unless

Otherwise Specified

•The Case Grows Forward With Temperture, Reducing Clearances

Aft of the Rotor and Increasing Clearances Forward of the Rotor

Figure 80. Critical Rotor Assembly Clearances (Minimum)

108
TABLE 11. — CRITICAL SPEED INFORMATION

Configuration

Disk

Blades

Splitter

Shroud

Weight

9.750 kg

(21.5 Ibm)

3.439 kg

(7.584 Ibm)

0.630 kg

(1.39 Ibm)

0.220 kg

(.485 Ibm)

Polar Moment of

Radius . Remarks

6.736 cm Centerline to ID flowpath

(2.652 in.)

17.333 cm ID flowpath to OD flowpath minus splitter ring inter!

(6.824 in.) section with blade. No fillet radii included.

13.818 cm Assumed full ring. No fillet radii were included.

(5.44 in.)

19.304 cm Assumed full ring (the shroud/airfoil volume is

(7.60 in.) fore duplicated): No fillet radii were included.

Inertia 0.5196 kg!cm!sec2

0.451 lb!in.!sec2

there!

No provisions for routing SG leads were made for the V/STOL blisk.

In conclusion, the V/STOL blisk design satisfies accepted criteria and there are no areas of

concern. The stress levels are in general low.

Exit Guide Vane (EGV)

The exit guide vane assembly shown in Figure 70 includes an integrally machined case with

38 vanes and a full ring splitter which extends from the rotor ТЕ aft through the diffuser, an OD

forward case, a variable LE fan flap and miscellaneous flap actuation components.

The EGV vane case was, like the IGV, designed to be integrally machined for cost savings

and made of AMS 5616. The splitter extends aft of the EGV ТЕ through the diffuser and is

located on a snap diameter and riveted into place. The vane case/splitter junction was located

close to the ТЕ to minimize the EGV raw material envelope and to maximize machining access

when fabricating the airfoils and installing instrumentation.

The forward OD case bolts to the vane case along the plane of the flap spindle centerlines.

Flanged steel bushings with Teflon bearing inserts are used to locate the EGV LE fan flaps. The

bushing flange bolts to the strut case OD permitting removal of the forward OD case. The

bushing and the flap sync arm are serialized for component matching, and like the IGV flaps,

require individual hand fitting of components during initial assembly to obtain the design

endwall gaps. The sync arm to flap spindle interface design is the same as the IGV and available

mono!bal l bearing assemblies were again used to reduce cost and procurement time.

The EGV fan flap spindle/sync arm locating flats were oriented so that the X!axis of the

airfoil as defined by the EMD is parallel to the sync arm centerline. The flap actuation system

and endwalls were designed to permit +10 deg rotation from the SLTO nominal or 0 deg

position. The ADP position is +3 deg.

A full 360 degree sync ring made of AMS 5665 with 12 Teflon support sliders which ride on

the strut case OD is used to actuate the EGV LE fan flap: Two existing 3.302 cm (1.300 in.)

stroke NASA actuation systems located approximately 180 deg apart as shown in Figures 70 and

71 move the sync ring. No mechanical stops to prevent over rotation of the sync ring (and the

damage resulting from vane end contact with the flowpath walls) have been provided. NASA is

responsible for setting actuator travel limits to prevent rig damage. Vane rotation will be

measured using existing NASA potentiometers mounted as shown in Figure 72.

109
The EGV ID and splitter are designed to accept the required LE (station 5) and ТЕ

(station 6) static pressure instrumentation listed in Table 13 (see Section VII). The fan OD

static pressures at the IGV LE and ТЕ are installed in the rotor and diffuser cases, respectively.

The core LE ID static pressure taps are designed as intersecting holes in a special "flange" to

eliminate the potential interference of routed hypo tubing and the installation of balance

weights on the aft flange of the rotor. Provision was also made for routing hypo tubing from the

ring ID cavity along the suction surface of the EGV core to the fan side of the splitter and

forward and aft to instrumentation stations 5 and 6. Where possible, instrumentation channels

were made similar to NASA design practice and large enough to accept 0.157cm (0.062 in.)

diameter hypo 6 tubing.

Provision for three spans of core LE Pt/Tt sensors at the centers of equal area are also

provided. Channels from the vane LE along the pressure surface to midchord and then radially

inward through the ID wall are detailed on the EGV vane case drawings. Five spans of fan flap

LE Pt/Tt sensors at the centers of equal areas are detailed on the EGV fan flap instrumentation

drawing No. HKJ 3400.

The results of the vibratory and stress analysis for the EGV vane case and LE flap are

presented in Section IV. Very low stress levels were predicted for all points analyzed including

the flap spindle with a 5 X nominal surge airload condition. The vane case and LE flap vibratory

analysis indicated no integral order resonance flutter response or other flow induced instabilities

are expected; however, it is recommended that the EGV LE flap be strain gaged and the stress

levels monitored during testing because P&WA has no practical experience for accurately

verifying the degree of stability of the LE flap configuration.

DISCHARGE DUCTS

The discharge extends from the EGV ТЕ to rig Station 114.839 and consists of those

components shown assembled to the EGV in HKJ 3382. The discharge ID, OD and splitter are

made of SAE 1010!1030 steel with an AMS 2485 black oxide finish. Instrumentation holes are

provided in these components at Stations 6 and 7 as defined by Table 13 of Section VII. The

forward end of ID case bolts to the EGV strut case; a 0.025 — 0.041 cm (0.010!0.016 inch) gap

exists between the aft end of the ID case and the bearing support structure. Per NASA design

practice, this gap was not sealed to prevent rotor discharge leakage from returning to the

flowpath upstream of the Station 7 radial probes. The splitter is riveted to the EGV and

centered downstream by locating shoulders on the four cross routed instrumentation tubes. The

OD case is a welded assembly with 2.032cm (0.800 in.) thick walls to permit the installation of

existing NASA circumferential traverse systems, the direct tapping on instrumentation holes

and the mounting of actuation system components. To insure correct circumferential alignment

when assembling these ducts to the EGV for line drilling the dowel pin, instrumentation tube

and Station 7 radial probe holes per HKJ 3382, "Top <£ " markings are provided instead of

offset holes. Two 1.27cm (0.500 in.) diameter threaded holes are provided at the top of the OD

case for lifting eyes.

The discharge splitter core side contour was designed to provide constant area flowpaths at

instrumentation Stations 6 and 7 and a uniform increase in area without exceeding the 7 degree

half-angle for diffusion through to the diffuser. An abrupt change or "dump" in the outer wall of

the core flowpath was incorporated aft of Station 7 to control the location of flow separation.

The discharge splitter fan side contour provides a constant area flowpath past Stations 6 and 7.

The cross routed instrumentation tubes are each designed to permit the routing of 25 0.165

cm (0.065 in.) diameter hypo tubes from the rig ID and have been aligned with the duct and core

side flow to minimize blockage at the aero design point. Plugs were designed for the Station 6 ID

110
and Station 7 OD and splitter radial discharge probe holes but none were designed for the four

circumferential traverse unit slots at Station 6.

Once the EGV and discharge OD, ID and splitter are assembled, line drilled for the

instrumentation tubes and radial probes and all the instrumentation installed and routed, it

should remain assembled.

DIFFUSER

The diffuser design goal was to provide a uniform increase in core and fan flow area from

the discharge exit at rig Station 114.839 (in.) to the existing NASA dual discharge valves and

exhaust collector. This was done by contouring the diffuser splitter core side in conjunction with

the existing bearing support contour to provide a uniform increase in area through to the core

side discharge valve without exceeding the 7!degree half!angle for diffusion. Similar consider!

ations determined the contour of the diffuser splitter fan side and the OD diffuser through to

the fan side discharge valve.

The diffuser splitter and the OD diffuser were designed of SAE 1010!1030 steel with a

black oxide finish. Two 0.030 cm (0.120 in.) deep channels are provided for routing four core OD

and four fan ID flowpath surface static pressure tapes from the discharge instrumentation

Station 7 aft along the splitter OD surface and through the exhaust collector wall via hollow

bolts as required by NASA design practice. The splitter forward end is notched to locate it

radially relative to the discharge splitter. No seal was incorporated to prevent leakage across this

interface. The OD diffuser contains no instrumentation.

ACTUATION SYSTEM COMPONENTS

The actuation system mechanical design goals were to: copy the NASA dual actuator

concept, use NASA actuators (part No. CF 853171) and rod ends (part No. HKJ3413), minimize

the number of new components, and design common parts. These goals were achieved for the

IGV core and fan ТЕ flaps and for the EGV LE fan flap as shown in Figure 71. Because of OD

case instrumentation and sync ring size considerations, the design of the case and sync ring

mount brackets were unique for each variable vane row. The case brackets were made of SAE

1010!1030 and the IGV fan flap sync ring bracket of AISI 410 stainless. The design of the

actuator rod end stand!offs was made of AISI 410 stainless for strength and of hex!stock to

permit wrenching; a groove/weld scheme was incorporated to permit safety wiring. The motor

guides were made of SAE 1010!1030 steel with a black oxide finish and can be formed from

0.48 cm (0.188 in.) flat plate as shown by HKJ 3402.

The two NASA gearmotor/jackscrew actuators located 180 deg apart to drive the variable

vane sync rings are rated at 315 Newtons (1400 Ibf) each. Their combined 630 Newtons (2800

Ibf) is considerable and if sync ring travel is not limited to 4.89, 2.34 and 1.19 cm (1.928, 0.923

and 0.469 inches) for the IGV core, fan and EGV fan flap either mechanically or electronically

over rotation will occur and the vane ends will be damaged when driven into the flowpath walls.

No mechanical or electrical stops were designed by P&WA and NASA is responsible for

incorporating a control/safety feature. The 6.04cm (2.380 inch) stroke actuator was used to vary

the IGV fan flap; 3.30 cm (1.300 inch) stroke actuators were used for the IGV core and EGV fan

flaps.

POSITION READ OUT COMPONENTS

The design objectives for the position read!out components were to provide dual read!out

capability for each vane row with existing NASA potentiometers and a minimum number of new

parts. This was achieved with the design of two mounting brackets and three shaft connectors

111
for the NASA potentiometer assemblies as shown in Figure 72. The mounting brackets are

fabricated from 0.480 cm (0.188 in.) thick SAE 1010-1030 .plate with a black oxide finish and

located 180 deg apart on the rig case for redundant measurements. The connectors for the IGV

case flap and the EGV fan flap potentiometers are made from hex-stock to permit wrenching

and can be safety wired once installed. The IGV fan flap mounting bracket is a "C" shaped subassembly made from 1/16 AISI410 steel with a machined collar for the potentiometer shaft. The

symmetry of the rotor case flange bolt hole patterns to the IGV and EGV airfoil spindles

permits relocating the brackets over other vanes should an instrumentation or other

consideration create a location conflict. The NASA potentiometer sub-assembly mounting

bracket hole pattern was not available and NASA must locate, drill and tap these holes into the

six P&WA designed brackets.

112
SECTION VII

AERODYNAMIC INSTRUMENTATION

The aerodynamic instrumentation summarized in Table 12 was requested by NASA and

incorporated with minor exceptions into the mechanical design of the V/STOL fan stage rig.

The type of instrumentation and its circumferential, axial and radial location within the core

and fan side flowpaths of the rig are presented in Table 13. The instrumentation is listed by

instrumentation station from the inlet (Station 1) to the discharge (Station 7) with angular

position increasing clockwise (the direction of rotation) when looking aft. Axial position is

referenced to the rotor/shaft interface arbitrarily identified as rig location 100.00 (in.). Figure 81

(L238020 sheet 1) includes: a flowpath cross section showing the axial location of instrumentation stations 1 through 7, a roll-out of both the core and fan outer and inner flowpath walls with

the circumferential and axial location of all instrumentation noted, axial sections of the rig at

instrumentation Stations 1 through 7 depicting the circumferential and spanwise distribution of

instrumentation, and recommended static pressure instrumentation installation schemes. The

system used to establish the title or "headers" assigned to the instrumentation is also explained

in Figure 81. The headers permit identification of: 1) the type of sensor, 2) the axial

instrumentation station number, 3) whether it is in the core or fan side flowpath, 4) its

circumferential or OD or ID wall position and 5) its radial location relative to the OD wall.

The radial traverse units shown at station 1, the EGV LE instrumentation at Station 5, the

circumferential traverse units and the core Pt/Tt radial probes at Station 6, and the fan/core

Pt/Tt radial probes at Station 7 are shown in phantom in Figure 81 because they are to be

provided by NASA. P&WA drawings detail the machining operations required to incorporate

NASA specified instrumentation but the design of probes and instrumentation installation

drawings were not contract requirements. Consequently, drawings of plugs are provided for case

holes where NASA designed probes are to be used. No plug drawings are provided for the holes

machined for existing NASA radial and circumferential traverse systems at Station 1 and 6.

The fan rig inlet instrumentation Station 1 is at axial location 89.530 (in.). It is located

7.315cm (2.88 in.) aft of the rig OD case/plenum (CF847546) interface to provide the 3.810cm

(1.500 in.) of upstream clearance required for existing NASA traverse mechanisms. Provision

was made for the installation of eight equally spaced OD and ID wall static taps and 4 LC Smith

radial traverse units (CD853946). The traverse units are approximately equally spaced

circumferentially and mid-gap to the leading edge of the IGV's. The eight OD and ID wall taps

are distributed between the traverse units and mid-gap to the IGV's.

The IGV inlet instrumentation Station 2 is located 0.610cm (0.240 inches) forward of the

strut LE and due to the 8 deg rearward cant of the IGV spindle centerline is not a radial plane.

The 8 OD, 16 splitter and 8 ID wall static taps are at axial stations 91.681, 92.850, and

93.242 (in.), respectively. Four of the eight static pressure taps on each flowpath surface are

approximately 90 deg apart and midgap between the IGV leading edges; the remaining four

static taps are similarly distributed circumferentially at 10%, 30%, 70% and 90% gap to

provide a cross-gap static pressure distribution. To facilitate installation, the 0.051 cm

(0.020 in.) diameter static pressure taps in the OD case at station 2 have been drilled at rearward

sloping angle to the flowpath of 60 deg. No measurement error is incurred with a 90 deg to 60

deg. installation angle unless the flow direction is reversed, causing a maximum +0.3 % q error

in Ps.

113
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134

Page intentionally left blank

Page intentionally left blank
The IGV exit/rotor inlet instrumentation Station 3 is located approximately midway

between the IGV fan and core flap trailing edges and the rotor leading edge and is therefore not

a radial plane. The 8 OD, 16 splitter and 8 ID wall static taps are at axial Stations 97.136, 96.560

and 96.361 (in,), respectively; and distributed similarly to those at the IGV LE. This placement

scheme is used for static pressure taps at all airfoil leading and trailing edge instrumentation

stations. The eight splitter ID and eight splitter OD wall static taps at both the IGV LE (Station

2) and the IGV ТЕ (Station 3) are routed along the splitter OD wall in pairs, through the splitter

in groups of four and along the pressure surface of the IGV core strut to the rig ID. The IGV and

splitter channels into which the hypo tubing for 32 static pressure taps must be buried have

been made to permit routing the 32 splitter taps and the 24 core ID wall taps from Stations 1, 2,

and 3 into the plenum through flange notches and holes provided in the forward ID flowpath

duct and the slip ring tube support assembly. NASA is responsible for the design and

fabrication of any umbilical system to route these lines out of the plenum and for grinding

routing notches into strut assembly CR846186 as shown in L237442!sheet 4.

Instrumentation Station 4 consists of 15 static pressure taps distributed axially over the

blade tip from 20% rotor axial chord upstream to 20% downstream. These taps are in two rows

at the blade tip section chord angle and spaced one blade tip gap apart as shown in section J!J

and detail X of Figure 81. The percent axial chord location of each tap is designated by the last

two spaces of the header, i.e., !2 denotes 20% upstream from the rotor LE, 02 denotes 20%

downstream from the LE and 12 denotes 120% downstream from the LE.

The rotor exit/EGV inlet instrumentation Station 5 is located 0.635 cm (0.250 in.)

upstream of the EGV LE at axial station 100.476 (in.). Eight static pressure taps are provided on

the fan OD, fan ID, core OD and core ID walls and are distributed relative to the EGV LE as

previously described. The fan ID and core OD or splitter wall taps must be routed through

channels machined into the fan side of the splitter and the EGV suction surface to the rig ID,

then aft and through the four cross routing instrumentation tubes. Once instrumented, the

EGV, OD and ID cases and the discharge splitter can remain assembled as a unit during rig

assembly and teardown. Also routed through the four cross!over tubes is the core LE Pt (vanes

#7, 16, 25 & 34) and LE Tt (vanes No 9, 18, 27 and 36) instrumentation (T4060685); i.e., three

spans of sensors are provided on 4 pairs of equally spaced vanes for a total of 12 Pt and 12 Tt

sensors. The Pt/Tt sensors are radially positioned at centers of equal areas at 15.7, 48.9, 64.3,

and 87.2% span for a total of 20 Pts and 20 Tts. This LE instrumentation is routed out of the

flowpath through channels in the flap pressure surface and spindle. A drawing (HKJ3400)

detailing machining, kiel head design and Pt/Tt installation has been provided.

The EGV exit instrumentation Station 6 is at axial location 104.630 (in.) or approximately

one chord length downstream of the EGV ТЕ. It was originally requested that this

instrumentation station be within a half chord length of the EGV ТЕ but to do this required

mounting the EGV LE flap sync ring forward of the spindle, thus eliminating the on!stand

removable rotor tip shroud design feature. The four fan duct circumferential/radial traverse

units (CF852670) located at axial station 104.820 (in.) are uniformly distributed circumferential!

ly and centered at 0% gap relative to the EGV ТЕ for an axial discharge flow condition. The

27.3 deg arc travel of the traverse units will permit the mapping of approximately 2.9 vane gaps

in each quadrant of the fan discharge. Eight fan OD static pressure taps located between the

traverse units provide the typical gapwise distribution. The eight fan ID wall taps are similarly

distributed for an axial discharge flow condition but more uniformly spaced circumferentially.

The eight core OD wall taps are located circumferentially near the fan ID splitter taps so they

can be routed in pairs along the splitter and the EGV core side suction surface to the rig ID.

This scheme minimizes the instrumentation machining required. The eight core ID wall taps at

Station 6 are circumferentially located opposite the OD wall taps, providing the typical gapwise

distribution for the 20 deg discharge flow condition. Eight equally spaced 3.175 cm (1.250 in.)

dia. holes in the core ID wall have been provided for the installation of 4 Pt and 4 Tt radial

137
probes. The probe hole centerline is axially 1.397 cm (0.550 in.) aft of instrumentation Station 6

so that the three radial sensors located at centers of equal area as listed in Table 13 will be on

approximately the same discharge plane as the wall statics. The eight holes are circumferentially

located to place the sensors at midgap to the EGV ТЕ for a 20 deg discharge flow condition.

Hole plugs (HKJ3375) which can be used in the fabrication of the probes have been detailed by

P&WA. NASA is responsible for the probe design and fabrication. If the eight probes are 0.318

cm (0.125 in.) wide and 1.778 cm (0.700 in.) long they will produce approximately 2.4% blockage

in the core discharge.

The four tubes at axial location 108.00 (in.) are for routing instrumentation from the rig ID

to the rig OD. Each is designed to accept 25 0.165 cm (0.065 in.) dia. instrumentation lines. If the

instrumentation scheme of Figure 81 is followed, each tube will carry 18 static pressures and 6

TC's. These 0.686 cm (0.270 in.) wide tubes detailed by P&WA (HKJ3374), contribute 3.14 and

2.12% blockage to the core and fan duct flow area when aligned at the flow discharge angle

shown.

The rig discharge instrumentation Station 7 is at axial location 112.339 (in.); approximate!

ly 6.350 cm (2.500 in.) forward of the rig/housing support interface (HKJ3371) as requested by

NASA. Four equally spaced sets of holes through the OD case and the splitter have been

provided for the installation of two Pt and two Tt radial probes. Plugs detailed by P&WA for

these holes (HKJ3377 and HKJ3376) may be incorporated into the NASA probe design. The

probe configuration shown in phantom in Figure 81 and listed in Table 13 provides five spans of

sensors for the fan duct and three spans of sensors for the core duct. The 16 Pt and 16 Tt sensors

are located at the center of equal areas and the core side probe shaft was aligned to account for

flow discharge angle change with discharge area increase. Four 0.51 cm (0.200 in.) wide probes

will produce 2.5% and 1.6% blockage of the core and fan duct discharge areas, respectively.

Four existing core ID wall static taps located circumferentially at 57, 147, 237, and 327 deg, and

axially at location 112.327 (in.) have been used. Four core OD and fan ID static pressure taps at

the same circumferential locations have likewise been incorporated to take advantage of the

routing scheme used by NASA. The use of 8 static taps on the splitter walls as originally

requested would have exceeded the capacity of the existing routing scheme and required facility

modification. Placement of the splitter static taps at axial location 113.177 (in.) or 2.129 cm

(0.838 in.) aft of the radial probe centerline was necessary to permit line drilling for the station 7

radial probes and disassembly of the EGV and discharge ducts as a subassembly without

disconnecting discharge instrumentation as explained in the rig design considerations of Section

V.

No provision was made for the installation of rotor strain gages. Recommended locations

are presented in Section IV. A drawing detailing the spindle hole required for routing strain

gages to the suction surface of the EGV LE fan flap (HKJ3397) is provided. No strain gage

installation was considered necessary for the integrally machined IGV and EGV case struts or

the IGV core and fan flaps.

138
SECTION VIII

FABRICATION

The purpose of this section of the design report is to provide the information required to

fabricate the NASA LeRC V/STOL fan rig since the P&WA fabrication effort (Phase III) was

terminated on 7 January 1983.

SPECIFICATIONS

The reference information required to support the fabrication and assembly of the rig, i.e.,

standard parts, materials specifications, manufacturing processes and drawing interpretations

was provided in a separate volume (Reference 5) delivered to NASA LeRC on 3 May 1983. The

information available in this volume is listed in Tables 14 through 17.

Part Number

AN 123629

AS 3066-12

AS 3068-10

AS 3168

HKJ 1655

HKJ 1656

HKJ 2515

HKJ 3413

MS 9111-20

MS 9226-05

MS 9276-10,-12

MS 9361-10.-12

MS 9386-023

MS 9390-430

MS 9565-08,

MS 9566-08,-ll,-12,-14

MS 9568-06

MS 9724

MS 9942-26.-32

MS 18064-13

MS 122030

MS 35338

MS 51023

MS 51038

MS 122032

MS 645231

MS 645232

MS 641619

MS 761787

MS 4022587

MS 4027818

TABLE 14.

Layout

L237442

L237441

L237442

54, 55

52, 53

57, 86, 46, 58

56, 60

L237441

L237441

L237441

L237442

— PARTS

Item Remarks

31

87

68

70 Sub for MS 9111-20

3 IGV Raw Matl

11 EGV Raw Matl

Rotor Forging

40 Heim Rod End Assy

90

42 Safety Wire

78

18

69

88

90 Sub for MS 9111-20

89

70

56 Sub for MS 122030 & 32

89 Sub for MS 18064-13

51

56

3

5

93 Sub for MS 9390-430

4

43

32

139
TABLE 15. — MATERIALS

AMS No. PWA No. AISI No. Other

3651

4117

4973

5060

5069

5070

1202 8.1.1.

Teflon

6061-T6 Al

SAE 1015

SAE 1018

SAE 1022

5504

5510

5599

5613

5616

5646

5663

5665

5666

410

347

5731

5735

5737

6322

7232

7240

7267

TABLE 16. — PROCESSES

Spec No. Title

AMS 2470 Anodic Treatment of Aluminum Alloys

AMS 2485 Black Oxide Treatment

PWA 6 Peening

PWA 11-15.-22 Heat Treatments

PWA 16-1.-2 Welding

PWA 19 Gold Nickel Alloy Braze

PWA 85 Silver Metal Braze

PWA 99 Surface Finishing of Holes

PWA 106 Grinding of Titanium and Titanium Alloys

PWA 255 Plasma Spraying (METCO 601)

PWA 310 Identification Marking

PWA 357 Rivets

PWA 362-1 Surface Texture

PWA 706 Brazing Filler Material

140
TABLE 17. — DRAWING INTERPRETATIONS

PWA Spec Title

DRM 14.13.1 Dimensions for Installation of Dowel Pins

DRM 15.1.1 Balancing of Rotating Parts

PWA 330 Drawing Interpretation for Airfoils

PWA 360 Drawing Interpretations

PWA 390 Computer File Airfoil Section

Coordinate Data and Fairing Procedures

ESTIMATED COST

An estimated cost or hours to fabricate column is provided in Tables 18 through 25. This

estimate is based on: vendor quotes for the IGV and EGV and the rotor bladed disk, vendor

costs for similar components escalated for inflation and factored for complexity, and estimates

of P&WA shop hours required to fabricate those components for which no comparable part cost

history is available: No P&WA labor or overhead rates were applied and no factors for quality

control, tooling, etc, were included. Separate cost-to-fabricate estimates are provided for the

slipring/inlet cases, IGV sub assembly, rotor and split case, EGV sub assembly, discharge/diffuser splitter and cases and the actuation/position indicating components in Tables 18

through 25. These tables include part name and number of units required and recommended

units purchased, material and vendor quote (if applicable) and may be used as a format for

recording fabrication costs.

C. PROCURED HARDWARE

Hardware procured in support of the fabrication effort prior to the 7 January 1983

termination notice and delivered to NASA LeRC on 2 February 1983, is listed in Table 26. Rotor

balance weights (Part Numbers 645231, 645232 and 761787) were to be provided from P&WA

stores but termination prevented their purchase and delivery.

The AMS4973 (PWA1202/8.1.1 titanium) rotor forging composition and mechanical

properties as determined by the Ladish Company were submitted to NASA LeRC on 5 April

1983 for review. Composition was as required with the exception of aluminum which exceeded

the 8.35% maximum by 0.17%. Mechanical properties were within specification and are

compared to the required properties in Table 27. The AMS5616 rings for the IGV and EGV

strut cases were supplied by L&B Engineering. L&B will provide lab test results if requested;

the rings were accepted by the P&WA materials control lab as adequate for experimental but

not flight use.

The following were P&WA considerations for designing specific rig components and

planning for their fabrication. Those vendors listed for particular machining operations are

P&WA source approved and have been found to be reliable and cost competitive.

141
RECOMMENDED VENDORS

IGV and EGV Strut Cases and Rotor Blisk

The IGV and EGV strut cases and the rotor blisk were designed to be integrally machined

for cost savings. The one-piece construction is less costly than individual components when

purchasing only one unit because there are fewer machined surfaces, less inspection time, no

welding/assembly tooling required and no assembly. In addition, dimensional control is

improved and the flowpath is continuous. Risk however, increases and care must be taken

during all phases of fabrication and assembly to prevent machining errors and damage due to

handling. No extra rotor forgings or IGV/EGV steel rings were purchased for scrap allowance

due to their high per unit cost — $23K and $6K, respectively. P&WA source approved vendors

capable of manufacturing these components include:

Atlas Tool Inc. Centura, Inc.

29880 Goresback Ave. 12411 Old Highway 99 N

Roseville, Michigan 48066 P.O. Box 769

(313)778-3570 (314)232-8858

Atlas tools uses numerically controlled machines that require airfoil coordinate tapes

provided by P&WA. Centura uses a pantograph/hand machining operation and tooling

simulating all airfoil passage contours.

IGV Core Flap and IGV/EGV Fan Flaps

These airfoils are conventional in design and require no unusual machining operations

with the exception of the 22.86 cm (9.000 in.) long 0.65cm (0.256 in.) diameter hole through the

IGV fan flap and spindle for the IGV core flap spindle. This configuration has not been required

for any P&WA airfoils. However, if a hole straight within the required tolerances is drilled first

and the airfoil machined relative to it, no fabrication problems should arise. Vendors with

P&WA source approval and capable of manufacturing conventional airfoils include:

Jarco Machine Products Jarvis Airfoils

P.O. Box 10 10 Glastonbury Turnpike

4373 Lilburn Industrial Way Portland, Conn. 06480

Lilburn, Georgia 30247 (203)342-3800

(404)921-5902

Mai Tool & Engineering New England Aircraft Products

291 Adams St. P.O. Box 83

Manchester, Conn. 06040 Spring Lane

(203)643-2473 Farmington, Conn. 06032

(203)677-1377

142
TRW Inc.

Compressor Components Division

P.O. Box 400

Route 11 Woodbine Lane

Danville, PA 17821

(717)271!7509

Instrumentation

The instrumentation channels required in the IGV and EGV strut cases, the IGV LE and

ТЕ splitter, the EGV ТЕ splitter and the EGV fan flap can be made by electrical discharge

machining (EDM). Qualified vendors include:

Electro Form Machining EDM Precision Products

P.O. Box 8927 5650 N.W. 9th Ave.

12095 N.W, 39th St. Ft. Lauderdale, Fla. 33309

Coral Springs, Fla. 33065 (305)771!1555

(305)755!2700

Installation of the EGV LE Pt/Tt instrumentation and fabrication of the EGV core

discharge and diffuser exit Pt/Tt radial probes can be done by:

H&B Tool & Engineers Projects Inc.

P.O. Box 717 P.O. Box 174

481 Sullivan Ave. 38 Addison Road

South Windsor, Conn. 06074 Glastonbury, Conn. 06033

(203)528!9341 (203)633!4615

Rotor Tip Shroud

The Metco 601 (PWA 255) rotor tip shroud material has been successfully installed on

P&WA rigs by:

Quantum Inc. General Plasma Assoc.

Barns Industrial Park 12 Thompson Road

Wallingford, Conn. 06492 East Windsor, Conn. 06002

(203)265!2811 (203)623!9901

Other

The remaining rig components require no special expertise to fabricate other than the

ability to meet tight concentricity, flatness, and dimensional tolerances. Qualified vendors

include:

Dexter Aero Space Caval Tool & Machine Co. Inc.

34260 James J. Pompo Dr. P.O. Box 11158

Eraser, Michigan 48020 275 Richard St.

(313)296!2611 Newington, Conn. 06111

(203)666!8414

143
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153
Page intentionally left blank

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SECTION IX

ASSEMBLY

The general assembly sequence for the IGV and the EGV/discharge ducts and installation

of the overall V/STOL model fan stage rig into the Lewis Research Centers W!8 test cell as

shown in Figure 82 are presented in this section. The rig was designed to require no special

tooling, and the assembly sequence presented herein is to explain the engineering philosophy of

the rig design. The IGV and EGV flaps were dimensionally toleranced to nominal and therefore

require hand fitting at assembly. This approach was taken to minimize endwall gaps.

INLET GUIDE VANE (IGV)

The IGV assembly includes installation of the fan and core ТЕ flaps, the aft OD and ID

cases, the splitter leading and trailing edges and the station 2 and 3 static pressure

instrumentation. The components required for the IGV assembly should be machined for

instrumentation but whether or not the IGV strut case is partially instrumented prior to

initiating the component fitting procedure presented below is at the discretion of the assembler.

It is important to remember that once the ТЕ splitter is riveted into place, the core flap cannot

be removed or installed.

1. Serialize the IGV core flaps (T4060682) 1 through 17 and install in the

IGV case splitter holes at the corresponding vane locations.

2. Install the aft ID shroud (T4060678) with 3 bolts (MS9942!26). Note the

circumferential location of the .instrumentation and metal stamp "Top

<£ " to facilitate reassembly.

3. The core flap should rotate ± 20 degrees freely and have some radial

movement — 0.010 to 0.023 cm (0.004 to 0.009 in.).

4. Load each flap toward the ID and rotate through the range of travel. There

should be 0.005 cm (0.002 in.) clearance between the vane end and the ID

wall at the closest approach. If necessary, remove vane end material until

0.005 cm (0.002 in.) clearance is achieved to insure a no-rub assembled

condition.

5. Align instrumentation channels and temporarily install the ТЕ splitter

(T4060679). Do not rivet.

6. Repeat step 4 loading the core flap toward the OD. This completes the core

flap fitting.

7. Serialize the IGV fan flaps (T4060681) 1 through 17 and install the

corresponding serial number steel bushing (HKJ3391) with the Teflon

thrust bearings (HKJ33480). If rotation of the spindle within the bearings

is stiff, enlarge the bearing ID.

8. Install the corresponding serial number sync arm (HKJ3395), tablock

washer (MS9276!12) and spindle nut (MS9361!12). Torque the nut to spec

but do not bend the tablock. Measure and record the thrust bearing to

sync arm gap (radial movement). If there is no radial movement and

155
rotation is again stiff, the required alteration will be explained in steps lie

and lid.

9. Slide the fan flap sub assemblies over the core flap spindles and bolt

(MS9566-08) the steel bushings (HKJ3391) to the IGV case (T4060676).

Torque to specification.

10. Install the aft case (T4060677) with 3 bolts (MS9942-32). Metal stamp

"Top ^ " to facilitate reassembly.

11. Check the fan flaps for the required rotation (+20 deg to —60 deg, -20 deg

is nominal, i.e., the sync arm is axial). Flap ends should clear the flowpath

walls by 0.005 cm (0.002 in.) at the closest approach when radially loaded

in the direction of the gap being measured.

(a) If the spindle/bushing subassembly has radial movement

0.003 to 0.015 cm (0.001 to 0.006 in.) as measured in step 8,

remove material from the OD flaps end until the required

clearance is achieved during rotation.

(b) Repeat step lla loading the flap toward the ID.

(c) If the spindle/bushing subassembly had no radial movement and touches the ID wall, remove material from the

surface of the Teflon thrust bearing nearest the flap

platform until some radial movement is possible (0.003 to

0.015 cm/0.001 to 0.006 in.). Reassemble. If the flap end

still contacts the ID wall when loaded radially inward,

remove material from the flap end until the required 0.005

cm (0.002 in.) clearance is achieved. Then load the flap

toward the OD and remove material from the flap end

until clearance is achieved.

(d) If the spindle/bushing subassembly had no radial movement and touches the OD wall, proceed as in step lie but

remove material from the surface of the Teflon thrust

bearing nearest the sync arm.

(e) If the spindle/bushing subassembly has radial movement

and flap end touches both the OD and ID walls, remove

material from both ends until the required clearance is

achieved.

(f) If the spindle/bushing subassembly has no radial movement and touches both the OD and ID walls, remove

material from the end of the Teflon thrust bearing nearest

the sync arm until radial movement is achieved and then

proceed as in step lie.

12. When the core and fan flap ends clear the walls and rotate freely, remove

the nut, tablock and sync arm from the fan flap spindle.

13. Press the Teflon sliders (HKJ3384) into the sync ring (HKJ3387) and trial

fit the sync ring assembly onto the IGV case (T4060676). If rotation is stiff,

156
reduce the slider thickness as required. Maintain concentricity between

the case and the sync ring (0.114 cm/0.045 in.) is the nominal gap.

14. Install sync arm pins (HKJ3388) and reinstall the fan sync arms, tablocks,

and nuts. Torque to specification. Do not bend the tablock. Check for sync

ring freedom of travel through the +20 deg to !60 deg rotation.

15. Repeat steps 13 and 14 for IGV core flap with items HKJ3384, HKJ3389,

HKJ3385, HKJ3394, MS9361!10 and MS9276!10. Sync arm rotation is

±20 degrees from axial. When smooth continuous rotation is possible, the

core flap fitting procedure is complete.

16. Install the remaining ID and OD aft case bolts (MS9942!26 and !32) and

torque to specification. Recheck clearances and travels.

17. If the station 2 and 3 static pressure hypo tubing is partially installed, the

ТЕ splitter (T4060679) could be permanently riveted (AN123629) into

place at this time per HKJ3415. However, if core flap disassembly is

necessary for any reason prior to test, do not rivet the ТЕ splitter.

18. The LE splitter (T4060680) could also be riveted (AN123629) at this time

(HKJ3415). However, its permanent addition further limits access to the

core strut where hypo tubing may have to be installed.

EXIT GUIDE VANE (EGV)/DISCHARGE DUCT ASSEMBLY

The EGV/discharge duct assembly includes temporary installation of the ТЕ splitter, OD

and ID cases for line drilling, fitting of the LE fan flap, disassembly for instrumentation

installation and reassembly. All instrumentation machining but no sensor installation should be

completed prior to initiating this assembly sequence.

1. Temporarily install the ТЕ splitter (HKJ3365) to the EGV case

(T4060684). Use rivets (AN123629) to align the close tolerance holes but

do not flare the rivets.

2. Install aft OD and ID cases (HKJ3366 and HKJ3367). Check axial stack

up of components per HKJ3382. Machine for dowel pins (641619) and

insert. Line drill for cross routing instrumentation tubes and station 7

discharge probes per HKJ3382.

3. Remove aft OD and ID cases and the ТЕ splitter and initiate LE fan flap

fitting procedure.

4. Serialize the EGV flaps (T4060686, HKJ3400, and HKJ3397) with strut

location as required per L238020 sheet 2 (Figure 82).

5. Install the corresponding serial number steel bushings (HKJ3393) with the

teflon thrust bearings (HKJ3392 and HKJ3383). If rotation of the spindle

within the Teflon bearings is stiff, enlarge the bearing ID.

6. Install the corresponding numbered sync arms (HKJ3396). Install tablock

(MS9276!12) and nut (MS9361!12). Torque to spec but do not bend

tablock. Measure and record the thrust bearing to sync arm gap (radial

157
movement). If there is no radial movement and rotation is now stiff, the

required alteration will be explained in step 8.

7. Place the EGV flap subassemblies into their respective locations in the

EGV core and install the forward OD case (T4060685) with 3 equally

spaced bolts (MS9566!08). Torque to specification.

8. Check the LE flap for ± 10 degrees of rotation. The flap ends should clear

the flowpath walls by 0.010cm (0.004 in.) at the closest approach; if not,

follow the fitting procedure presented for the IGV in steps lla through

llf.

9. When all flap ends clear the flowpath walls through maximum rotation,

remove the nut (MS9361!12) and tablock (MS9276!12) from the flap

spindle.

10. Press the Teflon sliders (HKJ3384) into the sync ring (HKJ3386) and trial

fit the assembly to the EGV case (T4060684). If rotation is stiff, reduce the

slider thickness as required. Maintain concentricity between the case and

the sync ring (0.114 cm/0.045 in. is the nominal gap).

11. Install the sync arm pins (HKJ3389) and reinstall the sync arms, tablock

washers and nuts. Torque the nut to specification but do not bend the

tablock. Recheck for sync ring freedom of travel. Install the remaining OD

case bolts (MS9566!08) and torque to specification. Recheck clearance and

travel.

12. This completes the flap fitting sequence. Disassemble the EGV. Install the

LE fan flap and core Pt and Tt sensors. Install the station 5 wall static taps

and route the Station 6 hypo tubing to the case/splitter interface.

13. Install the ТЕ splitter (HKJ3365) with rivets (AN123629). Complete the

routing and installation of the Station 6 wall static taps.

14. Reinstall the LE fan flaps and reassemble through installation of the sync

ring. Do not bend tablock.

15. Install aft OD and ID cases (HKJ3366 and HKJ3367) with bolts (MS9566!

08), align with dowel pins, torque to specification.

16. Install hole plugs (HKJ3375) or NASA provided core discharge probes and

the cross!routing instrumentation tubes (HKJ3374) with "0" rings

(MS9386!023) and route the ID instrumentation lines through to the case

OD. This completes the IGV/discharge duct assembly.

158
RIG ASSEMBLY/FACILITY INSTALLATION

Rig assembly and installation into the NASA LeRC W!8 test facility is outlined below. The

axial locations of component interfaces that must be inspected during assembly to insure

adequate rotor clearances are listed in Table 28 and shown in Figures 80 and 83.

1. Move the plenum to the maximum forward position to provide approxi!

mately 51cm (20 inches) assembly clearance between face 0 and face F of

Figure 83.

2. Install the rear housing support (HKJ3371), the OD diffuser (HKJ3372)

and the diffuser splitter (HKJ3370). Verify that the axial distance from

face 0 to face W is within tolerance. If it is not, measure EGV/discharge

assembly dimension from face R to face W and modify as required to

achieve the required face R to face 0 assembled dimension.

3. Install the EGV/discharge assembly over the bearing compartment/shaft

support and bolt into place.

4. Measure the axial distance from face О to face Q and from face 0 to face R.

If these dimensions are not within tolerance, running clearance between

the rotor and the EGV ID and splitter and between the rotor and the IGV

ID and splitter will not be correct and a rub could occur. Modify as

required to achieve the required dimensions. This completes the installa!

tion of the EGV/discharge duct assembly and the diffuser ducts.

5. Modify the plenum strut assembly to accept instrumentation lines

(L237442 sheet 4).

6. Install spacer CF851588 (0.838 cm/0.330 in. thickness) and the ID forward

case (HKJ3378). Measure the axial distance from face G to face F

(CF847546) as defined in Figure 83. If not as required, do not modify

spacer CF851588 until step 7.

7. Install the OD forward case (HKJ3373) and measure the axial distance

from face G to face H (Figure 83). If this distance is not 2.068 ± 0.005cm

(0.814 ± 0.002 in.), remove spacer CF851588 and modify as required.

8. Install the IGV assembly with bolts (MS9566!08 both OD and ID bolt

circles). Measure the axial distance from face L to face J. If this dimension

is not within tolerance, running clearance between the rotor and the IGV

ID and splitter will not be correct and a rub could occur. Modify as

required to achieve the required dimensions.

9. Route ID instrumentation forward into the plenum through the holes and

notches provided. Secure as necessary to prevent damage during installa!

tion of the slip ring support tube.

10. Install the balanced rotor assembly (4060073 with balance weights 645232,

761787 and 645231 as required), slip ring spacer (HKJ3398) and coupling

(CC848383) with tiebolts provided by NASA. Measure the axial gap

between the rotor and the EGV ID and splitter as shown in Figure 80.

These distances must be as required or a rub condition will exist during

rotation.

159
11. Mount the slip ring within the support tube (HKJ3379) with set screws

(MS1038!103). Axial placement of the slip ring within the tube is to be

determined by NASA. Insert this assembly through the center of the

plenum. Do not install the spacer (HKJ3381) or the tube support halves

(HKJ3380!1, !2) at this time.

12. Slide the slip ring/tube support assembly forward and attach to the rotor.

If necessary, bring the plenum aft to permit couplings.

13. Install 1/2 of the rotor case assembly (HKJ3369 installed within HKJ3368!

1 with bolts MS9565!08) to the EGV case (T4060685) with 6 bolts

(MS9566!08). Torque to specification.

CAUTION

Care must be taken during this operation or the

slip ring/coupling can be damaged. (Care must

also be taken during disassembly.)

14. Slowly move the plenum aft against the rotor case half installed in step 13.

As the plenum is moved aft, the slip ring support tube must be pulled into

the plenum.

15. Attach the rotor shroud to the IGV case with 6 bolts (MS9566!08). Torque

to specification. Measure the axial gap between the rotor and the IGV ID

and splitter as shown in Figure 80. These distances must be as required or

a rub condition will exist during rotation.

16. Measure the blade tip to tip shroud gap as shown in Figure 80. This

distance must be 0.96cm (0.038 in.) to maintain 0.051 cm (0.020 in.)

running clearance at 17762 rpm.

17. If all clearances are as required, loosen the rotor shroud to EGV and IGV

case bolts installed in step 13 and 15 and install the remaining 1/2 of the

rotor case assembly (HKJ3369 installed within HKJ3368!2). Install all

split line bolts (MS9566!2) and nuts (AS3068!10) and torque to specifica!

tion. Install and torque to specification all EGV and IGV to rotor case

bolts. This completes the basic IGV!rotor case!EGV/discharge assembly.

18. Measure the axial distance between face С and face A. Grind the spacer

(HKJ3381) to achieve this dimensional stack!up when assembled with the

tube support halves (HKJ3380!1, !2).

19. Install spacer (HKJ3381) and the tube support halves (HKJ3380!1, !2) as

shown in Figure 82 to fixture the slip ring support tube (HKJ3379) in

place. Route instrumentation lines through the notches provided. Torque

all bolts (MS9566!14 and AS3068!10) to specification.

20. The instrumentation (L238020 sheet 1), the actuation systems (L237442

sheet 2), and the vane angle potentiometers (L237442 sheet 3) are now

installed. This completes the V/STOL fan stage rig assembly.

160

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TABLE 28. — AXIAL LOCATION OF V/STOL FAN STAGE RIG

COMPONENT INTERFACES

Station/

Face Interface Description

Nominal

Axial Location

(Inches)

A Forward End!Slip Ring Tube Support

В Tube Support Aft End/Spacer (2.540cm/1.000 inch)

С Spacer/Plenum Strut Assembly

D Plenum Strut Assy/Spacer CF51588 (0.838cm/0.330 inch)

E NASA Spacer/Inlet ID Case Forward End

F Facility/Inlet OD Case

G Inlet ID CaseAGV Forward ID Case

H Inlet OD Case/IGV Forward OD Case

I Slip Ring !Support Tube Aft End

J IGV Aft OD Case/Rotor Case

К Slip Ring/Slip Ring Spacer (0.991cm/0.390 inch)

L IGV Aft ID Face

M Rotor Forward ID Flowpath Face

N Rotor/Slit Ring Spacer

0 Rotor/Shaft Interface

P Rotor Aft ID Flowpath Face

Q EGV Forward ID Face

R Rotor Case/EGV Forward OD Case

S EGV Aft OD Case/OD Discharge Case

Т EGV ID Case/ID Discharge Case

U EGV Splitter/Discharge Splitter (Aft Face)

V Discharge Splitter/Diffuser Splitter

W OD Discharge Case/Aft Mount Flange

X Aft Mount Flange/OD Diffuser

Y Aft End of OD Diffuser

Z Diffuser Splitter/Housing Support

54.139

62.639

63.639

74.699

75.029

86.649

91.342

92.156

93.139

96.236

96.800

96.966

96.986

97.190

100.000

100.203

100.233

100.971

102.451

102.471

103.751

113.005

114.839

115.959

123.550

132.895

163
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APPENDIX A

AIRFOIL AERODYNAMIC DESIGN SUMMARIES

The column headings, postscripts and superscripts for the airfoil aerodynamic design

summaries for the V/STOL rig IGV, Rotor and Stator (EGV) are identified below.

SL Streamline identification number 1!11, ID to OD

V Absolute Velocity

VM Meridional Component of Absolute Velocity

VO Tangential Component of Absolute Velocity

U Wheel Speed (See Note)

RHOVM Product of Static Density and Meridional Velocity

EPSI Streamline Slope Measured from Engine Axis

PO/PO Total Pressure Ratio, ТЕ to Inlet

TO/TO Total Temperature Ratio, ТЕ to Inlet

В Air Angle Relative to Axial Direction

M Mach Number

DIA LE Leading Edge Diameter

DIA ТЕ Trailing Edge Diameter

D FAC Diffusion Factor

OMEGA!B Loss Coefficient

LOSS!P Total Pressure Loss of Dynamic Head Rise

P02/P01 Total Pressure Ratio, ТЕ to LE

% EFF A Percent Adiabatic Efficiency — Total

% EFF P Percent Polytropic Efficiency — Total

PCT ТЕ Percent Span at Trailing Edge

WC1/A1 Inlet Specific Flow

NCORR Corrected Speed

WCORR Corrected Flow

Postscripts: 1 Leading Edge

2 Trailing Edge

Superscripts: ' Relative to Rotor

Note: Wheel speed is calculated assuming rotor is present at leading and trailing

edges. U = 0 for stators.

167
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APPENDIX В

AIRFOIL GEOMETRY

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APPENDIX С

AIRFOIL SECTION COORDINATES

The airfoil section coordinate printouts are explained in Table 41 and Figure 90.

TABLE 41. — AIRFOIL SECTION COORDINATE PRINTOUT FORMAT

PWA 390 REVISION В

CARD

1 thru 9

10

11

12

13*

14*

14A**

15*

ITEM

United Technologies Corporation Document Property Rights Notice

Computer file identification: FILENAME

FILETYPE

Identifying title (Title, stage, file category, EMD num! b e r . . . )

Total number of airfoil defining sections Number o: points along each surface Airfoil orientation indicator: 1 • Type I

2 » Type II

3 • Tvpe III •4 ' Type IV

Rotation Indicator (Blank or 0 ! Std P&UA;

1 ! Counter rotating)

Indicator for type of leading edge and

trailing edge: 0 or Blank ! circular arc LE and ТЕ

1 ! Elliptical LE, circular arcТЕ

2 ! Elliptical ТЕ, circular arc LE

3 ! Elliptical LE and ellipticalТЕ

First card of each airfoil defining section: Title identifying this section (e.g., "A!A")

"A" !dimension (basic distance locating the airfoil

section from a reference base plane)

Chord length (length of line from leading edze refer! ence point to trailing edge reference point)(B dim)

Coordinate of the stacking point along the G!axis (D dim)

Coordinate the stacking point along the F!axis (E dim)

Angle of the chord line measured from Plane Y (C dim)

Second card of each airfoil defining section: Leading edge radius or ellipse semi!major axis at LE (reference only)

F dimension td center of LE circle or ellipse С dimension to center of LE circle or ellipse Trailing edze radius or ellipse serai !major axis at ТЕ (reference only)

F dimension to center of ТЕ circle or ellipse

С dimension to center of ТЕ circle or ellipse

Semi !minor axis of ellipse at leading edge Angle* orienting ellipse at leading edge (ref only)

Semi!minor axis of ellipse at trailing edge Angle* orienting ellipse at trailing edge (ref only)

Third block of data for each airfoil defining section: (Four sets of airfoil surface coordinate data)

The first point is the point of tangency between the leading edge circle or ellipse and the airfoil surface.

The points continue along the surface to the last point which is the point of tangency between the trailing edge circle or ellipse and the airfoil

surface. Values of each set are written 8 to a card in the columns and format shown at the right. The four sets of data are: Set 1: All the F!coordlnate suction values Set 2: All the C!coordinate suction values Set 3: All the F!coordinate pressure values Set 4: All the G!coordinate pressure values

* Cards 13, 14, 14A, and 15!thru!As!Req'd repeat for the number of airfoil defining sections in the file.

** Card 14A appears in file only if the Indicator for the type of leading and trailing edge is 1, 2, or 3. There will be a card 14A for each card 14.

t Decimal degrees.

DATA

FIELD!

CODE '

1!80

1!8

11!18

1!64

1!5

6!10

11!15

16!20

26!30

4!10

11!20

21!30

31!40

41!50

51!60

1!10

11!20

21!30

31!40

41!50

51!60

1!10

11!20

21!30

31!40

1!10

11!20

21!30

31!40

41!50

51!60

61!70

71!80

FORMAT

CODE

20Л4

2A4

2A4

16A4

15

15

15

15

15

A4.A3

F10.4

F10.6

F10.6

F10.6

F10.5

F10.6

Fl.0.6

F10.6

F10.6

F10.6

F10.6

F10.6

F10.5

F10.6

F10.5

F10.6

F10.6

F10.6

F10.6

F10.6

F10.6

F10!.6

F10.6

195
PWA 390 REVISION В

TYPICAL AIRFOIL SECTIONS (TOP SECTION ORIENTATION)

SUCTION SURFACE

TYPE I

PLANE Y

SUCTION SURFACE

TYPE II

PLANE Y PLANE Y

SUCTION SURFACE

SUCTION SURFACE

TYPE III TYPE TV

These views (looking radially Inward) identify typical pressure and

suction surfaces and airfoil orientation data. Direction of artpws

for F & G coordinates indicate positive (+) values

FD 269064

Figure 90. Typical Airfoil Sections

196
CA131000.N22354E

TITLE

CASE!IMLET GUIDE VANE,FRONT (I.D.) CAT 1 P/N4060676 NO REV EMD4060676 NO REV

ENGINEER'S NAME

S.SALVAGGIO

OES JOB NO

31154B

_CHEJCKEH_BJL DATE APPROVED BY !DATE

R.L.WALLS

TO NO REV L/0 SH PART NO AIRFOIL EMD

237442 1 4060676 NO REV 4060676 NO PEV

830811 R.L.SANFORD 830811

С LAB

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

CASE!INLET GUIDE VANE, FRONT I.D.

7 50 3 0

A! A * 2.7000 1.149019 0

0.012053 0.012053 !0.000000 0

0.

0.

0.

0.

0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

1. !0.

!0.

!! !!0°.! !0.

!0.

!0.

001393

039888

173478

372987

587887

803590

018537

005625

040575

080810

110331

123678

121679

106164

001393

039888

173478

372987

587887

803590

018537

005625

040575

080811

110332

123680

121682

106167

0.

0.

0.

0.

0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

1. !0. !0.

!0.

!0.

!0.

!0.

!0.

003921

049655

194612

399749

614878

830467

056714

010313

045066

085004

112906

124049

120335

102348

003921

049655

194612

399749

614878

830467

056714

010314

045067

085005

112907

124050

120337

102351

0.006679

0.059538

0.215861

0.426626

0.641870

0.857344

0.014900

0.049216

0.088871

0.115136

0.124393

0.118786

0.006679

0.059538

0.215861

0.426626

0.641870

0.857344

!0.014900

!0.049216

!0.088872

!0.115137

!0.124395

!0.118789

CAT 1 P/N4060676 NO REV

.000000 !0.051240 90.00008

.102898 1.046122 !0,000002

0.009667

0.069647

0.237110

0.453502.

0

0

0

0

0

0

0

0

0

0

0

0

0

0

!0

!0

!0

!0

!0

!0

.668862

.884221

.019246

.052938

.092480

.117242

.124570

.117083

.009667

.069647

.237110

.453502

.668862

.884221

.019246

.052939

.092481

.117243

.124572

.117086

0

0

0

0

0

0

_o

0

0

0

0

0

0

0

0

0

0

0

!0

!0

!0

!0

!0

!0

.012999 0.016676

.090091 0.110766

.258358 0.292701

,480379 0,507256.

.695853

.911097

.023408

.059759

.095876

.119014

.124498

.115255

.012999

.090091

.258358

.480379

.695853

.911097

.023408

.059760

.095877

.119015

.124500

.115258

0.722845

0.937974

0. Q27266

0.065864

0.100901

0.120461

0.124192

0.113263

0.016676

0.110766

0.292701

0.507256

0.722845

0.937974

!0.027266

!0.065864

!0.100902

!0.120463

!0.124194

!0.113266

EMD4060676

0.020813

0.131555

0.319463

Q.,534133.

0.749837

0.964851

_Q,.030584_ 0.071334

0.104371

0.121742

0.123632

0.111088

0.020813

0.131555

0.319463

0.534133

0.749837

0 . 964851

!0.030584

!0.071334

!0.104372

!0.121744

!0.123634

!0.111091

NO REV

0.030236

0.152459

0.346225

0,561010

0.776713

0.991728

0.035814

0.076299

0.107497

0.122867

0.122783

0.108696

0.030236

0.152459

0.346225

0.561010

0.776713

0.991728

!0.035815

!0.076300

!0.107498

!0.122869

!0.122786

!0.108699

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060676, Sheet 1.

197
В!В

0.019485

0.003085

_ SU45.283

0.181221

0.383198

0.600720

г. 8600 1.168839 !0.000000 !0.010770 89.99991

0.019485 0.000000 0.105282 1.063558 0.000001

0.006124 0.009397 0.012904 0.016761 0.021086 0.

0.. 055335 _ 0^065505 _ 0.07579Д —О!096596_Я Л17519 __ 0

0.202611

0.410315

0.627954

0.224001

0.437432

0.655188

0.245507

0.464549

0.682422

0.267014

0.491783

0.709657

0.301963 0

0.519018 0

0.736891 0

5_161859.3__0.е45в2.7 _ 0.. 1.036066 1.074010

0.010521 0.015082 0.019276 0.023228 0.026967 0.030214 0

0.0't2629 0.046876 0.050800 0.054359 0.060962 0.066831 0.

0.081398 0.085480 _ (Ы>8Э248 _ Q.JJ92799 _ D_._Q96121 _ ОЛ01095_П

0.110419 0.112965 0.115211 0.117274 0.119071 0.120590 0

0.123330 0.124393 0.124770 0.124982 0.125024 0.124824 0

0.122732 0.121533 0.120116 0.118532 0.116801 0.114960 0

0.10 3367 0.10476 2

0.0030S5 0.006124 0.009397 0.012904 0.016761 0.021086 0

0.0'.5283 0.055335 0.065505 0.075791 0.096596 0.117519 0

0.181221 0.202611 0.224001

025762 0.035464

138675 _ 0.159948!

328963 0.356080

546252 0.573486

764125 0.791359

08957_

033011 0.038008

072135 0.077010

04509 _ ОЛ 07603.

121920 0.123054

124380 0.123686

112957 0.110750

0.012904 0.016761 0.021086

0.075791 0.096596 0.117519

0.245507 0.267014 0.301963 0.

_ 0.383193 0.410315 _ 0_._437432 _ О..464549_0,4_917е3._0...51901в _ Q..

0.600720 0.627954 0.655188 0.682422 0.709657 0.736891 0

0.818593 0.845827 0.873061 0.900295 0.927530 0.954647 0

1.036066 1.074011

. !0.010521 !0.015082 !0.019276!0.023228^!0.026967 !0.030214!0.

!0.042628 !0.046876 !0.050800 !0.054359 !0.060962 !6.066830!0

!0.081398 !0.085479 !0.089247 !0.092798 !0.096121 !0.101094 !0.

!0.110418 !0.112964 !0.115210 !0.117273 !0.119070 !0.120588 !0

!0.123878 !0.124392 !0.124769 !0. 124980 _jrO. 125023 !0.124823 !0.

025762 0.035464

138675 0.159948

328963 0.356080

546252 _ 0^.573486_

764125 0.791359

981764 1.008957

033011 !0.036008

072134 !0.077010

104508 !0.107602

121918 !0.123053

124379!0.123684. ! !0.122730

!0.108365

C!C

0.019287

0.

0.

0.

0.

0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.

"" ~~ oT

0.

0.

0.

1. !0.

!0.

!0.

! !0.

!0.

!0.

!0.

003518

047956

187745

39'i835

617453

840923

C63779

011106

042205

079821

103544

122879

124097

112860

CC3518

047956

187745

617458

840923

C63779

011106

042206

079822

122831

124100

112864

!0.121531

!0.104759

3.2350

0.019287

0.006770

0.058313

0.209658

0.422648

0.645391

0.868356

1.102787

0.015547

0.046287

0.083824

0.111131

0.123774

0.123325

0.109902

0.006770

0.058313

0.209658

0.422648

0.645391

0.868856

1.102786

!0.015547

!0.046287

!0.083824

!0.111132

!0.123776

!0.123328

!0.109906

!0.120113

1.204175

!0.00000

0.010262

0.068789

0.231692

0.450461

0.673324

0.896789

0.019564

0.050112

0.087563

0.113494

0.124333

0.122321

0.010262

0 . 068788

0.231692

0.450461

0.673324

0.896789

!0.019564

!0.050112

!0.087564

!0.113496

!0.124335

!0.122324

!0

!0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

!0

!0

!0

!0

!0

!0

.118530

.000000

.110263

.014116

.079384

.253725

.478274

.701257

.924602

.023527

.053546

.090998

.115559

.124662

.121121

.014116

.079384

.253725

.478274

.701257

.924601

.023527

.053546

.090998

.115561

.124664

.121123

!0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

!0.

!0.

!0.

!0.

!0.

!0.

116799

075280

093912

018331

100815

275759

506086.

729190

952414

027204

059921

094260

117479

124926

119742

018331

100815

275758

5C6086

729190

952414

027204

059921

094261

117481

124929

119745

!0.114957

90.00008

!0.000002

0

0

0

0

0

0

0

6

0

0

0

6

0

0

0

0

0

0

.022907

.122367

.311759

.533899

.757123

.980227

.030255

.065634

.099193

.119224

.125056

.118202

.022907

.122367

.311758

.533899

.757123

.980227

!0.030255

!0.065634

!0.099194

!0.119226

!0.125058

!0.118205

!0.112955

0.027724

0.144039

0.339451

0.561712

0.785056

1.008023

0.032896

0.070795

0.102624

0.120616

0.124947

0.116551

0.027724

0.144039

6.339451

0.561712

0.785056

1.008023

!0.032897

!0.070795

!0.102625

!0.120617

!0.124949

!0.116554

!0.110748

0.037720

0.165832

0.367143

0.589525

0.812989

1.035901

0.037709

0.075515

0.105722

0.121811

_ 0_. 124626

0.114791

0.037720

0.165832

0.367143 "

0.589524

0.812989

1.035901

!0.037710

!0.075515

!0.105723

!6.121813

!0.124629

!0.114794

I!

198
0!D

0.018419

0.003725

0.050188

0.193107

0.404523

0.631216

0.658895

1.086325

0.011106

0.041437

0.077634

0.106089

0.121052

0.124174

0.115852

3.6100

0.018419

0.007176

_Q_,.Q.65_za7_

0.215530

0.432860

0.659676

_Q,887355_

1.125576

0.015476

0.045354

0.081728

0.108690

0.122095

0.123796

0.113535

1.232330

0.000000

0.010996

.1ЫШ50

0.237953

0.461196

0.688136

_Q.915&15_

0.000000

0.113751

0.015063

'82348

0.260376

0.489533

0.716596

0.944275

0.156240 89.99991

1.118579 0.000002

0.019500 0.024307 0.029360 0.039712

0.282922 0.319760 0.347973 0.376186

0.517870 0.546206 0.574543 0.602880

0.745056 0.773515 0.801975 0.830435

0.972734 1.001069 1.029488 1.057906

0.019531 0.023420 0.026964 0.029887 0.

0.049061 0.052372 0.058509 0.064038 О,

0.085376 O.OS8731 !L.Q519J)Z CUBS6Z56 О.,

0.111086 0.113262 0.115164 0.116992 О,

0.122991 0.123627 0.123917 0.124190 0.

0.123213 0.122400 0.121399 0.120221 О,

032485 0.037128

069050 0.073642

,.1QQ14_2_ILJ.Q3252_

118599 0.119910

124389 0.124375

118892 0.117423

0

0

0

0

0

0

1

!0

!0

!0

!0

!0

!0

!0

0

0

0

0

0

0

0

1

0

.003725

.050188

.193107

.40'i523

.631216

.858895

.066325

.011106

.041437

.077833

.106068

.121051

.124172

.115849

E!E

.017657

.003836

.051522

.196941

.411742

.641975

.872836

.103628

.010968

0

0

0

0

0

0

1

!0

!0

!0

!0

!0

!0

!0

0

0

0

0

0

0

0

1

0

.007176

.060787

.215530

.432860

.659676

.887355

.125576

.015476

.045354

.081727

.108689

.122093

.123793

.113532

3.9850

.017657

.007475

.062314

.219651

.440474

.670833

.901693

.143790

.015340

0

0

0

0

0

0

!0

!0

!0

!0

!0

!0

1

!0

0

0

0

0

0

0

0

.010996

.071506

.237953

.461196

.688136

.915815

.019531

0

0

0

0

0

0

!0

.015063 0.019500

.082348 0.104278

.260376 0.282922

.489533 0.517670

.716596 0.745056

.944275 0.972734

.023420 !0.026964

.049061 !0.052372

.085375 !0.088730

.111085 !0.113261

.122989 !0.123625

.123210

.254807

.000000

.011365

.073230

.242486

.469206

.699690

.930551

.019258

!0

!0

0

0

0

0

0

0

0

0

.122398

.000001

.117210

.015632

.084271

.265321

.497938

.728548

.959409

.023151

!0.

!0.

!0.

!0.

!0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.024307 0.029360

0.126331 0.148508

0.319760 0.347973

0.546206 0.574543

0.039712

0.170807

0.376186

0.602880

0.773515 0.801975 0.830435

1.001069 1.029488 1.057906

!0.029887 !0.032485 !0.037128

058509 !0.064038 !0.069050 !0.073642

091902 !0.096755 !0.100148 !0.103251

115162 !0.116991 !0.118598 !0.119908

123915 !0.124188 !0.124387 !0.124373

121397

230170

137596

020275

106604

288156

526670

757405

988266

026663

!0.120218 !0.118889

90.00014

!0.000003

0.025169 0.030314

0.129063 0.151647

0.325671 0.354277

0.555402 0.584260

0.786263 0.815121

1.017051 1.045910

0.029482 0.032001

!0.117420

0.040855

0.174231

0.383010

0.613117

0.843978

1.074769

0.036511

0.040692

0.076288

0.104290

0.119959

0.124612

0.119021

0.003835

0.051522^

0.196941

0.411742

0.641975

0.872836

1.103627

!0.01098

!0.040692

!0.076289

!0.104293

!0.119962

!0.124616

!0.119027

0.044496

0.080094

0.106921

0.121067

0.124518

0.117042

0.007475

_0.i062314_

0.219650

0.440474

0.670833

.0..901693

"l". 14378 9~

!0.015340

!0.044496

!0.080096

0.048140

0.083688

0.109337

J). 122068

0.124262

0.011365

0.073230

0.242486

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п

200
CA131100.NZZ354B

TITLE

CASE-INLET GUIDE VANE.FRONT (O.D.) CAT 1 P/N4060676 NO REV EKD4060676 NO REV

_EHECKED_BL_ S.SALVAGGIO

OES JOB NO TD NO REV

31154B

R.L.WALLS

_DATE APPROVEO BY DATE

630811 R.L.SANFORD 830811

L/0 SH PART NO AIRFOIL EHD

237442 1 4060676 NO REV 4060676 NO REV

CUB

ASSOCIATED PART NUMBERS

ASSOCIATED EHD NUMBERS

ASSOCIATED COMPUTER FILES

CASE-INLET

11 50

AA-AA

0.024363

0.004702

0.051023

0.197033

0.416177

0.650498

0.835865

1.121197

0.014631

0.043772

0.092446

0.123454

0.150459

0.160409

0.159735

0.004918

0.051501

0.198165

0.417309

0.651630

0.886366

1.122196

-0.014437

-0.047669

-0.089318 - -0.122823

-0.142682

-0.150304

-0.143578

GUIDE VANE, 3 0

# 4.6800 1

0.024362 0

0.008235

0.061753

0.220060

0.445353

0.679804

0.915302

1.161263

0.019329

0.053381

0.097341

0.131987

0.152407

0.160774

0.158473

0.008451

0.062361

0.221192

0.446485

0.680936

0.916303

1 . 162257

-0.018994

-0.052133

-0.093912

-0.126066

-0.144423

-0.150970

-0.147062

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

_p

-0

-0

-0

-0

-0

-0

FRONT (O.D

.308396 -0

.000243 0

.012029

.072612

.243217

.474659

.709241

.944739

.023639

.057652

.101817

.135270

.154085

.161123

.012246

.073351 . 244349

.475791

.710243

.945740

.023184

.056264

.098097

.129063

.145829

.151147

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

. ) CAT 1 P/N4060676 NO REV

.002704 0.350610 89.37645

.152894 1.155608 0.005686

.016086

.083602

.266374

.503965

.738678

.974176

.027872

.061745

.106026

.138305

.155552

.161366

.016302

.084341

.267506

.505098

.739680

.975178

-0.027290

-0.060166

-0.102030 -0

-0

-0

.131829

.147043

.151177

0

0

0

0

0

1

0

0

0

0

0

0

0

0

0

0

0

1

-0

-0

-0

-0

-0

-0

.020404 0.024984

.105974 0.128477

.289532 0.328650

.533272 0.562578

.768116

.003609

.031880

.069000

.110009

.141150

.156864

.161403

.020620

.106844

.290664

.534404

.769117

.004608

.031186

.067114

.105753

.134393

.148115

.151007

0.797553

1.033006

0.035381

0.075621

0.116258

0.143787

0.158036

0.161261

0.025331

0.129478

0.329782

0.563710

0.798554

1.034005

-0.034692

-0.073416

-0.111556

-0.136766

-0.149082

-0.150668

END4060676

0.

0.

0.

0.

0.

1.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

0.

1.

-0. -0.

-0. -0. -0. -0.

029956

151242

357826

591885

826990

062403

038449

081687

120546

146215

159038

160946

030303

152243

358958

593017

827991

063402

037668

079153

115526

138944

149909

150171

NO REV

0.040424

0.174138

0.387002

0.621191

0.856427

1.091800

0.043918

0.087323

0.124630

0.148430

0.159848

0.160451

0.040771

0.175139

0.388134

0.622323

0.857429

1.092799

-0.042948

-0.084466

-0.119306

-0.140883

-0.150525

-0.149486

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060676, Sheet 1.

201
AB!AB

0.023767

0.004698

Q. 052135. 0.201473

0.424744

0.663734

JU2Q3524 0. 1.143270 1. 0.014514 0.

0.048727 0. 0.092444 0.

0.123256 0.

0.150214 0.

0.160121 0.

_u

0.004999 0.

0.053236 0.

0.202973 0. !Q.426377_0.

0.664968 0. 0.904624 0.

1.144365 1.

4.6800 1.33г395 !0.005988 0.«02113 89.10313

023765 0.00032* 0.153096 1.179393 0.005383

008296 0.012160 0.016291 0.020688 0.025485 0.030549 0.041209

063135 0^074385 0!005575—0!108488—0.131535_0!154714_0!178027.!

0.335622 0.365330

0.573946 0.603787

0.813603 0.843576

053394 L!083353

224919 0.246499 0.272076 0.295791

454584 0.484425 0.514265 0.544106

693708 0.723662 0.753655 0.783629

933497 0.963471 П. 993445 ]_0234

164739

019216 0.023505 0.027739 0.031751 0.035343 0.038369

053382 0.057676 0.061775 0.069013 0.075633 0.081686

.057252 0^1017.22 0^105952 Q .109952 0..11610S QJ.2Q409 0^124460!

131797 0.135058 0.136090 0.140905 0.143544 0.145964 0.148133

152182 0.153805 0.155229 0.156524 0.157737 0.158807 0.159651

160474 0.160849 0.161094 0.161142 0.161021 0.160740 0.160282

0.395037

0.633761

0.873550

0.043826

0.087265

008730 0.012728 0.016992 0.021522 0.026319 0.

064295 0.075485 0.086809 0.109722 0.132902 0.

226553 0.250132 0.273711 0.297424 0.337256 0.

.456035 0.4.85925_.0.515765 0.545606 0.575446 О. 694806 0.724782 0.754756 0.784729 0.814703 0.

934598 0.964571 0.994545 1.024530 1.054489 1.

185073

018921

031516 0.042309

156214 0.179527

366963 0.396670

605287_fl.635127! 844677 0.674650

084448 1.114407

!0.

!0.

!0.

0.

0.

0.

0.047379 !0.

0.033396 !0.

121570 !0.

1416.ft6_rlb

150201 !0.

143901 !0. AC!AC

0.022356

0.004785

0.055674

0.213479

.CL!44..857 7 fl

0.700028 0.

0.952318 0. 1.205076 1.

0.014404__CU

0.043930 0. 0.092629

0.128235

0.149990. 0.159969

0.159718

0.005328

ДЛ137614 0.

0.215978 6. 0.450517 0.

0.700990 0.

0.952162 0.

0.

0.

0.

0.

0.

0.

!0.. 051743 !0.055781 !0.059680 !0

092910 !0.097066 !0.100975 !0

124806 !0.127799 !0.130575 !0

143326_гО.,144.7.77_л0..146051_ О т.

150496 !0.150751 !0.150855 !0

147605

5.5000 1.397901 !0.014479 0

0.2 2 35.1, 0_,0 0 0571_0,154286 1

008700 0.012894 0.017368 0

067416 0.079297 0.091177 0

238216 0.263098 0.287978 0

47.9386 !L 5.11335._0.. 542784_0

731477 0.762926 0.794375 0

983907 1.015436 1.047043 1

247926

P_19182_0_._02356JZ 0_,.027917._.0

053589 0.057888 0.062007 0

097487 0.101961 0.106133 0

131745 0.134998 0.138006 0

151823 0.153548 0. !55057___0

160442 0.160687 0.160820 0

158691

009522 0.013856 0.

069356 0.081237 0.

_^0 .042756. 0.078377 !0.083594

0.114340 !0.118070

0.137748 !0.139709

Q .149211_тО. 149868! .150803 !0.150591 !0.150220 !0.149667

066469 !0.072698

104671 !0.110369

133152 !0.135562

240716

481826

732299

983611

.265458 0

0.513135 0

0.763608 0

1.015002 1

018470 0. 093258 __Q,

290337 0. 544444 0.

794917 О, 046470 1,

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,574232.. .825963

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156341,.

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1.203S07 1. 0.013925 !0.

0.046273 !0. ). 035439 !0.

ОЛ1761(Г!Ь. 0.135420 !0. 0.148933 !0.

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0.022806 !0. 0.054331 !0. .09.387_7_г

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6.124100 !6.

0.141985 !0. 0.149967 !0.

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101230

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88.26106

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0.139679

0.354649

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1.110256

_0.03552Z_ 0.075870

0.116220

0.143392

0.157478.

0.160890

0.028815

..0.142038

0.357009

0.607063

0.657815

.1.109405_

!0.034021 !0.070466

!0.106838

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0.032746

0.164139

0.385958

Q.637130

0.889141

1.141863

0.044070

0.188739

0.417268

0. 668579! 0.920729

1.173470

Л. 038570 5... 044020.. 0.081897 0.087512

0.120461 0.124452

0.145791 0.147982

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0.034407

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0.638372

0.889264

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0.669661

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0.036826

0.075866

0. 110682_

0.134334

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0.150832

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202
АО!АО

0.021643

0.004944

6.1100

0.021628

0.009161

1.454263 !0.020087

0.000809 0.155265

0.013670 0.018469

095S4g

0.692153

1.299042

0.023560

0.003734

0.029086 0.034904

0.146599 0.17

0.370860 0.403438

0.632499 0.665367

0.895446 0.928314

0.046830

0.436161

0.698235

0.961182

a _ 1.191504 I.gg440

0.223680 0.249567 0.275455 0.301342 0.327375

0.468884 0.501607 0.534330 0.567053 0.599776

0.731104 0.763972 0.796840 0.829709 0.862577

0.994051 1,fl2fe97fr 1^059881 1.092787 1.Iggfe9g

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II

206
CA131200.No23548

TITLE

FLAP!INLET GUIDE VANE,REAR (I.D.) CAT 1 P/N4060682 NO REV EHD4060682 NO REV

ENGINEER'S NAME

S.SALVAGGIO

DES JOB NO

31154B

-CHE£KEJUBY_ DATE APPROVED Bl

R.L.WALLS 830811 R.L.SANFORD

TO NO REV L/0 SH PART NO AIRFOIL EMD

237442 1 4060682 NO REV 4060682 NO REV

С LAB

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

ASSOCIATED COMPUTER FILES

FLAP!INLET

8 22

A!A

0.125000

0

0

0

0

0

0

0

0

0

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0

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0

0

0

0

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I !~0

!0

0

.062499

.364515

.743954

.108253

.004680

.048058

.062499

.364516

.743954

.103252

.084678

.048053

B!B

.125000

.066044

.382445

.782426

.110223

.085879

.043549

.066043

.332445

.782426

.110224

.005832

.043554

С!С

.125000

GUIDE VANE, REAR (I,

3 0

* 0.3180 1.004217

0.125000 0.000000

0.070921

0.411944

0.791379

0.111542

0.080089

0.043427

0.070922

0.411944

0.791379

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!0.080086

!0.043423

0.4800

0.125000

0.073998

0.432441

0.832419

0.113123

0.081193

0.043829

0.073997

0.432440

0.832419

!0.113123

!0.081196

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0.6410

0.125000

0.079942

0.459375

0.838802

0.112172

0 . 075524

0.038784

0. 079942

0.459376

0.838802

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1.055739

!0.000001

0 . 082446

0.482439

0.882412

0.113685

0.076535

0.039109

0.082445

0.482439

0.882411

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!0.076539

!0.039116

1.111037

!0.00000

.0.

0

0

0

0

0

0

0

0

0

0

) CAT 1

.000000

.024871

.127406

.506808

.886225

.107971

.070981

.034134

.127407

.506809

0.886226

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!0.070978

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0

0

0

0

0

0

0

0

0

0

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0

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.024982

.132492

.532441

.932404

.109576

.071901

.034387

.132491

.532439

.932404

.109577

.071905

.034394

.000000

.025008

P/N4060682 NO REV EMD4060682 NO REV

0.125000 89.99988

0.979346 0.000003

0

0

0

0

0

0

0

0

.174823

.554242

.933647

.103265

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0.933647

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.030757

.182486

.582441

.982396

.104874

.067271

.029657

.182485

.582441

.982396

.104876

.067275

.029664

.125000

.086029

0.222232

0.601674

0.981068

0.098469

0.061885

0.024814

0.222233

0.601674

0.981068

!0.098468

!0.061881

!0.024808

90.00026

!0.000004

0.232462

0.632441

1.032387

0.099981

0.062628

0.024925

0.232461

0.632440

1.032387

!0.099983

!0.062633

!0.024933

89.99985

0.000003

0

0

0

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.282452

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.095239

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0.317088

0.696529

0.089278

0.052687

0.317088

0.696530

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0.332450

0.732432

0.090576

0.053253

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For definition of this dimension — see drawing T4060682.

207
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0.401577

0.823689

0.112041

0.086948

0.04S947

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0.401578

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0.125000

0.076677

0.447094

0.923053

0.115282

0.039137

0.049643

0.076678

0.447094

0.923058

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E!E

0.125000

0.032075

0.494038

1.027457

0.117399

0.090972

0.050277

0.032075

0.494038

1.027456

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F!F

0.125000

0.090433

0.544900

1.138023

0.120125

0.093288

0.051034

0.090432

0.544899

1.133021

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0.125000

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0.597894

1.255424

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1.3990

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0.095426 0.089969 0. 08<W78

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0.096947 0.100679 0.104459

0.597894 _ О"бЯОПЙЬ П 7ftP

1.255424 1.337616 1.419809

!0.121811 !0.122426 !0.122567 "0.09542 6 "0.089968 "0.084478

H!H

0.125000

0.101779

2.3380 1.768634

0.125000 !0.000000

0.104660 0.107968

0.797050

0.079007 0.073582 0.068170 0.062766 0.057350

0.035700 0.030306 0.024913

0.186784 0.269076 0.351333 0.433540 0.515714

Ы> P.qPM.53 1.0QBfl4y ;b091Q4O Ч Л у ч р Ч?

1.502003 1.584197 1.666390

!0.119841 !0.116220 !0.111910 !0.106726 !0.101027

!0.079007 !0.073582 !0.068170 !0.062765 !0.057349 !0.03S6oy !n . P

!0.000000 0.125000

1.313622

0.122824

0.096762

1.399717 1.485813

0.123242 0.123340

0.091209 0.085632

0.046980 0.041477

0.025012

0.194206

_Я^31А2_ 1.571910

0.120919

0.080061

0.0359в

90.00012

1.743621 !0.000003

0.280408 0.366579

1

1.658008

0.117457

0.074538

0.452702 0.538784

055333 1.141АЗП I.gg

1.744105

0.113286

0.069027

0.108243

0.063521

0.102504

0.058008

0.101779 0.104860 0.107968 0.194206 0.280408 0.366579 0.452702 0.538784

0.624S65 0.710958 0.797050 0.883142 0.969238 1.055333 1.141430 1.227525

1.313622 1.399716 1.485813 1.571911 1.658008 1.744105

!0.091212 !0.085634 !0.080064 !0.074541 !0.069031 !0.063525 !0.058012

!0.046985 !0.041482 !0.035985 !0.030492 !0.025011

!0.096764

!0.052496

II

209
CA131300.N223545

TITLE

FLAP!INLET GUIDE VANE,REAR (O.D. ) CAT 1 P/W060681 NO REV EMD4060681 NO REV

ENGINEER'S NAME

"575ATVSGGIO

CHECKED BY DATE APPROVED BY DATE

R.L.WALLS

DES JOB NO

31154B

TD NO REV L/0 SH PART NO AIRFOIL EHD

237442 1 4060681 NO REV 4060681 NO REV

630811 R.L.SANFORD 830811

CLAB

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

ASSOCIATED COMPUTER FILES

FOP"INLET GUIDE VANE,REAR (O.D. ) CAT 1

11 27 3 0

A!A * 5.8980 1.760537 !0.001203

0.156000 0.155995 !0.001203 0.031246

ТГТ37757ОТШП З OTIW5ST557216 76T

0.541704 0.607791 0.673888 0.739989

1.070435 1.136514 1.202589 1.268657

1.598927 1.664968 1.731005

P/N4060681 NO REV EM04060681 NO REV

0.155995 89.45580

1.729291 0.000113

0.277529" OT3432S8~~574"09456~0.475604"

0.806082 0.872172 0.938262 1.004351

1.334721 1.400779 1.466833 1.532882

5Л5372707I55D056715555 ? ОП52375~

0.132608 0.127155 0.121818 0.116532

0.089384 0.083341 0.078245 0.072563

0.043343 0.037351 0.031312

U7I59755 0.146453 0.142559"^Л 37907"

0.111151 0.105736 0.100308 0.094868

0.066817 0.061020 0.055182 0.049290

0.135370

0.540845

1.069612

1.598684

!0.129010 !0.083238

!0.041132

0

0

0

—г

1

0

0

I ~V

0

0

0

.156000

.134311

.574343

.725766

.155797

.136358

7092085" .044046

.136912

.575980

0.139562 0.143771 0.210053

0.606905 0.672985 0.739079

1.135734 1.201859 1.267986

1.664835 1.730989

!ОЛ56214 !0.122884

!0.077833

!0.036081

b.4/UO

0.155994

0.137999

0.646326 !17222174! 1.797685

0.156176

0.130818

~~ff. 035359" 0.037728

0.140712

0.647768

!0.156240 !0.116987

!0.072463 !0.031086

~T7899Z5T

0.001311

0.141705

0.718315 —Г72 95135"

1.869597

0.156280

0.125369

0~.08055Z" 0.031355

0.144527

0.719574

!0 . !0. !0.

~тг

0.

0.

0.

0.

0. "ОТ

0.

0.

153395

111240

067120

d013ll

031170

213889

790310

35BC5S"

155068

119989

U /46 35

216532

791397

0

0

1

!и

!0

!0

и

1

0

0

1

0

0

и

0

0

.276290

.805174

.334119

.149670

.105495

.061828

0.342492

0.871273

1.400255

!0.145334 !0.099814

!0.056583

.15599$ 88.56271

.868085 0.000223

.286048 0.358177

.862296 0.934278

.438034

.152788

.114499 . UbOb4i

.288496

.863222

1.509975

0.149694

0.108949

0.06Z593

0.360424

0.935054

0

0

1

!U

!0

!0

0

1

1

0

0

~0

0

1

7468661"

.937380

.466393

0.474776

1.003492

1.532537

.Г40535 !0.135041

.094215 !0.088693

.051379 !0.046239

.430279

.006256

.581911

.146021

.103363

. 056471"

.432322

.006891

0.502337

1.078234

1.653841

0.141568

0.097753 !07050Z9I

0.504176

1.078737

* For definition of this dimension — see drawing T4060681.

210
150589

725576

153517

127183_

81330

0.040456

C!C

0.156000

1.222445

1.797470

0.153839

0.120988

T7138631

0.629181

1.275049

1.920508

ОЛ56887

0.140465

0.095509

0.045037

!0.07599 5 !0.035651

4.8440

0.155989

0.141304

0.709925

1.355761

2.001157

T7l5711<

0.134982

0.089575

0.038305

1.294309

1.869370

!0.153870

!0.114996

!0.07072 6 !0.03092 1

2.111616

0.001856

0.143984

0.790663

1.436464

2.081799

Л5719Т

0.129452

0.083539

0.031480

1.366174 1.438045 1.509921 1.581801 1.653687

!0.151195 !0.

0.109228!0. !о1065513 !0.

0.001856

_0^.031154_

01224912

0.871411

1.517158

147588

103476_

060362

155989

080463

305825"

952151

597840

•0.143309

l0..p978Q5_ !0.055288

87.19080

0.^00354

6;386~713~ 1.032881

1.678517

0.138567 !0.133190

0.092220 j;jK086729_ o".050272 !0.045328

0.467574 0.548403

1.113608 1.194331

1.759189 1.839852

0.156858 0.155291 0.152716 0.149416 0.145375

0.124021 0.118480 0.112830 0.107120 0.101352

0.077365 0.071066 0.064688 0.058229 0.051671

(17142324

0.631518

1.275574

1.920121

T7I45142

0.711984

1.356120

2.000715

0.147964О, 0.792462 О,

1.436672 1, 2.081317

228648 0.309294 0.389901 0.470477 0.551022

872962 0.953471 1.033984 1.114506 1.195035

517230 1.597798 1.678369 1.758945 1.839530

!0.153545 !0.131712 !0.15372"! !0.151191 "!6. 147561 !0.143317 !0.127017 !0.120580 !0.114285 !0.108281 !0.102383 !0.096562 !0.07980 5 !0.074447 !0.069182 !0.064037 !0.058990 !0.054046 !0.03979 0 !0.035237 !0.030788

0.138452 !0.133081

0.090859 !0.085279

0.049190 !0.044443

4T22515 2.329100 0.001835 0~7I5§ 95?55.79541

0.155989 0.001835 0.031184 2.297918 0.000463

0.147299 0.148555 0.238336 0.328118 0.417881 0.507622

0.776634 0.866248 0.955864 1.045477 1.135078 1.224672

03643

2.120144

0.157518

0.144773

i.493417

2.209638

0.157582

0.139467

1.582982

2.299124

0.157615

0.133913

1.672533 1.762072 1.851604 1.94112S

0.597336

1.314260

2.030639

0.158288

0.128432

157584

122881

~U7D79"357 07093216 67555933" 0.046156 0.038946 0.031624

0.149710 0.151312 0.152913

0.689469 0.778771 0.868075

0.080463 ТГГЯТЗвДЛ

0.155674 0.152839 0.149284

0.117149 0.111313 0.105391

"Б75б"?075 ОТ6502"1407053240"

0.242426 0.331905

0.957403 1.046745

ЧГСгТЛГ75935871755593 3 Г7572359 1.76ШЗ

119530 2.208978 2.298431

0.421344 0.510751

1.1360941.225451

1.651224 1794~бб53

!0.154038 !0.154077 !0.154065 !0.151536 !О

!0.126838 !0.120179 !0.113551 !0.107255!0 !07078124 !0.072727~^07057ч76 !0."552404!0 !0.039114 !0.034851 !0.030716

2.560472 0.003539 О

0.003547 0.031206 2

!071Ы357—07

0.945919 1.045022

1.738512

2.530357

"07159633—07159727—07159352—07I5IZ89—07

0.150843 0.145750 0.140156 0.134551 О

0.098076 0.091449 0.084579 О

0.039788 0.031803

"07ТБ92990725521 3 О"

E!E 3.6020

0.156137 0.156097 !07152293—ОЛ5514"Т

0.747685 0.846814

1.540420 1.639472

2.332466 2.431417

1.837537

147988 !0.143550 !0.138484

101164 !0.095181 !0.089339

0575S9 !0.052697 !0".б4"8037

155960 84.44019

529272 0.000616

— 0755ffff56 — 07559~295

144123 1.243209 1.342288

936543 2.035541 2.134527

0.104534

0.047628

—ОЛ5938Г

0.752005

1.541521

2.331735

KI55

_

5 — 07160378 0:i58064"

128905 0.123026 0.116992

077476 0.070234 0.062847

0.600131

1.314822

!"27630087 "

!0.132983

!0.083657

!5764!3516!

0 7555507"

1.441358

2.233502

'0715WO~

0.110826

0.055312

0.850658

1.640264

2.430547

0.949304

1.739016

2.529365 !07]

1.047971 1

1.837780 1

!0.124872 !0.118082 !0.111182 !0.104614 !О !0.075152 !0.069814 !0.064663 !0.059760 !О

!0.038066 !0.034241 !0.030590

"357055

146661

936556

I146410"

.098385

.055095

TJ. 455838 0.554587

1.245358 1.344069

2.035339 2.134130

"07653313! 1.442790

2.232928

0.092296 !0.086389 !0.080674

0.050582 !0.046235 !0.042061

211
F!F

0.156000

0.140280

0.801164

!ТГ576854~

2.551658

0.161623

0.158526

0.111077

0.049357

0.153065

0.806537

2. 9760

0.155868

0.141927

0.910686

1.786262

2.660938

0.161761

0.153806

0.104211

0.040701

0.154017

0.917358

176/9331

2.550914

!0.149571

!0.121713

1.788239

2.659900

•0.149579

!0.114892

2.800242

0.006404

0.143579

1.020166

2.770205

0.161844

0.148221

0.097117"

0.031857

0.154968

1.026158

1.897161

2.768893

!0.149569

!0.107742

0.006404

0.031095

0.253160

_1Л29637_

2.005030

0.164847

0.142429

0.155868

2.769157

0.362772

_

2.114383

0.166358

0.136613

0.089738 0.082036

0.263970

1.134974

0.372949

1.243819

2.006095 2.115043

!0.146941 !0.143463

!0.100857 !0.094433

83.10599

0.000780

0.472395

_1Л 348560

2.223722

0.166266

_ОЛ30536_

0.074144

0.481893

JL.352677_

2.223998

•0.139015

!0.088241

0.582008

1.458002

2.333049

0.164721

JB.124232_

0.066075

0.691598

1.567432

Z.442359

0.162078

0.117739

0.590799

1.461550

2.332962

!0.133725

!0.082280

0.057808

0.699679

1.570434

2.441936

•0.127923

•0.076564

!0.036591

G!G

0.156000

:

0.065847

•0.033335

2.3500

0.155894

!0.06035 1 !0.03031 4

3.047098

0.005760

!0.056165 !0.051820 !0.047685 !0.043759 !0.040064

0.125679 0.129T36 5TT33430

0.855243 0.975508 1.095726

1.816739 1.936859 2.056962

2.777138 2.897102 3.017049

0.158SCT5 OTI594TZОП 597ПГ

0.163451 0.159416 0.154194

0.116655 0.109514 0.102107

0.050970 0.041563 0.031920

0.005760 0.155894

0.0310263.016085

T725 36 7 7 О Л7397Г"

1.215921 1.336112

2.177046 2.297097

81.69252

0.000909

!07494296"

1.456292

2.417133

0.614627

1.576454

2.537152

0.734945

1.696602

2.657155

0.164Ё14 ОП672~2"2"ОПб"8523 0.168114" 0.148453 0.142635 0.136571 0.130176

0.094356 0.086157 0.077709 0.069042

0.166302

0.123524

0.060126

0.137056

0.862170

.818666

.776045

0.140698

0.981711

1.938295

2.895758

1.

2. :Г

0714"9098 !0.149407 !0.149448 !0.146421

!0.120053 !0.113090 !0.105652 !0.098333

!0.067662 !0.062434 !0.057515 !0.052962

!0.035264 !0.032549 !0.030111

177340 3.293350 0.006158

0.155878 0.006158 0.030990

0.144353 0.264058 0.383743 0.503398

1.101223 1.220741 1.340291 1.459864

2.057937 2.177594 2.297268 2.416950

3.015475

0.623015 0.742602

1.579451 1.699051

2.536641 2.656339

!0.14270 6 !0.1I5I22 !0.132641 !0.126514 "

!0.09154 2 !0.085133 !0.079016 !0.073189

•0.048872 !0.045073 !0.041523 !0.038254

H"H

0.156000

0.120575 0.124933

0.914885 1.045806

555—27139229:

.006879 3.137439

0.158922

0.129344 0.260187

1.176682 1.307526

23529—27353737" 267976

0.

3.267028

0.391074

1.438361

2.484409"

80.34ШГ

0.001059

0.522011

1.569184

T76T503B"

0.652974 0.783937

1.699987 1.830772 !2T74!5686—2Г876294—•

0.158111

0.169920

Л23553! 0.052948

0.132673

0.922537

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0.166517

0.042646

0.137019

1.052643

2.0937ZV2

3

0.159343

0.161713

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0.032035

0.1413S8

1.182721

0.165033 0.169561 0.172269 0.173082 0.172131

0.156037 0.150154 0.1440410.137493 0.130620

0.0911590.08252557072636 57562943"

0.401877 0.532090 0.662268 0.792412

1.442904 1.573035 1.703188 1.833357

"27755961—27875233"

0.271637

1.312796

2.35чТ5Т" 2.4844332.614694

005515 3.135800 3.266092

148106 !0.148567 !0.148646 !0.144955

117166 !0.110024 !0.102373 !0.094685 !OT05BUB4~^07TJ53Z99 !0.048949

!0.140889

•0.087534

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!0.033651 !0.031618 !0.029917

J!J 1.1080 3.561573 0.007008 0.155842

0.156000 0.155842 0.007008 0.031002 3.530594

0.111035—07TToZ»95 ОЛ2205В 07Z65T2I—0.406256

0.975235 1.117476 1.259678 1.401838 1.543981

2.112379 2.254428 2.396452 2.538449 2.680401

3.247941 3.389757 3.531545

"0.136065

"0.080919

0.130375 !0.123958

0.074661 !0.068783 !0.041623 !0.038750~0. D 36 0 27"

78.96831

0.001209

0.54S45S!

1.686113

2.822323

117590709 0.832975

1.828222

2.964223

1.970310

3.106095

212
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0.

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0.

2.

3.

156434

177465

131368

055199

134627

984162

114761

246368

0.157742

0.174735

0.123471

0.043883

0.13G250

1.125483

2.256165

3.387855

0.158430

0.170481

0.115155

0.032197

0.135929

1.266758

2.397584

3.529346

0

0

0

0

1

2

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.164960

.106375

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.408027

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0

0

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0

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0.176576

0.152779

0.086982

0.560170

1.690651

2.821937

0

0

0

0

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.146055

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.701536

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.963408

0.178932

0.138893

0.066157

0.842870

1.973373

3.104886

!0.145837

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!0.031879

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!0.10552 6

•0.052829

•0.030618

•0.146832

•0.097728

•0.048276

•0.029767

•0.142584 !0.137949 !0.132723 !0.126742 !0.120030

!0.089753 !0.082261 !0.075466 !0.069150 !0.063286

•0.044225 !0.040895 !0.038077 !0.035607 !0.033542

K"K

0.156000

0.105805

006697

0.7880

0.155769

0.111969

1.154838

3.699001

0.008483

0.118293

1.302943

0.008483

0.031098

0.266172

J 1.451006

2.191017 2.338957 2.486869 2.634753

3.373593 3.521267 3.666908

0.156265 0.157919 0.158775 0.167835

0.182306 0.179893 0.175922 0.170509

0.155769

3.667931

0.414179

_L599044_

2.782588

78.29906

0.001303

0.562254 0.710383

1.895076

0.858540

2.043057

2.930386 3.078156 3.225890

0.136365 0.128253

0.056804 0.044790

0.122150 0.128187

1.017211 1.164293

2.19405427341243

3.371948 3.519218

!0.143851 !0.144801

!0.109947 !0.102472

0.119678 0.110620

0.032335

0.134296 0.281500

1.311356 1.458406

"2.488448 2.635668

3.666491

!0.145007 !0.140571

!0.094621 !0.086533

0.174602

0.164480

0.100804

0.428688

1.605479

2.782907

0.179658

_ОЛ58203_

0.090423

0.575861

JL.J752588_ 2".930i59

0.182612

0.151385

0.183380

0.144073

0.079677 0.068444

0.723007

1.899725

0.870121

2.046880

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L!L 0.3000

0.156000 0.155765

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1.057465 1.214499

2.469730

3.722995

0.157136

0.186568

2.312912

3.566457

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0.188552

0.142483

0.058601

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3.909704 0.008562

0.008562 0.051354

0.115947 0.272614

1.371507 1.528467

2.626521 2.783278

3.879496

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!0.033348 !0.031797

0.133992

0.045938

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1.068274 1.224195

2.315852 2.471887

3.564484 3.720592

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77.31290

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0.429464 0.586384 0.743379 0.900417

1.685397 1.842313 1.999206 2.156070

2.939985 3.096647 3.253284 3.409885

0.124982 0.115415

0.032819

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1.380102 1.535995 1.691906 1.847857

2.627935 2.784000 2.940082 3.096175

3.876700

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0.176892 0.183232 0.187280 O.l59000"

0.171626 0.165155 0.158152 0.150539

0.105064 0.094041 0.082675 0.070832

07756391 0.912344"

2.003835 2.159834

3.252274 3.408378

!0.12223б"07П5265~

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213
CA135200.N2353FC

TITLE

BLADE * DISK (ID) CATEGORY 1 P/N40600Z3 NO REV EMD4060023 NO REV

ENGINEER'S NAHE

W.THOMAS

DES JOB NO

11154B

_CJHECKED_BI_ _DAIE AP_P.ROVED_B1_ JATE

R.L.HALLS

TD NO

830823 R.L.SANFORO 830823

REV L/0 SH PART NO AIRFOIL EMD С LAB

237441 I 4060083 NO REV 4060023 NO REV

ASSOCIATED PART NUMBERS

ASSOCIATED END NUMBERS

ASSOCIATED COMPUTER FILES

BLADE 8 DISK (ID)

11 50 1 1

AA!AA * 3.4500

0. 014551 0.003295

!0.010698

0.473803

0.958305

1.442801

1

2

2

0

1

1

1

1

0

0

0

0

0

1

1

2

2

0

0

__ _1

1

0

0

.927299

.411796

.896293

.018163

.071758

.465046

.472221

. 188463

.705379

.089495

.016673

.495634

.974595

.453553

.932511

.411469

.890428

.008452

.810744

.208414

.280931

.060184

.636875

.077967

0.049865

0.534366

1.018867

1.503364

1.987861

2.472358

2.956874

0.240534

1.143938

1.485087

1.451572

1.136909

0.634358

0.006511

0.076543

0.555504

1.034465

1.513423

1.992381

2.471339

2.950312

0.151404

0.876961

1.236693

1.268040

1.016436

0.573189

0.001533

CATEGORY 1

2.954144

0.014173

0.110427

0.594929

1.079429

1.563926

2.048423

2.532920

0.429979

1.206162

1.499480

1.425602

1.082984

0.561178

0.136413

0.615374

1.094335

1.573293

2.052251

2.531209

0.276732

0.936212

1.259975

1.249931

0.970220

0.507320

1

0

0

0

1

1

2

2

0

1

1

1

1

0

0

0

1

1

2

2

0

0

1

1

0

0

P/N4060023 NO REV

.046111 1.076665

.004120 2.953551

.170990

.655491

.139991

.624488

.108985

.593482

.577016

.263155

.508414

.394805

.026807

.486324

.196283

.675244

.154204

.633162

.112121

.591079

.386308

.991923

.277608

.227089

.921612

.439729

0

0

1

1

2

2

0

1

1

1

0

0

0

0

i

i

2

2

0

1

1

1

0

0

.231553

.716054

.200553

.685050

.169547

.654044

.700849

.316714

.511979

.360061

.967339

.410283

.256153

.735114

.214074

.693032

.171990

.650949

.486882

.045164

.289002

.200318

.869572

.370878

EMD4060023 NO REV

113.28210

0.004077

0

0

1

1

2

2

0

1

1

1

0

0

0

0

1

1

2

2

0

1

1

1

0

0

.292115 0.352678 0.413241

.776617 0.837179 0.897742

.261115 1.321677 1.382239

.745612 1.806174 1.866736

.230109

.714606

.809949

.365060

.509609

.322249

.905000

.332583

.316023

.794984

.273944

.752902

.231860

.710818

.578969

.094487

.294438

.170425

.814568

.300284

2.290671

2.775168

0.906532

1.406407

1.501736

1.281701

0.840311

0.252592

0.375894

0.854855

1.333814

1.812772

2.291730

2.770688

0.662618

1.138444

1.294533

1.137712

0.757171

0.227317

2.351233

2.835731

0.992664

1.439167

1.489044

1.236957

0.773793

0.171247

0.435764

0.914725

1.393683

1.872642

2.351600

2.830558

0.739341

1.175785

1.289845

1.100823

0.697953

0.152900

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060023.

214
0

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0

0

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1

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2

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0

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.013589

.007805

.456774

.921354

.385930

.850507

.315084

.779660

.017429

.893157

.308243

.421905

.226583

.764949

.099947

.016863

.476199

.935535 . 394870

.654205

.313540

.772876

.006180

.670549

.051039

.203253

.068078

.677150

.086089

СС!СС

.013447

.005673

.441626

.888924

.336221

.763517

.230813

.678109

.017880

.695645

.086339

.281152

.221371

.843850

.119126

.017473

.459546

.901620

.343690

.765760

.227830

.669900

.004319

.521996

.854979

.043743

.023276

.719518

.099809

0

0

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.837746

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.961875

.340206

.414157

.181684

.691274

.006762

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.992952

.452287

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.370957

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3.9050

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. 756454

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.956879

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.086879

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0.986950

0.004454

0.166412

0.6309.92! 1.095570

1.560146

2.024723

2.489300

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107.62515

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0.282557

_JUZ47136_ 1.211714

1.676291

2.140867

2.605444

0.340630 0.398702

!JU605209 0~£63281

1.269786 1.327858

1.734363 1.792435

2.198939 2.257011

2.663516 2.7?1ЧД8

0.321594 0.443619 0.552202 0.650300 0.738537 0.818738

1.023285 1.080720 1.135416 1.186283 1.232226 1.272320

JL36I539 1.3Ю572 1.A05Z66 1.416895 .1.!423260_ЬА24912

1.401098 1.383058 1.360516 1.333949 1.303362 1.267395

1.133451 1.081900 1.025604 0.965141 0.901220 0.834554

0.614225 0.534556 0.453021 0.368839 0.280964 0.190843

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0.274491 0.279325 0.283094 0.285730 0.287278 0.267814

0.283865 0.280647 0.276522 0.271534 0.265711 0.259007

0.233661 0.223485 0.212412 0.200471 0.187693 0.174111

0.126318 0.111459 0.093911 0.075456 0.055934 0.035713

219
СА135100.М2353вР

BLADE & DISK (00) CATEGORY 1

ENGINEER'S NAME

M.THOMAS

OES JOB NO

31154B

TITLE

Р/Ш060023 NO REV

_CHPCKED BY

EMD4060023 NO REV

PATE APPROVED BY !DATE.

TD NO REV

R.L.WALLS

L/0 SH

630823 R.L.SANFORD 830823

PART NO AIRFOIL EMD С LAB

4060023 NO REV 4060023 NO t? JL

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

_A.SSOCIATEJJ_COMBJ1E8JEILES_

BLADE & DISK (00)

11 50 1 1

A!A * 5.1300

0.009176 0.008296

0

0

0

1

1

2

2

0

0

0

0

0

0

0

0

0

0

1

1

2

2

!0

0

!_S!

! 00

0

0

.004030

.474072

.944113

.414152

.884192

.354231

.824270

.012045

.236064

.398012

.496670

.507890

.394325

.OS0737

.011689

.477594

.943500

.409404

.875308

.341212

.807117

.004605

.160591

.268325

.312756

.298901

.221362

.031363

0.

0.

1.

1.

1.

2.

2.

0.

0.

0.

0.

0.

0.

0.

0.

0.

1.

1.

1.

2.

2.

0.

0.

0.

0.

0.

0.

!0.

062785

532827

002868

472907

942946

412986

883037

043049

259908

413855

503459

501371

368858

018442

069927

535833

001738

467642

933546

399450

865362

018693

177453

277525

313983

293411

205350

003809

CATEGORY 1

2.884295

0.003921

0.121540

0.591582

1.061623

1.531662

2.001701

2.471741

0.073458

0.282284

0.428701

0.508992

0.492727

0.340428

0.128165

0.594071

1.059976

1.525880

1.991784

2.457688

0.041444

0.192943

0.285730

0.314304

0.286752

0.187679

P/N4060023 NO REV

0.327835 1.627526

0.014332 2,8727.3?.

0.180295 0.239051

0.650337 0.709092

1.120378 1.179132

_1_, 59041 7_1. 6491 72.

2.060456 2.119211

2.530496 2.589251

0.102782 0.131362

0.303711

0.442521

0.513148

0.481960

0.308603

0.186403

0.652309

1.118214

1.584118

2.050022

2.515926

0.063303

0.207562

0.292891

0.313732

0.278915

0.168049

0.324480

0.455322

0.515707

0.469342

0.272951

0.244642

0.710547

1.176452

1.642356

2.108260

2.574164

0.084519

0.221570

0.298958

0.312304

0.270167

0.146160

EM

64

!JU

0.

0.

1.

_!,.

2.

2.

0.

0.

0.

0.

0.

0.

0.

0.

1.

1.

2.

2.

0.

0.

0.

0.

0.

0.

104060023 NO REV

.38498

008478

297806 0.

767848 0.

237887 1.

ZQ7927_1.,

177966 2.

648005 2.

159007_0..

344425 0 .

467336 0.

516450 0.

454573 0.

232155

302880

768785

234690

700594

166498

63240 2. _

104888

234791

303900

310057

260276

121283

0.

0.

0.

1.

1.

2.

JL!

0.

0.

0.

0.

0.

0.

356561

826603

296642

766682_

236721

706760

185641_

363382

478403

515274

«7318_ 185594

361118

827024

292928

758832

224736

690640

124329

247047

307788

307060

248975

093253

0.415316

0.885358

1.355397

JL, 825437 _

2.295476

2.765515

.5.2.11311

0.381204

0.488266

0.512464

0.417239

0.134654

0.419356

0.885262

1.351166

1.817070

2.282974

2.748878

0.142886

0.258178

0.310711

0.303393

0.235999

0.062970

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060023.

220
в!в

0.008617

0.005534

0.4S9S58

0

1

1

2

2

0

0

0

0

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0

1

1

2

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.911468

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.421938

.904276

.006023

.091381

.149106

.162684

.140937

0.092726

0.007848

С!С

0.008191

0

0

1

1

2

2

3

0

0

0

0

0

0

0

0

0

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2

2

3

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.199568

.248192

.249814

.183753

.036402

.009216

.508694

.008171

.507645

.007118

.506592

.006065

.006551

.055712

.089302

.039626

.067454

.036725

.002150

0

0

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1

1

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0

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2

3

0

0

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0

0

0

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5.7600

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.066075

!S5J39a_

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.003365

^az666_ .972017

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.334761

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.233527

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2.976188

0.002370

0.126615

0.610939

1.095262

1.579583

2.063905

_2^54в226_

0.051073

0.189094

Л^гв526_в 0.338245

0.322575

0.213996

0.130826

0.613168

1.095509

1.577847

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.964586

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J.J65_34_ .084374

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6.2000

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.024553

.130879

.207412

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.569026

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0.109586

0.156932

0.159858

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3.078791

0.001555

0.131311

0.632008

1.132704

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2.634787

0.039312

0.141989

0.214750

0 . 253959

0.240025

0.155201

0.134086

0.633563

1.133039

1.632513

2.131987

2.631460

0.011381

0.066852

0.092733

0.085363

0.060626

0.027508

0.196792

0.010067

0.167156

JO. £71479

1.155802

1.640123

2.124445

2-608767

0.070386

0.203310

0.295040

0.340780

0.314453

0.192531

0.191118

0.673460

1.155801

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0.159770

0.157603

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0.009737

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0.694595

1.195291

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2.697373

0.053589_ 0.152647

0.221575

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0.196520

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1.195474

1.694947

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1.382776

1.882250

2.381723

2.881197

0.042619

0.083828

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0.045278

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221
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0.007849

0.006667

1.040641

1.557624

2.074607

g. 591590

6.7000 3.180624

0.007806 0.000821

0.071291 0.135914

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1.105264 1.169887

1.622247 1.686870

2.139230 2.203853

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0.061743 1.768847 44.33003

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0.200538 0.265161 0.329784 П.717524 O.Tflglufl Q.ftuA77)

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0.394408 0.459031

1.428378

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0.132003 0.136430 0.140522^

0.159947

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0.156113 0.157836 0.159132

0.154228 0.151074 0.147251

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0.104764 0.110907 0.116724 0.122158

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1.610202

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3.290259

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0.674730

1.209287

1.743640

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3.213S62

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0.089497

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0.542264

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1.611239

2.145724

2.680209

2.211575

2.746128

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0.016350

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0.088488

0.072464

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225
CA135500.N236347

TITLE

CASE!EXIT GUIDE VANE CATEGORY 1 P/N4060684 NO REV EMD4060686 NO REV

ENGINEER'S NAME _СИЕСКЕО_ВХ_

S.SALVAGGIO

DES JOB NO

31154B

R.L.WALLS

DATE APPROVED BY DATE

830823 R.L.SANFORD 830823

TD NO REV L/0 SH PART NO AIRFOIL EMD

1 4060684 NO REV 4060686 NO REV

С LAB

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

CASE!EXIT GUIDE VANE (OD) CAT 1 P/N4060686 NO REV EMD4060686 NO REV

9 25 3 0

A A!A A # 5.9900 0.947383 0.014523 !0.058385 93.25330

0.044623 0.043558 !0.009690 0.008933 0.936689 !0.002055

0

0

0

0

0

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!0

0

0

0

0

0

!0

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!0

!0

0

0

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0

0

0

!d?

1!0

0

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.053174

.346007

.642298

.937332

.033885

.012288

.021533

.006702

.034722

.337850

.642725

.941129

.053430

.089163

.074819

.010649

AB!AB

.045481

.053992

.345294

.637353

.933399

.034893

.009854

.019307

.006871

.035778

.337236

0.089447

0.382947

0.679348

0.026193

!0.015498

!0.02006 9

0.072243

0.376032

0.680555

!0.06057 3

!0.090181

!0.06954 4

6.1350

0.044461

0.090096

0.382022

0.676691

0.027459

!0.012988

!0.018013

0.073109

0.375190

0

0

0

0

!0

!0

0

0

0

!0

!0

!0

0

!0

0

0

0

0

!0

!0

0

0

.125820

.419932

.716375

.018992

.018136

.018090

.109889

.414227

.718277

.067032

.090400

.063549

.943361

.009578

.126292

.418793

.713509

.020493

.015566

.016224

.110564

.413153

0.162307

0.456954

0.753359

0

!0

!0

0

0

0

!0

!0

!0

0

0

0

0

0

0

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!0

0

0

.012393

.020190

.015445

.147672

.452419

.755849

.072634

.089811

.056676

.014315

.008950

.162597

.455598

.750290

.014111

.017590

.013809

.148147

.451110

0

0

0

0

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0

0

0

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0

0

0

0

0

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0

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.198883

.494004

.790281

.006304

.021648

.012033

.185554

.490589

.793227

.077519

.086417

.048819

.059365

.934636

.198983

.492430

.787016

.008215

.019056

.010657

.185822

.489045

0.235550

0.531074

0.827135

0.000791 !0.022511 !0.007957

0.223529

0.528720

0.830428

!0.081619

!0.086205

!0.040156

92.09789

!0.001993

0.235454

0.529280

0.823683

0.002865

!0.019957

!0.006875

0.223584

0.526941

0.272298

0.568152

0.863953

!0.004152

!0.02276 9

!0.003561

0.261581

0.566796

0.867539

!0.084928

!0.083184

!0.031114

0.272000

0.566139

0.860317

!0.001941

!0.020289

!0.00278 7

0.261417

0.564784

0.309119

0.605230

0.900721

!0.00852 2

!0.02243 2

0.001228

0.299693

0.604802

0.904479

!0.087448

!0.079379

!0.021408

0.308616

0.603000

0.896905

!0.006185

!0.02006 4

0.001683

0.299307

0.602558

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060686 Sheet 1.

226
0.640253 0.677859 0.715361 0.752718 0.789892 0.826893 0.863808 0.900563

0.937042

!0.054222 !0.

!o.oeaggg !о. !0.073516 !О

!0.010614

AC!AC

0.046966 О

0.055181 О

0.344433 О

0.636453 О

0.927841

061061 !0.067244 !0.072605 !0.077271 !0.081170 !0.084305 !0.086663

089098 !0.089190 !Q.QS85Q9 !0.087059 !0.084fllft !О. Пв1796_Ы) .0780?^

,068319 !0.062421 !0.055672 !0.047973 !0.039482 !0.030620 !0.021126

6.3880 0.937718 0.013540 !0.061166 89.81125

091082 0.127058 0.163125 0.199261 0.235466

380851 0.417305 0.453791 0.490302 0.526831

672984 0.709501 0.745992 0.782443 0.818848

0.271733 0.308058

0.563370 0.599913

0.855228 0.891572

0.036995 О,

!О.ОС4489 !0.

!0.014716 !О,

0.007064

030177 0.023767 0.017883 0.012439 0.007480 0.002996 !0.000990

007512 !0.010057 !0.012103 !0.013644 !0.014671 !0.015182 !0.015193

013769 !0.012372 !0.010428 !0.007830 !0.004658 !0.001203 0.002600

0.037943 0.

0.336846 О,

0.636875 О,

0.931230

075023 0.112199 0.149479 0.186836 0.224265

374426 0.412012 0.449590 0.487146 0.524665

674125 0.711284 0.748312 0.785172 0.821876

0.261751 0.299283

0.562134 0.599541

0.858502 0.894989

!0.05533 7 !О,

!0.08544 7 !0.

!0.07052 3 !О,

!0.010535

061474 !0.067006 !0.071792 !0.075940 !0.079372 !0.082106 !0.084128

086085 !0.086022 !0.085248 !0.083743 !0.081503 !0.078536 !0.074872

065511 !0.059859 !0.053410 !0.046059 !0.037962 !0.029519 !0.020503

AD!A D 7.0690

0.051061 0.050266

0.059335 0.094957

0.345930 0.382023

0.634917 0.671071

0.923677

0.041276 О

0.003262 О

0.933312 0.013334 !0.065562 87.41484

!0.00897 4 0.008991 0.924471 !0.001637

0.130643 0.166399 0.202214 0.238085 0.274006 0.309973

JL.418097LJLt454197. PJ.49.Q318_JLl526455_0_t562604 ftJzS8760_

0.707219 0.743354 0.779465 0.815550 0.851619 0.887669

!0.00794 0 !О,

0.007319

0.042432 О,

0.338498 О,

0.029260 0.023888 0.018919 0.014382 0.010255 0.006548

0.002068 !0.004106 !0.005725 !0.006921 !0.007680 !0.008010

007478 !0.006657 !0.005379 !6.003558 !0.001266 0.001253 0.004033

035076

000387

079211

375667

0.116076

0.412834

0.635135 0.671994

0.926678

!0.05943 0 !0.

!0.08439 5 !0.

!0.06768 8 !О

!0.01035 3

АЕ"АЕ

0.055076 О

,064809

J>§4609_

,062894

0.153019 0.190027 0.227091 0.264199 0.301338

!Ол44??вв_р .487116_0_, 524207 Q_. 561246 0.598231_

0.708733 0.745318 0.781847 0.818274 0.854511 0.890707

!0.06955 8 !0.073657 !0.077112 !0.079908 !0.082051 !0.083545

!0.084183 !0.083093 !0.0813.46_;0.078?27_=0_.07582р_лО.072118_

!0.05727 0 !0.050698 !0i043829 !0.036440 !0.028168 !0.019719

0.064448 0.

0.349226 О,

0.636565 О,

_0^9^3741

0.043358 0.

О.С0370 7 0.

!0.00841 2 !О,

7.7500 0.933396 0.015108 !0.069438 87.92320

JL54JU6_j^ Л110241 О .J090_0_4_JO.924562. !0.001743

099814 0.135254 0.170767 0.206346TM Ь.241986~

385060 0.420928 0.456826 0.492747 0.528687

672526 0.708481 0.744421 0.780335 0.816220

0.277683

0.564641

0.852090

0.313432

0.600602

0.887939

037325 0.031215 0.025539 0.020293 0.015483 0.011117 0.007192

000660 !0.001953 !0.004124 !0.005865 !0.007162 !0.008007 !0.008420

007988 !0.007180 !0.005883 !0.004003 !0.001633 0.000959 0.003828

_ _ . "о". 045266 О,

0.340739 0.

0.636753 О,

J>..9269_23_ _

!0^064602 !О,

!0.090378 !О,

!0.071920 !0.

!0.010432

081946

377847

673479

07022Т~

090513

066604

0.118723 0.155586

0.414951 0.452038

0.710107 0.746600

0.192521 0.229517 0.266561 0.303639

0.489094 0.526106 0.563060 0.599945

0.782927 0.819105 0.855212 0.891182

!0.075171 !0.079429 !0.083007 !0.085896 !0.088083 !0.089570

!0.089959 !0.088705 !0.086744 !0.084068 !0.080691 !0.076631

!0.060654 !0.053926 !0.046350 !0.038092 !0.029531 !0.020413

227
AF!AF 8.4310 0.933227 0.017050 !0.073327 68.66858

0.059096 0.057947 !0.011593 0.009023 0.924399 !0.001866

0.069672 0.104762 0.139933 0.175183 0.210507 0.245899 0.281356 0.316869

О . 352433 0.368041 0.4g3fe89 0.459369 О.йЧЧП?^ tU530805_JU566548__J[U6023ai_

0.633056 0.673809 0.709555 0.745286 0.780990 0.816663 0.852320 0.887950

0.923524

0.046328 0.039462 0.033019 0.027034 0.021492 0.016413 0.011797 0.007643

0.003948 0.000711 !0.002061 ^0.004386^(LQQ6254!tLJQ7628 !0.008542 !О.ООЙ997

!0.009016 !0.008611 !0.007798 !0.006470 !0.004548 !0.002125 0.000526 0.003514

0.007114

0.048037 0.084607 0.121287 0.158068 0.194926 0.231852 0.268829 0.305843

0.342879 0.379923 0.416960 0.453977 0.490957 0.527687 Л.56475Л__0^601540_

0.638241 0.674847 0.711337 0.747670 0.783812 0.819790 0.855696 0.891436

0.926915

!0.069852 !0.075675 !0.080854 !0.085262 !0.088958 !0.091914 !0.094136 !0.095612

!0.096383 !0.096440 !0.095760 !0.094340 !0.092175 !0.089270 !0.085627 !0.081288

!0.076258 !0.070579 !0.064204 !0.056979 !0.048857 !0.040034 !0.030923 !0.021185

!0.010531

AG!AG 9.1120 0.933944 0.019146 !0.077107 89.51190

0.063011 0.061645 !0.013048 0.009029 0.925130 !0.001958

0.109706 0.144639 0.179670 0.214780 0.249966 0.285221 0.320540

0.391340 0.426807 0.462311 0.497844 0.533400 0.568974 0.604557

0.675729 0.711306 0.746866 0.782395 0.817892 0.853373 0.888825

0.074858

0.355916

0.640144

0.924221

0.048562

0.004075

!0.009650

0.007025

0.041352 0.034559 0.028291 0.022483 0.017155 0.012309 0.007952

0.000677 !0.002239 !0.004678 !0.006635 !0.008109 !0.009096 !0.009604

!0.009266 !0.008439 !0.007046 !0.005017 !0.002484 0.000269 0.003360

0.050642

0.345190

0.640253

0.927792

0.087101 0.123731 0.160479 0.197309 0.234213 0.271172 0.308170

0.382218 0.419237 0.456231 0.493182 0.530073 0.566891 0.603622

0.676779 0.713172 0.749382 0.785384 0.821210 0.856962 0.892519

!0.07509 1

!0.10242 0

!0.08046 1

!0.01056 5

•0.081143 !0.086545 !0.091089 !0.094910 !0.097933 !0.100187 !0.101676

•0.102404 !0.101615 !0.100038 !0.097658 !0.094483 !0.090543 !0.085864

•0.074388 !0.067561 !0.059827 !0.051171 !0.041813 !0.032173 !0.021846

0.649451

0.935186

0.049372

!0.000895

!0.014991

0.006718

0.051340

0.35033JL

0.649584

0.939288

!0.081616

!O._113882_

!0~.О 8974 8

!0.010825

AJ!AJ

АН!АН 9.7930 0.945083 0.023153 !0.080523 94.57854

0.067126 0.065231 !0.015839 0.009058 0.936314 !0.002269

0.081153 0.116044 0.151042 0.186173 0.221405 0.256741 0.292167 0.327675

0.363255 0.398899 0.434597 0.470343 0.506126 __0.541938 0.577769 0.613609

0.685285 0.721105 0.756896 0.792640 0.828339 0.864015 0.899642

0.041171 0.033443 0.026336 0.019752 0.013752 0.008308 0.003430

!0.004663 !0.007664 !0.010492 .И).012540 !0.014004 !0.Р1_4в84_г0_.р15210_

!Oioi4264 !0.013013 !O.oiil0l"!0.6o8458!0.005255 !0.001801 0.002110

0.088750 0.125810 0.163010 0.200324 0.237734 0.275221 0.312760

0.387913 0.425488 0.463032_ 0.500524 0.537943_ 0.575271 0.612490

"0.686546 0.723325 0.759867 6.796151 0.832234 0.868228 67903938

_ _ _

0.' 083884"

0.369363

0.659944

0.949499

!0.088674 !0.094920 !0.100272 !0.104769 !0.108363 !0.111058 !0.112888

!О .113999!0.113217 !0.111513 !0.108892 !0.105382 !J. 1010.PJL.!P._0?5788 :

0".082937 !0.075207 !0.6б6423 !6.056629!0.046113 !0.035299 !0.023585

10.0000 0.959543 0.026222 !0.080849 99.13867

0.Р6591р_7р_.01в035_0.0p902p__0.95088 2 !.002520

6.119092 0.154440 СГ.~189956 6.225601 0^261379 0.29727~5

0.405531 0.441767 0.478060 0.514398 0.550768 0.587157

0.696320 0.732670 0.768971 0.805201 0.841365 0.877498

0.333275

0.623553

0.913558

228
0.047893

!0.00806 8

!0.031029

0.006394

0.038670 0.030001 0.028045 0.014692 !0.012135 !0.015540 !0.018274 !0.020325 !0.020646 !0.018816 !0.016199 !0.012736

0.008018 0.002004 !0.003350 !0.021694 !0.022380 !0.022418

•0.008634 !0.004261 0.000663

0.050517

0.354613

0.660064

0.9540Ы

0.088018 0.125703

0.392988 0.431358

0.697720 0.735144

0.163566 0.201577

0.469698 0.507978

0.772274 0.809102

0.239716 0.277954 0.316261

0.546178 0.584273 0.622242

0.845709 0.882214 0.918347

!0.06461 2

!0.12273 2

!0.09767 2

!0.01096 1

!0.092761 !0.100011 !0.106270 !0.111547

!0.12309 9 !0.122455 !0.120773 !0.118059

!0.09026 8 !0.081775 !0.072079 !0.061289

!0.115812 !0.119071 !0.121373

!0.114381 !0.109751 !0.104167

!0.049767 !0.037925 !0.025005

229
CA135300.N236340

TITLE

CASE!EXIT GUIDE VANE CATEGORY 1 P/N4060684 NO REV EMD4060684 NO REV

ENGINEER'S NAME _CHECKEQ_fiY_ (1PPROVED BY .DATE

S.SALVAGGIO

DES JOB NO

31154B

R.L.WALLS

TO NO

JAIE.

830833 R.L.SANFORD 830823

REV L/0 SH PART NO AIRFOIL EMD С LAB

237442 1 4060684 NO REV 4060684 NQ REV

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

ASSOCIATED COMPUTER FILES

CASE!EXIT GUIDE VANE (ID) CAT

9 50 3 0

A! A * 5.0750 1.996845

0.005857 0.005748 0.001129

0

0

0

0

1

1

1

0

0

0

0

0

0

0

0

0

0

0

1

1

1

!0

0

0

1 " 0

0

0

0

.004546

.328901

.653257

.977612

.301961

.626309

.950658

.006862

.073720

.127905

.159694

.158300

.115334

.025849

.006579

.329814

.653050

.976285

.299520

.622754

". 945989

.004669

.040507

.074620

.087787

.079S52

.052314

.001776

0

0

0

1

1

1

1

0

0

0

0

0

0

0

0

0

0

1

1

1

1

0

0

0

0

0

0

!0

.045090

.369446

.693801

.018156

.342505

.666853

.991226

.015351

.081543

.133259

.161530

.155285

.106748

.011106

.046983

.370219

.693454

.016689

.339924

.663158

.986400

.001111

.045730

.077548

.087865

.077580

.047301

.006297

0.085635

0.409990

0.734346

1.058700

1.383048

1.707397

0.023921

0.088958

0.138252

0.162872

0.151603

0.097440

0.087388

0.410623

0.733859

1.057094

1.380328

1.703563

0.006967

0.050585

0.080141

0.087598

0.074975

0.041936

1 P/N4060684 NO REV El

0.095421 1.094944

0.009054 1.988166

0

0

0

1

1

1

0

0

0

0

0

0

0

0

0

1

1

1

0

0

0

0

0

0

.126179

.450535

.774890

.099243

.423592

.747940

.032468

.096114

.142842

.163688

.147256

.087436

.127792

.451028

.774263

.097498

.420732

.743967

.012782

.055207

.082366

.087001

.072029

.036221

0.166723

0.491079

0.815435

1.139787

1.464135

1 . 788484

0.040976

0.103097

0.147003

0.163917

0.142267

0.076761

0.168196

0.491432

0.814667

1.137902

1.461137

1.784371

0.018543

0.059676

0.084193

0.086101

0.068770

0.030161

ЛО

6

!ft

0

0

0

1

1

1

0

0

0

0

0

0

0

0

"о

1

1

1

0

0

0

0

0

0

4060684 NO REV

3.91212

.002584

.207268 0.247812

.531624 0.572168

.855979 0.896524

...180330_U?20874_ .504679 1.545222

.829027 1.869571

.049367 0.057630

.109829

.150829

.163499

.136610

.065230

.208601

.531836

.855072

.178307

.501541

.824776

.024203

.063914

.085625

.084926

.065184

.023662

0.116230

0.154276

0.162396

0.130255

0.052712

0.249005

0.572241

0.895476

1.218711

1.541945

1.865180

0.029750

0.067841

0.086685

0.083499

0.061256

0.016667

0.288357

0.612712

0.937068

1,_261417 ...

1.585766

1.910114

O.J065752 . „

0.122230

0.157258

0.160665

0.123172

0.039490

0.289410

0.612645

0.935881

1.259115

1.582350

1.905584

0.035181

0.071387

0.087398

0.081816

0.056969

0.009323

This dimension is the radial distance from the rig centerline to the airfoil manufacturing cross section. For further

definition, see drawing T4060684 Sheet 1.

230
0

0

0

0

0

1

1

1

0

0

0

0

0

0

0

0

0

0

0

1

1

1

!0

0

0

0

0

0

!0

0

0

0

0

0

1

1

1

0

0

0

0

0

0

0

0

0

0

0

1

1

1

!0

0

0

1 0

0

0

!0

в!в

.005914

.004862

.329052

0

0

0

.653242 0

.977432 1

.301620 1

.625309 1

.949997 1

.006710 0

.С59096 0

.100754 0

.126567

.127153

.093703

.022393

.006453

.329794

.653135

.976476

.299809

.623143

.946477

.005004

.025679

.046791

.053555

.047181

.029767

.001944

С!С

.005934

.004887

.323965

.653043

.977121

.301195

.625269

.949343

.006689

.056247

.091167

.111281

.110184

.080521

.019999

.006458

.329831

.653203

.976576

.299948

.623320

7946692

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0.080660 0.070434 0.059599 0.048051

0.128399 0.168686 0.208973 0.249259

0.450692 0.490979 0.531265 0.571552

0.772985 0.813272 0.853558 0.893845

1 . 095276^ .135562 1 .175848^1. 216134_

1.417564 1.457850 1.498137 1.538423

1.739853 1.780139 1.820425 1.860711

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0.097388

0.084607

0.057822

0.023141

0.130141

0.009018

0.124562

0.447133

0.769705

1.092277

1.414847

1 . 737418

0.058510

0.152870

0.201478

0.210878

0.177107

0.098817

0.128906

0.449829

0.770752

1.091675

1.412598

1.733521

0.036594

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0.530060

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0.117688

0.134228

0.116508

0.078658

0.025371

0.245526

0.568098

0.890670

1.213241

1.535811

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0.100172

0.176121

0.209646

0.203566

0.153013

0.057967

0.249252

0.570175

0.891098

1.212021

1.532944

1.853867

0.069181

0.121994

0.133465

0.112763

0.072772

0.017708

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1.257519

1.581066

0.094627

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0.289546

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1.578709

1.900997

0.061435

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0.285847

0.608419

0.930991

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1.576133

1.898704

0.112331

0.182523

0.211193

0.199663

0.143601

0.042869

0.289368

0.610291

0.931213

1.252136

1.573059

1.893982

0.078326

0.125623

0.132194

0.108703

0.066665

0.009876

234
CA135400.N236344

FLAP!EXIT GUIDE

ENGINEER'S

S.SALVA6SIO

DES JOB NO TD

VANE

JJAMJE.

N0

CATEGORY 1

R.L

TITLE

P/N4060686 NO REV EMD4060686 NO REV

CHECKED BY DATE APPROVFn RY

.HALLS

REV L/0 SH PART

237442 1 4060686

NO

NO

830823 R.L.SANFORO

AIRFOIL EMD

REV 406068ft NO REV

*

DATF

830825

С LAB

1

ASSOCIATED PART NUMBERS

ASSOCIATED EMD NUMBERS

FLAP!EXIT GUIDE VANE (OD) CAT 1 P/N4060686 NO REV EMD4060686 NO REV

9 25

A"A

0.005995

0.006768

0.348754

0.692639

1.031895

0.004483

!0.03292 7

!0.02398 3

0.032779

3 О

# 4.1650

_ОЛ10582р_

0.049237

0.391758

0.735439

1.086260

!0.001436

0.091785

0.434787

0.778153

!0.011126 1.042488 64.81732

(Ы)45165 _ 1.042486_Hl.J3.11127

0.134440 0.177166 0.219969

0.477819 0.520845 0.563857

0.262847

0.606837

0.820768 0.863271 0.905650 0.947888

0.305781

0.649767

0.989975

!0.00246 0 !0.008895 !0.014580 !0.019701 !0.024131 !0.027766 !0.030674

!0.03448 2 !0.034961 !0.034782 !0.034100 !0.032762 !0.030654 !0.027702

!0.019524 !0.014309 !0.008333 !0.001607 0.005861 0.014093 0.023057

O.OC4018

0.348649

0.702613

1.053209

!0.00715 3

"0.08965 2

"0.10608 0

!0.05500 1

0.046269 0.088712 0.131489 0.174488

0.392617 0.436754 0.480986 0.525265

0.746887 0.791084 0.835179 0.879143

0.217708 0.261164 0.304822

0.569586 0.613935 0.658287

0.922951 0.966578 1.010004

•0.020653 !0.033536 !0.045260 !0.056145 !0.066112 !0.074994 !0.082826

•0.095493 !0.099869 !0.103184 !0.105782 !0.107505 !0.108187 !0.107681

•0.103414 !0.099668 !0.094874 !0.089006 !0.082061 !0.074062 !0.065034

в"в

О.ОС603 4

О.С0675 7

_0.348275_

' О7б~91330

1.030180

0.004595

!0^ ^2 9382

!0.'020117 "

0.034041

0.004120

0.348219

4.0200 1.085256 !0.010858 1.040486 64.71063

0.005875 !0.001374 0.046067 1.040486 !0.010858

0.049205 0.091724 0.134337 0.177011 0.219750 0.262550 0.305396

JO.!391179 0.434103__р_.477р28_.0д519947 0.562851 _0.6Р5725_ 0.648554

0".734041 07776675 0.819219 0.8б1б63 0.903992 0.94б197 0.988263

•0.001784 !0.007675 !0.012845 !0.017485 !0.021489 !0.024762 !0.027378

!p.030700_ :0,.03mi_^p.p3p774_j:p.030023 !0.028662_!р_.026570_!0.023695

!6.015834 !0.010835 !o"."o05i27 0.6oi286 0.008414 "Ь.016247 0.024792"

0.046369

0.392054

0.088795 0.131525 0.174460 0.217603 0.260966 0.304515

0.436052 0.480139 0.524272 0.568448 0.612652 0.656861

For definition of this dimension, see drawing T4060686.

235
0.701049 0.745192 0.789267 0.833250 0.877116 0.920839 0.964398 1.007771

1.050939

!0.007147 !О !0.087073 !О

!0.103856 !О

!0.055723

С!С

0.006067

0.006713

0.347339

0.689079

l.Qg?15a

0.004715 !О !0.024841 !О

!0.014651 !О

0.043954

0.10046

101417 !0.097951 !0.093465

.020172 !0.032606

09gSg4 !0,097153

0.054498 !0.064157 !0.072776 !0.080395

03092!0.104874 !0.10SA44 !0_1Q5?79

0.087953 !0.081403 !0.073838 !0.065273

3.7670 1.083549 !0.010467 1.037067 64.45766

,049093 0.091532

,390087 0.432846

,731641 0.774135

0.134050 0.176619 0.219239

0.475604 0.518354 0.561089

0.816553 0.858886 0.901122

0.261906 0.304610

0.603795 0.646461

0.943252 0.985267

000967 !0.006191

025825 !0.025903

010539 !0.005785

!0.010732 !0.014772 !0.018213 !0.021029 !0.023216

•0.025446 !0.024548 !0.023089 !0.020960 !0.018129

!0.000390 0.005639 0.012307 0.019621 0.027562

0.004211 0.046388 0.088724 0.131341 0.174145 0.217147

0.347248 0.390885 0.434676 0.478552 0.522472 0.566433

0.698402 0.742346 0.786233 0.830040 0.873747 0.917332

0.260353 0.303730

0.610423 0.654420

0.960773 1.004052

!0.007121 !0.019651 !0.031629 !0.042567 !0.052753 !0.062060 !0.070372 !0.077738 !0.084222 !0.089851 !0.094103 !0.097398 !0.100021 !0.101841 !0.102695 !0.102476 !0.101267 !0.099089 !0.095948 !0.091858 !0.086801 !0.080783 !0.073806 !0.065887

!0.057034

D!D 3.0860 1.088389 !0.011241 1.037717 63.23354

0.006088 0.005944 !0.001314 0.051904 1.037716 !0.011240

0.006716 0.049115 0.091572 0.134104 0.176683 0.219314 0.261991 0.304701

0.347432 0.390179 0.432932 0.475680 0.518418 0.561139 ^).603830 U.6.464Z8..

0.689077 0.731618 0.774095 0.816499 0.858820 0.901052 0.943183 0.985209

1.027123

0.004725

!0.023035 !0.011237

0.039571

0.004177

0.34722L_ 0.698970

1.040511 !0.007140

!0.000769 !0.005784 !0.010125 !0.013984 !0.017214 !0.019763 !0.021675

!0.023762 !0.023601 !0.022930 !0.P21823_r.P.,J)201Zl_^0.J»l 186JL^0^014863_ !0.006992 !0.002134 0.003313 0.009373 0.016030 0.023296 0.031146

0.046299

0.390916

0.088590

0.434774

0.743002

0.131183 0.173982 0.216993 0.260230 0.303652

JL4je719_.0.522712_.Q_.56675fl <Ь610820 О.,654902_ 0.786978 0.830879 0.874681 0.918366 0.961910 1.005298

!0.020141 !0.032578 !0.043939 !0.054499 !0.064154 !0.072747 !0.080345

!0.087053 !0.092875!0.097313 !0.100790 !0.103592_rO.10557в_^0.106584_r0.106489

!0.105387 !O.io"3291 !0.100238 !0.096238 !0.091285 !0.085377 !0.078510 !0.070720 !0.062010

Е!Е

0.006143

0.006805

0.343830

0.691947

1.030933

0.004639 !0.025454

!0.012808

0.041709. "о'.'ооЧозз' 0.348608

0.703366

_1.;055614_ !0.007278

!0.094108

!0.114403 !0.067666

2.4050 1.097693 !0.013005

JL! РР5970_^д. 001447 О ._Q56Q62_ 0.049327 0.091928 0.134618

0.391766 0.434710 0.477648

0.734652 0.777282 0.819827

1.043159 62.04858

_1.043159 !p.0130Q4_

0.177363, 6.220165

0.520575 0.563482

0.862277 0.904620

0.263020 0.305912

0.606354 0.649178

0.946848 0.988955

!0.001367 !0.006786 !0.011449 !0.015586 !0.019078 !0.021843 !0.023955 !0.026266 !0.026117 !0.025377 !0.024171 !0.022374 !0.019889 !0.016682

!0.008277 !0.003093 0.002745 0.009245 0.016407 0.024215 0.032655

0.046222 0.088598 0.131333 0.174316 0.217544 0.261030 0.304734

0.392628 0.436838 0.481155 0.525529 0.569955 0.614421 0.658901

0.747791 0.792150 0.836423 0.880581 0.924599 0.968456 1.012136

•0.021388 !0.034905 !0.047244 !0.058688 !0.069168 !0.078518 !0.086791

•0.100492 !0.105372 !0.109187 !0.112266 !0.114458 !0.115594 !0.115556

•0.112184 !0.108909 !0.104612 !0.099270 !0.092880 !0.085458 !0.077057

236
F!F 1.7240 1.107209

0.006074 0.005874 !0.001545

0.006779 0.049426 0.092169

Q.3501S3 0.393290 O.u^f.4^7

0.69*837 0.737708 0.780493

1.03*760

0.00**61 !0.002088 !0.007979 !0.0283*4 !0.029261 !О.Й29134

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0.043602

0.003831 0.046049 0.088500

0.3*9931 0.394307

0.707982 0.752835 0.797614

1.063161 !0.00726 5 !0 !0.101670 !0

0.015017 1.048826 60.95474

0.060283 1.048826 !0.015016

0.135017 0.177931 0.220912 0.263952 0.307037

0!425528 0!522208 0!565814 П,б0в«в0 0.651691

0.823182 0.865759 0.908214 0.950539 0.992724

0.013051 !0.017534 !0.021326 !0.024372 !0.026690

0^028390 !0.097191 !И. 094911 !П.nggSOS !n.moi.fcg

0.001734 0.008727 0.016424 0.024808 0.033872

ОЛ31374 0.174535 0.217976

gfc 0.1?!>Яй17 O.S73260

0.842292 0.886838 0.931222

0.261723 0.305725

n^fclplfcfc tt"663082"

0.975416 1.019405

!0.12397 9 !О

!0.07357 0

6!6

0. 005926 О

022599

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121603

0.037274

O.. 113973

0.050666 !0.063101 !0.074515 !0.084700 !0.093716

0.118126!=Д^214а7!=Д!Л23.900_=Д.Л25189!!^0^125198 !0.118088 !0.113463 !0.107711 !0.100815 !0.092793 !0.083717

0.0066*6 О

0.351473 О

0.697875 О

1.0420 1.116944 !0.017182 1.054828 59.89813

6*4*9 l.OSuagft !0.017182

0.135344 0.178437 0.221607

0.481561 0.524911 0.568235

0.826737 0.869451 0.912021

049*37

39*829

740934

0.092328

0.438198

0.783893

0.264848 0.308142

0.611515 0.654732

0.954437 0.996695

0.00*229

!0.031559

!0.017659

0.0*5239

!0.002836 !0.009255 !0.014783 !0.032631 !0.032564 !0.031802

!0.012476 !0.006522 0.000206

!0.019674 !0.023830 !0.027162 !0.029716

!0.030497 !0.028507 !0.025716 !0.022081

0.007720 0.016006 0.025051 0.034809

0.003554

0.351215

0.712847

1.071176

0.045762 0.088264 0.131268

0.395972 0.440979 0.486147

0.758168 0.803*03 0.8*8521

0.174626 0.218303

0.531394 0.576711

0.893485 0.938258

0.262326 0.306651

0.622081 0.667*72

0.982812 1.027124

!0.0071* 6

!0.109587

!0.133569 !0.07952 3

!0.023851 !0.039786 !0.054312 !0.117151 !0.123038 !0.127542

•0.131434 !0.127660 !0.122683

!0.067749 !0.080110 !0.091172 !0.100956

!0.131173 !0.133783 !0.135205 !0.135236 !0.116465 !0.108999 !0.100312 !0.090478

H!H 0.3620 1.133053 !0.021501 1.067714

0.006617 0.006275 !0.002099 0.068785 1.067714

0.007626 0.050516 0.093566 0.136816 0.180196

0.354830 0.398659 0.442522 0.486385 0,530234_

0.705069 0.748520 0.791831 О.в349"в1 0.877946

1.0*7635

0.00*379

!0.0*352 4

!0.029S8 0

0.0**289

0.003597

0.35*5*9

0.723717

1.С33104

!О.С0815 0 !0.12621 0

!0.153165 !0.037194

J"J

JL 006823

O.~6b8032

O.J57331

0.711296

1.055502

!0.00*821 !0.013239 !0.020559 !0.027074

Л!М5306_^0 ._Q457_Q3_rO..J)45176_=.0.043979

!0.023863 !6.016908 !о'.009020 !0.000174

0.045942 0.088554 0.131781 0.175512

0.400125 0^446007 O..j*92127_ 0.538355

0.770086 6.816237 6.862228 0.908003

!0.027116 !0.045468 !0.062319 !0.077819

!Р. IS^e??^ 1.14.1765_Г0..146819_Г0.150780

!0.149926 !6.145219 !ОЛ39101 !6.131539

0.1550 1.143979 !0.024464 1.078399

О ..р06388__!0. 002397__0..р69995_ 1.078398

6.051010 6.094170 6.i37608 0.181239

0.401604 0.445938 0.490283 0.534618

0.755144 0.798809 0.842268 0.885*85

60.08163 !0.021501

0.223701 0.267327 0.311046

_0I5.74054_0J 61Шв_0., 661496. 0.920702 0.963239 1.005553

!0.032683 !0.037263 !0.040845

!0.04196.7^0.038984_г_0^ 034930_ 0.009627 0.020345 0.031905

0.219649 0.264238 0.309231

0.584665_ 0.631039 0.677434

0.953514 6.998718 1.043591

!0.092119 !0.104939 !0.116263 !0.153603 !0Л5504_9_гО.154894

!0.122518 !6.112068 !0.100285

60.44670

!0.024464

0.225041 0.269016 0.31з£26

0.578920 0.623158 0.667293

0.928426 0.971077 1.013438

237
0.004225 -0.006708

-0.054503 -0.057051

-0.041838 -0.035211

..0.041680

0.003431 0.045882

0.357114 0.403388

0.732739 0.779811

0.016891 -0.025816 -0.033754 -0.040677 -0.046396 -0.050968

0.058019 -0.057766 -0.056721 -0.054746 -0.051664 -0.047340

0.027469 -0.018645 -0.008702 0.002372 0.014516 0.027633

0.088605 0.132009 0.176049 0.220564

0.449999 0.496918 0.543971 0.591116

0.826715 0.873401 0.919811 0.965873

0.265612 0.311160

0.638328 0.685558

1.011542 1.056799

-0.008545 -0.029258 -0.049402 -0.068024 -0.085098 -0.100887 -0.115082 -0.127582

-0.138497 -0.147967 -0.155583 -0.161016 -0.165133 -0.168016 -0.169409 -0.169035

-0.166857 -0.162967 -0.157406 -0.150245 -0.141470 -0.131024 -0.118974 -0.105457

-0.090501

238
APPENDIX D

DEFINITION OF SYMBOLS

AR Aspect ratio, H/b

A1 Point midway between leading edge and first captured Mach wave emanation

point on suction surface

A/A* Critical area ratio

a Acoustic velocity, m/s or ft/a

BPR Bypass ratio

b Chord, centimeters or inches

CAIJO Ratio of front camber!to!chord/total camber!to!chord

C, Lift coefficient for 63 series airfoil b

TV«. • * * 1 д. ! Г

D,, Diffusion factor =1 !! л !! 1! (77+ ri+1) V, a

E Excitation per rotor revolution

H Average blade height, cm or in.

ID Inner diameter, cm or in.

i, inc Incidence = angle between inlet air direction and tangent to blade reference

station, degrees

LER Leading edge radius, cm or in.

LMT Location of maximum thickness, % of chord

M Mach number

MCA Multiple circular arc

N Rotor speed, radians/s or rpm

OD Outer diameter, cm or in.

P Pressure, n/cm2 or psia

r Radius, cm or in.

SM Stall margin, %

Т Temperature, °R or °F

t Maximum blade thickness, cm or in.

239

3937C
t/b Maximum thickness!to!chord ratio

TER Trailing edge radius, cm or in.

U Rotor wheel velocity, m/s or ft/s

V Air velocity, m/s or ft/s

W Mass flowrate, kg/s or Ibm/s

Z Loss coefficient; axial direction

a Air angle, angle between absolute velocity vector and axial direction, degrees

/8 Air angle, angle between relative velocity vector and axial direction, degrees

•y Ratio of specific heats

Д Difference

5 Ratio of total pressure to NASA standard sea level pressure of 10.1315 N/cm2

(14.694 psia)

5° Deviation angle, degrees

17 Efficiency, %

!Л! Hub!to!tip ratio

в Ratio of total temperature to NASA standard sea level temperature of 518.7°R;

turning angle, degrees

в* Camber angle, degrees

p Density, kg/m3 or lbm/ft3

Ф Meridional flow angle, angle between axial velocity vector and centerline,

degrees

a Solidity, ratio of chord to spacing; stress, kg/cm2 or lbf/in.2

т Blade spacing, cm or in.

a; Total pressure loss coefficient

240
Tangential Direction

Axial RLE

Direction

Transition т

 ' Point

Front

Camber

F

of Maximum Thickness

Chord, b

Rotation

Figure 91. Multiple Circular Arc Airfoil Definitions

Axial Direction

FD 270101

241
MCA INCIDENCE DEFINITION

First Captured

Mach Wave

Suction

Surface

Tangential

A' Station Note: Incidence is Measured

at The A' Station.

FD 270102

Figure 92. Multiple Circular Arc Incidence Definitions

М2д = Constant

z

(Loss)

! Т

al a

Min

Stall a 2

<*R «n

Choke

Incidence (i) = amin " «2

»R " «L = Range of Loss Bucket Incidence for Low Loss

FD 268993

Figure 93. Cascade Loss Bucket Definitions

242
Subscripts

ad Adiabatic

CH Chord

F Front

L Left, stall limit of cascade loss bucket

LE Leading edge

M Meridional

0 Stagnation values

R Right, choke limit of cascade loss bucket

S Static

SS Suction surface

Т Total

ТЕ Trailing edge

в Tangential direction

1 Rotor inlet station

2 Stator inlet station

3 Stator exit station

Superscript

* Metal geometry

— Mean value

' Relative to rotor

243
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APPENDIX E

DEFINITION OF VARIABLES

Absolute Mach number:

M =

Static Temperature:

Т, т, =

1 + ' „ М

!1 ,„2

Acoustic Velocity:

a =

Absolute Velocity:

V = Ma

Axial Component of Absolute Velocity:

Vz

 = v

\/sec24> + tan2/?

Meridional Component of Absolute Velocity:

V m = Vz sec<t>

Tangential Component of Absolute Velocity:

V9 = Vz tanj3

Radial Component of Absolute Velocity:

V, = Vz tan0

Absolute Air Angle (meridional plane):

! tan!

245
Wheel Speed:

IIND U constant

Tangential Component of Relative Velocity:

v; ! и! vs

Relative Air Angle:

V = tan ]!TTv ~ (axial plane) z

V = tan" 1!??— {meridional plane)

Relative Velocity:

V = V.secJ'

Relative Mach Number:

M' = M cos/?'

Relative Total Pressure:

Pt' ! P

Relative Total Temperature:

Tt' ! T8[l + ^~! (M')2]

Pressure Ratio:

exit Pt PR Inlet Pt

Turning:

в' = inlet j8' ! exit j8' (rotor)

6' = inlet |8 ! exit J3 (stator)

246
Loading Parameter:

АР

Loss Coefficient:

exit ideal?,'! exit Pt' w/ = inlet P/! inlet Ps

 (rotor)

inlet Pt ! exit Pt u = inlet Pt! inlet Ps

 (stator)

where:

ideal P/ = inlet P / ll + !^ZlJ .!xit.

U

' ; ' Ч 2 !ygR inlet Tt

7

Г г _ (inlet radius \21

L ^ exit radius ' J

Loss Parameter:

Т Р' ! "'

 cos

 ( exit

LP ^) / * \ To (rotor)

_ и cos (exit g) . , , . ! 1^! (stator)

Diffusion Factor:

r>F ! i e x i t V ' , exit DV/ ! inlet DV/ U

* ~ l inlet V + ( e x i t D + inlet D)<nnlet V <KotorJ

r>F exit V (inlet DV,! exit DVa U1< : inlet V + ( e x i t D + inlet D) crinlet V

Adiabatic Efficiency:

ad

 TR! 1

Deviation Angle:

5° = /32* ! j82 (rotor)

6° = «3* — «3 (stator)

247
Surge margin:

I bM — L '"^surge ~~ ""• given point' /"''•given point Jat constant flow ^ 100

Carter's Rule

Deviation System

Carter's rule plus an experience factor was used to obtain deviation for both the rotor and stator.

The form of Carter's rule that was used is given below:

(A0-i)me V V

Carter's rule deviation (Ml 14 deviation) =

m =0.92 (a/b)2 +0.002/32

a _ distance to m a x i m um camber point from leading edge

b ~ chord

Delta deviation = M114 deviation - actual deviation

248
REFERENCES

1. AIAA-80-1248, "Part Span Variable Inlet Guide Vanes for V/STOL Fan

Thrust Modulation," V. L. Reed and P. W. Schneider, General Electric Co.

2. AIAA-81-1362, "Effect of a Part-Span Variable Inlet Guide Vane on the

Performance of a High-Bypass Turbofan Engine," G. A. Bobula, R. H.

Soeder, and L. A. Burkhardt, NASA LeRC

3. NASA CR-72787, "Evaluation of Range and Distortion Tolerance for High

Mach Number Transonic Fan Stages - Volume I," K. R. Bilwakesh,

General Electric Co., January 1971

4. NACA TN 3959, "Cascade Investigation of a Related Series of 6-PercentThick Guide-Vane Profiles and Design Charts," J. C. Dunavant, Langley

Aeronautical Laboratory, May 1957

5. Technical Proposal FP 78-108, "V/STOL "A" Fan Demonstration Program - Volume 1," Pratt & Whitney Aircraft Group, March 1978

249
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