Karen Kontos, Don Weir, and Dave Ross

Honeywell, Phoenix, Arizona

Quiet High Speed Fan II (QHSF II): Final Report

NASA/CR—2012-217451

August 2012

21–12567
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientific and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Officer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of significant 

scientific and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Karen Kontos, Don Weir, and Dave Ross

Honeywell, Phoenix, Arizona

Quiet High Speed Fan II (QHSF II): Final Report

NASA/CR—2012-217451

August 2012

21–12567

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NAS3–01136, Task number 2
Acknowledgments

This work was funded under NASA Contract NAS3-01136 from the NASA Glenn Research Center. The author would like to express 

thanks for the many useful contributions of the staff at the NASA Glenn Research Center. Special thanks go to John Gazzaniga 

who served as the Technical Monitor, and to Joe Grady who provided support from the Quiet Aircraft Technology program office. 

Significant contributions to this report were made by many individuals, but the author would like to specifically recognize the 

contributions of those who have served as focal points in key technical areas throughout the QHSF II program: Joe Panovsky for 

aeroelastic tool development and design, Nick Nolcheff and Dave Hanson for aerodynamics, Srinivas Chunduru for mechanical 

design, Lysbeth Lieber for rotor-strut interaction analyses, Glen Schroering for rig hardware fabrication and instrumentation, Bill 

Schuster for acoustic analyses, and Bruce Bouldin for aerodynamic installation analyses.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by NASA technical management OR expert reviewer(s).

This report contains preliminary findings, 

subject to revision as analysis proceeds.
TABLE OF CONTENTS

Page

1. ACKNOWLEDGEMENTS 1

2. INTRODUCTION 1

2.1 Motivation 1

2.2 Description of Work 1

2.2.1 Aerodynamic and Mechanical Design 1

2.2.2 Aeroelastic Analysis 2

2.2.3 Acoustic Analysis 2

2.2.4 Blade and Vane Fabrication 2

2.2.5 Rig Modifications 3

3. AEROELASTIC TOOL VALIDATION 3

3.1 TURBO Modeling for QHSF I 3

3.1.1 TURBO Grid 4

3.1.2 Inlet and Exit Profiles 6

3.1.3 QHSF I TURBO Solution 6

3.1.4 Static Deflections 8

3.2 Tip Clearance Sensitivity 10

3.3 Mode Shape Sensitivity 12

3.4 Boundary Condition Sensitivity 16

3.5 Part Speed Geometry Sensitivity 16

3.6 Comparison With QHSF I Data 17

4. EVALUATION OF QHSF I TEST DATA 27

4.1 Evaluation of Performance Differences Between the 18” and 22” Rig Tests

 of the QHSF I 27

4.2 Evaluation of the Acoustic Results of the 22” Rig Test of the QHSF I 30

4.2.1 V072 Validation 31

4.2.2 Broadband Noise Source 33

4.2.3 Comparison With CFD 33

4.3 Evaluation of the Rotor/Strut Interaction 45

5. QHSF II DESIGN 55

5.1 Approach for the QHSF II Design 55

5.2 Rotor Stacking Design of Experiments 56

5.3 Aeroelastic Verification of the Case 14a Rotor Design 65

5.4 Justification for the Use of the TURBO Evaluation of the Case 14a Rotor 75

5.5 Final Rotor Optimization 78

5.6 Stator Design 84

NASA/CR—2012-217451 iii
TABLE OF CONTENTS (Cont)

Page

5.7 SOURCE3D and V072 Studies for QHSF II Stator DOE I 89

5.7.1 Calibration of SOURCE3D With Straight-Lean Stators 90

5.7.2 Comparison of SOURCE3D and V072 at 62 Percent Speed 91

5.7.3 Modification of PREV072 Calculation of YRD 94

5.7.4 Comparison of QHSF I Cases With QHSF II 95

5.7.5 Adjustment for Low Cutoff Ratio 96

5.7.6 Comparison of Rotor Loss Profiles 98

5.8 Final QHSF II Design 98

5.8.1 Aerodynamic Performance 99

5.8.2 Mechanical Performance 104

5.8.3 Aeroelastic Performance 116

5.8.4 Acoustic Performance 116

5.9 Further Revisions to the Stator Design 116

5.10 Modifications to the Baseline II Stators 119

5.11 Analysis of the Rotor-Strut Interaction With the Baseline II and QHSF II Stators 122

6. RIG MODIFICATIONS 128

6.1 Overview 128

6.2 Front Frame 132

6.3 Rotating Group 133

6.4 Rotating Stator Assembly 134

6.5 Safety Review 136

6.6 Rig Speeds 139

6.7 Nozzle Sizing 139

6.8 Instrumentation 140

6.8.1 Accelerometers 140

6.8.2 Boundary Layer Rakes 140

6.8.3 Capacitance Probes 140

6.8.4 Comb Rakes 140

6.8.5 Distortion Rakes 141

6.8.6 Kulites 141

6.8.7 Static Pressures 142

6.8.8 Strain Gages 143

6.9 Distortion Screens 150

6.10 Model Assembly 155

6.11 Modification of Rotating Group After Initial Assembly 158

7. NEW TECHNOLOGY 161

8. SUMMARY AND CONCLUSIONS 162

NASA/CR—2012-217451 iv
TABLE OF CONTENTS (Cont)

Page

8.1 Aeroelastic Tool Evaluation 162

8.2 QHSF I Data Evaluation 162

8.3 QHSF II Design 163

8.4 Rig Modifications 163

9. REFERENCES 165

Appendix

I Instrumentation (17 pages)

NASA/CR—2012-217451 v
LIST OF FIGURES

Page

Figure 1. Flutter Boundary Based on Rig Testing and the Predicted Speed Lines From TURBO. 3

Figure 2. Grid of the QHSF I Used for the TURBO Analyses. 5

Figure 3. Inlet Profile for Total Pressure at 85% Speed. 7

Figure 4. ANSYS Model of Full Blade. 9

Figure 5. Coordinate Systems for Static Blade Deflections - the View Is Radially Inward. 9

Figure 6. The Steady TURBO Analysis of the QHSF I in NASA 22" Rig Size for the 85%

Speed Line Using Actual Blade Shape and Tip Clearance Is Compared to the

Fan Map Extrapolated From 18" Rig Data. 10

Figure 7. Change in Aerodynamic Damping as a Function of Tip Clearance. These Results

Were Obtained From the Analysis of the 18” Rig at 85% Speed for the 2 Nodal

Diameter Forward Traveling Wave. 11

Figure 8. Comparison of the Static Pressure Fields of the Baseline I Fan and the QHSF I.

The Difference in Pressure Across the Blade Tip Is Significantly Larger for the

Baseline I Fan and, as a Result, This Design Is More Sensitive to Changes in Tip

Clearance. These Sections Are Slightly Below the Blade Tip, at 85% Speed,

Near Stall Conditions. 12

Figure 9. The QHSF I ANSYS Airfoil Only Analysis With Root Fixed in All Directions

Calculated a First Mode Frequency of 342 Hz. 13

Figure 10. The QHSF I ANSYS Airfoil and Platform Analysis With the Dovetail Fixed

 in All Directions Calculated a First Mode Frequency of 316 Hz. 13

Figure 11. The QHSF I ANSYS Airfoil and Platform Analysis With the Dovetail Fixed

 in Local Normal Direction Calculated a First Mode Frequency of 304 Hz. 14

Figure 12. The QHSF I ANSYS Airfoil, Platform & Disk Analysis With Disk Cyclic

 Symmetry for the Nodal Diameter = 2 Case Calculated a First Mode

 Frequency of 285 Hz. 14

Figure 13. Results of the TURBO Mode Shape Study Show Little Sensitivity to the

 Assumptions Used for Calculation of the Mode Shapes. 15

Figure 14. The Full Blade Model Showed Little Difference in Aerodynamic Damping

 as Compared to the Airfoil Only Model. 15

Figure 15. Changes to the Inlet and Exit Pressure Profiles Have Only a Minor Effect

 on the Predicted Damping. 16

Figure 16. Effect of Part-Speed Geometry on Steady Solutions and Flutter Boundary. 17

Figure 17. The TURBO Calculation Near Stall for the 100% Speed Line Shows Good

 Convergence on the Aerodynamic Damping. 18

Figure 18. The TURBO Analysis Identified the Nodal Diameter Wave With Minimum

 Damping. 19

Figure 19. TURBO Predicted the 85% Speed Line Instability Point With Reasonable Accuracy. 19

Figure 20. Distribution of Aerodynamic Damping on Blade Surfaces. 20

Figure 21. TURBO Damping Results for the QHSF I on the 75% Speed Line. 21

Figure 22. Summary of Stability Predictions for the 85% and 75% Speed Lines. 23

Figure 23. The TURBO-AE Steady Calculation at 100% Speed Has Been Added to the

 Performance Summary Map. 24

NASA/CR—2012-217451 vi
LIST OF FIGURES (Cont)

Page

Figure 24. The TURBO-AE Analysis Shows That the Minimum Damping Occurs

 for the 2 Nodal Diameter Pattern Near Stall at 100% Speed. 25

Figure 25. Adjusting of the TURBO-AE Calculated 100% Speed Line to Match the

 Measured Speed Line at Peak Efficiency Shows That the Correct Flutter

 Prediction Is Maintained. 26

Figure 26. A Comparison of the 18” and 22” Rig Data Shows Differences in the Fan

 Stage Work Performed. 28

Figure 27. Detailed Examination of the 100% Speed Line Shows That the 18” QHSF I

 Reached a Higher Choked Mass Flow Than the 22” QHSF I. 29

Figure 28. Results of the 22" Rig Testing Showing Dramatic Differences in Noise Levels for

 a 1500 ft Fly Over at Matched Thrust Conditions From the Three Fan Configurations. 31

Figure 29. At Supersonic Tip Speeds, the Primary Noise Reduction Was in the Blade

 Passage Tone (13831 Rig RPM, 131 Degrees From the Inlet). 31

Figure 30. Comparison With Measured Narrow Band Data Shows That V072 Underestimated

 the Tone Noise Reduction at 2x and 3x the Blade Passage Tone at 55.5% Speed

 in the Forward Arc. 32

Figure 31. Comparison With Measured Narrow Band Data Shows That V072 Underestimated

 the Tone Noise Reduction at 2x and 3x the Blade Passage Tone at 55.5% Speed

 in the Aft Arc. 32

Figure 32. Unknown Broadband Noise Source Must Be Identified and Eliminated for

 Redesign of the QHSF (Data at 61 Degrees With Barrier). 34

Figure 33. LDV Axial Velocity Data Taken Downstream of the QHSF I Show Flow

 Separation at Low RPM That Is Reduced at Higher Values of RPM. 35

Figure 34. Location of the LDV Planes Relative to the Trailing Edges of the Baseline I

 and QHSF I Rotors. 35

Figure 35. The Wake Structure at the LDV Plane for the Baseline I and QHSF I Rotors,

 at 81.4% and 90.1% Corrected Fan Speed. 36

Figure 36. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 81.4% Corrected

 Fan Speed. 37

Figure 37. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 90.1% Corrected

 Fan Speed. 40

Figure 38. A Good Comparison Is Seen Between the Measured and Calculated Rotor

 Wakes for the Baseline I Fan at a Typical Cutback Takeoff Condition. 43

Figure 39. A Good Comparison Is Seen Between the Measured and Calculated Rotor

 Wakes for the Baseline I Fan at a Typical Full Power Takeoff Condition. 44

Figure 40. A Good Comparison Is Seen Between the Measured and Calculated Rotor

 Wakes for the QHSF I at a Typical Cutback Takeoff Condition. 44

Figure 41. A Good Comparison Is Seen Between the Measured and Calculated Rotor

 Wakes for the QHSF I at a Typical Full Power Takeoff Condition. 45

Figure 42. Acoustic Modal Measurements in the Aft Fan Duct With a Rotating Rake

 Show Significant Rotor Strut Tones for the Baseline I Fan. 46

Figure 43. CFD Models for the Rotor/Strut Interaction Study Modeled 26 Vanes, 5 Struts,

 and the Split Flow Path. 47

NASA/CR—2012-217451 vii
LIST OF FIGURES (Cont)

Page

Figure 44. A Periodic Boundary Condition Was Used to Model the Total 360 Degree Flowfield. 48

Figure 45. An Unstructured Grid Was Used to Model the Rotor/Strut Interaction Flowfield

 (QHSF I). 48

Figure 46. Pressure Coefficient Contours Between the Rotor and Stator Have Been

 Produced From the Fluent® CFD Analysis for the Baseline I Fan. 49

Figure 47. Pressure Coefficient Contours Between the Rotor and Stator Have Been

 Produced From the Fluent® CFD Analysis for the QHSF I. 50

Figure 48. Data Planes Were Selected for Comparison of Circumferential Static Pressure

 Profiles (Baseline I Fan). 50

Figure 49. Data Planes Were Selected for Comparison of Circumferential Static Pressure

 Profiles (QHSF I). 51

Figure 50. Circumferential Cuts Show the Relative Positions of the Baseline I and QHSF I

 Vanes at Various Radii (QHSF I Shown in Black). 51

Figure 51. Data Comparisons Emphasize Two Radii Near the Vane Shroud (QHSF I Shown). 52

Figure 52. Comparison of Circumferential Pressure Distributions at R-0.375m (QHSF I

 Data Shifted in Angle and Level to Align With Baseline I Data). 53

Figure 53. Comparison of Circumferential Pressure Distributions at R=0.350m (QHSF I

 Data Shifted in Angle and Level to Align With Baseline I Data). 54

Figure 54. An Interdisciplinary Process Has Been Defined for the Design of the QHSF II. 56

Figure 55. Four Parameters Define the Rotor Blade Stacking for DOE 1. 57

Figure 56. A Go-Forward Blade Stacking Has Been Defined to Meet All Acoustic,

 Mechanical, Aerodynamic, and Aeroelastic Requirements. 62

Figure 57. The Bird Strike Criterion Put a Restriction on the Design Space for Rotor DOE 1. 62

Figure 58. The Strength of the Shock in Front of the Blade Leading Edge at 90% Span

 Was Shown to Decreased With Increasing Forward Sweep. 63

Figure 59. Comparison of the Go-Forward Rotor Blade Design to the QHSF I. 63

Figure 60. QHSF II, Flutter Boundaries for DOE 1 Cases in Baseline II Tested Map. 64

Figure 61. The Logarithmic Sum of the Sound Power Levels for the Rotor-Stator Interaction

 Noise at 89% Speed Shows Significant Variation for Range of DOE Parameters. 65

Figure 62. Steady-State Solutions for Pressure Ratio (pr) and Mass Flow Rate (m) at 89%

 Speed for the Case 14a Rotor Show a Consistent Trend. 66

Figure 63. Damping Extrapolation as a Function of Mass Flow Rate (m) Is Shown at 89%

 Speed for the Case 14a Rotor. 66

Figure 64. Damping Extrapolation With Pressure Ratio (pr) Is Shown at 89% Speed

 for the Case 14a Rotor. 67

Figure 65. Frequency Sensitivity: Damping Extrapolation With Mass Flow Rate (m)

 Is Shown at 89% Speed for the Case 14a Rotor. 68

Figure 66. Frequency Sensitivity: Damping Extrapolation With Pressure Ratio (pr)

 Is Shown at 89% Speed for the Case 14a Rotor. 68

Figure 67. Twist-to-Flex Ratio Sensitivity: Damping Extrapolation With Mass Flow

 Rate (m) Is Shown at 89% Speed for the Case 14a Rotor. 69

Figure 68. Twist-to-Flex Ratio Sensitivity: Damping Extrapolation With Pressure Ratio

 (pr) Is Shown at 89% Speed for the Case 14a Rotor. 70

NASA/CR—2012-217451 viii
LIST OF FIGURES (Cont)

Page

Figure 69. Analysis of the Case 14a Rotor Blade With TURBO Shows That the Blade

 Is Unstable Near the Operating Line at an Intermediate Speed Condition. 71

Figure 70. The Steady Flow Results From Inviscid TURBO Show the Expected Change

 in Pressure Ratio and Flow Characteristics From the Viscous TURBO Results. 71

Figure 71. A Satisfactory Convergence History Was Obtained With the Viscous TURBO

 Analysis. 72

Figure 72. The Minimum Damping for the Inviscid Analysis of the 70% Speed Line

 Occurs Between 0 and 2 Nodal Diameters. 72

Figure 73. The Inviscid TURBO Results Show a Slight Improvement in Stability Over

 the Viscous Results. 73

Figure 74. Mach Number Contours in the Blade Passage Show a Region of Flow

 Separation in the Case 14a Rotor Design. 74

Figure 75. Mach Number Contours Near the Blade Suction Surface Show a Region

 of Flow Separation in the Case 14a Rotor Design. 74

Figure 76. The Aerodynamic Damping as a Function of Mass Flow Rate (Mass) Shows

 the Small Effect of the Separation Region on the Blade Stability. 75

Figure 77. Evaluation of the Reduced Frequency of the Case 14a Rotor Blade Does Not

 Identify the Instability at 70% Speed. 77

Figure 78. Evaluation of the Reduced Frequency of the Case 14a Rotor Blade Does Not

 Identify the Instability at 70% Speed. 78

Figure 79. Comparison of the ADPAC Computational Grid for the Split-Stator

 Configuration With the Streamline Pattern. 79

Figure 80. A Comparison of the TURBO and ADPAC Mach Number Contours Shows

 No Significant Flowfield Differences. 80

Figure 81. A Comparison of the TURBO and ADPAC Mach Number Contours Shows

 No Significant Flowfield Differences. 81

Figure 82. Preliminary ADPAC Results Show That the Case 14a Rotor Has the Potential

 to Meet the Design Point Flow and Pressure Ratio by Adjusting the Mean

 Line Angle Distribution. 82

Figure 83. The ITER07 Case Shows Mode 2/3E Crossing at 100% RPM. 83

Figure 84. A Design Study Completed to Restore Adequate Frequency Margin Suggested

 Sloped Attachment Solution. 83

Figure 85. The Campbell Diagram for the ITER07 Rotor Blade With Sloped Attachment

 Shows Adequate Design Margin. 84

Figure 86. Comparison of the Stator Lean Profiles (YSD Parameter) for the 15 Cases

 of Stator DOE 1. 85

Figure 87. These Four Circumferential Lean Distributions Were Used for Stator DOE 2. 85

Figure 88. The Design Approach for Stator DOE 2 Is to Apply Non-Linear Sweep to

 Take Maximum Advantage of the Optimum Lean. 86

Figure 89. These Three Axial Sweep Distributions Were Used for Stator DOE 2. 87

Figure 90. Summary of the Total Rotor/Stator Interaction Tone Sound Power Reduction

 at a Typical Aircraft Approach Condition Due to the Elements of the QHSF II. 88

NASA/CR—2012-217451 ix
LIST OF FIGURES (Cont)

Page

Figure 91. The Go-Forward Design for the QHSF II Stator Is a Full Span Vane With

 Nonlinear Axial Sweep and Tangential Lean. 88

Figure 92. Preliminary Mode Shapes for the QHSF II Stator Vane Show a Potential

 Mode 2 Flutter Problem. 89

Figure 93. A Split-Stator Fan Flow Path Was Used for the Initial Stator DOE Tone

 Noise Calculations. 89

Figure 94. SOURCE3D Results for QHSF II at 62% Speed, With Straight-Leaned Stators. 90

Figure 95. SOURCE3D Results for QHSF II at 77% Speed, With Straight-Leaned Stators. 91

Figure 96. SOURCE3D Results for QHSF II at 89% Speed, With Straight-Leaned Stators. 91

Figure 97. V072 Results for QHSF II at 62% Speed, With Straight-Leaned Stators. 92

Figure 98. Comparison of the SOURCE3D Results Relative to Unleaned Stator. 93

Figure 99. Comparison of V072 Results Relative to Unleaned Stator. 93

Figure 100. V072 Results Based on the Revised Specification of YRD. 94

Figure 101. Comparison of V072 Results Relative to Unleaned Stator, With Revised

 Specification of YRD. 95

Figure 102. Comparison of the QHSF I and QHSF II Cases With the Baseline I

 and Baseline II Cases. 96

Figure 103. Comparison of QHSF I and QHSF II Cases Adjusted for Low Cutoff Ratio. 97

Figure 104. Variation in PWL for QHSF II With Stator Lean at 62% Speed, Adjusted

 for Low Cutoff Ratio. 97

Figure 105. Radial Loss Distributions From AXCAPS. 98

Figure 106. QHSF II Reduces Noise Through Rotor Sweep and Stator Sweep and Lean. 99

Figure 107. The QHSF II Rotor Meets Pressure Ratio and Efficiency Goals Set for the Program. 99

Figure 108. The QHSF II Stage Meets Pressure Ratio and Efficiency Goals Set for the Program. 100

Figure 109. The Results of the ADPAC Analyses Show the Differences in Mach Number

 Contours on the Suction Side of the Blade Between the Baseline II and

 QHSF II Rotor. 100

Figure 110. The Results of the ADPAC Analyses Show the Differences in Mach Number

 Contours at the Rotor Exit Between the Baseline II and QHSF II Rotor. 101

Figure 111. The Results of the ADPAC Analyses Show the Differences in Mach Number

 Contours on the Suction Side of the Vane Between the Baseline II and

 QHSF II Stator. 101

Figure 112. The Results of the ADPAC Analyses Show the Differences in Mach Number

 Contours at the Vane Exit Between the Baseline II and QHSF II Stator. 102

Figure 113. A Quick Analysis of the Stator Separation Indicated a Small Reduction in

 Efficiency for the QHSF II Stator. 102

Figure 114. Mach Number Contours From the TURBO Viscous Analysis at 89%

 Speed Show That Shock Capture Was Not Achieved for the QHSF II Fan. 103

Figure 115. Mach Contours From the TURBO Viscous Analysis Show the Shock

 Positions for the 89% Speed Condition. 104

Figure 116. A Sloped Attachment Was Designed for the QHSF II Rotor. 104

Figure 117. Honeywell Applied State-of-the-Art Finite Element Modeling Techniques

 in the Mechanical Analysis of the QHSF II Rotor Blade and Attachment. 105

NASA/CR—2012-217451 x
LIST OF FIGURES (Cont)

Page

Figure 118. QHSF II Fan Blade Stress Levels at Aerodynamic Design Point Are Within

 Honeywell Design Experience. 106

Figure 119. QHSF II Fan Blade Attachment Stress Levels at Aerodynamic Design Point

 Are Within Honeywell Design Experience. 106

Figure 120. Disk Analysis at Aerodynamic Design Point Show Principal Stress Well

 Below NASA Burst Margin Criteria. 107

Figure 121. The Blade Weight and Center of Gravity (CG) Location Relative to the

 Intersection of the Axis of Rotation and the Stacking Axis Are Provided

 for Rig Structural and Dynamic Analyses. 107

Figure 122. The Campbell Diagram for the QHSF II Rotor Shows Adequate Frequency

 Margin for the Three Primary Vibration Modes. 108

Figure 123. The QHSF II Rotor Blade Has a Complex Vibration Modal Structure. 109

Figure 124. The Direction for the Maximum Deflections of the QHSF II Rotor

 Is in the Circumferential Direction. 109

Figure 125. The Analysis of the QHSF II Rotor Blade With NOSAPM Shows

 That the Blade Will Meet the Bird Strike Criteria. 110

Figure 126. QHSF II Rotor – Partial Assembly. 111

Figure 127. QHSF II Fan Blade – Comparison of Holography and Finite Element Analysis. 113

Figure 128. The Campbell Diagram for the QHSF II Vane Shows That No Vibration

 Issues Are Expected. 114

Figure 129. The First Six Mode Shapes for the QHSF II Rig Vane Were Calculated

 With ANSYS®. 115

Figure 130. Viscous TURBO Results Show Reduced Risk for 4.5 Degree Sloped

 Attachment. 116

Figure 131. The Stator Separation Was Reduced Through Increased Stator Lean

 and Metal Angle Changes. 117

Figure 132. The QHSF II Stator Redesign Maintained Good Throat Area Margin. 117

Figure 133. Mach Number Contours on the Vane Suction Surface Show the Reduction

 of the Flow Separation Near the Shroud. 118

Figure 134. Trailing Edge Mach Contours Show Reduced Stator Separation for the Q2a Design. 118

Figure 135. The Results of the CFD Analysis Indicate an Improvement of 0.3% in Peak

 Efficiency. 119

Figure 136. The Baseline II Vane Has Less Circumferential Lean Than the QHSF II Vane. 120

Figure 137. The Baseline II Vane Has a Very Similar Loading Distribution as the QHSF II

 Vane at the Design Point. 120

Figure 138. QHSF II Stator and Baseline II Stator Have Nearly Identical Pressure Ratio. 121

Figure 139. QHSF II Stator and Baseline II Stator Have Nearly Identical Efficiency. 121

Figure 140. A Baseline II Stator Set Was Designed and Fabricated to Match the QHSF II

 Rotor for Study of Rotor/Strut Interaction Effects. 122

Figure 141. Four Stator/Strut Clocking Positions Were Analyzed for Both the Baseline II

 and QHSF II Stators. 123

Figure 142. The Flow Path Model for Rotor/Strut Interaction Analyses Consisted of a

 Periodic Sector of 5 Stators and 1 Strut. 124

NASA/CR—2012-217451 xi
LIST OF FIGURES (Cont)

Page

Figure 143. The Unstructured Triangular Surface Mesh, Shown Here Applied to the

 Bypass Strut, Formed the Basis for the Tetrahedral Volume Mesh. 124

Figure 144. Axial and Radial Locations at Which Static Pressure Data Were Processed

 in the Region Upstream of the Stators. 125

Figure 145. Circumferential Static Pressure Distributions at R = 7.48 Inches and X = -5.71

 Inches Show Differences in Pulse Amplitude and Shape With Strut Clocking.

 Unbowed Baseline II Stator Case Is Shown on the Left; QHSF II Stator Case

 Is on the Right. 125

Figure 146. Circumferential Static Pressure Distributions at R = 8.66 Inches and X = -4.92

 Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator

 Case. Also, Differences Are Seen in Pulse Amplitude and Shape With Strut

 Clocking. Unbowed Baseline II Stator Case Is Shown on the Left; QHSF II

 Stator Case Is on the Right. 126

Figure 147. Circumferential Static Pressure Distributions at R = 9.84 Inches and X = -4.92

 Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator Case.

 Also, Differences Are Seen in Pulse Amplitude and Shape With Strut

 Clocking. Unbowed Baseline II Stator Case Is Shown on the Left;

 QHSF II Stator Case Is on the Right. 127

Figure 148. Circumferential Static Pressure Distributions at R = 9.84 Inches and X = -4.13

 Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator

 Case. Also, Pressure Pulse Shapes for the Baseline II Stators Show the

 Influence of the Stator Pressure Pulses More Than QHSF II Stators, at the

 Same Axial Position. Unbowed Baseline II Stator Case Is Shown on the

 Left; QHSF II Stator Case Is on the Right. 128

Figure 149. The Major Differences Between the QHSF I and QHSF II Rigs Are the

 Hub-to-Tip Ratio, Rotating Stator Set, and Frame Modifications. 129

Figure 150. Schematic Diagram of the Rig Installed in the 9x15 Wind Tunnel. 130

Figure 151. The Fan Distortion Measurement Configuration Is Shown for the 22" QHSF II Rig. 131

Figure 152. The Performance Configuration Is Shown for the 22" QHSF II Rig. 131

Figure 153. The Acoustic Configuration Is Shown for the 22" QHSF II Rig. 131

Figure 154. The Front Frame Design for QHSF II Was Modified to Incorporate 5 Comb

 Rakes for Stage Performance Measurements. 132

Figure 155. QHSF II Fan Frame on the NASA 22” Dummy Drive Rig. 132

Figure 156. Changes to the Rotor System Design for QHSF II Include a New Disk,

 Torque Sleeve, and Spinner. 133

Figure 157. The Maximum Stress Level in the Spinner Was Computed to Be 13 ksi. 134

Figure 158. The Rotating Stator Concept Allows for Variable Positioning of the Stator

 Relative to the Struts in the 22” QHSF II Rig, as Well as Facilitate Stage

 Performance Measurements. 135

Figure 159. The Final Rotating Stator Actuation System Uses One Actuator and Horizontal Bar. 135

Figure 160. Average Blade Root and Airfoil Root Section Stress Calculation at 16402 RPM. 137

Figure 161. The Predicted QHSF II Operating Line Shows That the Existing QHSF I

 Nozzle Will Accommodate the QHSF II Fan. 139

NASA/CR—2012-217451 xii
LIST OF FIGURES (Cont)

Page

Figure 162. Boundary Layer Rake Shown in QHSF II Fan Case Plug (18 Degree Location). 140

Figure 163. Aft Looking Forward View of Fan Frame and QHSF II Stator, With One

 Comb Rake Shown. 141

Figure 164. Inlet Distortion Rake Design for QHSF II. 141

Figure 165. QHSF II Fan Case Plug Kulites. 142

Figure 166. Aerodynamic Performance Measurements on QHSF II Fan Rig. 143

Figure 167. Strain Gage Locations Were Defined to Measure 7 Major Vibrational Modes. 144

Figure 168. The Top View of Blade Shows the Strain Gage Locations and the Relative

 Positions of the Engine and Dovetail Axes. 144

Figure 169. The Fifth Strain Gage Is Located on the Blade Dovetail. 145

Figure 170. QHSF II Fan Blade – Strain Gages. 145

Figure 171. A Strain Gage Was Also Mounted on the Fan Disk to Monitor the

 Mechanical Behavior. 146

Figure 172. The Side View of Vane Shows the Strain Gage Locations Relative to the Shroud. 149

Figure 173. QHSF Rig Circumferential Distortion Screens. 151

Figure 174. QHSF Rig Tip Radial Distortion Screen. 152

Figure 175. Left Engine 30-Knot Crosswind Distortion Screen. 153

Figure 176. Right Engine 30-Knot Crosswind Distortion Screen. 154

Figure 177. QHSF II Fan Frame on the NASA 22” Dummy Drive Rig. 155

Figure 178. QHSF II Fan Stator and Rotating Stator Actuation Assembly on the NASA

 22” Dummy Drive Rig. 155

Figure 179. NASA Wind Tunnel Installation of QHSF II Stator Assembly, Shown With

 Outer Fan Case. 156

Figure 180. Aft View of NASA USB Drive Rig/QHSF II During Wind Tunnel Installation. 156

Figure 181. Configuration for Operability Testing Includes Screen Rotator Assembly

 and Long “Tomato Can” Inlet. 157

Figure 182. QHSF II Tunnel Installation, in Aerodynamic Performance Measurement

 Configuration. 157

Figure 183. QHSF II (a) in Far-Field Acoustic Measurement Configuration (b). 158

Figure 184. QHSF II Fan Disk Modified for Larger Bore Diameter to Allow Clearance

 for the Top Hat. 159

Figure 185. QHSF II Aft Spinner Modification Was Made to Allow Clearance

 for the Dynamic Balance Top Hat. 159

Figure 186. QHSF II New Aft Spinner Maximum Principal Stress -- New Design

 Maintains Adequate Margin of Safety. 160

NASA/CR—2012-217451 xiii
LIST OF TABLES

Page

Table 1. Speeds of Interest for NASA Rig. 4

Table 2. Tip Clearances at Speeds of Interest for NASA Rig. 5

Table 3. Inlet Profiles for Total Pressure and Radial Flow Angle at 85% Speed. 7

Table 4. Static Tip Deflections as a Function of Speed and Gas Loads. 10

Table 5. The Summary of Aeroelastic Results Identify the Differences Between

the 18" and 22" Rigs. 27

Table 6. Effect of Geometric Scaling on Hot Blade Shape. Table Lists Displacements at

LE and TE of Tip for Several Loading Alternatives. The Displacements for Each

Rig Size, When Normalized by Tip Radius Are Identical for Corresponding Cases. 30

Table 7. The Available Corrected Fan Speeds for the LDV Measurements and CFD

Analyses Were Matched as Closely as Possible. 43

Table 8. Summary of CFD Models Run in the Rotor/Strut Potential Interaction Study. 47

Table 9. Summary of the Cases for DOE 1. 58

Table 10. Acoustic Y-Factors for DOE 1. 59

Table 11. Aeroelastic Y-Factors for DOE 1. 59

Table 12. Aerodynamic Y-Factors for DOE 1. 60

Table 13. Mechanical Y-Factors for DOE 1. 60

Table 14. Summary of the Sensitivity of the Four Key Parameters to the Rotor Blade Stacking. 61

Table 15. Tone Sound Power Results of Stator DOE 2 From SOURCE3D at 61.7%

 Corrected Fan Speed. 87

Table 16. All of the Mechanical Design Requirements Were Met for the QHSF II. 105

Table 17. QHSF II Fan Blade -- ARS Bench Test Frequencies. 111

Table 18. QHSF II Fan Blade -- Comparison of Finite Element Analysis, Acoustic Ring

 Signature, and Holography Test. 112

Table 19. LCF Life for C-250 Fan Disk. 136

Table 20. New QHSF II Hardware Satisfies NASA Factor of Safety Requirements. 137

Table 21. QHSF II Fan Rotor Margin of Safety Calculation. 138

Table 22. QHSF II Physical and Corrected Design Speed for Key Operating Points. 139

Table 23. The Strain Gages Are Described for the QHSF II Rotor Blade. 143

Table 24. Strain Gage Locations Are Identified for the QHSF II Rotor Blade. 143

Table 25. Summary of Maximum Allowable Strains. 146

Table 26. Allowable Strains for Modes 5 and 6, Gage Location #5. 147

Table 27. Allowable Strains for Modes 1 and 2, Gage Location #1. 147

Table 28. Allowable Strains for Modes 1, 2, and 3, Gage Location #2. 148

Table 29. Allowable Strains for Modes 4, 5, and 6, Gage Location #3. 148

Table 30. Allowable Strains for Modes 5 and 6, Gage Location #4. 149

Table 31. QHSF II Stator Strain Gage Limits. 150

Table 32. Recommended Distortion Screens for the QHSF II Rig Test. 150

Table 33. Summary of Each Innovation and Applicable Section of This Report. 161

NASA/CR—2012-217451 xiv
REVOLUTIONARY AERO-SPACE ENGINE

RESEARCH (RASER)

TASK ORDER NO. 2

QUIET HIGH SPEED FAN II (QHSF II)

FINAL REPORT

1. ACKNOWLEDGEMENTS

This work was funded under NASA Contract NAS3-01136 from the NASA Glenn

Research Center. The author would like to express thanks for the many useful contributions of

the staff at the NASA Glenn Research Center. Special thanks go to John Gazzaniga who served

as the Technical Monitor, and to Joe Grady who provided support from the Quiet Aircraft

Technology program office. Significant contributions to this report were made by many

individuals, but the author would like to specifically recognize the contributions of those who

have served as focal points in key technical areas throughout the QHSF II program: Joe

Panovsky for aeroelastic tool development & design, Nick Nolcheff and Dave Hanson for

aerodynamics, Srinivas Chunduru for mechanical design, Lysbeth Lieber for rotor-strut

interaction analyses, Glen Schroering for rig hardware fabrication and instrumentation, Bill

Schuster for acoustic analyses, and Bruce Bouldin for aerodynamic installation analyses.

2. INTRODUCTION

2.1 Motivation

Honeywell Engines, Systems & Services (Honeywell), building on the technology

developed under the Quiet High Speed Fan contract (AOI 14, NAS3-27752) (Reference 1), has

designed and fabricated an advanced forward swept fan with the following characteristics:

1. Reduced noise at supersonic tip speeds, in comparison to current state-of-the-art fan technology

2. Improved aeroelastic stability within the anticipated operating envelope

3. Aerodynamic performance consistent with current state-of-the-art fan technology

2.2 Description of Work

2.2.1 Aerodynamic and Mechanical Design

Honeywell has completed a new design for the QHSF II (lower hub/tip ratio and higher

specific flow than the Baseline I fan). Aerodynamic and mechanical design studies were

conducted to allow for study of alternative fan blade and vane geometry.

Honeywell conducted an analytical Design of Experiments (DOE) of potential blade and vane

modifications with the goal of identifying a design that provides improved vibration response

while maintaining or improving aerodynamic and acoustic performance. The DOE included:

modifying the incidence of the forward-swept fan blades to improve aeroelastic response,

examining the sensitivity in performance with reduced levels of forward sweep at the rotor LE tip,

and modifying the geometry of 3-D aft-swept vanes including removal of scallop to improve

performance and reduce noise. A design was selected based on the results of the DOE study.

NASA/CR—2012-217451 1
Detailed aerodynamic analysis was performed at design and off-design speeds, using 2-D

axisymmetric streamline curvature analysis of the overall fan stage (rotor, stator, front frame), and

3-D viscous flow analyses for both isolated rotor and stator airfoils as well as combined stage

designs. Mechanical analyses of the redesigned fan blade attached to the existing disk and stator

vane was performed. An analysis of the rotor-stator-strut interaction was performed to assess the

risk of strut-induced fan rotor forced vibration. A comparison of the predicted aerodynamic

performance of the Quiet High Speed Fan II (QHSF II) to the original Quiet High Speed Fan

(QHSF I) and the corresponding baseline fans was performed.

Honeywell performed CFD analysis of the Baseline II and QHSF II stator & strut

configurations at different stator clocking angles to determine the influence of stator

configuration on the rotor flow field. The aerodynamic design of the Baseline II stator was

modified as necessary to efficiently run with the QHSF II rotor.

2.2.2 Aeroelastic Analysis

Prior to the initiation of the QHSF II design and in collaboration with NASA, an aeroelastic

tool calibration study was conducted using the QHSF I result, which established the best

practices for flutter prediction using the TURBO aeroelastic analysis code. A baseline

aerodynamic solution for the QHSF I using the Average Passage / NASA (APNASA) CFD code

was prepared. A tip clearance sensitivity study was conducted. In collaboration with NASA,

Honeywell conducted a mode shape sensitivity study. NASA conducted a grid sensitivity study

and an inlet/exit boundary condition sensitivity study. Honeywell defined the part speed

geometry for the QHSF I, and NASA performed 3-D unsteady aerodynamic analyses of the

QHSF I along two speed lines (part speed and near design speed) using TURBO. NASA

assessed the impact of rig size differences on the aeroelastic response. In collaboration with

NASA, Honeywell prepared a summary of the tool calibration study results.

In collaboration with NASA, Honeywell conducted an aeroelastic evaluation of the DOE

design modifications and an aeroelastic analysis of the final redesigned configuration, and

provided a comparison of the predicted aerodynamic damping and flutter boundary with the

QHSF I and with the Baseline I.

2.2.3 Acoustic Analysis

Honeywell evaluated the effect of the DOE blade design modifications on the rotor-stator

interaction tone duct mode noise levels at several part speed conditions using SOURCE3D. Predictions

of the interaction tone noise levels of the final design using SOURCE3D/TfaNS, as well as

comparisons to the QHSF I and the Baseline I at the same operating conditions, were also prepared.

2.2.4 Blade and Vane Fabrication

Honeywell generated solid models and detail drawings of the QHSF II fan, stator, and stator

housing rings. Twenty-seven fan blades (22 + 5 spares) were machined from bar stock.

Acoustic ring signature and holography data on the blades to determine natural frequencies and

strain gage locations were produced. Honeywell installed strain gages on the blades using

traditional type strain gages. The wiring was run to dog-bone connectors on the front of the

blades. NASA mated the instrumented blades with the rotor and completed the strain gage

wiring into the existing hub and rotor.

NASA/CR—2012-217451 2
Honeywell fabricated a set of 55 (50 plus 5 spares) stator vanes each for both the QHSF II

and the Baseline II designs. Honeywell also fabricated the stator hub and tip mounting rings and

assembled the stators.

2.2.5 Rig Modifications

Honeywell designed and procured additional 22” rig hardware as required by the new

Baseline II fan design. This hardware included the fan disk, spinner, front frame, and additional

stator assembly hardware. Honeywell made all modifications to existing rig hardware as

necessary to accommodate the new flow path. The front frame of the rig was modified to

accommodate hot wire/film probes to make flow measurements downstream of the vanes.

NASA provided the hot wire/film probes and supporting instrumentation.

3. AEROELASTIC TOOL VALIDATION

3.1 TURBO Modeling for QHSF I

The QHSF I design consists of 22 forward-swept inserted blades with a moderate aspect

ratio and state-of-the-art aerodynamic performance and operability (Reference 1 and 2). During

rig testing in a high-speed wind tunnel, the fan performed well at design speed (100 percent

speed), and was successfully throttled to the stall line. However, large vibratory responses due to

flutter were encountered just above the sea level static (SLS) operating line at several part-speed

conditions. The flutter mode was identified as the fundamental bending mode of the airfoil, in a

2 nodal diameter (ND) forward-traveling wave (FTW) pattern. The experimentally determined

flutter boundary is depicted in Figure 1.

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Rotor Map

TURBO at 100% Speed

TURBO at 85% Speed

TURBO at 75% Speed

100%

85%

75%

SLS Op Line

ALT Op Line

Predicted Stall Line

Measured Flutter

Boundary

Figure 1. Flutter Boundary Based on Rig Testing and the Predicted Speed Lines From

TURBO.

NASA/CR—2012-217451 3
To help understand the reasons for the occurrence of flutter in this design, a detailed

computational assessment was undertaken using the TURBO code developed by Mississippi State

University (Reference 3 and 4) and NASA Glenn Research Center (Reference 5). TURBO is a 3D,

time-accurate unsteady, Reynolds-Averaged Navier-Stokes code with the ability to model rotating

or stationary blade rows, rotor-stator interaction, and blade motion. The approach followed for the

flutter analysis in TURBO consists of introducing the blade motion by physically deforming the

surfaces of the blade as defined by the mode shape at the prescribed vibratory frequency

(Reference 5). Phase lag boundary conditions (Reference 6) are used, which allow any desired

nodal diameter to be calculated with a single passage. The resulting aerodynamic work per cycle is

converted into a more meaningful damping value to determine stability of the blade motion. If the

damping is positive, the motion is stable. If the damping is negative, the motion is unstable and

flutter can occur. In this paper, damping values are reported in terms of critical damping ratio, ζ.

All structural analyses were conducted using the commercial finite element program ANSYS

(Reference 7).

An aeroelastic tool calibration study was conducted using the QHSF I rotor geometry. The

test conditions were at an inlet total pressure of 14.3 psi and an inlet total temperature of 73°F.

Steady and unsteady flow computations were conducted for a total of three speed lines. The

correlation effort included 100 percent, 85 percent, and 75 percent speed lines. Recall that no

flutter was encountered during any testing at design speed, while the rig tests indicated

significant flutter response just slightly toward the stall side of the operating line at the two lower

speeds. Choosing these three speeds allows predictions along both stable and unstable speed

lines to be evaluated. Table 1 summarizes speeds of interest.

Table 1. Speeds of Interest for NASA Rig.

% Speed Corrected

RPM

Physical

RPM

100 15357 15572

85 13053 13236

75 11518 11679

3.1.1 TURBO Grid

The baseline grid for TURBO was generated by MMESH and then smoothed using a Poisson

algorithm in CURVE2. The geometry used for this grid was based on the fully hot (100 percent

speed) airfoil shape. Later the grid was refined to be consistent with the actual speed line analyzed.

MMESH is a grid-generation code developed at NASA Glenn. It takes the airfoil geometry

and generates the three-dimensional computational grids for APNASA code (Reference 8).

CURVE2 is a Honeywell developed code. It takes the shear-H grid from MMESH and creates a

smooth and nearly orthogonal computational grid by solving Poisson equation for TURBO code.

Values for the tip clearance are based on measurements taken during the NASA 22” rig testing.

NASA provided the results, and values of interest are summarized in Table 2. There is significant

skew in the clearance, and modifications were made to the grid generator to allow the actual gap to

be modeled. However, the initial model uses the average of the leading and trailing edge gaps. Note

that the full physical clearance is modeled. Four cells (radially) are used for the tip gap.

NASA/CR—2012-217451 4
Table 2. Tip Clearances at Speeds of Interest for NASA Rig.

Location 85% Speed 75% Speed

Lead Edge 0.046 0.053

Mid Chord 0.033 0.040

Trail Edge 0.015 0.021

Average 0.031 0.037

The grid size was limited to the maximum size that would run on Honeywell workstations.

Parameters that describe the resulting grid are:

• # axial cells (ni)= 121

• # radial cells (nj) = 51

• # tangential cells (nk) = 39

• leading edge (ile) = 32

• trailing edge (ite) = 86

• blade tip (jtip) = 47.

Following the practice recommended by Chen et al. (Reference 9), a utility program was used

to initialize the TURBO solutions by mapping an existing APNASA solution on a fine grid. This

initial solution is for a different speed, and actually for the 18” rig size. But this solution is still a

much better starting point than uniform flow, and so the convergence of the TURBO steady

solution should be much faster. Note that the original aerodynamic grid files and solutions are

flipped relative to the hardware (i.e., direction of rotation is opposite). To ensure consistency

between the CFD grid and the FEM mesh, the TURBO grid and initial solution were flipped after

the mapping procedure using a second utility program. This consistency of orientation (or rotation

direction) is required to allow mapping of mode shapes and other quantities between the

aerodynamic and mechanical domains. This final grid is shown in Figure 2.

Figure 2. Grid of the QHSF I Used for the TURBO Analyses.

NASA/CR—2012-217451 5
3.1.2 Inlet and Exit Profiles

The inlet and exit profiles of pressure, temperature, and flow angles which define the

boundary conditions for the steady CFD analyses are based on a combination of NASA rig

measurements, Honeywell rig measurements, and calculations from APNASA. The stations for

the inlet profile are a combination of spans from the NASA measurements and spans in the

APNASA input. The NASA data provides stations near the tip to specify the total pressure

losses, while APNASA provides those near the hub to account primarily for changes in radial

flow angle. The transition occurs just above 80 percent span.

The results from measurements in the NASA rig provided the inlet total pressure profile

near the tip, accounting for losses due to the inlet (bellmouth, etc). At the hub, no measurements

were available, so a computational estimate of the boundary layer as calculated by APNASA for

the analysis of the 18” rig is used. This hub boundary layer extends 1 percent into the flow field.

Two additional stations, at 0.5 percent and 1 percent span, were added to the APNASA stations

to permit this profile to be specified. At other spans, a constant value for total pressure was

assumed.

The inlet total temperature and tangential flow angle were assumed to be constant. The

radial flow angle was obtained from APNASA input. All the stations from the APNASA input

were used up to about 80 percent span, since the angle is nonzero up to this span. A value of 0.5

degrees was specified at 86.61 percent span to act as a transition. Beyond this span, the radial

angle was specified to be zero, and the span locations correspond to the NASA measurements.

The exit profile for the static pressure was based on work done earlier for the 18” rig.

Ideally, there would be measurements from the NASA tests, but no data was available between

the rotor and the stator. Data at the hub and case was available from the 18” rig tests. It was

assumed that this profile would be the same for both rigs. The work conducted previously

determined that the ratio of the tip static pressure to the hub static pressure was 1.1635. This was

found to be in reasonably good agreement with a TURBO analysis run using the assumption of

radial equilibrium. However, the experimental value was used for all TURBO analyses of the

NASA rig. These data are summarized in Table 3 and Figure 5.

3.1.3 QHSF I TURBO Solution

Most of the background information for the TURBO analysis has been specified previously.

However, key features are noted here. The initial TURBO steady analysis was set up following

Honeywell current best practices. This case provided a baseline for determining sensitivities to

alternative modeling approaches.

This first set of analyses focused on the 85 percent speed line, or 13236 RPM (physical). A

constant tip clearance of 0.031” was used, corresponding to 85 percent speed. Exit conditions

for the first analysis were arbitrarily selected to have the hub static pressure equal to the average

inlet total pressure. The tip static pressure was then determined by multiplying this value by

1.1635, to provide a profile consistent with experimental measurements. The initial flow field

was specified from a mapping of an existing APNASA solution. The TURBO grid and initial

flow field were then flipped to align the CFD grid with the hardware and the FEM mesh.

NASA/CR—2012-217451 6
Table 3. Inlet Profiles for Total Pressure and Radial Flow Angle at 85% Speed.

85% Speed

Inlet Radial

Radius Span Pt (norm) Pt (psi) Pt (Pa) Angle

1.900 0.0000 0.9434 13.491 93014.6 25.9995

0.0050 0.9853 14.090 97145.7 25.5192

0.0100 1.0000 14.300 98595.0 25.0388

0.0837 1.0000 14.300 98595.0 17.9586

0.1480 1.0000 14.300 98595.0 14.4816

0.1963 1.0000 14.300 98595.0 12.5103

0.2447 1.0000 14.300 98595.0 10.8811

0.2803 1.0000 14.300 98595.0 9.8474

0.3095 1.0000 14.300 98595.0 9.0772

0.3353 1.0000 14.300 98595.0 8.4482

0.3613 1.0000 14.300 98595.0 7.8559

0.4564 1.0000 14.300 98595.0 5.9715

0.5181 1.0000 14.300 98595.0 4.9274

0.5745 1.0000 14.300 98595.0 4.0614

0.6267 1.0000 14.300 98595.0 3.3188

0.6754 1.0000 14.300 98595.0 2.6660

0.7215 1.0000 14.300 98595.0 2.0797

0.7652 1.0000 14.300 98595.0 1.5506

0.8070 1.0000 14.300 98595.0 1.0716

9.666 0.8661 1.0000 14.300 98595.0 0.5000

9.804 0.8815 0.9999 14.299 98585.2 0.0000

9.946 0.8973 0.9998 14.297 98575.3 0.0000

10.099 0.9144 0.9996 14.294 98555.6 0.0000

10.257 0.9320 0.9993 14.290 98526.0 0.0000

10.398 0.9477 0.9990 14.286 98496.5 0.0000

10.552 0.9649 0.9834 14.063 96958.4 0.0000

10.710 0.9825 0.9461 13.529 93280.8 0.0000

10.788 0.9912 0.9302 13.302 91713.1 0.0000

10.828 0.9957 0.9140 13.070 90115.9 0.0000

10.867 1.0000 0.8800 12.584 86763.6 0.0000

QHSF Inlet Total Pressure Profile for NASA Rig

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

0.85 0.90 0.95 1.00 1.05

Total Pressure

Span (%)

Figure 3. Inlet Profile for Total Pressure at 85% Speed.

NASA/CR—2012-217451 7
TURBO in the steady mode was then used to converge the flow field to the new boundary

conditions. This solution used 1000 iterations at a CFL number of 1000, utilizing the k-ε

turbulence model. The solution converged well, but there were some (very) minor oscillations in

the mass flow versus iteration that was damped out. The overall results from the solution are:

( )( )

9156.0

501.1 9707.0

457.1 p

s/lbm53.87042.12.2

2

1749.381869.38

m

R

corr

=η

= =

⎟ = ⎠

⎞ ⎜

⎝

⎛ + & =

(1)

3.1.4 Static Deflections

The tip deflections at speeds of interest were calculated for the QHSF I in NASA rig size

(22” diameter). These deflections were determined using nonlinear, large deflection static

analysis in ANSYS. The model, shown in Figure 4, includes pressure loads and temperatures

based on earlier CFD analysis at the Aerodynamic Design Point (ADP), or 100 percent speed.

Note that the gas loads were determined for an inlet total pressure of 12.5 psi, corresponding to

the Honeywell rig test, while the inlet total pressure in the NASA rig was 14.3 psi. These

calculations may be repeated with gas loads that reflect the true inlet pressure and the actual

speed, and perhaps even the position along the speed line. The results presented here are best

taken as an indication of the importance of each of these influences.

The deflections at the blade tip are summarized in Table 4, with the coordinate systems

shown Figure 5. The values tabulated are averages for the tip section of deflections along the

chord (uξ), normal to the chord (uη), and rotation (φ). These average values are based on the

displacement of the nodes at the leading and trailing edges. For reference, the tip true chord is

4.586 inches, and the tip stagger (as measured from the axial direction) is -58.9°.

A total of 5 cases were run. First, the effect of the gas loads was isolated. Then, analyses at

100 percent and 85 percent speed were conducted, both with and without gas loads. From the

results, note that the stagger increases with speed. The effect of the gas loads is fairly small,

only about +0.2° (at either speed), and acts to reduce the amount of change in the stagger.

Relative to the fully hot shape at 100 percent, the blade rotates +0.4° at the 85 percent speed

condition, and this is in the direction of reducing stagger.

NASA/CR—2012-217451 8
Figure 4. ANSYS Model of Full Blade.

y

φ

x

LE

TE

η

ξ

Figure 5. Coordinate Systems for Static Blade Deflections - the View Is Radially Inward.

A new steady TURBO solution for the 22” rotor blade for part speed analysis was

developed. The solution grid is for the actual blade-shape at 85 percent speed (rather than fully

hot) and includes the skewed tip gap based on the NASA measurements. A grid file, steady

TURBO solutions for fac=1.00, steady input files, unsteady input files, and mode shape files for

several nodal diameter patterns were provided to NASA. Figure 6 shows a comparison of the

calculated 85 percent speed line performance as compared to the measured performance in the

18” rig.

NASA/CR—2012-217451 9
Table 4. Static Tip Deflections as a Function of Speed and Gas Loads.

Phys Speed % Speed Gas Loads uξ (in) uη (in) φ (deg)

0 0 ADP 0.068 -0.177 1.166

15572 100 no -0.090 0.325 -2.720

15572 100 ADP -0.073 0.290 -2.573

13236 85 no -0.084 0.296 -2.390

13236 85 ADP -0.062 0.251 -2.182

1.0

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

2.0

2.1

40.0 50.0 60.0 70.0 80.0 90.0 100.0 110.0

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Final Rotor Map

Stability Line

TURBO 85% - Fully Hot, Avg Gap

TURBO 85% - Part Speed Geom, Skewed Gap

105%

95%

100%

90%

85%

80%

75%

70%

N1c=50% 55% 60% 65%

97.5%

Figure 6. The Steady TURBO Analysis of the QHSF I in NASA 22" Rig Size for the 85%

Speed Line Using Actual Blade Shape and Tip Clearance Is Compared to the

Fan Map Extrapolated From 18" Rig Data.

3.2 Tip Clearance Sensitivity

The importance of tip clearance effects on the predicted stability boundary was quantified

for the QHSF I correlation effort. Related studies on the Baseline I fan indicated a significant

effect of tip clearance on aerodynamic damping, and so it was initially thought to be crucial to

model the gap accurately for the present correlation effort.

During the initial analysis of the QHSF I in the 18” rig size, a detailed parametric study of the

effects of the tip clearance was performed. While the study was conducted at the 18” rig size

rather than the 22” size, the chosen 85 percent speed line is consistent with the analysis for the 22”

rig. Three different tip gaps were used: the nominal physical value of 0.039”, a tighter gap of

0.010”, and a nearly limiting case of 0.002”. Note that the latter two values were chosen arbitrarily

to span the tip gaps of interest. Unexpectedly, this study indicated little effect of the tip gap on

NASA/CR—2012-217451 10
either the steady flow field or the aerodynamic damping. The damping at two steady conditions

along the 85 percent speed line is presented in Figure 7. The drop in damping as the tip gap is

tightened is fairly significant for the peak efficiency condition. There is still a noticeable drop at

near stall conditions, but the change in damping is only slightly more than 0.1 percent. The range

of gaps considered is considerably larger than would actually need to be considered.

A second result is obtained indirectly from the work done to date on the 22” rig. On the

original “100 percent Geometry” analyses, the average tip clearance calculated from

experimental measurements was used. For the subsequent “85 percent Geometry” analyses, the

skewed tip gap as actually measured in the rig was accurately modeled. Results from these sets

of analyses show only minor differences, as has been previously documented. The conclusion is

that the skew in the tip gap has little influence. While this conclusion is somewhat contaminated

by the fact that both the blade geometry as well as the tip gap were changed at the same time, the

chance that each is a significant influence and just happen to cancel each other out is remote.

The conclusion is that the tip gap must be modeled with reasonable accuracy for the QHSF

I, but it is not a strong driver on the blade’s stability. Note that this conclusion does not agree

with the stability trends of other fan blades analyzed recently at Honeywell, such as the Baseline

I fan. The root cause for the difference may lie in the basic geometry of the QHSF I design.

Figure 8 compares the static pressure field for the Baseline I fan and the QHSF I. Both of these

plots are for a radial station near the tip (but below the tip gap) at near stall conditions. As the

plots indicate, the pressure across the blade tip is significantly less in the QHSF I design due to

the shock still being captured in the passage due to the forward sweep. It is this pressure

difference that will be the driving mechanism for tip leakage flows, and determine the

importance of the tip gap. With less of a pressure difference, the QHSF I has a lower sensitivity

to clearance effects.

0.0

0.5

1.0

1.5

0.000 0.010 0.020 0.030 0.040

Clearance Gap (in)

Damping (%)

Peak Efficiency

Near Stall

Figure 7. Change in Aerodynamic Damping as a Function of Tip Clearance. These Results

Were Obtained From the Analysis of the 18” Rig at 85% Speed for the 2 Nodal

Diameter Forward Traveling Wave.

NASA/CR—2012-217451 11
Baseline I QHSF I

Figure 8. Comparison of the Static Pressure Fields of the Baseline I Fan and the QHSF I.

The Difference in Pressure Across the Blade Tip Is Significantly Larger for the

Baseline I Fan and, as a Result, This Design Is More Sensitive to Changes in Tip

Clearance. These Sections Are Slightly Below the Blade Tip, at 85% Speed,

Near Stall Conditions.

3.3 Mode Shape Sensitivity

The sensitivity of the TURBO results to the assumptions used is the mode shape

calculations was evaluated. The first mode vibration shapes with the following hub boundary

conditions were calculated with ANSYS:

• Root fixed in all directions

• Dovetail fixed in all directions

• Dovetail fixed in local normal direction

• Disk cyclic symmetry

The results of each of the analyses are presented in Figure 9 to Figure 12.

NASA/CR—2012-217451 12
Figure 9. The QHSF I ANSYS Airfoil Only Analysis With Root Fixed in All Directions

Calculated a First Mode Frequency of 342 Hz.

Figure 10. The QHSF I ANSYS Airfoil and Platform Analysis With the Dovetail Fixed in

All Directions Calculated a First Mode Frequency of 316 Hz.

NASA/CR—2012-217451 13
Figure 11. The QHSF I ANSYS Airfoil and Platform Analysis With the Dovetail Fixed in

Local Normal Direction Calculated a First Mode Frequency of 304 Hz.

Figure 12. The QHSF I ANSYS Airfoil, Platform & Disk Analysis With Disk Cyclic

Symmetry for the Nodal Diameter = 2 Case Calculated a First Mode Frequency

of 285 Hz.

NASA/CR—2012-217451 14
The effects of changes in mode shape on TURBO aerodynamic damping was examined.

Figure 13 shows the results for the viscous, 85 percent speed, near-stall, 2 nodal diameter

forward travelling wave case using 5 different mode shape models from ANSYS. Very small

differences in the calculated aerodynamic damping are seen. Figure 14 shows that a similar

trend can be seen for calculations at all inter-blade phase angles. During design iterations, it will

be possible to use the simpler airfoil-alone model to calculate mode shapes for TURBO

aeroelastic analysis.

0.12%

0.22%

0.23%

0.22%

0.17%

0.00% 0.10% 0.20% 0.30%

Aero-damping (%)

Baseline blade

model

Linear geometry

effects

Normal dovetail

constraint 

Cyclic-symmetry

analysis

Airfoil-only

model

Figure 13. Results of the TURBO Mode Shape Study Show Little Sensitivity to the

Assumptions Used for Calculation of the Mode Shapes.

Inviscid, 85% Speed, NS, Mode 1

0.00

1.00

2.00

3.00

4.00

5.00

6.00

7.00

8.00

9.00

-11 -9 -7 -5 -3 -1 1 3 5 7 9 11

Nodal Diameter

Damping %

Airfoil only

Blade only

Figure 14. The Full Blade Model Showed Little Difference in Aerodynamic Damping as

Compared to the Airfoil Only Model.

NASA/CR—2012-217451 15
3.4 Boundary Condition Sensitivity

Because previous studies have shown a high sensitivity to inlet total pressure profile effects

and especially the shroud inlet boundary layer (Reference 10), care was taken in the analysis to

accurately capture this profile and account for changes with speed. To determine the

significance for the present fan design, the analysis was re-run with a uniform inlet profile and

also by specifying radial equilibrium at the exit rather than the experimentally-obtained exit

profile of static pressure. In Figure 15, the dashed line indicates the damping at different

positions along the 85 percent speed line using the experimental profiles. The triangles denote

solutions with other assumptions for the inlet and exit profiles, and are seen to lie very near the

dashed line. These results indicate that the profile changes do affect the damping to a small

degree, but are attributed primarily to the change in position on the fan map.

0.000

0.002

0.004

0.006

0.008

0.010

0.012

0.014

0.016

76 78 80 82 84 86 88

Inlet corrected flow

Inlet/Exit profile

Other boundary conditions

fac=1.01

(peak-eff)

fac=1.08

(near-stall)

uniform inlet

radial equilibrium

uniform inlet & 

radial equilibrium

Figure 15. Changes to the Inlet and Exit Pressure Profiles Have Only a Minor Effect on

the Predicted Damping.

3.5 Part Speed Geometry Sensitivity

One detail of the analysis that was initially thought to be crucial for accurate predictions of

the flutter boundary was the use of part speed geometry. The airfoil geometry was updated at

each speed to reflect the actual deflections due to the speed and other loads. This refinement is

in contrast to simply using the fully hot design shape for all speeds. To evaluate whether this

refinement is necessary, the analysis at 85 percent speed was repeated using the fully hot (100

percent speed) geometry. Results demonstrate that the effect on the damping calculation is small

over the entire ND range. The fan map shown in Figure 16 indicates that the change in steady

solutions is also fairly minor, while the resulting extrapolation of the flutter boundary is slightly

affected but not substantially. The impact of using part-speed geometry is minimal and is of

importance only in cases that require the very highest accuracy in the prediction of the flutter

boundary, such as a final design verification or a correlation effort.

NASA/CR—2012-217451 16
Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Rotor Map

TURBO at 85% Speed - Fully Hot Geometry

TURBO at 85% Speed - Partspeed Geometry

100%

85%

75%

SLS Op Line

ALT Op Line

Predicted Stall Line

Measured Flutter

Boundary

Predicted Flutter

for Fully Hot Geometry

Predicted Flutter

for Partspeed Geometry

Figure 16. Effect of Part-Speed Geometry on Steady Solutions and Flutter Boundary.

3.6 Comparison With QHSF I Data

A TURBO part-speed geometry analysis, with the 85 percent speed geometry and skewed

tip gap, was completed for 85 percent speed line for the QHSF I rotor. Figure 17 shows the

convergence of aerodynamic damping with cycle count for the near-stall condition for various

nodal-diameter modes. It can be seen from the picture that all modes were adequately

converged.

Figure 18 shows a plot of the aerodynamic damping as a function of the nodal diameter

number of the traveling wave. It can be seen that the location of the minimum damping is

consistent with the measured results and the use of the 85 percent vs. 100 percent hot shape

geometry had little effect on the results.

Figure 19 shows the 18”-fan operating map scaled to 22” size, the predicted steady speed

line from TURBO at 85 percent, and the projected location of aeroelastic instability from

TURBO. The predicted instability agrees well with the measured instability line.

To help understand the reasons for the change in damping when approaching stall, the

distribution of damping on the surface of the blade was considered. Figure 20 shows the damping

distribution at the least stable nodal diameter, ND (+2) at 85 percent speed. Correlating this plot

with that for Mach number distribution, we find that regions of significant damping are strongly

tied to shock location. At peak efficiency (PE) conditions, the passage shock runs roughly from

mid-chord of the pressure surface to near the trailing edge of the suction surface of the adjacent

blade. This is demonstrated in Figure 20a by the Mach number contours in the blade-to-blade view

at 90 percent span. (The shock locations have been highlighted in this plot.) From Figure 20b, a

NASA/CR—2012-217451 17
significant region of positive damping is associated with the shock on the pressure surface, while

Figure 20c shows a region of negative damping associated with the shock on the suction surface.

For near-stall (NS) conditions, Figure 20d gives the shock location at 90 percent span. Figure 20e

shows that as the shock moves forward the positive damping region follows it and is now at the

blade leading edge. The suction surface plot in Figure 20f indicates that the de-stabilizing region

has moved forward from the trailing edge to mid-chord, and has grown in strength relative to PE

conditions. It is interesting to note that the loss of stability in throttling from PE to NS is due to

nearly equal changes in damping on each surface. Similar plots for the most stable ND (11), not

included here for brevity, indicate the de-stabilizing regions of Figure 20 become stable, while

stable regions become more stable and larger in extent.

0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

0.10

0 2 4 6 8 10 12 14 16

cycle

0ND_zeta 2ND_zeta

4ND_zeta 6ND_zeta

11ND_zeta -6ND_zeta

Aero-damping

Figure 17. The TURBO Calculation Near Stall for the 100% Speed Line Shows Good

Convergence on the Aerodynamic Damping.

NASA/CR—2012-217451 18
0.000

0.002

0.004

0.006

0.008

0.010

0.012

-10 1 23 4 5

no da l dia me te r

ne a r-sta ll 85% ge ome try

ne a r-sta ll 100% ge ome try

Figure 18. The TURBO Analysis Identified the Nodal Diameter Wave With Minimum

Damping.

1.0

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

2.0

2.1

40.0 50.0 60.0 70.0 80.0 90.0 100.0 110.0

Inlet Corrected Flow (lbm/s)

Ro

tor

Pre

ss

ure

Rat

io

Final Rotor Map

Stability Line

TURBO 85% - Fully Hot, Avg Gap

TURBO 85% - Part Speed Geom, Skewed Gap

Flutter Boundary for Fully Hot Geom

Flutter Boundary for Part Speed Geom

105%

95%

100%

90%

85%

80%

75%

70%

N1c=50% 55% 60% 65%

97.5%

Figure 19. TURBO Predicted the 85% Speed Line Instability Point With Reasonable

Accuracy.

NASA/CR—2012-217451 19
 (a) PE, Mach at 90% Span (b) PE, Pressure Surface (c) PE, Suction Surface

 (d) NS, Mach at 90% Span (e) NS, Pressure Surface (f) NS, Suction Surface

Figure 20. Distribution of Aerodynamic Damping on Blade Surfaces.

The QHSF I rotor aeroelastic behavior at 75 percent speed was calculated with TURBO.

Inviscid analysis was used to screen the important inter blade phase angles (IBPA) and modes.

Viscous analyses were performed at the IBPAs where the damping was low in the inviscid

analysis. Results are shown in Figure 21 for both the peak efficiency (PE) and near stall (NS)

points. A reasonable prediction of flutter boundary was obtained, though not as good as 85 percent

speed. Mode 1 was predicted as critical, but mode 3 damping was always extremely low.

Stabilizing De-Stabilizing

NASA/CR—2012-217451 20
0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.09

-11 -9 -7 -5 -3 -1 1 3 5 7 9 11

Nodal Diameter

Damping

PE, Inviscid

NS, Inviscid

PE, Viscous

NS, Viscous

(a) Mode 1 (first bending)

0.0000

0.0001

0.0002

0.0003

0.0004

0.0005

0.0006

0.0007

-11 -9 -7 -5 -3 -1 1 3 5 7 9 11

Nodal Diameter

Damping

PE, Inviscid

NS, Inviscid

PE, Viscous

NS, Viscous

(b) Mode 3 (first torsion)

Figure 21. TURBO Damping Results for the QHSF I on the 75% Speed Line.

Figure 22 provides a summary of the stability line calculations for the two part speed cases.

The 100 percent speed line performance map was generated with TURBO-AE by the NASA

Glenn staff. The calculated map differences are consistent with differences observed for stage data

between the 22-inch and 18-inch scale rigs as shown in Figure 23. These results are consistent

with the previous analysis from the 75 percent and 85 percent speed lines and Honeywell’s

experience. This difference in not considered an issue considering the analysis accuracy.

NASA/CR—2012-217451 21
The flutter prediction for 100 percent speed line was completed with TURBO-AE by the

NASA Glenn staff. Mode 1 analyzed for several nodal diameter patterns. It was determined that

the 2 nodal diameter pattern was least stable near stall, whereas the 0 nodal diameter was least

stable near the peak efficiency point as shown in Figure 24. Extrapolation of the data showed

that the predicted flutter boundary was beyond the stall line, consistent with the measured data.

To determine if the difference in the predicted vs. measured flow made a difference in the flutter

prediction, the speed line calculated by TURBO-AE at 100 percent speed was shifted to match at

the peak efficiency point (see Figure 25). Even after accounting for the shift in performance

map, the predicted flutter boundary lies beyond stall line.

NASA/CR—2012-217451 22
2.0

1.91.81.71.61.51.41.31.21.11.0

40 50 60 70 80 90 100 110

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Rotor MapSLS Op LineALT Op LinePredicted Stall LineHON Stability LineNASA Stability LineTURBO 85% - Fully Hot Geom with Avg GapTURBO 85% - Partspeed Geom with Skewed GapTURBO 85% - Inviscid with Fully Hot GeomTURBO 75% - Partspeed Geom with Skewed GapTURBO 75% - Inviscid with Partspeed GeomPredicted Flutter Boundary Points

105%

95%

100%

90%

85%

80%

75%

70% 65% 60% 55% N1c=50%

97.5%

Figure 22. Summary of Stability Predictions for the 85% and 75% Speed Lines.

NASA/CR—2012-217451 23
TURBO Analysis of QHSF Flutter

Analysis Based on 22" NASA Rig

2.12.01.91.81.71.61.51.41.31.21.11.0

40 50 60 70 80 90 100 110

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Rotor MapSLS Op LineALT Op LinePredicted Stall LineHON Stability LineNASA Stability LineTURBO 85% - Fully Hot Geom with Avg GapTURBO 85% - Partspeed Geom with Skewed GapTURBO 85% - Inviscid with Fully Hot GeomTURBO 75% - Partspeed Geom with Skewed GapTURBO 75% - Inviscid with Partspeed GeomPredicted Flutter Boundary PointsTURBO 100% - Old gridTURBO 100%

105%

95%

100%

90%

85%

80%

75%

70% 65% 60% 55% N1c=50%

97.5%

Figure 23. The TURBO-AE Steady Calculation at 100% Speed Has Been Added to the Performance Summary Map.

NASA/CR—2012-217451 24
Variation of aero-damping with nodal diameter

0.080.070.060.050.040.030.020.010.00

-11 -9 -7 -5 -3 -1 1 3 5 7 9 11

nodal diameter

aero-damping peak efficiency Near St all

Figure 24. The TURBO-AE Analysis Shows That the Minimum Damping Occurs for the 2 Nodal Diameter Pattern Near

Stall at 100% Speed.

NASA/CR—2012-217451 25
TURBO Analysis of QHSF Flutter

Analysis Based on 22" NASA Rig

2.12.01.91.81.71.61.51.41.31.21.11.0

40 50 60 70 80 90 100 110

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

Rotor MapSLS Op LineALT Op LinePredicted Stall LineHON Stability LineNASA Stability LineTURBO 85% - Fully Hot Geom with Avg GapTURBO 85% - Partspeed Geom with Skewed GapTURBO 85% - Inviscid with Fully Hot GeomTURBO 75% - Partspeed Geom with Skewed GapTURBO 75% - Inviscid with Partspeed GeomPredicted Flutter Boundary PointsTURBO 100% - Old gridTURBO 100%

105%

95%

100%

90%

85%

80%

75%

70% 65% 60% 55% N1c=50%

97.5%

Figure 25. Adjusting of the TURBO-AE Calculated 100% Speed Line to Match the Measured Speed Line at Peak

Efficiency Shows That the Correct Flutter Prediction Is Maintained.

NASA/CR—2012-217451 26
4. EVALUATION OF QHSF I TEST DATA

4.1 Evaluation of Performance Differences Between the 18” and 22” Rig Tests of the QHSF I

An attempt was made to derive a rotor-only performance map of the 22” QHSF I rig, since the

rotor performance was not directly measured in the test. It was thought that data from the 18” rig

could be used to estimate the rotor-only performance from the 22” test. An assessment of the

differences between the QHSF I aerodynamic performance in the Honeywell 18” rig (Reference 1)

and the NASA 22” rig (Reference 11) tests was conducted. This assessment was motivated by the

differences in aeroelastic performance of the 18” and 22” rig tests of the QHSF I as summarized in

Table 5.

Table 5. The Summary of Aeroelastic Results Identify the Differences Between the 18" and

22" Rigs.

Speed HON (18" diam) NASA (22" diam)

50% Mode at 860 Hz (system umbrella mode?) response >1000 

ue. Occurs near stall.

Reached predicted stall line.

55% Mode at 860 Hz (system umbrella mode?) response >600 

ue. Occurs above op line.

Reached stress limits before hitting predicted stall line. No 

component of flutter (all forced response / SFV).

60% Responses in 125-350 ue range in Modes 2 & 3. Reached stress limits before hitting predicted stall line. No 

component of flutter (all forced response / SFV).

65% Mode 2 NSV predominates. Reached stress limits before hitting predicted stall line. No 

component of flutter (all forced response / SFV).

70% Mode 2 NSV up to 900 ue near op line at 737 Hz. Increase 

to 72% speed resulted in rapid onset of flutter in Mode 1, 

with amplitudes exceeding 2000 ue at 350 Hz.

Reached stress limits before hitting predicted stall line. No 

component of flutter (all forced response / SFV).

75% Mode 1 flutter up to 2000 ue on op line in 2 ND FTW. At 

lower pressure ratio, modes 1, 2 and 3 all exhibit moderate 

levels of NSV (125-450 ue).

Mode 3 (926 Hz) NSV identified during data reduction, 3 ND 

FTW. Levels are low, up to 130 ue-SA. Forced response 

from 1/rev up to 200 ue.

80% Mode 1 flutter just above op line, in 2 ND FTW. Levels are in 

500 ue range at steady state data point, higher transiently.

Mode 1 flutter identified during data reduction, 2ND FTW. 

Levels are low, up to 70 ue-SA. Overall signal dominated by 

forced response from 1/rev, up to 180 ue.

85% Mode 1 flutter above op line. 2 ND FTW. Levels reach 700 

ue at 385 Hz at steady state data point, higher transiently.

Mode 1 flutter observed at 308 Hz, up to 497 ue-SA. 2 ND 

FTW. Forced response from 1/rev up to 240 ue.

90% Mode 1 flutter above op line. 2 ND FTW. Reached stress limits before hitting predicted stall line. No 

component of flutter (all forced response / SFV).

95% Mode 1 flutter above op line. 2 ND FTW. Reached predicted stall line.

100% Reached predicted stall line. Reached predicted stall line.

Figure 26 is the full map of the work characteristics of the 18" and 22" QHSF I fan rigs and

Figure 27 is a detail of the 100 percent speed line. The work characteristics are different

between the two fans, which verifies the rotor is setting the choke flow. It is possible that as the

back-pressure was lowered, the rotor work, efficiency, and pressure ratio became low enough to

send the stator enough corrected flow that it choked as well. Unfortunately, the stator choke

behavior cannot be proven from the data.

The performance differences are consistent with aeroelastic differences, which may suggest

a different hot running shape between the 18” and 22” blades. Also, it is noted that for the two

rigs to operate at the same operating line, the 18" rig is running further from choke and most

likely at higher incidence levels. This difference in incidence levels may explain the changes in

aeroelastic behavior.

NASA/CR—2012-217451 27
There does not appear to be an effective way to get to the 22" rotor-only performance by

extracting the stator performance from the overall stage performance. Stator loss buckets are

typically defined based on incidence or inlet corrected flow. Stator loss data is not available for

the 22" rig, and it would be questionable to assume that the 18” and 22” vanes are the same and

back out the rotor from the stage. It would have to be assumed that the loss buckets are a

function of exit corrected flow (instead of inlet conditions) and there is no way to quantify this

error. Also, the stator performance is a function of the span-wise distribution of loss. It is likely

that since the rotor performance is different, the stator inflow was not the same. It would be a

significant effort, using several non-quantified assumptions, to derive a 22" rotor-only

performance map.

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

0.2

0.22

0.24

40 50 60 70 80 90 100 110

Wc (22 in scale)

DTOT

22" Rig Data From NASA

18", speed corrected & scaled to 22"

50 55

55.9 60

65

70

75

80

85

90

95

97.5

100

105

102

Figure 26. A Comparison of the 18” and 22” Rig Data Shows Differences in the Fan Stage

Work Performed.

NASA/CR—2012-217451 28
0.18

0.19

0.2

0.21

0.22

0.23

0.24

0.25

86 88 90 92 94 96 98 100 102

Wc (22 in scale)

DTOT

22" Rig Data From NASA

18", speed corrected & scaled to 22"

Figure 27. Detailed Examination of the 100% Speed Line Shows That the 18” QHSF I

Reached a Higher Choked Mass Flow Than the 22” QHSF I.

In order to resolve the question of geometric scaling and potential “non-linear” effects on the

hot shape deflection pattern of the blade, a full blade ANSYS model was run. Both the 22” and

18” rig sizes were assessed, with the geometry scaled using an available command in ANSYS. All

analyses were at 100 percent speed, with the speed for the two rig sizes adjusted by the inverse of

the geometric scale factor. Both linear and nonlinear analyses were conducted, as well as cases

with and without gas loads and temperatures (i.e., speed only). The linear cases were run primarily

as a check on the scaling operation, since these results must scale by definition. Cases without gas

loads were run with uniform room temperature. Cases with gas loads were run with the

aerodynamic design point (ADP) pressure distribution and a radial temperature profile

corresponding to these conditions. Note that the identical pressure distribution and temperature

profile was applied to both rig sizes.

The results for six analyses are tabulated in Table 6. The physical displacements (the

magnitude of the displacement, in inches) at the tip leading edge (LE) and trailing edges (TE) are

provided for each run, and then these are normalized by the tip radius of each rig size. The

resulting normalized values for corresponding loadings are identical. While only the tip

displacements are summarized here, other locations on the airfoils also exhibit the same behavior.

NASA/CR—2012-217451 29
Table 6. Effect of Geometric Scaling on Hot Blade Shape. Table Lists Displacements at LE

and TE of Tip for Several Loading Alternatives. The Displacements for Each Rig

Size, When Normalized by Tip Radius Are Identical for Corresponding Cases.

Rig Size Gas

Loads

Temps Solution Disp at

Tip LE, in

Disp at Tip

TE, in

Norm

Disp LE

Norm

Disp TE

22 no RT linear 2.1233 1.2328 .1930 .1121

22 no RT nonlinear 0.4428 0.2410 .0403 .0219

22 ADP ADP nonlinear 0.4004 0.2068 .0364 .0188

18 no RT linear 1.7131 0.9946 .1930 .1121

18 no RT nonlinear 0.3573 0.1944 .0403 .0219

18 ADP ADP nonlinear 0.3230 0.1669 .0364 .0188

The conclusion from this study is that the identical hot shape will result from identical cold

shapes if the only change is a geometric scaling. This result is true even when pressure and

temperature loading is included.

Note that the relationships between physical and corrected conditions change under

different ambient conditions. The deflection of the blade is driven by the physical conditions.

So if the two rigs were run to the same corrected conditions at different ambient conditions, there

would be a difference in the loads. This difference in ambient conditions, though, would have to

be quite large to significantly affect the hot shape.

Also note that this study did not attempt to address the issue of whether the actual hardware

used in each rig deflects as intended. The deflection of each size could be affected significantly

by the conformance of those blades to the intended (and analyzed) nominal shape. In order to

conduct such a study, detailed geometric measurements of representative sample blades in each

scale would be needed, and then new ANSYS models constructed.

4.2 Evaluation of the Acoustic Results of the 22” Rig Test of the QHSF I

An assessment of the noise data from the QHSF 22” rig wind tunnel test was conducted.

The purpose of this assessment was to identify the acoustic benefits (and problems) of the QHSF

I design and to validate the design process for the QHSF II design. Figure 28 shows the

summary of the results of the QHSF I far field noise measurements (Reference 12). Also shown

on the figure are the Baseline I results and the later measurements of the Baseline I fan rotor and

the QHSF I stator.

The QHSF I successfully reduced interaction tone noise for both rotor-stator and rotor-strut

interaction. This tone reduction is responsible for up to 6 EPNdB reduction at higher fan tip

speeds. Figure 29 presents a spectrum comparison showing the effect.

NASA/CR—2012-217451 30
70

80

90

100

700 800 900 1000 1100 1200 1300 1400 1500

Fan Tip Speed, fps

EPNdB, dB

Baseline I Fan

QHSF I

Baseline I Rotor/QHSF I Stator

Figure 28. Results of the 22" Rig Testing Showing Dramatic Differences in Noise Levels for a

1500 ft Fly Over at Matched Thrust Conditions From the Three Fan Configurations.

70

75

80

85

90

95

100

105

100125160200250315400500630800

1000125016002000250031504000500063008000

10000125001600020000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I

 * 1*BPF

** 2*BPF

*

**

Figure 29. At Supersonic Tip Speeds, the Primary Noise Reduction Was in the Blade

Passage Tone (13831 Rig RPM, 131 Degrees From the Inlet).

4.2.1 V072 Validation

A check was performed to see if V072 is a reliable prediction tool. The V072 predictions

were assumed to be conservative for the original design because of the inaccurate loss profile of

the Baseline I fan. Actual reductions in tones were greater than those predicted, and the general

trends predicted by V072 were confirmed by test data. Figure 30 shows the results for the

forward arc and Figure 31 shows the results for the aft arc.

NASA/CR—2012-217451 31
• Data indicate that 1*BPF tone is due to RotorStrut interaction

• Measured reductions at 2*BPF and 3*BPF are

greater than predicted, which is consistent with

low Baseline I losses in V072

• V072 trends are basically correct

Inlet 1*BPF Cutoff

Inlet 2*BPF -5 dB

Inlet 3*BPF +1 dB

V072:

0

5.9 Hz Narrowband Data - With Wall

8516 RPM - 31 degrees

60

70

80

90

100

110

120

130

1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I

∆ = -9.4

∆ = -6.4

5.9 Hz Narrowband Data - With Wall

8516 RPM - 50 degrees

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I

∆ = -1.3

∆ = -2.3

0

5.9 Hz Narrowband Data - With Wall

8516 RPM - 70 degrees

60

70

80

90

100

110

120

130

1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I

∆ = -7.3

∆ = -4.7

Figure 30. Comparison With Measured Narrow Band Data Shows That V072

Underestimated the Tone Noise Reduction at 2x and 3x the Blade Passage Tone

at 55.5% Speed in the Forward Arc.

• Data indicate that 1*BPF tone is due to Rotor-Strut interaction

• Measured reductions at 2*BPF and 3*BPF are greater than

predicted, which is consistent with low Baseline I losses in V072

• V072 trends are correct

Exit 1*BPF Cutoff

Exit 2*BPF -6 dB

Exit 3*BPF -3 dB

V072:

5.9 Hz Narrowband Data

8516 RPM - 131 degrees

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I

∆ = -11.6

∆ = -5.0

Figure 31. Comparison With Measured Narrow Band Data Shows That V072

Underestimated the Tone Noise Reduction at 2x and 3x the Blade Passage Tone

at 55.5% Speed in the Aft Arc.

NASA/CR—2012-217451 32
4.2.2 Broadband Noise Source

An unknown broadband noise source was identified in the QHSF I data that must be

identified and eliminated for the redesign as shown in Figure 32. The LDV data in Figure 33

was taken during the test and shows a flow separation on the QHSF I rotor blades at 9510 RPM.

The data at higher rotor speeds show little separation.

4.2.3 Comparison With CFD

The first study performed was a comparison of the measured wake structures of the

Baseline I and QHSF I rotors at the LDV plane, at 81.4 percent and 90.1 percent corrected fan

speed. The position of the LDV plane downstream from the rotor trailing edge is shown for both

rotors in Figure 34.

The wake structure in the flow path cross-section was compared graphically, using the

FIELDVIEW program. As shown in Figure 35, at both speeds, the wakes from the Baseline I

and the QHSF I rotors have a similar slope throughout the inner span region. However, in the

outer span region, the QHSF I wake displays more tangential lean. This increase in lean is a

result of the increased distance between the rotor trailing edge and the LDV plane in the outer

span region, due to the forward sweep of the QHSF I rotor. The wake rotates further tangentially

through this additional axial distance, in effect having more lean than the wake in the inner span

region, at the LDV plane. The aft sweep and lean of the stator leading edge further enhance this

effective “lean” of the rotor wake. As a result, the QHSF I rotor wake traverses the stator

leading edge much more slowly than does the Baseline I rotor wake. This behavior serves to

reduce the rotor-stator interaction tone noise.

In addition to examining the graphical representation of the wake structure using the LDV

data, a more detailed study of the wake profiles was conducted with the LDV measurements.

Wake width and depth were compared for the Baseline I and QHSF I rotors at 81.4 percent and

90.1 percent corrected fan speed, for three selected span-wise positions (38 percent, 57.4 percent,

and 78.8 percent of the rotor trailing edge span). The wake profiles were normalized by the free

stream resultant velocity, and were shifted tangentially, to overlap for comparison purposes. The

wakes at 81.4 percent and 90.1 percent corrected fan speed are shown in Figure 36 and Figure 37,

respectively. At 38.0 percent span, the Baseline I and QHSF I wakes display very similar

profiles, because the distance from the trailing edge to the LDV plane is essentially identical.

Moving out to 57.4 percent span, the depth of the wake is greatly reduced for the QHSF I rotor

and the width is increased, due to the increased distance from the rotor trailing edge. This trend

continues at 78.8 percent span; however, while the wake is only slightly evident at 81.4 percent

speed, it maintains more strength at 90.1 percent speed.

Clearly, the QHSF I rotor wake exhibits less strength over the outer portion of the span at

the LDV plane, due to the increased distance from the rotor trailing edge. This behavior is

further enhanced at the stator leading edge, and serves to reduce the impact of the rotor-stator

interaction tone noise.

NASA/CR—2012-217451 33
60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSFI Fan

(a) 8516 RPM (b) 9510 RPM

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

70

80

90

100

110

120

130

140

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

(a) 10,646 RPM (b) 11,572 RPM

70

80

90

100

110

120

130

140

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

70

80

90

100

110

120

130

140

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

(a) 12,500 RPM (b) 13,000 RPM

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

60

70

80

90

100

110

120

130

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

Sound Pressure Level, dB

Baseline I Fan

QHSF I Fan

(a) 13,342 RPM (b) 13,831 RPM

Figure 32. Unknown Broadband Noise Source Must Be Identified and Eliminated for

Redesign of the QHSF (Data at 61 Degrees With Barrier).

NASA/CR—2012-217451 34
(a) axial position of LDV survey

I I I

(b) 9510 RPM (c) 12,500 RPM (d) 13,831 RPM

Figure 33. LDV Axial Velocity Data Taken Downstream of the QHSF I Show Flow

Separation at Low RPM That Is Reduced at Higher Values of RPM.

Figure 34. Location of the LDV Planes Relative to the Trailing Edges of the Baseline I and

QHSF I Rotors.

NASA/CR—2012-217451 35
Baseline I Fan and QHSF I

QHSF I Baseline I Fan

(a) 12500 RPM

Baseline I Fan and QHSF I

QHSF I

Baseline I Fan

(b) 13831 RPM

Figure 35. The Wake Structure at the LDV Plane for the Baseline I and QHSF I Rotors, at

81.4% and 90.1% Corrected Fan Speed.

NASA/CR—2012-217451 36
LDV Wake Profiles

38.0% TE Span (Radius = 7.55 Inches)

81.4% Corrected Fan Speed (12,500 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec)

Baseline I FanQHSF I

Figure 36. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 81.4% Corrected Fan Speed.

NASA/CR—2012-217451 37
LDV Wake Profiles

57.4% TE Span (Radius = 8.55 Inches)

81.4% Corrected Fan Speed (12,500 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec)

Baseline I FanQHSF I

Figure 36. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 81.4% Corrected Fan Speed (Cont).

NASA/CR—2012-217451 38
LDV Wake Profiles

78.8% TE Span (Radius = 9.65 Inches)

81.4% Corrected Fan Speed (12,500 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec) Baseline I Fan

QHSF I

Figure 36. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 81.4% Corrected Fan Speed (Cont).

NASA/CR—2012-217451 39
LDV Wake Profiles

38.0% TE Span (Radius = 7.55 Inches)

90.1% Corrected Fan Speed (13,800 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec)

Baseline I FanQHSF I

Figure 37. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 90.1% Corrected Fan Speed.

NASA/CR—2012-217451 40
LDV Wake Profiles

57.4% TE Span (Radius = 8.55 Inches)

90.1% Corrected Fan Speed (13,800 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec) Baseline I Fan

QHSF I

Figure 37. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 90.1% Corrected Fan Speed (Cont).

NASA/CR—2012-217451 41
LDV Wake Profiles

78.8% TE Span (Radius = 9.65 Inches)

90.1% Corrected Fan Speed (13,800 RPM)

0.90.80.70.6

1

1.21.1

0 5 10 15 20 25 30 35 40 45 50

LDV Pitchwise Location

Resultant Velocity (ft/sec) Baseline I Fan

QHSF I

Figure 37. Comparison of Baseline I and QHSF I Rotor Wake Profiles at 90.1% Corrected Fan Speed (Cont).

NASA/CR—2012-217451 42
The LDV data were also compared to the CFD data. Comparisons of the overall wake

structure at same downstream plane were made for both the QHSF I and Baseline I rotors. The

operating line points were analyzed at the corrected fan speeds shown in Table 7. The CFD

predictions were taken from the DAWES analyses performed as part of QHSF I design activity.

In the inner span region at the LDV measurement plane, the slopes of the wakes of the QHSF I

and Baseline I rotors are similar. In the outer span region, the QHSF I wakes display more

tangential lean at LDV plane. This difference is due to the increased distance between rotor

trailing edge and the LDV plane resulting from the forward sweep of the QHSF I rotor. The

wakes rotate further tangentially through the additional axial distance, producing more lean than

the wakes in the inner span region. The CFD predictions of the rotor wake structure at the LDV

plane appear to be in good qualitative agreement with the LDV data as shown in Figure 38 to

Figure 41.

Table 7. The Available Corrected Fan Speeds for the LDV Measurements and CFD

Analyses Were Matched as Closely as Possible.

Baseline I Fan QHSF I

LDV Data CFD Analyses LDV Data CFD Analyses

81.4% 81.4% 80%

85%

90.1%

(Sideline)

90.1%

(Sideline)

90%

Baseline I Fan Baseline I Fan

(a) LDV - 81.4% Speed (b) CFD – 85.0% Speed

Figure 38. A Good Comparison Is Seen Between the Measured and Calculated Rotor

Wakes for the Baseline I Fan at a Typical Cutback Takeoff Condition.

NASA/CR—2012-217451 43
Baseline I Fan Baseline I Fan

(a) LDV – 90.1% Speed (b) CFD – 85.0% Speed

Figure 39. A Good Comparison Is Seen Between the Measured and Calculated Rotor

Wakes for the Baseline I Fan at a Typical Full Power Takeoff Condition.

QHSF I

(a) LDV – 81.4% Speed (b) CFD – 80.0% Speed

Figure 40. A Good Comparison Is Seen Between the Measured and Calculated Rotor

Wakes for the QHSF I at a Typical Cutback Takeoff Condition.

NASA/CR—2012-217451 44
QHSF I

(a) LDV – 90.1% Speed (b) CFD – 90.0% Speed

Figure 41. A Good Comparison Is Seen Between the Measured and Calculated Rotor

Wakes for the QHSF I at a Typical Full Power Takeoff Condition.

4.3 Evaluation of the Rotor/Strut Interaction

During QHSF I testing, rotor-strut interaction tones were observed at 90 percent speed for

the Baseline I fan, but were not present for the QHSF I fan (Reference 13). Figure 42 shows

typical circumferential mode results from those measurements. Both the Baseline I fan and the

QHSF I have 22 fan blades and 10 struts. Therefore, the expected rotor/strut circumferential

modes are

(... -18, -8, 2, 12, 22, ...)

It was proposed that the difference in stators might have impacted strut-induced pressure

disturbances at the rotor exit. A rotor-stator-strut interaction study was conducted to

• identify pressure disturbances upstream of the stator that could be produced by the strut

• identify any differences in pressure distribution from one strut passage to the next, due to

the differing relative position of stators vs. struts for each strut passage over 180 degrees

• qualitatively assess any differences between the pressure distributions upstream of the

Baseline I and QHSF I stators

NASA/CR—2012-217451 45
-32 -30 -28 -26 -24 -22 -20 -18 -16 -14 -12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32

100

105

110

115

120

125

130

135

140

PWL, dB

Circumferential Mode Order, m

Total tone PWL = 142.4 dB

(a) Baseline I Fan

-32 -30 -28 -26 -24 -22 -20 -18 -16 -14 -12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32

Circumferential Mode Order, m

Total tone PWL = 127.1 dB

(b) QHSF I

Figure 42. Acoustic Modal Measurements in the Aft Fan Duct With a Rotating Rake Show

Significant Rotor Strut Tones for the Baseline I Fan.

CFD analyses were performed with the Fluent® CFD code for the two fan configurations:

• Baseline I Stator + Strut

• QHSF I Stator + Strut

The strut geometry was identical for both studies. The upstream boundary conditions were

taken to be the QHSF I rotor exit conditions from 18” rig test data at 90 percent corrected speed

on the standard sea level operating line. Predicted static pressure fields were examined upstream

of stators. The actual configuration of both fans was 52 stator vanes and 10 struts; therefore, by

NASA/CR—2012-217451 46
applying periodic boundary conditions on a 26 vane and 5 strut model, a full flow field was

evaluated. Diagnostic runs were also made with an approximate model consisting of 5 vanes and

1 strut. Table 8 provides a summary of the analysis runs performed.

Table 8. Summary of CFD Models Run in the Rotor/Strut Potential Interaction Study.

Model Cell count Comment

1.5 M 26 stator, 5 strut

3.3 M 26 stator, 5 strut

4.9M 26 stator, 5 strut, 1 adaptation on Ps gradient Baseline I

1.9 M 5 stator, 1 strut

1.9 M 26 stator, 5 strut

3.3 M 26 stator, 5 strut

5.4M 26 stator, 5 strut, 1 adaptation on Ps gradient QHSF I

2.2 M 5 stator, 1 strut

Figure 43 shows the CFD geometries that were used for the 5 strut models. Periodic

boundary conditions were applied to be consistent with the full 360-degree fan configuration as

shown in Figure 44. The unstructured grid used to perform the analysis is shown in Figure 45.

Stators

26 equally spaced

Core

struts

Bypass

struts

Bypass exit

Pressure outlet

Massflow outlet BC

Core exit

Pressure outlet

Massflow outlet BC

Fan exit

Pressure Inlet

Profiles for

Total pressure

Total Temperature

Flow Direction

Stators

26 equally spaced

Core

struts

Bypass

struts

Bypass exit

Pressure outlet

Massflow outlet BC

Core exit

Pressure outlet

Massflow outlet BC

Fan exit

Pressure Inlet

Profiles for

Total pressure

Total Temperature

Flow Direction

(a) Baseline I Fan (b) QHSF I

Figure 43. CFD Models for the Rotor/Strut Interaction Study Modeled 26 Vanes, 5 Struts,

and the Split Flow Path.

NASA/CR—2012-217451 47
Periodic

Surface

Periodic surface is a planar surface

Splits the edge stator

Figure 44. A Periodic Boundary Condition Was Used to Model the Total 360 Degree

Flowfield.

Figure 45. An Unstructured Grid Was Used to Model the Rotor/Strut Interaction

Flowfield (QHSF I).

NASA/CR—2012-217451 48
The contour plots in Figure 46 and Figure 47 show that there is more static pressure

variation at the rotor exit and in the mid-region between the rotor and stator for the Baseline I fan

than the QHSF I. To further visualize the variation in static pressure between the two fans, the

static pressure results were plotted at several data planes. Figure 48 and Figure 49 show the

location of the data planes for the two fans. Figure 50 show overlays of circumferential cuts of

the two-stator systems to demonstrate the difference in the stator configurations for the two fans.

exitfansexitfant

exitfant s

PP

PP Cp −⋅ −⋅

−⋅

−

− = ( )

(a) Rotor Exit (b) Mid Region

Figure 46. Pressure Coefficient Contours Between the Rotor and Stator Have Been

Produced From the Fluent® CFD Analysis for the Baseline I Fan.

NASA/CR—2012-217451 49
exitfansexitfant

exitfant s

PP

PP Cp −⋅ −⋅

−⋅

−

− = ( )

(a) Fan Exit (b) Mid Region

Figure 47. Pressure Coefficient Contours Between the Rotor and Stator Have Been

Produced From the Fluent® CFD Analysis for the QHSF I.

-0.290m

-0.252m

-0.225m

-0.200m

Data Planes

Rhub = 0.219m

Rshd = 0.386m

Figure 48. Data Planes Were Selected for Comparison of Circumferential Static Pressure

Profiles (Baseline I Fan).

NASA/CR—2012-217451 50
-0.290m

-0.252m

-0.225m

-0.200m

Data Planes (equidistant with Baseline I Fan)

Rhub = 0.219m

Rshd = 0.386m

Figure 49. Data Planes Were Selected for Comparison of Circumferential Static Pressure

Profiles (QHSF I).

Hub R=0.240m R=0.300m

R=0.350m R=0.375m Shroud

Figure 50. Circumferential Cuts Show the Relative Positions of the Baseline I and QHSF I

Vanes at Various Radii (QHSF I Shown in Black).

NASA/CR—2012-217451 51
Significant differences in the stator geometries occur out near the tip. Data comparisons

were performed at the radii shown in Figure 51. Figure 52 and Figure 53 show the

circumferential variation of the static pressure for the QHSF I and Baseline I fans at 3 different

axial positions for the two radii. It is clear from the data that the Baseline I fan has more static

pressure variation than the QHSF I. This evidence suggests that the cause of the rotor strut

interaction tones is the rotor responding to the variation of the potential pressure field of the

struts.

(a) R=0.350m (b) R=0.375m

Figure 51. Data Comparisons Emphasize Two Radii Near the Vane Shroud (QHSF I

Shown).

NASA/CR—2012-217451 52
17.30

17.35

17.40

17.45

17.50

17.55

17.60

17.65

17.70

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

X = -0.200m ∆=0.05

17.20

17.22

17.24

17.26

17.28

17.30

17.32

17.34

17.36

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

X = -0.225m ∆=0.02

TDC 17.18

17.20

17.22

17.24

17.26

17.28

17.30

17.32

17.34

17.36

17.38

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

∆=0.02

X = -0.252m

Baseline I Stator

QHSF I Stator

Figure 52. Comparison of Circumferential Pressure Distributions at R-0.375m (QHSF I

Data Shifted in Angle and Level to Align With Baseline I Data).

NASA/CR—2012-217451 53
17.10

17.15

17.20

17.25

17.30

17.35

17.40

17.45

17.50

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

X = -0.200m ∆=0.05

17.16

17.18

17.20

17.22

17.24

17.26

17.28

17.30

17.32

17.34

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

X = -0.225m ∆=0.02

TDC 17.24

17.26

17.28

17.30

17.32

17.34

17.36

17.38

17.40

17.42

-1.50 -1.00 -0.50 0.00 0.50 1.00 1.50 2.00

Pitchwise Position (rad)

Static Pressure (psi)

X = -0.252m ∆=0.02

Baseline I Stator

QHSF I Stator

TDC

Figure 53. Comparison of Circumferential Pressure Distributions at R=0.350m (QHSF I

Data Shifted in Angle and Level to Align With Baseline I Data).

NASA/CR—2012-217451 54
5. QHSF II DESIGN

5.1 Approach for the QHSF II Design

The QHSF II design will be developed using an analytical Design of Experiments (DOE) to

define an optimum rotor and stator system. The interdisciplinary process being used is described

in Figure 54. The DOE outline is:

• Blade Forward Sweep

− Axial and tangential sweep components coupled

− Key outputs: Acoustics, Aeroelastics, Aerodynamics

• Blade Tangential Lean

− Optimized independent of selected axial sweep

− Key outputs: Acoustics, Aeroelastics, Mechanical

• Blade Thickness Distribution

− Key outputs: Aeroelastics, Mechanical, Aerodynamics

• Stator Optimization

− Sweep & Lean

− Key outputs: Acoustics, Aerodynamics, Mechanical

• Rotor Incidence

− Key outputs: Aerodynamics, Aeroelastics

NASA/CR—2012-217451 55
Mechanical

Aero Acoustic

Solutions

Geometry Input

Xcg distribution

Ycg distribution

Tmax location

Thickness

Tip slope

Beta angles 

Chord

Thickness 

DOE

Analysis

• Design of Experiments (DOE) achieves a solution

in less time/cost which satisfies multiple criteria.

• DOEs are selected to determine main effects and

interactions between factors.

Aeroelastic

Figure 54. An Interdisciplinary Process Has Been Defined for the Design of the QHSF II.

5.2 Rotor Stacking Design of Experiments

The first DOE for the QHSF II fan rotor determined the blade forward sweep stacking. The

configurations are assumed to have equal axial and radial center of gravity (Xcg and Ycg) offsets

to facilitate evaluating multiple configurations. Four parameters are used to define the blade

stacking as shown in Figure 55. Table 9 summaries the 25 cases for DOE 1.

A number of dependent variables (Y-factors) have been defined for DOE 1. These

quantities will be used to evaluate the merits of the various configurations in the DOE and point

to the possible go-forward designs for the DOE 2. Table 10 to Table 13 summaries the factors

for DOE 1.

NASA/CR—2012-217451 56
QHSF II DOE 1 - Geometry Generation Description

DOE Defines Ycg = Xcg; 5th Order Polynomial Curve Fit

-0.4

0

1.61.20.80.4

2

0 2 4 6 81

0

PSPAN/10 (simulates av

g blade hei

ght in inches (9.815 vs 10.0))

XCG (inches)

case 4 case 10 case 12case 13 case 16 case 8

DOE Parameter 1XCG at 45% span+0.2 in, -0.1 in

XCG at 5% spanLinear interpolationfrom 0% to 45% 

DOE Parameter 2Slope from 45% span to 65% span defined based on delta from slope from 0% to 45%+7 degrees, -7 degrees

DOE Parameter 3Slope from 65% span to 75% span defined based on delta from slope from 45% to 65%+7 degrees, -3 degrees

DOE Parameter 4Slope from 75% span to 100% span defined based on delta from slope from 65% to 75%+10 degrees, -2 degrees

+7 degrees

-7 degrees

Figure 55. Four Parameters Define the Rotor Blade Stacking for DOE 1.

NASA/CR—2012-217451 57
Table 9. Summary of the Cases for DOE 1.

HIGHLIGHTED CASES ARE PART OF THE PARTIAL FACTORIAL RUN FIRST 

StdOrder RunOrder CenterPt Blocks xcg45 delslope65 delslope75 delslope100

1 1 1 1 -0.1 -7 -3 -2

2 2 1 1 0.2 -7 -3 -2

3 3 1 1 -0.1 7 -3 -2

4 4 1 1 0.2 7 -3 -2

5 5 1 1 -0.1 -7 7 -2

6 6 1 1 0.2 -7 7 -2

7 7 1 1 -0.1 7 7 -2

8 8 1 1 0.2 7 7 -2

9 9 1 1 -0.1 -7 -3 10

10 10 1 1 0.2 -7 -3 10

11 11 1 1 -0.1 7 -3 10

12 12 1 1 0.2 7 -3 10

13 13 1 1 -0.1 -7 7 10

14 14 1 1 0.2 -7 7 10

15 15 1 1 -0.1 7 7 10

16 16 1 1 0.2 7 7 10

17 17 0 1 0.05 0 2 4 CENTER POINT

18 18 2 -0.1 0 2 4

19 19 2 0.2 0 2 4

20 20 2 0.05 -7 2 4

21 21 2 0.05 7 2 4 AXIAL

22 22 2 0.05 0 -3 4 POINTS

23 23 2 0.05 0 7 4

24 24 2 0.05 0 2 -2

25 25 2 0.05 0 2 10

NASA/CR—2012-217451 58
Table 10. Acoustic Y-Factors for DOE 1.

Y-factor Data type 

delz 50 Axial distance RTE – VLE midspan

delz 100 Axial distance RTE – VLE tip

dB des Interaction noise sound power level

eff swp des Rotor LE effective sweep

% c 50 des Passage shock loc. Design Nc, 50% span 

% c 60 des Passage shock loc. Design Nc, 60% span 

% c 70 des Passage shock loc. Design Nc, 70% span 

% c 80 des Passage shock loc. Design Nc, 80% span 

% c 90 des Passage shock loc. Design Nc, 90% span 

% c 95 des Passage shock loc. Design Nc, 95% span 

dB 89 Interaction noise sound power level

eff swp 89 Rotor LE effective sweep

% c 50 89 Passage shock loc. 89% Nc, 50% span 

% c 60 89 Passage shock loc. 89% Nc, 60% span 

% c 70 89 Passage shock loc. 89% Nc, 70% span 

% c 80 89 Passage shock loc. 89% Nc, 80% span 

% c 90 89 Passage shock loc. 89% Nc, 90% span 

% c 95 89 Passage shock loc. 89% Nc, 95% span 

Table 11. Aeroelastic Y-Factors for DOE 1.

Y-factor Data type 

Damping - extrapolated to stall line scalar for each speedline 

Damping - massflow at flutter boundary scalar for each speedline 

Damping - flutter margin relative to PE scalar for each speedline 

Damping - critical nodal diameter scalar for each speedline 

Reduced frequency, mode 1 spanwise distrib for each steady condition

Reduced frequency, mode 2 spanwise distrib for each steady condition

Reduced frequency, mode 3 spanwise distrib for each st eady condition

Twist/flex ratio spanwise distrib for each steady condition

Incidence spanwise distrib for each steady condition

Relative Mach spanwise distrib for each steady condition

Optional:

Shock location (as % chord) spanwise distrib for each steady condition

Shock strength (as delta p) spanwise distrib for each steady condition

Separations (location, size) spanwise distrib for each steady condition

NASA/CR—2012-217451 59
Table 12. Aerodynamic Y-Factors for DOE 1.

Y-factor Data type Wc Inlet Wc at Peak Effy

Eff Rotor Peak Effy

PR Rotor PR at Peak Effy

Ws50 Wennerstrom shock loss @ 50% span

Ws70 Wennerstrom shock loss @ 70% span

Ws80 Wennerstrom shock loss @ 80% span

Ws90 Wennerstrom shock loss @ 90% span

Ws100 Wennerstro m shock loss @ 100% span

Table 13. Mechanical Y-Factors for DOE 1.

Y-factor Data type psmax max pressure surface stress

ssmax max suction surface stress

plemax max leading edge PS stress 4-100% span

slemax max leading edge SS stress 4-100% span

umax tip max tip deflection

u tiple tip leading edge deflection

freq 1 frequency margin mode 1

freq 2 frequency margin mode 2 

freq 3 frequency margin mode 3

fec 1 placement of 2/rev crossing mode 1

fec 2 placement of 4/rev crossing mode 2

fo dam fold-over damage

An extensive evaluation of the acoustic, aerodynamic, mechanical and aeroelastic

properties of the 25 configurations was performed. The data were all collected into the

MINITAB® software to perform regression, analysis of variance, and sensitivity studies. The

rotor blade stacking was particularly sensitive to five key parameters:

• Total sound power in the first two harmonics of the rotor/stator interaction noise at 89

percent speed as calculated by SOURCE3D (RSI_89_T)

• Maximum Mach Number relative to the blade leading edge at 89 percent speed as

calculated from the mean streamline method (max mrn)

• Fold over damage to the blade from a bird impact at 100 percent speed as calculated

by the NOSAPM program (FO Damag)

• Total fan weight flow at the design point as calculated by the inviscid TURBO-AE (Wc)

• Maximum leading edge stress at 100 percent speed as calculated with ANSYS

Maximum leading edge stress was later dropped from the analysis when it was determined

that it could be easily controlled with small changes in Ycg. Table 14 shows a summary of the

NASA/CR—2012-217451 60
sensitivity study from MINITAB®. The results show that the optimum acoustic, mechanical,

aerodynamic, and aeroelastic design is one that has the stacking initially aft for the inboard part

of the blade, and then curving forward at the tip, corresponding to Case 14. Figure 56 shows the

blade stacking profile as compared to the most aft swept (Case 1), the most forward swept (Case

16), the QHSF I, and the Baseline I sweep distribution. It can be seen from the figure that the

slope of the blade near the tip approaches the original QHSF1 design, but the blade as less

forward sweep inboard to meet mechanical requirements. Figure 57 shows how the fold over

damage criterion had a significant impact on the optimum design selection. Figure 58 shows the

decrease in the leading edge shock strength as the forward sweep is increased.

The go-forward design of the rotor was selected to be Case 14a. Case 14a is the optimum

configuration from DOE I that was subsequently optimized for Ycg offset in ANSYS to

minimize leading edge stress and blade normal Mach number. Figure 59 shows an overlay of the

QHSF I fan and the Case 14a fan.

Table 14. Summary of the Sensitivity of the Four Key Parameters to the Rotor Blade

Stacking.

Hi

Lo 0.52118

D

Optimal

Cur

d = 0.93723

Targ: 245.0

Wc

d = 0.92051

Targ: 0.850

FO Damag

d = 0.10594

Minimum

max mrn 

d = 0.46851

Minimum

RSI_89_T

y = 242.6147

y = 0.8782

y = 1.2341

y = 131.7204

-2.0

10.0

-3.0

7.0

-7.0

7.0

-0.10

0.20

xcg45 delslope65 delslope75 delslope100

[-0.10] [-6.9996] [7.0] [10.0]

NASA/CR—2012-217451 61
QHSF II DOE 1 - HALF FACTORIAL W/ CENTER POINT

DOE Defines Ycg = Xcg

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.5

2.0

2.5

0 2 4 6 810

Avg blade height in inches

XCG (inches)

AS907 BLOCK 2 & QHSF I

case 14

case 1

case 16

Figure 56. A Go-Forward Blade Stacking Has Been Defined to Meet All Acoustic,

Mechanical, Aerodynamic, and Aeroelastic Requirements.

Foldover damage vs RSI TN @ 89% Nc

0

0.5

1

1.5

2

2.5

3

126 128 130 132 134 136 138 140

RSI TN

%F

O

Foldover damage vs Max Mrnormal @ 89% Nc

0

0.5

1

1.5

2

2.5

3

1 1.1 1.2 1.3 1.4

Max Mrnormal

%F

O

CASE 14 BASELINE

Violates Bird

Ingestion Criteria

Violates Bird

Ingestion Criteria

Figure 57. The Bird Strike Criterion Put a Restriction on the Design Space for Rotor DOE 1.

NASA/CR—2012-217451 62
(a) Case 1 (b) Case 14 (c) Case 16

Figure 58. The Strength of the Shock in Front of the Blade Leading Edge at 90% Span

Was Shown to Decreased With Increasing Forward Sweep.

Side View Fwd Looking Aft

QHSF I - Blue Case 14a - Red

(at engine scale)

Figure 59. Comparison of the Go-Forward Rotor Blade Design to the QHSF I.

The TURBO code in the inviscid mode was used to evaluate the aeroelastic stability of the

25 configurations for DOE 1. The surprising result was that the instability point was relatively

insensitive to the blade rotor blade stacking. Figure 60 shows the variation in the calculated

instability point for the 25 configurations.

To assess the acoustic impact of changes in rotor stacking, SOURCE3D predictions were

generated for the 25 DOE configurations at 89 percent and 100 percent speed. Figure 61 shows

the overall sound power at the blade pass and twice blade pass tone for each of the

configurations at 89 percent speed.

NASA/CR—2012-217451 63
Inlet Corrected Flow (lbm/s)

Bypass Pressure Ratio

Predicted

Cruise

TakeOffBlock IIop line Stall, Block II, Corr.case1case2case3case4case5case6case7case8case9

case10case11case12case13case14case15case16case17case18case19case20case21case22case23case24case25

Baseline II Rt_ datamatch

110%

105%

100%

95%

90%

85%

80%

70%

Figure 60. QHSF II, Flutter Boundaries for DOE 1 Cases in Baseline II Tested Map.

NASA/CR—2012-217451 64
125.0 

130.0 

135.0 

140.0 

123456789

10111213141516171819202122232425

Baseline II

Figure 61. The Logarithmic Sum of the Sound Power Levels for the Rotor-Stator

Interaction Noise at 89% Speed Shows Significant Variation for Range of DOE

Parameters.

5.3 Aeroelastic Verification of the Case 14a Rotor Design

Confirmation runs for the Case 14a rotor with TURBO in the viscous mode. The

aeroelastic behavior of the rotor at 60 percent, 70 percent, 89 percent, and 100 percent was

determined to verify the stability of the design throughout the fan operating range.

The steadystate solutions, including the results from both NASA and Honeywell, are shown

in Figure 62. All NASA and Honeywell runs are TURBO viscous results at 89 percent speed.

There are 3 points on the map based on Honeywell’s results: Pexit=14.0, 14.7, and 14.9. The

points at Pexit=14.0 and 14.9 are fully converged, but Pexit=14.7 was only run 500 iterations.

Both Honeywell and NASA results show a consistent trend along the speed line.

For the unsteady analyses, Honeywell used values of Pexit of 14.7 and 14.9 while NASA

used Pexit of 14.28 and 14.84. NASA ran 12 vibratory cycles and Honeywell ran 20 cycles

starting from the steady state solutions. All NASA and Honeywell unsteady runs are fully

converged according to the convergence history of damping. The extrapolation to obtain the

mass flow at zero damping, given in Figure 63, shows some difference between Honeywell and

NASA results. In the large part, this is because the damping is relatively high (all above 0.5), so

that the extrapolation to zero damping is quite sensitive. If the damping is small (below 0.2), the

extrapolation will not be this sensitive. Figure 64 shows the damping extrapolation with pressure

ratio. The pressure ratio is not as sensitive as the mass flow because at near stall conditions the

speed line is almost flat on the mass flow vs. pressure ratio map (Figure 62).

NASA/CR—2012-217451 65
mass

pr

190 195 200 205 210 1.4

1.5

1.6

1.7

1.8

NASA

Honeywell

Figure 62. Steady-State Solutions for Pressure Ratio (pr) and Mass Flow Rate (m) at 89%

Speed for the Case 14a Rotor Show a Consistent Trend.

m

%damping

150 160 170 180 190 200 210 0

0.25

0.5

0.75

1

NASA (Base)

Honeywell (Base)

Figure 63. Damping Extrapolation as a Function of Mass Flow Rate (m) Is Shown at 89%

Speed for the Case 14a Rotor.

NASA/CR—2012-217451 66
pr

%damping

1.55 1.6 1.65 1.7 1.75 1.8 0

0.25

0.5

0.75

1

NASA (Base)

Honeywell (Base)

Figure 64. Damping Extrapolation With Pressure Ratio (pr) Is Shown at 89% Speed for

the Case 14a Rotor.

Sensitivity studies were conducted with TURBO to determine the influence of reduced

frequency and twist to flex ratio on the flutter boundary. The geometry of Case 14a at 89 percent

speed was used as the test case. Two operating points, Pexit=14.7 and Pexit=14.9 (fac=1.05 and

1.0643), were used to predict the flutter boundary. Six nodal diameters (0, 2, 4, 6, -2, and –4)

were run for each operating point to obtain the least stable nodal diameter.

The reduced frequency of the Case 14a rotor at 87.5 percent span is 0.3423 and 0.3442 for

Pexit values of 14.7 and 14.9, respectively. For frequency sensitivity analyses, two reduced

frequency cases are run: half and double Case 14a. All other parameters including pressure ratio

and mode shape remain the same. The damping plotted in Figure 65 indicates that it is more

stable when the reduced frequency is increased. The extrapolation of the damping with pressure

ratio given in Figure 66 shows a similar situation.

NASA/CR—2012-217451 67
m

%damping

150 160 170 180 190 200 210 0

0.25

0.5

0.75

1

Honeywell (Base)

0.5*Freq.

2*Freq

Figure 65. Frequency Sensitivity: Damping Extrapolation With Mass Flow Rate (m) Is

Shown at 89% Speed for the Case 14a Rotor.

pr

%damping

1.55 1.6 1.65 1.7 1.75 1.8 0

0.25

0.5

0.75

1

Honeywell (Base)

0.5*Freq.

2*Freq

Figure 66. Frequency Sensitivity: Damping Extrapolation With Pressure Ratio (pr) Is

Shown at 89% Speed for the Case 14a Rotor.

NASA/CR—2012-217451 68
Two new mode shape files were created for the twist-to-flex ratio sensitivity analyses. The

twist-to-flex ratio of zero is pure bending at 87 percent span while the twist-to-flex ratio of

infinity is pure torsion. The amplitude of the movement of airfoil at each span-wise location is

similar to the Case 14a, which has a twist-to-flex ratio of approximately 0.4. With reduced

frequency fixed, Figure 67 and Figure 68 show that it is more stable when twist-to-flex ratio is

small. This aeroelastic sensitivity study has quantified the sensitivity of aerodynamic damping to

changes in the frequency and the mode shape, which will be useful in assessing the design tradeoff during the QHSF II design effort.

m

%damping

150 160 170 180 190 200 210 0

0.25

0.5

0.75

1

Honeywell (Base)

Twist/Flex=0

Twist/Flex=inf.

Figure 67. Twist-to-Flex Ratio Sensitivity: Damping Extrapolation With Mass Flow Rate

(m) Is Shown at 89% Speed for the Case 14a Rotor.

NASA/CR—2012-217451 69
pr

%damping

1.55 1.6 1.65 1.7 1.75 1.8 0

0.25

0.5

0.75

1

Honeywell (Base)

Twist/Flex=0

Twist/Flex=inf.

Figure 68. Twist-to-Flex Ratio Sensitivity: Damping Extrapolation With Pressure Ratio

(pr) Is Shown at 89% Speed for the Case 14a Rotor.

Figure 69 shows a summary of the aeroelastic analysis of the Case 14a rotor. A qualitative

assessment of the predicted instability points produced the stability line on the figure.

The viscous TURBO analysis indicates that the fan will have a potential stability problem at

70 percent speed. The near stall steady-flow results at 70 percent speed show flow separation on

the suction surface at about 75 percent span. The Case 14a fan was rerun using TURBO in the

inviscid mode to compare to viscous solution obtained from NASA. It was first verified that the

steady inviscid solution did not show any flow separation. Figure 70 shows the mass flow rate

versus pressure ratio map from the TURBO steady solutions for the 70 percent speed line. As

expected, the inviscid TURBO solution predicts higher flow with the same pressure ratio (or higher

pressure-ratio with the same flow) than the viscous solution. For each operating point, TURBO

unsteady flutter solutions were run for nodal diameters of 0, 2, 4, 6, and –2. Typical unsteady

convergence histories of the damping are shown in Figure 71. The unsteady runs are fully

converged for a total of 20 vibratory cycles from the steady state solutions. The resulting damping

versus nodal diameter comparison is given in Figure 72. The minimum damping occurs between

nodal diameters 0 and 2. The damping versus mass flow rate comparison of the inviscid results

with the viscous results is given in Figure 73. The inviscid results show that instability point occur

at a slightly lower mass flow rate than viscous results, but the design still becomes unstable in the

operating rage below stall. This small shift in the stability point is probably due to the viscous

solution showing a separation which inviscid solution did not capture.

NASA/CR—2012-217451 70
Rotor Pressure Ratio

Inlet Corrected Flow (lbm/s)

Increasing Speed

SS PS

SS PS

SS PS

SS PS

Op Line

TURBO Steady

TURBO Unsteady

Case 14a Predicted Flutter Boundary

Stall line

Stable

Neutral

Unstable

Damping

SS = Suction Side

PS = Pressure Side

Figure 69. Analysis of the Case 14a Rotor Blade With TURBO Shows That the Blade Is

Unstable Near the Operating Line at an Intermediate Speed Condition.

1.0

1.1

1.2

1.3

1.4

1.5

110 120 130 140 150 160 170 180 190 200 210

Inlet Corrected Flow (lbm/s)

Rotor Pressure Ratio

TURBO 70% Nc

TURBO_INV_70%

Figure 70. The Steady Flow Results From Inviscid TURBO Show the Expected Change in

Pressure Ratio and Flow Characteristics From the Viscous TURBO Results.

NASA/CR—2012-217451 71
0.00

0.50

1.00

1.50

2.00

2.50

3.00

0 5 10 15 20

Cycle

Damping

0 ND

2 ND FTW

4 ND FTW

6 ND FTW

Figure 71. A Satisfactory Convergence History Was Obtained With the Viscous TURBO

Analysis.

0.0

0.5

1.0

1.5

2.0

2.5

3.0

-4 -20 2 4 6 8

Nodal Diameter

% Damping

Figure 72. The Minimum Damping for the Inviscid Analysis of the 70% Speed Line

Occurs Between 0 and 2 Nodal Diameters.

NASA/CR—2012-217451 72
mass flow rate

damping

140 145 150 155 160 165 170 175 180 0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

0.55

0.6

Viscous

Inviscid

Figure 73. The Inviscid TURBO Results Show a Slight Improvement in Stability Over the

Viscous Results.

The TURBO results for Case 14a showed a small region of flow separation (Figure 74 and

Figure 75). It was decided to continue the analysis with rotor ITER07, on which the rotor

performance was closer to design intent and no separation was seen (Figure 76). The previous

Case 14a TURBO analyses were conducted with mode shapes from an airfoil-only ANSYS

model. It was discovered that there was a significant change in the mode shapes and frequencies

for Case 14a with the attachment (unlike the results for QHSF I). It appears that rotor blade

ITER07, with blade/attachment mode shapes and frequencies, is a stable configuration.

NASA/CR—2012-217451 73
(a) Case 14a (b) ITER07

Figure 74. Mach Number Contours in the Blade Passage Show a Region of Flow

Separation in the Case 14a Rotor Design.

(a) Case 14a (b) ITER07

Figure 75. Mach Number Contours Near the Blade Suction Surface Show a Region of Flow

Separation in the Case 14a Rotor Design.

NASA/CR—2012-217451 74
mass

damping

140 150 160 170 180 0

0.1

0.2

0.3

0.4

0.5

0.6

14GoFwd, Viscous, 70%Nc

14GoFwd, Inviscid, 70%Nc

Iter07, Inviscid, 70%Nc

Figure 76. The Aerodynamic Damping as a Function of Mass Flow Rate (Mass) Shows the

Small Effect of the Separation Region on the Blade Stability.

5.4 Justification for the Use of the TURBO Evaluation of the Case 14a Rotor

The TURBO code was a key element in the design of QHSF II. The QHSF I design relied

on the empirical guidelines available at that time to determine the acceptability of the fan blade

design from a stability standpoint. The stability assessment was based on the consideration of

reduced frequency, defined as

V

b k ω =

where

k = reduced frequency

b = half-chord (true chord/2)

ω = circular frequency

V = reference velocity

Honeywell has defined critical reduced frequency values based on experience to assure a

stable design; the reduced frequency must be above the critical values of

k > 0.165 for pure bending

k > 0.80 for pure torsion

NASA/CR—2012-217451 75
For these calculations, the parameters are based on values at 75 percent span for design

point conditions. (In the QHSF I final report, an equivalent parameter called the “flutter

parameter” was used. Honeywell has since adopted the industry standard “reduced frequency”

and the criteria have been updated accordingly.)

For the QHSF I, the reduced frequency was calculated for each mode. Then each mode was

classified as either a bending mode or a torsion mode so that the reduced frequency could be

compared to the appropriate criterion. This classification was based on a calculation of the twistto-flex ratio, which quantifies the amount of torsion in a particular mode and is defined by

h

bα

=ψ

where

ψ = twist-to-flex ratio

b = half-chord (true chord / 2)

α = angular deflection in mode shape (pitching motion)

h = translational deflection in mode shape (plunging motion).

The value of the twist-to-flex ratio at the 75 percent and 95 percent spans was determined.

For modes with small values of twist-to-flex ratio, the bending criterion was applied; for larger

values, the torsion criterion was used. For example, mode 1 had a twist-to-flex value at each

span of approximately 0.4, and the mode was classified as a bending mode. The reduced

frequency was about 0.3, which is above the criterion of 0.165 for pure bending modes, and so

the blade should be stable based on this empirical guideline. Other modes similarly met the

appropriate guidelines.

Obviously, this approach was not successful in QHSF I, as flutter was encountered for

mode 1 at part speed conditions near the stall line. There were a number of contributors to the

breakdown of the design criteria. Note that in the empirical approach outlined above, there is no

consideration of engine speed, operating condition along a speed line, incidence, shock location,

contribution from tip effects, etc. All of these are known from experimental data and previous

computational studies to be important contributors to the actual flutter behavior. Flutter is also

known to be very sensitive to mode shape. The mode shape was considered only to calculate

twist-to-flex to determine the criterion to be used, and the classification was based on judgement.

(Note that if mode 1 had been classified as a torsion mode, flutter would have been expected.)

At the time, there was no criterion for modes with intermediate values of twist-to-flex ratio.

The QHSF I experience is an example that highlights the need for advanced computational

tools such as TURBO-AE, and demonstrates that the use of the tool is crucial to have a

successful redesign effort in the QHSF II program. With these advanced tools, the effects that

are known to be important can be addressed directly. The entire blade is included, rather than

just a single representative span. The actual mode shape calculated by finite element analyses is

used. And the steady flow field is based on the actual speed and operating conditions.

The potential benefits of using TURBO-AE are clear, given that it successfully predicts the

QHSF I design to be unstable. If the tool had been available during the original design, the

NASA/CR—2012-217451 76
flutter would have been predicted and changes could have been incorporated during the design

phase to eliminate the problem. A second benefit has already occurred during the QHSF II

design work. We had assumed up to this time that the forward sweep of the QHSF I was a

significant de-stabilizing influence. However, our results have clearly shown that this is not true,

and forward sweep actually has a small benefit for flutter. This insight has had a significant

effect on our design approach.

TURBO-AE was applied to predict the aeroelastic stability of the Case 14a rotor and

determined that it was not stable on the 75 percent speed line near stall. This result would not

have been determined from analysis of reduced frequency and twist-to-flex ratio. Figure 77

shows a plot of the reduced frequency as a function of incidence. Two points are shown on the

figure for each speed: one at peak efficiency and one near the stall line. There is nothing in this

data that signals that the 70 percent speed line should be unstable as compared to the other speed

lines. A similar conclusion is reached from the reduced frequency vs. twist-to-flex ratio plot

shown in Figure 78.

Incidence

Reduced Frequency (

ωb/V)

0 2 4 6 8 10 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0

100%

89%

70%

60%

Stable

Unstable

Figure 77. Evaluation of the Reduced Frequency of the Case 14a Rotor Blade Does Not

Identify the Instability at 70% Speed.

NASA/CR—2012-217451 77
Figure 78. Evaluation of the Reduced Frequency of the Case 14a Rotor Blade Does Not

Identify the Instability at 70% Speed.

In summary, the reduced frequency limit is a useful guideline to distinguish blade designs

and/or modes that are either very stable (so that further detailed analysis is not necessary) or very

unstable (so that a significant design change is needed to eliminate flutter) from marginally

stable designs. As such, it can be a very effective screening tool. Recent experience indicates

that modern fan designs result in blades where factors such as mode shape and operating

conditions must be taken into account. In these cases, the use of an advanced computational tool

such as TURBO-AE is crucial in properly determining the stability of the design. In the QHSF

II, we are using TURBO-AE as part of a Design of Experiments approach to help identify the

factors that are de-stabilizing the blade. TURBO-AE is needed in the detailed blade design

process to define an aeroelastically stable fan design.

5.5 Final Rotor Optimization

An ADPAC 3-D viscous CFD model of the Case 14a rotor and the Baseline II stator was

developed to do detailed analyses of the fan stage. This model was used to optimize the rotor

blade thickness and incidence distribution.

The ADPAC full-stage model is comprised of two rotor blocks, one core stator block, one

core duct/strut block, one bypass stator block and one bypass duct/strut block. This split-flow

modeling technique, using separate core and bypass stream throttle pressures, has been shown to

be necessary to properly establish the prescribed fan bypass ratio using current CFD codes.

Otherwise, the rotor passage shock in the tip region will not be properly located for the

Twist-to-Flex Ratio (αb/h)

Reduced Frequency (

ωb/V)

0 0.25 0.5 0.75 1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

100%

89%

70%

60%

87.5% span

75% span

Stable

Unstable

NASA/CR—2012-217451 78
aerodynamic design reference conditions. The complex grid structure for the full-stage ADPAC

model introduced a new complication to the post-processing analysis due to the highly skewed

grid surfaces in the rotor (see Figure 79). Grid surfaces could no longer be assumed to

approximate streamline surfaces through the rotor passage, prompting a modification to the postprocessing code that provides flow properties along quasi-streamlines.

Figure 79. Comparison of the ADPAC Computational Grid for the Split-Stator

Configuration With the Streamline Pattern.

Since the CFD analysis of the rotor for DOE 1 was performed in TURBO, it was desirable

to compare the TURBO and ADPAC analysis results for Case 14a to ensure that the transition to

a different software tool did not change the performance of Case 14a. Figure 80 shows a

comparison of the Mach number contours at two different radial positions for the two codes.

Figure 81 shows a comparison of the Mach number contours near the suction and pressure

surfaces for the two codes. Little difference is seen for the two models.

NASA/CR—2012-217451 79
(a) TURBO – 60.1% span (b) ADPAC – 60.1% span

(a) TURBO – 80.0% span (b) ADPAC – 80.0% span

Figure 80. A Comparison of the TURBO and ADPAC Mach Number Contours Shows No

Significant Flowfield Differences.

NASA/CR—2012-217451 80
(a) TURBO – Suction Surface (b) ADPAC – Suction Surface

(a) TURBO – Pressure Surface (b) ADPAC – Pressure Surface

Figure 81. A Comparison of the TURBO and ADPAC Mach Number Contours Shows No

Significant Flowfield Differences.

To begin the process of optimizing the Case 14a rotor aerodynamic design, several rotor

airfoil models with different mean line angle distributions were constructed using the streamline

curvature/airfoil generator code. Vibration characteristics of the resultant blades (i.e., airfoil plus

attachment) were analyzed. CFD analyses of the airfoils were performed on the candidate mean

line distributions.

NASA/CR—2012-217451 81
In response to the decision to return to the full-span stator design of QHSF I, the ADPAC

model was then modified to include the full span stator design. After completion of the revision,

the rotor evaluations were resumed.

For the Case 14a rotor aerodynamic design, several rotor airfoil models with different mean

line angle distributions were analyzed with the ADPAC code to optimize rotor performance.

Airfoil changes include modifications to incidence, angle passage area distribution, and turning,

such that rotor performance was brought closer to design point objectives. These initial ADPAC

analyses show that design flow can be achieved as shown in Figure 82.

1.80

1.85

1.90

1.95

2.00

216 218 220 222 224

Total Mass Flow, lb/s

Rotor Pressure Ratio

Rotor Goal

QHSFII,case14a,APNASA

QHSFII,case14a,ADPAC

ITER03, ADPAC

ITER04, ADPAC

Figure 82. Preliminary ADPAC Results Show That the Case 14a Rotor Has the Potential

to Meet the Design Point Flow and Pressure Ratio by Adjusting the Mean Line

Angle Distribution.

A total of 25 configurations were analyzed with ADPAC (referred to as ITER01 to ITER25).

A Campbell diagram (shown in Figure 83) of design iteration ITER07 indicated a mode

2/3E crossing at 100 percent RPM. A study was completed to assess tradeoffs between blade

and attachment weight for optimum mechanical performance. Figure 84 illustrates the ITER07

status relative to the mechanical design goal. Based on the results of this study, a sloped

attachment was selected, which results in adequate frequency margin with minimal aeroelastic

risk. Figure 85 confirms that the design goals have been achieved with the sloped attachment.

NASA/CR—2012-217451 82
1E 

2E 

3E 

10E 9E 8E 7E 6E 5E 4E 

0

250

500

750

1000

0 1600 3200 4800 6400 8000 9600 11200 12800 14400 16000

Speed [rpm]

Frequency [Hz]

1st Bending

2nd Bending

Figure 83. The ITER07 Case Shows Mode 2/3E Crossing at 100% RPM.

2.9

3.0

3.1

3.2

3.3

3.4

-10% -5% 0% 5% 10%

Delta CF Load, %

Mode 2 Frequency, /Rev

Goal (10% Margin)

3E Resonance

Straight

Attachment

Sloped

Attachment

Oct Status

Design

10 deg Slope

Status Design

BC

Effect

BC

Effect

Figure 84. A Design Study Completed to Restore Adequate Frequency Margin Suggested

Sloped Attachment Solution.

NASA/CR—2012-217451 83
10E 9E 8E 7E 6E 

5E 

4E 

3E 

2E 

1E 

0

250

500

750

1000

1250

1500

0 1600 3200 4800 6400 8000 9600 11200 12800 14400 16000

Speed [rpm]

Frequency [Hz]

QHSF II (Iter18h)

QHSF I (22" Rig) 

15621 RPM

(ADP)

First Bending Mode

Second Bending Mode

First Torsion Mode

15842 RPM

(MDP)

Figure 85. The Campbell Diagram for the ITER07 Rotor Blade With Sloped Attachment

Shows Adequate Design Margin.

5.6 Stator Design

15 split-span stator vane configurations were defined for the first stator DOE as shown in

Figure 86. The axial sweep was kept constant for this evaluation. The acoustic and mechanical

evaluations were completed to determine the optimum stator for the Case 14a rotor. However,

after extensive evaluation of the SOURCE3D and ANSYS results, it was determined that the

conservative design at the hub (leaning against the direction of fan rotation) to prevent suction

side flow separation was a significant negative influence on the acoustic results. Three of the

DOE 1 cases (1, 5, and 7) that had significant tip sweep in the direction of rotation were carried

forward into stator DOE 2 (see Figure 87). The lean distribution labeled “Case 1 Mod” is the

Case 1 profile that has been adjusted to have zero lean at the hub. This profile was added to

access the impact of the suction side lean on the tone noise reduction. It was also decided to

perform DOE 2 with a 50 vane stator instead of the 70 vane stator of DOE 1 to reduce the

broadband noise levels and have a stator count similar to the QHSF I (52 vanes).

NASA/CR—2012-217451 84
-1.5

-1

-0.5

0

0.5

1

1.5

10.5 11.5 12.5 13.5 14.5 15.5

Case 1

Case 2

Case 3

Case 4

Case 5

Case 6

Case 7

Case 8

Case 9

Case 10

Case 11

Case 12

Case 13

Case 14

Case 15

15 Degree Lean

-15 Degree Lean

-25 Degree Lean

25 Degree Lean

YSD = Non-radial variation

at the stator leading edge

Direction of

Rotor Rotation

Figure 86. Comparison of the Stator Lean Profiles (YSD Parameter) for the 15 Cases of

Stator DOE 1.

-1.5

-1

-0.5

0

0.5

1

10.5 11 11.5 12 12.5 13 13.5 14 14.5 15 15.5 16

Radius

YSD

DOE 1 Case 1 YCG DOE 1 Case 5 YCG

DOE 1 Case 7 YCG DOE 1 Case 1 Mod YCG

Figure 87. These Four Circumferential Lean Distributions Were Used for Stator DOE 2.

NASA/CR—2012-217451 85
To overcome the negative impact on noise of the vane suction side lean at the hub, both

non-linear sweep in the axial direction and non-linear lean on the tangential direction were

explored in DOE 2. The design concept is to apply sweep where the vane cannot be leaned and

apply lean where the vane cannot be swept. The concept is shown in Figure 88.

0

1

2

3

4

5

0 0.5 1 1.5 2

XSD (in)

R - Rhub (in)

13 Degrees Linear

30-18-5 Degrees

45-20-0 Degrees

0

1

2

3

4

5

1 0.5 0 -0.5 -1 -1.5 -2

YCG Shift (in)

Large sweep is used

to offset acoustically

adverse lean at hub

 Small sweep at tip is

offset by large lean.

Figure 88. The Design Approach for Stator DOE 2 Is to Apply Non-Linear Sweep to Take

Maximum Advantage of the Optimum Lean.

Two nonlinear axial sweeps were defined for stator vane DOE 2 in addition to the nominal

linear sweep as shown in Figure 89. Constraints on the vane position at the hub and shroud for

the 50-vane configuration limited the linear sweep to 13 degrees.

NASA/CR—2012-217451 86
-2

-1.5

-1

-0.5

0

10.5 11 11.5 12 12.5 13 13.5 14 14.5 15 15.5 16

Radius

XSD

Straight 13 Degree Sweep

30-18-5 4th Order Sweep

45-20-0 4th Order Sweep

Figure 89. These Three Axial Sweep Distributions Were Used for Stator DOE 2.

Based on the acoustic results, Case 109 (see Table 15) was chosen as the go-forward

design. Case 109 has the 45-20-0 4th order sweep profile and the Case 1 tangential lean profile

from DOE 1.

Table 15. Tone Sound Power Results of Stator DOE 2 From SOURCE3D at 61.7%

Corrected Fan Speed.

Sweep 1 = Sweep 1 = Sweep 1 =

Straight 30-18-5 45-20-0

YCG = Bypass Stator Case 104 Case 108 Case 112

DOE 1 Case 1m -2.0 -2.6 -2.6

YCG = Bypass Stator Case 101 Case 105 Case 109

DOE 1 Case 1 -1.3 -1.9 -2.2

YCG = Bypass Stator Case 102 Case 106 Case 110

DOE 1 Case 5 -0.1 0.1 0.4

YCG = Bypass Stator Case 103 Case 107 Case 111

DOE 1 Case 7 0.5 1.5 1.9

The decision was made for the QHSF II to return to the full span stator configuration since

the optimum core and bypass stator counts were both 50. Confirmation runs were performed to

ensure that the full span stator preserved the tone noise reduction benefit obtained with the split

span stator design of DOE II. Figure 90 shows the summary of the analysis. SOURCE3D

predicts a 3 dB reduction in tone sound power for the QHSF II go forward design as compared to

the Baseline II rotor at 61.7 percent speed. Figure 91 shows a two view drawing of the goforward vane design.

NASA/CR—2012-217451 87
-3 dB

-1 dB

-1 dB

Reduction

below

 baseline

baseline

QHSF II rotor / Baseline II stator

QHSF II rotor / QHSF II stator

Baseline II rotor / Baseline II stator

Baseline II rotor / QHSF II stator

Figure 90. Summary of the Total Rotor/Stator Interaction Tone Sound Power Reduction

at a Typical Aircraft Approach Condition Due to the Elements of the QHSF II.

Side View Fwd Looking Aft

Figure 91. The Go-Forward Design for the QHSF II Stator Is a Full Span Vane With

Nonlinear Axial Sweep and Tangential Lean.

A confirmation analysis of the go-forward stator mechanical design was conducted with

ANSYS. As expected, the vane has atypical vibration mode shapes and the flutter parameters

are not all within the Honeywell design experience (see Figure 92).

NASA/CR—2012-217451 88
Flutter Param.=0.6492 Flutter Param.=0.8948

(a) Mode 1 (b) Mode 2

Figure 92. Preliminary Mode Shapes for the QHSF II Stator Vane Show a Potential Mode

2 Flutter Problem.

5.7 SOURCE3D and V072 Studies for QHSF II Stator DOE I

The fan tone noise calculations for the first stator DOE were performed using an engine

scale fan flowpath as shown in Figure 93. The fan had a forward-swept rotor and a split-span

stator. The tone noise predictions were performed using the SOURCE3D program, which is part

of NASA’s TFaNS fan tone noise prediction tool, and the V072 program, an earlier tool upon

which the SOURCE3D program is based.

0

2

4

6

8

10

12

14

16

18

-8 -6 -4 -2 0 2 4 6 8 10 12 14

Axial Distance

Radius

Figure 93. A Split-Stator Fan Flow Path Was Used for the Initial Stator DOE Tone Noise

Calculations.

NASA/CR—2012-217451 89
5.7.1 Calibration of SOURCE3D With Straight-Lean Stators

To calibrate the results of the SOURCE3D program, tone noise predictions for the bypassportion of the QHSF II fan were examined at three speeds (62 percent, 77 percent, and 89 percent).

The stators of DOE I were replaced by a set of 5 stators with straight lean of +25 degrees, +15

degrees, 0 degrees - straight radial, -15 degrees, and –25 degrees. Positive lean angles were

defined as lean in the direction of rotor rotation. (Note that the SOURCE3D/V072 variable YSD is

negative, for vane lean in the direction of rotor rotation.) The behavior of these cases was expected

to follow the trend of increased tone noise as the stator was leaned against rotation.

The SOURCE3D results showed different trends at each speed, as presented in Figure 94 to

Figure 96. Depending on which circumferential modes dominated, the sound power level was

either nonlinear (62 percent speed), constant (77 percent speed), or linear (89 percent speed) with

stator lean. At 77 percent and 89 percent speed, the 1*BPF (m = 22) and 2*BPF (m = 44) rotorlocked modes had constant PWL for all stator leans. At all speeds, the 2*BPF (m = -26)

circumferential mode showed nonlinear behavior with lean. Also, the 3*BPF (m = -4) mode was

nonlinear with lean, at 62 percent speed.

All radial modes were constant for the rotor-locked circumferential modes. All other radial

modes demonstrated non-linear variation with stator lean. It was the combination of these radial

modes that determined the overall behavior of the PWL trend with stator lean.

90.0

95.0

100.0

105.0

110.0

115.0

120.0

125.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 94. SOURCE3D Results for QHSF II at 62% Speed, With Straight-Leaned Stators.

NASA/CR—2012-217451 90
110.0

115.0

120.0

125.0

130.0

135.0

140.0

145.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 1*BPF

Inlet - 2*BPF

Exit - 1*BPF

Exit - 2*BPF

Figure 95. SOURCE3D Results for QHSF II at 77% Speed, With Straight-Leaned Stators.

110.0

115.0

120.0

125.0

130.0

135.0

140.0

145.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 1*BPF

Inlet - 2*BPF

Exit - 1*BPF

Exit - 2*BPF

Figure 96. SOURCE3D Results for QHSF II at 89% Speed, With Straight-Leaned Stators.

5.7.2 Comparison of SOURCE3D and V072 at 62 Percent Speed

Review of the QHSF I Design Report showed that the V072 analyses performed for the

QHSF I stator at 55.9 percent speed predicted a fairly linear variation of PWL with stator lean,

unlike the 62 percent speed SOURCE3D predictions. To determine if V072 was consistent with

NASA/CR—2012-217451 91
the predictions from SOURCE3D for the QHSF II, the set of straight-lean stators was studied at

62 percent speed, using V072 with input files generated by PREV072 (a preprocessor to V072

developed at Honeywell).

The V072 input files were generated with all 24 streamlines from AXCAPS, in contrast to

the SOURCE3D input files, which used a subset of 17 streamlines. The straight-lean stator

geometry was copied from the SOURCE3D input files, and additional data was added to account

for the additional streamlines. One difference between the input files concerned the

representation of YRD, the circumferential offset of the rotor trailing edge, relative to the radial

direction. The scheme for calculating YRD for a split-span stator in PREV072 resulted in bypass

YRD values that were not zero-based at the flow splitter radius. It was not clear that this was

necessary within the SOURCE3D/V072 algorithms; however, the SOURCE3D input had used

zero-based YRD distributions.

Comparison of the SOURCE3D and V072 results at 62 percent speed (Figure 94 and Figure

97, respectively) showed that the V072 overall PWL values were higher. Also, the V072 results

showed less non-linearity with stator lean compared to SOURCE3D. This behavior was even

more apparent when the differences in PWL for the leaned stators relative to the unleaned stators

were compared for SOURCE3D and V072. As shown in Figure 98 and Figure 99, the behavior of

the V072 case was more consistent with the expected trends. In general, the variation in results

between SOURCE3D and V072 was of a similar magnitude to the variations due to stator lean.

90.0

95.0

100.0

105.0

110.0

115.0

120.0

125.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 97. V072 Results for QHSF II at 62% Speed, With Straight-Leaned Stators.

NASA/CR—2012-217451 92
-6

-4

-2

0

2

4

6

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 98. Comparison of the SOURCE3D Results Relative to Unleaned Stator.

-6

-4

-2

0

2

4

6

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 99. Comparison of V072 Results Relative to Unleaned Stator.

NASA/CR—2012-217451 93
5.7.3 Modification of PREV072 Calculation of YRD

The circumferential offset of the rotor trailing edge, relative to the radial direction (YRD)

was being computed in PREV072 inconsistently with its usage in V072. YRD was computed in

PREV072 using the offset angle relative to the hub and the local radius at the rotor trailing edge.

Although this produced an accurate circumferential offset, the V072 program did not use YRD in

the same way. In V072, the offset angles that were obtained from YRD were computed using

the reference radii (i.e., the RADIUS array specified in the input file), which in this case was at

the bypass stator leading edge.

PREV072 was modified to compute YRD based on the reference radius. In addition, for

bypass-only analyses, the offset was recomputed to have a zero-base at the flow splitter. Results

of this modification are shown in Figure 100. Only a minimal change in PWL was seen for the

modification to YRD. This was even more apparent in Figure 100, where trends for the

differences in PWL relative to the unleaned stator remained similar, with the revised input files.

This would seem to indicate that the zero- and non-zero-based YRD distributions result in

essentially the same rotor wake behavior at the stator leading edge.

As a check of the validity of the original SOURCE3D input, the new correctly computed

YRD distribution for the V072 input was compared with the YRD distribution used in

SOURCE3D. There was a significant difference in the two distributions. Although this

discrepancy in rotor trailing edge circumferential offset was constant across all cases examined,

it may have contributed to the disparity in the SOURCE3D and V072 results.

90.0

95.0

100.0

105.0

110.0

115.0

120.0

125.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 100. V072 Results Based on the Revised Specification of YRD.

NASA/CR—2012-217451 94
Delta Power Level at 62% Speed

(V072 w/New Zero-Based YRD)

-6

-4

-2

0

2

4

6

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 101. Comparison of V072 Results Relative to Unleaned Stator, With Revised

Specification of YRD.

5.7.4 Comparison of QHSF I Cases With QHSF II

After modifying PREV072, new V072 analyses were performed for the Baseline I fan and

QHSF I, along with the Baseline II straight stator and a QHSF II leaned stator case. Speeds were

not directly comparable; however, all cases were near Approach. The Baseline I and QHSF I

cases were at 55.9 percent Speed, the Baseline II case was at Approach, and the QHSF II fan

with the +15 Degree leaned stator was at 62 percent Speed. Results are shown in Figure 102.

General sound power level trends of the Baseline I, QHSF I, Baseline II, and QHSF II fans

indicated that the QHSF II was comparable to the Baseline II fan, and Baseline II fan had louder

tones than the old baseline and QHSF I.

Comparison of the QHSF I case relative to the Baseline I fan showed somewhat different

trends (varying by several dB) than those presented in the QHSF I Design Report (Reference 1).

In the final report, the differences were generally reported to be much greater. A contributing

factor may be the underestimate of the radial loss distribution specified for the Baseline I rotor in

the original QHSF I evaluations. This original loss model was replaced by the QHSF I loss

distribution for the current V072 analysis of the Baseline I fan.

NASA/CR—2012-217451 95
80.0

85.0

90.0

95.0

100.0

105.0

110.0

115.0

120.0

Baseline I (55.9%

Speed)

QHSF I (55.9%

Speed)

Baseline II APP

QHSF II +15 Deg

(62% Speed)

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

(6140 RPM)(6140 RPM)

(5843 RPM)(6618 RPM)

Figure 102. Comparison of the QHSF I and QHSF II Cases With the Baseline I and

Baseline II Cases.

5.7.5 Adjustment for Low Cutoff Ratio

It was noted that some of the radial modes had cutoff ratios very close to 1.0, along with

very large predicted sound power levels. Because the accuracy of the cutoff ratio calculation in

V072 was not well established, there remained some question as to the validity of the sound

power levels for these modes. To determine the impact of the modes with low cutoff ratio, the

V072 results were adjusted by computing the sound power levels after discarding any modes

having cutoff ratios less than or equal to 1.1. The primary impact of this adjustment was on the

QHSF II case, which was reduced significantly in overall sound power level, as shown in Figure

103. Compared to the unfiltered results shown in Figure 102, this sound power level indicates a

substantial decrease in fan tone noise for the QHSF II stator with +15 degrees of straight lean.

To determine the impact on sound power level with stator lean, results for the 5 straightlean stator cases at 62 percent speed were filtered to remove any mode with a cutoff ratio less

than or equal to 1.1. The results of this analysis are shown in Figure 104. The cutoff ratio

filtering impacted only the 2*BPF modes, reducing them significantly, relative to the unfiltered

modes shown in Figure 100. As a result, the overall sound power levels were also substantially

reduced for the filtered predictions.

NASA/CR—2012-217451 96
V072 Power Levels

(Adjusted for Low Cutoff Ratios)

80.0

85.0

90.0

95.0

100.0

105.0

110.0

115.0

Baseline I (55.9%

Speed)

QHSF I (55.9%

Speed)

Baseline II APP

QHSF II +15 Deg

(62% Speed)

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

(6140 RPM)(6140 RPM)

(5843 RPM)(6618 RPM)

Figure 103. Comparison of QHSF I and QHSF II Cases Adjusted for Low Cutoff Ratio.

90.0

95.0

100.0

105.0

110.0

115.0

120.0

125.0

Lean=25Lean=15

Lean=0

Lean=-15Lean=-25

LOGSUM

Inlet - 2*BPF

Inlet - 3*BPF

Exit - 2*BPF

Exit - 3*BPF

Figure 104. Variation in PWL for QHSF II With Stator Lean at 62% Speed, Adjusted for

Low Cutoff Ratio.

NASA/CR—2012-217451 97
5.7.6 Comparison of Rotor Loss Profiles

Before reaching any conclusions concerning the QHSF II stator V072 analyses, additional

issues were considered. One area of concern was the rotor loss distribution for the QHSF II.

Rotor loss distributions for the QHSF I at 55.9 percent speed, the Baseline II fan at Approach,

and QHSF II at 62 percent speed were compared, as shown in Figure 105. The QHSF I and

Baseline II profiles appeared to be similar; however, the QHSF II distribution was quite low in

the outer span region.

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0 10 20 30 40 50 60 70 80 90 100

Percent Span

Loss

QHSF I - 55.9% Spd

Baseline II Approach

QHSF II - 62% Spd

QHSF II Core QHSF II Bypass

Figure 105. Radial Loss Distributions From AXCAPS.

5.8 Final QHSF II Design

The final aerodynamic design of the rotor and stator for the QHSF II was selected to be

Case 18h (which was derived from Rotor ITER18 and Stator Case 109). Case 18h is the

optimum configuration resulting from a series of analytical DOEs that were subsequently further

optimized for reduced mechanical stress and improved aerodynamic performance. Figure 106

depicts the geometry of the final QHSF II design relative to the baseline engine.

NASA/CR—2012-217451 98
FLA side FLA side

Baseline II QHSF II rotor / QHSF II stator

Iter18h

Figure 106. QHSF II Reduces Noise Through Rotor Sweep and Stator Sweep and Lean.

5.8.1 Aerodynamic Performance

Figure 107 shows the aerodynamic performance of the QHSF II rotor at 100 percent

corrected fan speed. Figure 108 shows the aerodynamic performance of the QHSF II stage at

100 percent corrected fan speed. The QHSF II stage meets pressure ratio and efficiency goals.

Suction-side Mach number contours are shown in Figure 109 for the QHSF II and Baseline II

rotors. Figure 110 shows the Mach number contours at the rotor exit. The rotors show very

similar aerodynamic performance.

1.75

1.80

1.85

1.90

1.95

2.00

205.0 210.0 215.0 220.0 225.0

Fan Inlet Flow, lb/s

Rotor Pressure Ratio

Baseline II Rotor

QHSF II Goal

QHSF II Design

0.86

0.88

0.90

0.92

0.94

0.96

205.0 210.0 215.0 220.0 225.0

Fan Inlet Flow, lb/s

Rotor Efficiency

Baseline II Rotor

QHSF II Goal

QHSF II Design

Figure 107. The QHSF II Rotor Meets Pressure Ratio and Efficiency Goals Set for the

Program.

NASA/CR—2012-217451 99
1.70

1.75

1.80

1.85

1.90

1.95

205.0 210.0 215.0 220.0 225.0

Fan Inlet Flow, lb/s

Stage Pressure Ratio

Baseline II Fan

QHSF II Goal

QHSF II Design

0.84

0.86

0.88

0.90

0.92

0.94

205.0 210.0 215.0 220.0 225.0

Fan Inlet Flow, lb/s

Stage Adiabatic Efficiency

Baseline II Fan

QHSF II Goal

QHSF II Design

Figure 108. The QHSF II Stage Meets Pressure Ratio and Efficiency Goals Set for the

Program.

 QHSF II Rotor Baseline II Rotor

Figure 109. The Results of the ADPAC Analyses Show the Differences in Mach Number

Contours on the Suction Side of the Blade Between the Baseline II and QHSF

II Rotor.

NASA/CR—2012-217451 100
 QHSF II Rotor Baseline II Rotor

Figure 110. The Results of the ADPAC Analyses Show the Differences in Mach Number

Contours at the Rotor Exit Between the Baseline II and QHSF II Rotor.

Figure 111 shows the suction side Mach number contours for the QHSF II and Baseline II

stators. Figure 112 shows the Mach number contours at the vane exit. There is a small amount

of flow separation introduced in the QHSF II stator design relative to the Baseline II design. The

separation was estimated to be worth approximately 0.5 point in efficiency as shown in Figure

113 and has an undetermined noise impact judged to be of low risk.

 QHSF II Stator Baseline II Stator

Figure 111. The Results of the ADPAC Analyses Show the Differences in Mach Number

Contours on the Suction Side of the Vane Between the Baseline II and QHSF

II Stator.

NASA/CR—2012-217451 101
 QHSF II Stator Baseline II Stator

Figure 112. The Results of the ADPAC Analyses Show the Differences in Mach Number

Contours at the Vane Exit Between the Baseline II and QHSF II Stator.

0

20

40

60

80

100

1.75 1.8 1.85 1.9

Overall Stage Pressure Ratio

Stator Exit Span ~ 

%

Nominal Pr Profile

Modified Pr Profile

Modified Pr Profile Assumes 

No Defect in

Region of Stator Flow

Separation

0

20

40

60

80

100

70 80 90 100

Overall Stage Efficiency ~ %

Stator Exit Span ~ 

%

Nominal Eff Profile

Modified Eff Profile

Stator Pressure Ratio 

"Defect" Results in 

0.5-point Loss in 

Stage Efficiency

Figure 113. A Quick Analysis of the Stator Separation Indicated a Small Reduction in

Efficiency for the QHSF II Stator.

NASA/CR—2012-217451 102
As part of the original design goals for the QHSF I program, an attempt was made to adjust

the rotor shock position at the critical takeoff condition so that the shock would be totally

captured in the blade passage. Unfortunately, the multidiscipline optimization process of DOE 1

led to a design that did not achieve shock capture. Figure 114 shows Mach number contours for

a typical take-off condition at four spanwise radii on the blade. It can be seen from the figure

that the shocks are not contained in the blade passage. Figure 115 shows the Mach number

contours near the pressure and suction side of the blade surface. The spanwise variation in shock

position is clearly shown.

(a) 60.1% span (b) 68.3% span

(c) 80.0% span (d) 92.1% span

Figure 114. Mach Number Contours From the TURBO Viscous Analysis at 89% Speed

Show That Shock Capture Was Not Achieved for the QHSF II Fan.

NASA/CR—2012-217451 103
(a) Suction Surface (b) Pressure Surface

Figure 115. Mach Contours From the TURBO Viscous Analysis Show the Shock Positions

for the 89% Speed Condition.

5.8.2 Mechanical Performance

A sloped attachment of 4.5 degrees was selected for the QHSF II rotor design as shown in

Figure 116. This slope is adequate to put the blade-out loads equal to Baseline II engine levels,

meet frequency goals, and provides adequate flutter margin. A state-of-the-art finite element

model of the rotor blade and attachment was used for the mechanical analysis. Table 16 is a

summary of the mechanical design status.

4.5° Ramp Angle

10° Broach Angle

Figure 116. A Sloped Attachment Was Designed for the QHSF II Rotor.

NASA/CR—2012-217451 104
Figure 117. Honeywell Applied State-of-the-Art Finite Element Modeling Techniques in

the Mechanical Analysis of the QHSF II Rotor Blade and Attachment.

Table 16. All of the Mechanical Design Requirements Were Met for the QHSF II.

Blade Weight

• Meets target weight requirements (less than Baseline II)

Blade Modal Characteristics

• Adequate frequency margins at 100 percent Speed for 1E to 4E distortion

• Fundamental mode crossing speeds are within Honeywell experience

Medium Bird Ingestion

• Calculated blockage is with in Honeywell experience

Blade Stress

• Results are within Honeywell design experience

Fan Disk Burst Margin

• Relative to NASA criteria, 30 percent above burst margin requirement

• Relative to Honeywell criteria: 53 percent margin, required > 25 percent margin

Fan Stator Vane

• Airfoil-only vibration analysis shows adequate flutter margin

Figure 118 shows blade stress levels at the aerodynamic design point (defined as a tip speed

of 1506 ft/s, corresponding to a corrected speed of 15,621 rpm on the 22” rig). All stresses are

within Honeywell design experience. Figure 119 shows that stress levels in the sloped blade

attachment are well balanced and relatively low in the retention area. The attachment minimum

neck stresses redistribute, but are within experience for a frictionless condition. The axial contact

stress for the limiting (frictionless) condition is 33.8 Ksi. Figure 120 shows the results of 3-D disk

wedge analyses. Results indicate that the disk has adequate burst margin relative to NASA criteria

as shown (mechanical design point is defined as physical speed of 15,842 rpm on the 22” rig).

NASA/CR—2012-217451 105
Figure 118. QHSF II Fan Blade Stress Levels at Aerodynamic Design Point Are Within

Honeywell Design Experience.

Figure 119. QHSF II Fan Blade Attachment Stress Levels at Aerodynamic Design Point

Are Within Honeywell Design Experience.

SS

PS

QHSF II (μ=0.1) 90.2

81.66 (90.2) Ksi

90.6 (87.9) Ksi

91.5 (89) Ksi

85 (30) Ksi

QHSF II (μ=0) 91.5

Numbers in parenthesis

 are for μ=0.1 case, refer

to previous chart

for contour plot

Numbers in parenthesis

 are for μ=0.1 case, refer

to previous chart

for contour plot

Case

Axial 

Load (lbs)

Contact 

Stress (ksi)

μ = 0 2810 33.8

μ = 0.1 832 10.0

Material Ti-6-4

Min. Yield at 200 F = 106.2 ksi

QII yield safety factor = 1.16

NASA/CR—2012-217451 106
Disk Material: Maraging Steel C-250

Ultimate Strength (200° F), Ftu = 239 ksi

BURST MARGIN

NASA Criteria : Avg Hoop < 0.47 Ftu (=112 ksi)

Average Hoop @ aero design point: 93.4 ksi

Average Hoop @ mech design point: 96 ksi (< 112 ksi)

Figure 120. Disk Analysis at Aerodynamic Design Point Show Principal Stress Well Below

NASA Burst Margin Criteria.

To support the determination of dynamic loads on the rig and potential blade out loads, the

blade weight and center of gravity locations were estimated. Figure 121 shows the calculated

values for the airfoil (A/F) only, the release blade (largest portion lost in a blade-out event), and

the total blade for the QHSF II.

Axial CG Radial CG Weight, lb

A/F Only -0.327 7.019 0.55

Release Blade -0.145 6.166 0.78

Total Blade -0.127 5.895 0.86 

1.728 1.827

Figure 121. The Blade Weight and Center of Gravity (CG) Location Relative to the

Intersection of the Axis of Rotation and the Stacking Axis Are Provided for

Rig Structural and Dynamic Analyses.

NASA/CR—2012-217451 107
Modal analysis of the final QHSF II rotor blade design was performed and is summarized

by the Campbell Diagram in Figure 122. The characteristics of the first tree mode shapes are

shown in Figure 123. The first bending mode is placed between first and second rotational

harmonics at the mechanical design point with a 40 percent frequency margin. The second mode

is placed between the third and fourth rotational harmonic with a 12.6 percent frequency margin.

The third mode was placed between the fourth and fifth rotational harmonic with an 8 percent

frequency margin.

Figure 124 shows the ANSYS® calculation of the maximum deflections of the QHSF II

rotor blade at 100 percent Speed. The figure shows that the direction of principal motion is

circumferential. Figure 125 shows that the QHSF II rotor blade meets the bird strike criterion.

10E 9E 8E 7E 6E 

5E 

4E 

3E 

2E 

1E 

0

250

500

750

1000

1250

1500

0 1600 3200 4800 6400 8000 9600 11200 12800 14400 16000

Speed [rpm]

Frequency [Hz]

QHSF II (Iter18h)

QHSF I (22" Rig) 

15621 RPM

(ADP)

First Bending Mode

Second Bending Mode

First Torsion Mode

15842 RPM

(MDP)

Figure 122. The Campbell Diagram for the QHSF II Rotor Shows Adequate Frequency

Margin for the Three Primary Vibration Modes.

NASA/CR—2012-217451 108
(a) First Bending (b) Second Bending (c) First Torsion

Figure 123. The QHSF II Rotor Blade Has a Complex Vibration Modal Structure.

LE

TE

PS

SS

Figure 124. The Direction for the Maximum Deflections of the QHSF II Rotor Is in the

Circumferential Direction.

NASA/CR—2012-217451 109
LE Blockage = 0.729%

TE Blockage = 0.407%

LE Blockage = 0.729%

TE Blockage = 0.407%

Deformed Shape

Figure 125. The Analysis of the QHSF II Rotor Blade With NOSAPM Shows That the

Blade Will Meet the Bird Strike Criteria.

Figure 126 shows the completed rotor blades being installed in the fan disk. The boundary

conditions for the finite element analysis for the QHSF II fan blade assumed the dovetail contact

surfaces to be fully fixed in all degrees of freedom. This type of boundary condition was

calibrated for fan blades with a beaver tooth, which restrained (in the axial direction) the blade

dovetail at the forward end of the dovetail. The QHSF I fan blade has no beaver tooth and thus

does not have this additional constraint on the forward side of blade dovetail. The blade is

retained axially by the feature added on the pressure side of the shank and the mating tab on the

fan disk post. Since the mating retention feature was not incorporated on the broach block, the

finite element model boundary conditions were adjusted to match the bench test condition. In

the analysis the fan blade was fixed normal to the dovetail contact plane, and a group of nodes

were fixed parallel to the contact plane near the dovetail axial center.

A set of 22 QHSF II fan blades were acoustic ring (ARS) tested. The fan blades were

mounted on the broach block (P/N R3563132-1) and secured in the slot with the mounting bolts.

A bolt torque of 500 in-lb. was applied to ensure the dovetail was securely mounted and the

mating surfaces were fully in contact. Results of the test for the first 10 modes are documented

in Table 17. There is good agreement with the predicted frequencies of the finite element model.

NASA/CR—2012-217451 110
Figure 126. QHSF II Rotor – Partial Assembly.

Table 17. QHSF II Fan Blade--ARS Bench Test Frequencies.

Blade S/N Mode 1 Mode 2 Mode 3 Mode 4 Mode 5 Mode 6 Mode 7 Mode 8 Mode 9 Mode 10

Analysis 234 654 1158 1638 2136 2463 2570 2798 3455 3876

02 238 664 1154 1634 2132 2407 2549 2783 3428 3727

10 236 662 1168 1650 2138 2465 2610 2828 3495 3603

11 238 665 1171 1655 2141 2453 2587 2825 3482 3714

12 239 668 1174 1659 2141 2459 2601 2829 3484 3724

13 240 670 1164 1652 2131 2426 2580 2812 3452 3703

14 241 674 1168 1662 2144 2460 2579 2842 3475 3747

15 240 669 1165 1654 2134 2438 2576 2828 3466 3724

16 240 670 1166 1654 2133 2424 2574 2820 3461 3706

17 241 670 1168 1656 2137 2434 2572 2825 3468 3696

18 240 670 1166 1655 2136 2437 2572 2826 3463 3738

19 241 672 1168 1662 2141 2444 2584 2828 3470 3710

20 240 671 1166 1658 2134 2436 2575 2828 3463 3741

21 241 672 1167 1657 2135 2424 2579 2821 3459 3746

22 240 668 1167 1654 2138 2459 2582 2835 3464 3736

23 240 670 1168 1659 2141 2462 2573 2830 3469 3741

24 240 668 1166 1651 2138 2432 2573 2822 3466 3739

25 240 669 1163 1658 2135 2428 2579 2829 3466 3742

26 240 670 1165 1657 2137 2444 2574 2825 3469 3735

27 239 667 1165 1660 2134 2453 2576 2833 3470 3742

29 239 668 1168 1660 2138 2452 2584 2841 3479 3603

30 240 669 1163 1651 2127 2432 2582 2821 3458 3745

Min 236 662 1154 1634 2127 2407 2549 2783 3428 3603

Max 241 674 1174 1662 2144 2465 2610 2842 3495 3747

Mean 240 669 1166 1655 2136 2441 2579 2825 3467 3717

Std Dev 1.2 2.7 3.8 6.0 4.0 15.6 11.7 11.8 13.2 41.0

NASA/CR—2012-217451 111
Holography bench testing was conducted on the QHSF II fan blade using the same type of

set-up that was used in the ARS bench test. The results are documented in Table 18 where the

holography frequencies are compared to the average ARS frequencies of the 21 blades, as well as

to the analytically predicted frequencies. The data shows there is good agreement between the

bench test and finite element analysis frequencies. Figure 127 shows mode shapes (for modes 1-

6, respectively) obtained from holography test and finite element analysis. There is good

agreement between the holography bench test and analytically predicted mode shapes. The close

agreement of the results indicates the boundary condition used in the finite element analyses is

consistent with the conditions of the bench test.

Table 18. QHSF II Fan Blade -- Comparison of Finite Element Analysis, Acoustic Ring

Signature, and Holography Test.

ARS** Frequency (Hz)

Mode

Analysis*

(Hz) Average Min Max Std Dev

Percent

Difference

Holography***

(Hhz)

Percent

Difference

1 234 240 236 241 1.2 2.5 236 1.0

2 654 669 662 674 2.7 2.2 659 0.8

3 1158 1166 1154 1174 3.8 0.7 1160 0.2

4 1638 1655 1634 1662 6.0 1.0 1639 0.0

5 2136 2136 2127 2144 4.0 0.0 2137 0.1

6 2463 2441 2407 2465 15.6 -0.9 2428 -1.4

7 2570 2579 2549 2610 11.7 0.4 2562 -0.3

8 2798 2825 2783 2842 11.8 1.0 2798 0.0

9 3455 3467 3428 3495 13.2 0.3 3483 0.8

10 3876 3717 3603 3747 41.0 -4.3 3736 -3.8

* Fixed normal to contact surface + a row of nodes fixed parallel to contact surface at

 mid-dovetail

**Acoustic ring frequency of a set of 21 fan blades fixed in broach block

***Holography test of a single blade fixed in broach block

NASA/CR—2012-217451 112
Mode 1 -236 Hz Mode 1 - 233 Hz Mode 2 -659 Hz Mode 2 - 654 Hz

(a) Mode 1 (b) Mode 2

Mode 3 - 1160 Hz Mode 3 - 1158 Hz Mode 4 - 1639 Hz Mode 4 - 1638 Hz

(c) Mode 3 (d) Mode 4

Mode 5 - 2137 HZ Mode 5 - 2136 HZ Mode 6 - 2428 Hz Mode 6 - 2463 Hz

(e) Mode 5 (f) Mode 6

Figure 127. QHSF II Fan Blade – Comparison of Holography and Finite Element Analysis.

NASA/CR—2012-217451 113
For the rig configuration, the stator vane is made from SS355 material. The metal vane is

brazed to the outer shroud and thus in the finite element model it is fully fixed at the outer shroud.

At the hub, the vane is positioned in a pre-cut slot, and constrained in the radial direction by an

aluminum ring. The hub shroud itself is on rollers that allow the stator assembly to rotate about the

engine axis. In the finite element model, the vane is fixed in radial and axial directions at the hub.

The resulting reduced frequencies for the first two modes are above the design criteria. The

Campbell Diagram for the stator vane is shown in Figure 128 and the predicted mode shapes are

shown in Figure 129. The frequencies for the first 6 vane modes are given below.

Mode Frequency, Hz

1 807.2

2 3067.2

3 3562.8

4 3755.

5 5132.7

6 5508.7

2 2 E fan ro t o r

10E 

9E 

8E 

7E 

6E 

5E 

4 E 

3 E 

2 E 

1E 

M o d e 1 

M o d e 2

M o d e 3

M o d e 4

M o d e 5

M o d e 6 

100% MDP

0

1000

2000

3000

4000

5000

6000

0 1600 3200 4800 6400 8000 9600 11200 12800 14400 16000

Speed [rpm]

Frequency [Hz]

Figure 128. The Campbell Diagram for the QHSF II Vane Shows That No Vibration

Issues Are Expected.

NASA/CR—2012-217451 114
Mode 1, reduced frequency=0.65 Mode 2, reduced frequency=2.45 Mode 3

Mode 4 Mode 5 Mode 6

Figure 129. The First Six Mode Shapes for the QHSF II Rig Vane Were Calculated With ANSYS

®.

NASA/CR—2012-217451 115
5.8.3 Aeroelastic Performance

Viscous TURBO-AE analysis of the final design was performed. The predicted flutter

boundary is shown in Figure 130 for the various blade attachment concepts. For the current 4.5

degree sloped attachment design, the aeroelastic “pinch point” at 70 percent speed is predicted to

be stable based on the viscous results.

5.8.4 Acoustic Performance

Acoustic analysis was performed to verify the performance of the Case 18h design. In

summary, noise reduction on QHSF II is anticipated to be similar to that demonstrated on the

QHSF I. Noise is reduced primarily by decreased rotor-stator and rotor-strut interactions that are

obtained by the geometric features of the rotor and stator.

1.0

1.2

1.4

1.6

1.8

2.0

40 50 60 70 80 90 100

Inlet Corrected Flow (lbm/sec)

Rotor Pressure Ratio

60%

70%

80%

90%

100%

Straight Att

 Flutter Boundary

Slope 6 Att

Flutter Boundary

Slope 4.5 Att

70% Nc

Pinch Point

(Estimated)

Figure 130. Viscous TURBO Results Show Reduced Risk for 4.5 Degree Sloped

Attachment.

5.9 Further Revisions to the Stator Design

Additional aerodynamic design efforts were applied to examine potential solutions to the

separation in Case 18h and resulting efficiency loss. Figure 131 shows stator lean and metal

angle profiles for the Baseline II, Case 18h, and a new study case for the stator separation

elimination, labeled Q2a. A more aggressive lean, with compensating changes to the leading

edge metal angle are combined to reduce flow separation and could also offer further noise

reduction. Aerodynamic results for QHSF II and “Q2a” are shown in Figure 133 and Figure 134.

While the separation has not been completely eliminated, it has been significantly reduced, with

an estimated efficiency improvement of 0.3 percent at the peak point, as shown in Figure 135.

The improvements to the flow field appear to be significant, indicating that the efficiency

improvement may be somewhat under predicted by the CFD results.

NASA/CR—2012-217451 116
0

20

40

60

80

100

-0.6 -0.4 -0.2 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6

Tangential stack, Ycg, inches

% span

Q2a

QHSFII

Baseline II

0

20

40

60

80

100

40 45 50 55 60 65 70

Inlet Metal Angle, degrees

% Span

Q2a

QHSFII

Baseline II

-10 -8 -6 -4 -2 0

Exit Metal Angle, degrees

Q2a

Baseline II

Figure 131. The Stator Separation Was Reduced Through Increased Stator Lean and

Metal Angle Changes.

0

20

40

60

80

100

0.95 1.05 1.15 1.25 1.35 1.45

A/A*

% Span

Q2A

QHSF II

Baseline II

Figure 132. The QHSF II Stator Redesign Maintained Good Throat Area Margin.

NASA/CR—2012-217451 117
(a) QHSF II – Q2a (b) QHSF II – Case 18h (c) Baseline II

Figure 133. Mach Number Contours on the Vane Suction Surface Show the Reduction of

the Flow Separation Near the Shroud.

QHSFII_old QHSFII_new

Figure 134. Trailing Edge Mach Contours Show Reduced Stator Separation for the Q2a

Design.

NASA/CR—2012-217451 118
0.875

0.880

0.885

0.890

0.895

0.900

0.905

0.910

1.76 1.78 1.80 1.82 1.84 1.86 1.88 1.90 1.92

Stage Pressure Ratio

Stage Adiabatic Efficiency

QHSFII_old

QHSFII_new

Figure 135. The Results of the CFD Analysis Indicate an Improvement of 0.3% in Peak

Efficiency.

5.10 Modifications to the Baseline II Stators

A program augmentation was received on August 11, 2003 to make hardware for a set of

Baseline II stators. The aerodynamic design of the Baseline II vane needed to be modified to run

effectively behind the QHSF II rotor.

Figure 136 shows a solid model comparison of the QHSF II stator design to the modified

Baseline II stator. The new stator has geometry similar to the Baseline II stator, with nearly

identical performance to the QHSF II design as indicated by the Mach No. Profile in Figure 137,

the pressure ratio comparison in Figure 138, and the adiabatic stage efficiency comparison in

Figure 139.

NASA/CR—2012-217451 119
SHROUD

HUB

Moderate-Bow

QHSFII Vane

Existing QHSFII

Vane

Suction 

Surface

Suction 

Surface

Pressure 

Surface

Pressure 

Surface

Vanes are bowed so the 

pressure surface faces the 

hub and shroud.

Positive Ycg shift moves the

vane in the direction of the 

suction surface. Negative 

Ycg shift moves the vane in 

the direction of the pressure

surface.

Figure 136. The Baseline II Vane Has Less Circumferential Lean Than the QHSF II Vane.

QHSFII_Final Modified Baseline II

0.0

1.5Mach No.

Flowpath splitter effect

Vane LE

Vane TE

Slightly less lowmomentum flow

Figure 137. The Baseline II Vane Has a Very Similar Loading Distribution as the QHSF II

Vane at the Design Point.

NASA/CR—2012-217451 120
1.45

1.50

1.55

1.60

1.65

1.70

1.75

1.80

1.85

1.90

1.95

185.0 190.0 195.0 200.0 205.0 210.0 215.0 220.0 225.0

Engine-Size Fan Inlet Flow, lbm/s

Stage Pressure Ratio

QHSFII_Final

QHSFII Goal

QHSFII_moderate bow

N1c=100%

N1c=85%

Figure 138. QHSF II Stator and Baseline II Stator Have Nearly Identical Pressure Ratio.

0.86

0.87

0.88

0.89

0.90

0.91

0.92

0.93

0.94

0.95

0.96

0.97

185.0 190.0 195.0 200.0 205.0 210.0 215.0 220.0 225.0

Engine-Size Fan Inlet Flow, lbm/s

Stage Adiabatic Efficiency

QHSFII_Final

QHSFII Goal

QHSFII_moderate bow

N1c=85%

N1c=100%

Figure 139. QHSF II Stator and Baseline II Stator Have Nearly Identical Efficiency.

NASA/CR—2012-217451 121
Figure 140 shows the final Baseline II stator fabricated for the 22” QHSF II Rig.

 

Figure 140. A Baseline II Stator Set Was Designed and Fabricated to Match the QHSF II

Rotor for Study of Rotor/Strut Interaction Effects.

5.11 Analysis of the Rotor-Strut Interaction With the Baseline II and QHSF II Stators

The purpose of the Rotor-Strut Interaction Analysis was to further understand the role of

the stator in that interaction. In particular, the study focused on the impact on flow behavior of

the stator shape (lean and bow) and the pitchwise alignment (or circumferential clocking) of the

stators relative to the struts.

Stator/strut flow predictions were performed for 8 flowpath configurations, including 4

different stator/strut clocking positions, for both the Baseline II and QHSF II stators. For purposes

of the analyses, the struts were clocked relative to the stators, with clocking angles of 0.0, 1.8, 3.6,

and 5.4 degrees. The 4 stator/strut clocking positions are illustrated in Figure 141, for both the

Baseline II and QHSF II stators. The strut geometry was identical for both stators. Inlet flow

conditions for the stator/strut analyses were taken from an axisymmetric flow analysis prediction

of QHSF II rotor exit conditions at the 85 percent corrected speed, SLS operating line point.

The stator/strut flow predictions were performed using the Fluent® CFD analysis program.

Each of the flowpath models contained an annular periodic sector composed of 5 stators and 1

strut, as shown in Figure 142. The computational mesh consisted of an unstructured tetrahedral

volume mesh constructed from a triangular surface mesh. The triangular surface mesh on the

strut is shown in Figure 143. The computational meshes for the Baseline II cases had

approximately 3.8 million cells; the meshes for the QHSF II cases totaled approximately 4.2

million cells. These meshes were considerably more dense than those used in the QHSF I rotorstrut interaction analyses. The solutions were performed using Version 6.1.23 of Fluent®, and

employed the segregated, implicit solver, with the realizable k-epsilon turbulence model using

non-equilibrium wall functions.

NASA/CR—2012-217451 122
Evaluation of the flow analyses focused on the predicted static pressure fields upstream of

the stators. Static pressure data were processed at various axial and radial locations in the

upstream flowfield, as illustrated in Figure 144.

Figure 145 - Figure 148 present the circumferential distributions of static pressure at

selected locations resulting from the flow analyses of all 8 stator/strut configurations. Based on a

review of these figures, the following observations may be made:

• At all sampling locations upstream of the stator, a strut-induced static pressure

disturbance or pulse is evident. The pulses decay with increasing distance upstream

from the strut.

• Lean and bow of the QHSF II stator appear to affect the strut pressure pulses at outer

span radial locations. The pulse amplitude appears higher for the unbowed Baseline II

stator cases. In addition, as seen in Figure 148, the unbowed Baseline II stator cases

show evidence of the stator pressure pulses superposed over the strut pulse; in

contrast, the QHSF II cases show very little evidence of stator pressure pulses.

• Clocking effects appear to be evident with both stators. Differences in pulse

amplitude and peak shape appear to correlate with stator/strut clocking.

It may be concluded from the Rotor/Strut Interaction Study that the stator shape and pitchwise

alignment relative to the struts does have an impact on the static pressure distribution upstream

of the stators.

Unbowed Baseline II Stator

QHSF II Stator

0.0 deg 1.8 deg 3.6 deg 5.4 deg

Q_00

U_00

Q_18 Q_36 Q_54

U_18 U_36 U_54

Figure 141. Four Stator/Strut Clocking Positions Were Analyzed for Both the Baseline II

and QHSF II Stators.

NASA/CR—2012-217451 123
Figure 142. The Flow Path Model for Rotor/Strut Interaction Analyses Consisted of a

Periodic Sector of 5 Stators and 1 Strut.

Figure 143. The Unstructured Triangular Surface Mesh, Shown Here Applied to the

Bypass Strut, Formed the Basis for the Tetrahedral Volume Mesh.

NASA/CR—2012-217451 124
X = -6.50 in

X = -5.71 in

X = -4.92 in

X = -4.13 in

+θ

R = 10.24 in

R = 9.84 in

R = 8.66 in

R = 7.48 in

R = 6.22 in

FLA

Figure 144. Axial and Radial Locations at Which Static Pressure Data Were Processed in

the Region Upstream of the Stators.

17.3

17.4

17.5

17.6

17.7

17.8

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

Unbowed Stator - U_00

Unbowed Stator - U_18

Unbowed Stator - U_36

Unbowed Stator - U_54

17.3

17.4

17.5

17.6

17.7

17.8

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

QHSF II Stator - Q_00

QHSF II Stator - Q_18

QHSF II Stator - Q_36

QHSF II Stator - Q_54

Figure 145. Circumferential Static Pressure Distributions at R = 7.48 Inches and X = -5.71

Inches Show Differences in Pulse Amplitude and Shape With Strut Clocking.

Unbowed Baseline II Stator Case Is Shown on the Left; QHSF II Stator Case

Is on the Right.

NASA/CR—2012-217451 125
17.9

18.0

18.1

18.2

18.3

18.4

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

Unbowed Stator - U_00

Unbowed Stator - U_18

Unbowed Stator - U_36

Unbowed Stator - U_54

17.9

18.0

18.1

18.2

18.3

18.4

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

QHSF II Stator - Q_00

QHSF II Stator - Q_18

QHSF II Stator - Q_36

QHSF II Stator - Q_54

Figure 146. Circumferential Static Pressure Distributions at R = 8.66 Inches and X = -4.92

Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator

Case. Also, Differences Are Seen in Pulse Amplitude and Shape With Strut

Clocking. Unbowed Baseline II Stator Case Is Shown on the Left; QHSF II

Stator Case Is on the Right.

NASA/CR—2012-217451 126
18.5

18.6

18.7

18.8

18.9

19.0

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

Unbowed Stator - U_00

Unbowed Stator - U_18

Unbowed Stator - U_36

Unbowed Stator - U_54

18.5

18.6

18.7

18.8

18.9

19.0

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

QHSF II Stator - Q_00

QHSF II Stator - Q_18

QHSF II Stator - Q_36

QHSF II Stator - Q_54

Figure 147. Circumferential Static Pressure Distributions at R = 9.84 Inches and X = -4.92

Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator

Case. Also, Differences Are Seen in Pulse Amplitude and Shape With Strut

Clocking. Unbowed Baseline II Stator Case Is Shown on the Left; QHSF II

Stator Case Is on the Right.

NASA/CR—2012-217451 127
18.4

18.5

18.6

18.7

18.8

18.9

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

Unbowed Stator - U_00

Unbowed Stator - U_18

Unbowed Stator - U_36

Unbowed Stator - U_54

18.6

18.7

18.8

18.9

19.0

19.1

-25.0 -20.0 -15.0 -10.0 -5.0 0.0 5.0 10.0 15.0 20.0

Pitchwise Position (deg)

Static Pressure (psi)

QHSF II Stator - Q_00

QHSF II Stator - Q_18

QHSF II Stator - Q_36

QHSF II Stator - Q_54

Figure 148. Circumferential Static Pressure Distributions at R = 9.84 Inches and X = -4.13

Inches Show Pulse Amplitude Is Higher for the Unbowed Baseline II Stator

Case. Also, Pressure Pulse Shapes for the Baseline II Stators Show the

Influence of the Stator Pressure Pulses More Than QHSF II Stators, at the

Same Axial Position. Unbowed Baseline II Stator Case Is Shown on the Left;

QHSF II Stator Case Is on the Right.

6. RIG MODIFICATIONS

6.1 Overview

Several modifications to the existing QHSF rig were required to accommodate the QHSF II

design. Key mechanical differences include a reduced hub/tip ratio, the addition of a rotating

stator set & actuation system, and fan frame modifications to accommodate additional

instrumentation. An overlay of the QHSF I and QHSF II rigs is shown in Figure 149. Figure

150 is a schematic diagram of the rig installed in the wind tunnel.

A set of distortion screens and a screen rotator device were provided to complete additional

mechanical and operability testing of the QHSF II rig in the NASA wind tunnel. Figure 151

shows the inlet distortion screen rotator that has been proven in prior rig testing at Honeywell.

Figure 152 shows the rig in the performance test configuration, which has been modified for

rotation of stator set during rig operation. Figure 153 shows the rig in the acoustic configuration.

Changes to acoustic configuration are minimal. The inlet liner has been changed by NASA from

fiberglass to aluminum. During far-field acoustic testing of the QHSF II, the rotating stator

actuation system was removed, eliminating the need for any modifications to the nacelle.

NASA/CR—2012-217451 128
QHSF - dashed line

QHSF II - solid line

Figure 149. The Major Differences Between the QHSF I and QHSF II Rigs Are the Hub-to-Tip Ratio, Rotating Stator Set,

and Frame Modifications.

NASA/CR—2012-217451 129
Fan Stacking Axis

UHB Drive Rig

HSF Model Installed with Bellmouth Inlet and Variable Fan Exit Nozzle (VFEN)

End of Acoustic Boxes

End of Test Section

Traversing Microphone Mechanism Enclosure

Test Section Centerline

Cruciform Rake

Start of Test Section

Start of Acoustic Boxes

(Total Temperatures 21.5” Apart)

(Total Pressures 16.0” Apart)

Flow

335.9”

344.0”

178.7”

89.25”

30.0”

59.4”

10.0” 4.0”

21.5” 16.0”

176.0”

67.0” 121.0”

207.55”

27.0”

Figure 150. Schematic Diagram of the Rig Installed in the 9x15 Wind Tunnel.

NASA/CR—2012-217451 130
Figure 151. The Fan Distortion Measurement Configuration Is Shown for the 22" QHSF II Rig.

Figure 152. The Performance Configuration Is Shown for the 22" QHSF II Rig.

Figure 153. The Acoustic Configuration Is Shown for the 22" QHSF II Rig.

NASA/CR—2012-217451 131
6.2 Front Frame

Figure 154 shows the modifications to the front frame design for the QHSF II rig and

Figure 155 shows the new QHSF II aluminum frame mounted on the 22” dummy drive rig.

10 struts cored for instrumentation routing

5 comb rakes (w/ removable plugs in core

and bypass)

Figure 154. The Front Frame Design for QHSF II Was Modified to Incorporate 5 Comb

Rakes for Stage Performance Measurements.

Figure 155. QHSF II Fan Frame on the NASA 22” Dummy Drive Rig.

NASA/CR—2012-217451 132
6.3 Rotating Group

Figure 156 shows the modifications that were made to accommodate the new fan design.

The torque sleeve was modified to accommodate the new flow path that was introduced due to

the lower hub-to-tip ratio of the QHSF II. The spinner was re-designed to match the new

flowpath and to match a more current engine design. The design of the disk was modified to

implement the new blades with the sloped attachment. The mechanical design of the disk is

presented in Section 5. Figure 157 shows the results of the stress analysis of the aft spinner

modification. The maximum stress level was well below acceptable limits established by

Honeywell design practice.

Torque sleeve OD

reduced by .485 in

due to change in 

flow path

4.50 sloped blade 

attachment on disk

New spinner

Figure 156. Changes to the Rotor System Design for QHSF II Include a New Disk, Torque

Sleeve, and Spinner.

NASA/CR—2012-217451 133
(a) axial (b) circumferential

Figure 157. The Maximum Stress Level in the Spinner Was Computed to Be 13 ksi.

6.4 Rotating Stator Assembly

A rotating stator concept was proposed to allow detailed flow measurements behind the

stator and to allow the stator positions to be clocked relative to the strut positions. A drawing of

the dual-actuator system is shown in Figure 158. This concept allows for mechanical rotation of

the stators in both the acoustic and performance configurations of the 22” rig.

Figure 158 shows the original design concept for the rotating stator assembly with 2

actuators and 2 horizontal rods. The design load capability for each actuator is 1000 lb. The

expected maximum total aerodynamic and mechanical load is expected to be 650 lb. During

assembly, it was determined that the opposing actuators, as designed, could potentially bind each

other. It was decided that since one actuator had enough authority to rotate the stators, only one

actuator and horizontal rod would be used. Figure 159 shows a photograph of the final

configuration.

NASA/CR—2012-217451 134
Figure 158. The Rotating Stator Concept Allows for Variable Positioning of the Stator

Relative to the Struts in the 22” QHSF II Rig, as Well as Facilitate Stage

Performance Measurements.

Figure 159. The Final Rotating Stator Actuation System Uses One Actuator and

Horizontal Bar.

NASA/CR—2012-217451 135
6.5 Safety Review

Honeywell’s safety analyses for the QHSF II rig are supplemental to those performed on

the entire QHSF rig structure (Reference 1), and are focussed only on the previously described

mechanical differences in the QHSF II rig. In general, all changes to the QHSF II mechanical

structure resulted in either identical or better margins of safety relative to the previous design.

Stress levels of the QHSFII fan disk meet Honeywell design requirements. Table 19 shows

the calculated LCF life for the QHSF II fan disk with maraging steel C-250 properties using

Honeywell’s Browse material database. The maximum test speed analyzed represents the

highest speed the rig is expected to achieve as part of the planned testing, and is equivalent to

105 percent of the aerodynamic design speed of the QHSF II.

Table 20 and Figure 160 show the results of the QHSF II airfoil stress calculations at the

maximum test speed of 16,402 rpm.

Factor of Safety margins for the new rotating QHSF II hardware are summarized in Table

21, and satisfy the NASA requirements as shown. Details of the calculations are summarized in

Table 21.

Table 19. LCF Life for C-250 Fan Disk.

Condition Max Test Speed Trip Speed

Speed (RPM) 16402 16700

LCF Life with Min Material

Properties (cycles) 9071 7553

LCF Life with Average

Fatigue Properties (cycles) 22177 18456

Temperature (Deg F) 200 200

Stress Ratio 0 0

Stress Range (KSI) 186 193

Component

CF Radial

Load (lb)

Section

Area (in^2)

Average Stress

(ksi)

Blade root 38844 1.003 39

Airfoil root 29288 0.774 38

NASA/CR—2012-217451 136
A

A

Fr=29288 lbf

Area=0.774 in2

Section A-A

Fr=38,844 lbf

Area=1.003 in2

B

B

Section B-B

Figure 160. Average Blade Root and Airfoil Root Section Stress Calculation at 16402 RPM.

Table 20. New QHSF II Hardware Satisfies NASA Factor of Safety Requirements.

QHSF QHSF II at

15621 rpm

QHSF II at

16402 rpm

NASA Req

Yield Margin for Blade Airfoil 1.2 1.2 1.1 1.1

Yield Margin for Blade

Attachment

n/a 1.2 1.1 1.1

Burst Margin (NASA Criterion) 1.5 (ultimate) 1.2 1.1 1.1(1.5

Ultimate)

Yield Margin for Spinner 1.9 (ultimate)

(fwd spinner)

3.9 3.7 1.1 (1.5

Ultimate)

NASA/CR—2012-217451 137
Table 21. QHSF II Fan Rotor Margin of Safety Calculation.

Factor of Safety Met?

Component Qty Material

Fty Yield

(ksi)

Ftu

Ult.

(ksi)

Fsu Ult.

Shear

(ksi)

Max

Stress

(ksi)

Avg.

Stress

(ksi)

FS

Yield

FS

Ult.

FS

Shear

1.1y

1.5u

1.5y

3.0u

3.0y

5.0u

Comments andSource

Disk/Hub 1

Maraging

Steel C250 230 @200°F 239 @200°F 186 103 1.2 1.6 Yes @16402 rpm, avg tangentail stress

Blade Root 22 Ti-6-4 106.2 @200°F 117 @200°F 99 39 1.1 2.1 Yes @16402 rpm, Avg section stressBlade Root 22 Ti-6-4 106.2 @200°F 117 @200°F 90 35 1.2 2.3 Yes @15621 rpm Avg section stress

Blade airfoil 22 Ti-6-4 106.2 @200°F 117 @200°F 97 38 1.1 2.2 Yes @16402 rpm Avg section stressBlade airfoil 22 Ti-6-4 106.2 @200°F 117 @200°F 88 34 1.2 2.4 Yes @15621 rpm Avg section stress

NASA/CR—2012-217451 138
6.6 Rig Speeds

Table 22 defines key physical operating speeds and their relationship to the corrected

aerodynamic design speed of the QHSF II fan.

Table 22. QHSF II Physical and Corrected Design Speed for Key Operating Points.

Physical Speed* %N1 re QHSF II

QHSF II Aero Design Point (1485 ft/s

corrected tip speed req)

15621 100%

QHSF II Mechanical Design Point

(from 1506 ft/s mech tip speed req.)

15842 101%

Max Speed in QHSF II Test Plan 16402 105%

22" Rig Mechanical Speed Limit 16850 108%

* Mech. Speed calculations are applicable to 59 deg. Std. Day

6.7 Nozzle Sizing

The predicted exit flow (Wc) for the QHSF II matches the QHSF along most of the

operating line (including Approach and Cutback acoustic conditions), but is slightly lower near

the Sideline condition (~90% fan speed, Nc). The nozzle may be slightly large for the QHSF II;

but the small differences shown in Figure 161 are not critical to the acoustic measurements.

10

20

30

40

50

60

70

20 30 40 50 60 70 80 90 100 110

% Nc

E

xit

Wc (lb

m/sec)

QHSF II Sea Level Standard

QHSF 1 Sea Level Standard

QHSF I Cruise

Figure 161. The Predicted QHSF II Operating Line Shows That the Existing QHSF I

Nozzle Will Accommodate the QHSF II Fan.

NASA/CR—2012-217451 139
6.8 Instrumentation

The complete instrumentation list for the QHSF II rig is presented in Appendix I. A

general description of the rig instrumentation by Honeywell appears below. NASA-provided

instrumentation not described in this section (but included in Appendix I) include rig speed & all

rig mechanical integrity, instrumented performance bellmouth, rotor exit radial survey probes,

light probes, core flow rakes, and the LDV measurement system. Also included in Appendix I

are detailed descriptions and figures of the fan case plug instrumentation.

6.8.1 Accelerometers

Three accelerometers mounted at top dead center on the fan frame will provide continuous

vibration data (in the vertical, axial, and horizontal directions).

6.8.2 Boundary Layer Rakes

Five rakes with 10 pressure elements each are used to determine the magnitude of the inlet

boundary layer (these rakes are identical to the rakes used in QHSF I testing, as described in

Reference 1 and Appendix I). The boundary layer rake is pictured in Figure 162.

Figure 162. Boundary Layer Rake Shown in QHSF II Fan Case Plug (18 Degree Location).

6.8.3 Capacitance Probes

Four capacitance probes spaced equally spaced around the fan at rotor leading edge, midspan, and trailing edge are used to measure rotor clearance.

6.8.4 Comb Rakes

Five rakes consisting of 14 pressure and temperature elements each (9 bypass and 5 core)

will be used to measure stage performance at traversing positions across the vane passage

(accomplished by clocking of the stator set). The comb rakes mimic the trailing edge stator

geometry as shown in Figure 163.

NASA/CR—2012-217451 140
Figure 163. Aft Looking Forward View of Fan Frame and QHSF II Stator, With One

Comb Rake Shown.

6.8.5 Distortion Rakes

Ten rakes with six total pressure elements will be used to measure radial inlet distortion

effects, forward of the fan rotor (Figure 164).

Figure 164. Inlet Distortion Rake Design for QHSF II.

6.8.6 Kulites

Four kulites in a fan case plug and an additional kulite in the fan case spanning 1 strut

passage will be used to evaluate strut potential field at the rotor. An additional 10 kulites placed

diagonally across the plug will be used to evaluate rotor shock position. Figure 165 shows the

kulite locations in the fan case plug.

NASA/CR—2012-217451 141
Figure 165. QHSF II Fan Case Plug Kulites.

6.8.7 Static Pressures

Numerous statics (as defined in Appendix I) are located throughout the rig at critical

locations on the hub and shroud, including core, bypass, and vane leading edge measurement

planes. Figure 166 depicts locations of the static pressure measurements (PS) acquired during

aerodynamic performance mapping. (Total temperature (TT) and total pressure (PT)

measurement points also shown.)

NASA/CR—2012-217451 142
PT/TT Combo

PS

 PT

Figure 166. Aerodynamic Performance Measurements on QHSF II Fan Rig.

6.8.8 Strain Gages

A total of twenty strain gages were mounted on critical stress areas of 4 rotor blades (5 per

blade), as shown below in Table 23, Table 24, and Figures 167-170. Two strain gages were

placed on the fan disk, as shown in Figure 171. Two strain gages were placed on each of two

QHSF II stator vanes as shown in Figure 172.

Table 23. The Strain Gages Are Described for the QHSF II Rotor Blade.

GAGE RATIOS 90 75 80 90 90 90 90 90 80 90 GAGE LOCATIONS

Freq(Hz) -> 370 886 1256 1872 2309 2418 2709 3032 3619 3798 Axial Radial Angle

Mode #--> 1 2 3 4 5 6 7 8 9 10 (in) (in) (deg)

Gage #

1 PS 96 79 2 41 5 17 6 6 3 2 -1.604 -0.586 3.820

2 SS 29 65 89 6 11 7 10 1 30 3 1.603 0.519 6.130

3 SS 3 41 29 100 93 81 8 92 7 39 -1.392 -1.534 9.792

4 SS 1 15 32 63 7 20 97 26 58 58 -0.462 0.009 9.849

Gage Size = 3.2E-2

Table 24. Strain Gage Locations Are Identified for the QHSF II Rotor Blade.

Global Cartesian Origin at Ref. Pt (TE Tip)

x-along engine axis

Origin at Ref. Point (TE Tip)

x-along dovetail

S/G Loc X Y Z X Y Z X Y Z

1 -1.604 -0.586 3.820 2.202 -2.561 -6.495 -2.608 -2.150 6.495

2 1.603 0.519 6.130 -1.004 -1.456 -4.185 0.741 -1.619 4.185

3 -1.392 -1.534 9.792 1.990 -3.509 -0.523 -2.564 -3.120 0.523

4 -0.462 0.009 9.849 1.060 -1.966 -0.466 -1.950 -2.367 -0.434

Ref. Pt

(TE Tip)

0.599 1.975 10.315 0.000 0.000 0.000 0.000 0.000 0.000

NASA/CR—2012-217451 143
X

Z

X

Z

Figure 167. Strain Gage Locations Were Defined to Measure 7 Major Vibrational Modes.

x-axis aligned

with engine Axis

x-axis aligned 

with dovetail

S/G Loc #1

S/G Loc #2

S/G Loc #3

S/G Loc #4

Top View

Figure 168. The Top View of Blade Shows the Strain Gage Locations and the Relative

Positions of the Engine and Dovetail Axes.

NASA/CR—2012-217451 144
Section 

A-A

A

A

0.423 in

0.030 in

Strain Gage

Location #5

4.5 degrees

Figure 169. The Fifth Strain Gage Is Located on the Blade Dovetail.

Figure 170. QHSF II Fan Blade – Strain Gages.

NASA/CR—2012-217451 145
2 Gages, 30 degrees Apart

Oriented Circumferentially

0.478 in.

Figure 171. A Strain Gage Was Also Mounted on the Fan Disk to Monitor the Mechanical

Behavior.

Strain gage strain limits are set based on Goodman diagram data for Ti-6-4 MA at 190°F. The

maximum allowable strain on each gage is summarized below in Table 25. Tables 26 – 30 describe

in detail the allowable strain levels at each gage location, for various critical vibration modes.

Table 25. Summary of Maximum Allowable Strains.

Gage

Location

Max Allowable

Strain (p-p)

1 4500

2 5780

3 5000

4 6440

5 4500

Disk 1000

NASA/CR—2012-217451 146
Table 26. Allowable Strains for Modes 5 and 6, Gage Location #5.

Engine Gage Loc 5 Allowable

Order Mode RPM Freq (hz)

Mean

Stress (ksi) Gage Ratio Sig-Alt (ksi) Strain (p-p)

4 1 3100 244 3.13 0.98 53.19 6240

3 1 5400 252 9.49 0.98 50.76 5955

2 1 8200 272 21.88 0.98 46.02 5399

1 16634 352 90.02 0.98 8.68 1018

6 2 7000 702 15.94 0.99 48.79 5723

5 2 8800 724 25.20 0.99 45.21 5304

4 2 11500 766 43.03 0.99 38.32 4495

2 16634 895 90.02 0.99 8.77 1029

Table 27. Allowable Strains for Modes 1 and 2, Gage Location #1.

Engine Gage Loc 1 Allowable

Order Mode RPM Freq (hz)

Mean

Stress(ksi) Gage Ratio Sig-Alt (ksi) Strain (p-p)

4 1 3100 244 1.30 0.96 52.79 6193

3 1 5400 252 3.94 0.96 51.80 6077

2 1 8200 272 9.09 0.96 49.88 5851

1 16634 352 37.40 0.96 39.27 4606

6 2 7000 702 6.62 0.79 41.80 4904

5 2 8800 724 10.47 0.79 40.62 4765

4 2 11500 766 17.88 0.79 38.33 4497

2 16634 895 37.40 0.79 32.31 3791

NASA/CR—2012-217451 147
Table 28. Allowable Strains for Modes 1, 2, and 3, Gage Location #2.

Engine Gage Loc 2 Allowable

Order Mode RPM Freq (hz)

Mean

Stress(ksi) Gage Ratio Sig-Alt (ksi) Strain (p-p)

4 1 3100 244 0.01 0.29 16.09 1888

3 1 5400 252 0.03 0.29 16.09 1888

2 1 8200 272 0.06 0.29 16.09 1887

1 16634 352 0.25 0.29 16.07 1885

6 2 7000 702 0.04 0.65 36.06 4231

5 2 8800 724 0.07 0.65 36.06 4230

4 2 11500 766 0.12 0.65 36.04 4228

3 2 17800 895 0.29 0.65 36.00 4223

10 3 7000 1167 0.04 0.89 49.38 5793

9 3 7800 1170 0.06 0.89 49.38 5792

8 3 8800 1173 0.07 0.89 49.37 5792

7 3 10100 1178 0.09 0.89 49.36 5791

6 3 11950 1195 0.13 0.89 49.35 5789

5 3 14450 1204 0.19 0.89 49.33 5787

4 3 18475 1232 0.31 0.89 49.29 5782

Table 29. Allowable Strains for Modes 4, 5, and 6, Gage Location #3.

Engine Gage Loc 3 Allowable

Order Mode RPM

Freq

(hz)

Mean

Stress(ksi) Gage Ratio Sig-Alt (ksi) Strain (p-p)

10 4 10300 1717 2.26 1.0 54.62 6407

9 4 11650 1748 2.90 1.0 54.37 6378

8 4 13300 1773 3.77 1.0 54.03 6338

7 4 15650 1826 5.23 1.0 53.46 6271

14 5 9050 2112 1.75 0.93 50.98 5981

13 5 9770 2117 2.04 0.93 50.88 5968

12 5 10600 2120 2.40 0.93 50.75 5953

11 5 11575 2122 2.86 0.93 50.58 5933

10 5 12730 2122 3.46 0.93 50.36 5908

9 5 14180 2127 4.29 0.93 50.06 5872

8 5 16000 2133 5.46 0.93 49.63 5822

14 6 10500 2450 2.35 0.81 44.21 5186

13 6 11300 2448 2.72 0.81 44.09 5173

12 6 12200 2440 3.18 0.81 43.95 5156

11 6 13280 2435 3.76 0.81 43.77 5134

10 6 14575 2429 4.53 0.81 43.52 5106

9 6 16100 2415 5.53 0.81 43.21 5069

8 6 18000 2400 6.91 0.81 42.77 5017

NASA/CR—2012-217451 148
Table 30. Allowable Strains for Modes 5 and 6, Gage Location #4.

Engine Gage Loc 4 Allowable

Order Mode RPM Freq (hz)

Mean

Stress(ksi) Gage Ratio Sig-Alt (ksi) Strain (p-p)

14 5 9050 2112 0.40 1.00 55.34 6492

13 5 9770 2117 0.47 1.00 55.32 6489

12 5 10600 2120 0.55 1.00 55.29 6485

11 5 11575 2122 0.66 1.00 55.24 6481

10 5 12730 2122 0.80 1.00 55.19 6474

9 5 14180 2127 0.99 1.00 55.11 6465

8 5 16000 2133 1.26 1.00 55.01 6453

14 6 10500 2450 0.54 1.00 55.29 6486

13 6 11300 2448 0.63 1.00 55.26 6482

12 6 12200 2440 0.73 1.00 55.22 6477

11 6 13280 2435 0.87 1.00 55.16 6471

10 6 14575 2429 1.04 1.00 55.09 6463

9 6 16100 2415 1.27 1.00 55.00 6452

8 6 18000 2400 1.59 1.00 54.88 6438

4.400in.

2.552 in.

S/G Loc 1

S/G Loc 3

• All gages on the concave side with the gage center 0.100 in. from the edge.

Figure 172. The Side View of Vane Shows the Strain Gage Locations Relative to the

Shroud.

NASA/CR—2012-217451 149
Fan stator vane strain gage strain limits are set based on Goodman diagram data for S355

Stainless Steel at 75°F, as shown in Table 31.

Table 31. QHSF II Stator Strain Gage Limits.

Gage

Location

Max Allowable

Strain (p-p)

1 3520

3 3520

6.9 Distortion Screens

A set of classical, tip radial, and complex distortion screens (described in Table 32) are

recommended distortion screens for testing on the QHSF II rig. The screen selections were made

after considering the unique design of the QHSF II and reviewing the Honeywell’s 18” rig test

data of the Baseline II distortion screens. Honeywell has tested all the recommended distortion

screens previously on the Baseline II fan rig, with the exception of the tip radial distortion

screens. Given the unique nature of the fan tip design of the QHSF II, the tip radial distortion

screens were included in this test.

Table 32. Recommended Distortion Screens for the QHSF II Rig Test.

Screen # MPR Description Wire

Dia.

Grid

Size

Porosity Dist.

Level

1 0 Backer 0.080 1.016 84.8%

104 1 1E Classical 0.079 1.017 85.1% 4%

112 1 1E Classical 0.054 0.334 70.1% 12%

115 1 1E Classical 0.018 0.085 62.9% 15%

204 2 2E Classical 0.079 1.017 85.1% 4%

212 2 2E Classical 0.054 0.334 70.1% 12%

304 3 3E Classical 0.079 1.017 85.1% 4%

312 3 3E Classical 0.054 0.334 70.1% 12%

404 4 4E Classical 0.079 1.017 85.1% 4%

412 4 4E Classical 0.054 0.334 70.1% 12%

503 0 Tip Radial 0.062 0.606 80.5% 3%

506 0 Tip Radial 0.055 0.314 68.0% 6%

901 Complex Left Eng. 30 Kt x-wind See Figure 3

902 Complex Right Eng. 30 Kt x-wind See Figure 4

Table 32 also indicates what screen material was used to build the distortion screens and the

resulting screen porosity. Sample circumferential distortion screens are shown in Figure 173.

Figure 174 shows a tip radial distortion screen. Figure 175 and Figure 176 show the complex

crosswind distortion screens for the left and right engines, respectively. The OD and ID values

quoted in Figure 173 and Figure 174 reflect the diameter of the QHSF II rig test screen holder.

NASA/CR—2012-217451 150
Figure 173. QHSF Rig Circumferential Distortion Screens.

NASA/CR—2012-217451 151
OD = 10.99 in

ID = 8.97 in

Figure 174. QHSF Rig Tip Radial Distortion Screen.

NASA/CR—2012-217451 152
30 Knot Left Cross-Wind

Left Engine @ Takeoff Power

A

B

C

D

B

C

Squares/Inch Porosity Wire Dia. (in) Notes

Region A -- 84.6 -- *Backer Screen

Region B 4.5 78.6 0.025

Region C 8.5 58.5 0.028

Region D 10.0 34.8 0.041

Aft Looking Forward

78.9

56.3

33.2

0.028

0.025

0.035

4.0

10.0

12.0

Figure 175. Left Engine 30-Knot Crosswind Distortion Screen.

NASA/CR—2012-217451 153
A

B

C

D

B

C

30 Knot Left Cross-Wind

Right Engine @ Takeoff Power

Aft Looking Forward

Squares/Inch Porosity Wire Dia. (in) Notes

Region A -- 84.6 -- *Backer Screen

Region B 4.5 78.6 0.025

Region C 10.0 59.3 0.023

Region D 11.0 37.9 0.035

B

0.025

0.018

0.035

81.0

60.8

33.2

4.0

12.0

12.0

Figure 176. Right Engine 30-Knot Crosswind Distortion Screen.

NASA/CR—2012-217451 154
6.10 Model Assembly

Overall assembly of the QHSF II rig is as described in Reference 1. Figures 179 – 184

depict phases of the QHSF II rig assembly and wind tunnel installation as described below.

Figure 177. QHSF II Fan Frame on the NASA 22” Dummy Drive Rig.

Figure 178. QHSF II Fan Stator and Rotating Stator Actuation Assembly on the NASA

22” Dummy Drive Rig.

NASA/CR—2012-217451 155
Figure 179. NASA Wind Tunnel Installation of QHSF II Stator Assembly, Shown With

Outer Fan Case.

Figure 180. Aft View of NASA USB Drive Rig/QHSF II During Wind Tunnel Installation.

NASA/CR—2012-217451 156
Figure 181. Configuration for Operability Testing Includes Screen Rotator Assembly and

Long “Tomato Can” Inlet.

Figure 182. QHSF II Tunnel Installation, in Aerodynamic Performance Measurement

Configuration.

NASA/CR—2012-217451 157
Figure 183. QHSF II (a) in Far-Field Acoustic Measurement Configuration (b).

6.11 Modification of Rotating Group After Initial Assembly

NASA discovered an issue with rotor bore size and the “top hat” of the dynamic balance.

The top hat was not a feature that had been included in Honeywell’s QHSF drawings that were

used as the basis of the QHSF II design. Three parts were identified that would require

modification for a larger bore diameter to accommodate the top hat are: 1) fan disk, 2) aft

spinner, and 3) torque sleeve nut.

Mechanical design evaluation of the modified hardware was completed. Figure 184 and

Figure 185 show the drawings of the modified fan disk and new aft spinner. The modification of

the disk has no impact on the hoop strength of disk bore or the peak static stress at the fan disk

dovetail slot. Therefore, there will be no change to the previously calculated factors of safety for

the fan disk.

Stress analysis results of the new aft spinner are shown in Figure 186. Shifting the inner

segment forward (instead of radial on the original design) attenuated the radial stress through the

slot hole. The overall peak stress of the modified design (55.3 ksi) is higher than the previously

calculated value of 43 ksi. Nevertheless, the redesigned aft spinner has adequate margin of

safety. The recalculated values are shown in the table.

NASA/CR—2012-217451 158
Original Design

Modified Design

Figure 184. QHSF II Fan Disk Modified for Larger Bore Diameter to Allow Clearance for

the Top Hat.

Original Design

Modified Design

Figure 185. QHSF II Aft Spinner Modification Was Made to Allow Clearance for the

Dynamic Balance Top Hat.

NASA/CR—2012-217451 159
Figure 186. QHSF II New Aft Spinner Maximum Principal Stress -- New Design

Maintains Adequate Margin of Safety.

Factor of Safety

Met?

Component Qty Material

Fty

Yield

(ksi)

Ftu Ult.

(ksi)

Fsu

Ult.

Shear

(ksi)

Peak

Stress

(ksi)

Avg.

Stress

* (ksi)

FS**

Yield

FS***

Ult.

FS

Shear

1.1y

1.5u

1.5y

3.0u

3.0y

5.0u

Aft Spinner 1

Stainles

s Steel

17-4

154

@75F

172

@75F 55.3 12.7 2.8 9.5 Yes Yes

 *Avg stress = average tangential stress

** FS Yield = (PeakStress)/(FtyYield)

** FS Ult = [Avg Stress/(0.7*FtuUlt)]

Peak Stress =

55,274 psi

Avg Tang. Stress =

12,078 psi

NASA/CR—2012-217451 160
7. NEW TECHNOLOGY

This final report has identified all nonpatentable discoveries, innovations, and computer

code improvements, and all patentable inventions that were developed or discovered during the

performance of the contract. In summary, there have been a number of innovations throughout

the design, hardware, and test phases of the QHSF II program. Such innovations include the

TURBO-AE aeroelastic tool improvement, the multi-disciplinary analytical design of

experiments approach to the QHSFII fan design, and the rotating stator capability added to the

NASA 22" rig. To date, two innovations have been identified as being patentable: the advanced

swept stator and the sloped disk attachment feature. Table 31 summarizes each innovation and

references the applicable section of this report. Possible secondary applications of the reported

new technology are also identified.

Table 33. Summary of Each Innovation and Applicable Section of This Report.

QHSF II 22"

Rig Project

Section Innovation Benefit Secondary

Application

Patentable

Design 3.0 Aeroelastic Tool Calibration Improved design accuracy/higher

risk design trades possible

Future Fans no

Design 5.1 Multi-disciplinary DOE

design approach

Optimized design that meets all

acoustic, aero, & mechanical

requirements

Future Fans no

Design 5.6 Optimized Non-Linear

Stator

Elimination of rotor-strut

interaction noise

Future Fans yes

Design 5.8.2 Sloped fan disk attachment Uniform stress balance, reduced

weight

Future Fans yes

Hardware/

Instrumentation/

Test

6 All aero measurements

(including operability, inlet

distortion, synchronous

vibration) on NASA 22" Rig

Eliminate need for 2nd rig,

reduces test time, hardware fab

cost, improves data quality

Future Fans

in NASA

9X15

facility

no

Hardware 6.4 Rotating Stator System for

NASA 22" Rig

Variable stator position allows

investigation of fan source noise

evaluation, and investigation of

aero/acoustic optimum stator

positions

Future Fans

in NASA

9X15

facility

no

Test 6.8.4 Stage exit rakes with stator

TE geometry

Improved measurement quality

required for advanced stator

Future fans

with

advanced

swept stator

no

It has been shown that the shank configuration (such as the buttress and/or the dovetail

slope) significantly affects the fan blade mode shape, which in turn is a key driver for flutter

behavior. The dovetail slope provides the means to get the desired flutter margin (by controlling

the mode shape) while meeting other constraints. This approach is very different from the prior

approaches used at Honeywell and other companies, which eliminated flutter problems by

increasing the frequency, changing the loading distribution, or clipping the geometry.

TURBO-AE would be applicable to any system that flutters, such as LPT blades and isolated

wings. This approach could also be used to reduce the vibration level in a more-typical (Campbell

Diagram type) forced response.

NASA/CR—2012-217451 161
8. SUMMARY AND CONCLUSIONS

8.1 Aeroelastic Tool Evaluation

Results from a detailed study of flutter encountered on a forward-swept fan were

summarized and compared to experimental data. Overall, the correlation of the computational

results is in good agreement with experimental observations. The blade was correctly predicted

to be free of flutter up to the stall line at 100 percent speed. At 85 percent and 75 percent speeds,

flutter was predicted just above the operating line, and well before stall was encountered. This is

again consistent with the rig test, and the predicted and measured flutter boundaries in terms of

the fan map are in good agreement. The correct critical nodal diameter was also predicted.

Inviscid analyses were shown to provide predictions of damping in remarkably good

agreement with the viscous results. The key benefit of utilizing an inviscid assumption is that

the computations require only 2-5 percent of the computational time required for full viscous

analyses. Potential uses include identification of the critical nodal diameter, trade studies, and

sensitivity assessments. The conclusions from the results provided here are that the primary

drivers for fan flutter are dominated by shock structures and are inviscid in nature, and the

damping change along the speed line is tied principally to the change in shock position.

Several additional studies were conducted to further investigate the flutter behavior, and to

better understand sensitivities. Changes in the aerodynamic damping predictions due to

variations in blade shape due to speed changes, inlet and exit pressure profiles, tip clearance, and

mode shapes were all evaluated. Surprisingly, none of these changes had a very significant

effect on the damping prediction. The position on the fan map gave a consistently good

indication of the stability for these effects.

One shortcoming in this correlation effort is that the flutter boundary had to be extrapolated

for the viscous analyses at part-speed conditions because “numerical” stall was en-countered.

This is a limitation of the “steady” solver, i.e., not related to the blade motion, and is a common

limitation of CFD codes. Note that TURBO was able to reach the measured stall boundary at

100 percent speed, and this is very encouraging. However, the ability to reach the actual stall

line would provide much more confidence in the ability of the unsteady analysis to accurately

predict the flutter boundary.

8.2 QHSF I Data Evaluation

A study was completed to identify and explain any differences between the QHSF I 18-inch

and 22-inch measured aerodynamic and aeroelastic performance. This study concludes that the

differences in performance would be consistent if the two blade sets differed in hot shape. This

difference would have altered blade incidence, thereby explaining differences in both aeroelastic

behavior and aerodynamic performance. The possibility that the two blade sets, identical in

design except for scale, differed in cold shape was considered unlikely and not further

investigated in this study.

Acoustic data from the QHSF I 22” rig test was conducted. The primary noise reduction

was achieved by the significant reductions in both rotor-stator and rotor-strut interaction, and is

responsible for up to 6 EPNdB noise reduction at higher tip speeds. Less acoustic benefit was

achieved at lower tip speeds, and is attributable to a flow separation on the QHSF I rotor blades

NASA/CR—2012-217451 162
that occurred only at the lower fan speeds. Comparisons of acoustic results with CFD analyses

and LDV measurements are in good agreement, and show the reduced strength of the QHSF I

rotor wake relative to the Baseline I. The forward swept blade increases the distance between

the rotor trailing edge and stator leading edge, which contributes to the reduction of rotor-stator

interaction noise. The rotor wake structure displays a tangential lean due to forward swept

geometry, this effect is enhanced by addition lean and sweep of the stator leading edge. As a

result, the QHSF I rotor wake traverses the stator leading edge more slowly than the Baseline I

stator, and serves to further reduce the rotor-stator interaction tone noise.

Additional noise reduction was achieved by the virtual elimination of rotor-strut interaction

tones at some speeds. A rotor-strut interaction CFD study was conducted to assess any

differences between pressure distributions upstream of the Baseline I and QHSF I stators.

Significant differences in the stator geometry occur near the tip, and results show that the

Baseline I fan has more static pressure variation than the QHSF I, suggesting that the cause of

rotor-strut interaction tones is the rotor responding to the variation in potential pressure field of

the struts.

8.3 QHSF II Design

The QHSF II design was developed based on the experience of the QHSF as well as other

recent Honeywell product fan design experience. The design process relied heavily upon use of

analytical Design of Experiments (DOEs) to define the optimum rotor and stator system to

achieve all acoustic, aerodynamic, aeroelastic, and mechanical performance goals. This series of

analyses included the selection of blade forward sweep, blade tangential lean, blade thickness

distribution, rotor incidence, and stator sweep and lean optimization. The DOE process allowed

the rapid assessment of these key design features relative to all of the interdisciplinary design

goals, as well as an understanding of the interaction and sensitivity of key design parameters.

Design tool inputs to the DOEs included ADPAC for aerodynamics, ANSYS for stress, V072

and SOURCE3D for acoustics, and TURBOAE for aeroelastics.

The final design selection of the QHSF II includes a forward-swept blade with reduced

sweep at the tip and additional sweep at lower spans relative to the QHSF I, and a full span stator

with an optimized, non-linear sweep and lean. These features analytically demonstrate the

simultaneous achievement of all the NASA program and internal Honeywell goals.

Subsequent to completion of the QHSF II design, a program augmentation was received to

fabricate a set of Baseline II stators. The “moderate bow” stators were modified to run

effectively behind the QHSF II rotor. The Baseline II stators are expected to have identical

aerodynamic performance (pressure ratio and efficiency) as the QHSF II design. Relative to

noise, a rotor-strut interaction study concludes that the stator shape and pitchwise alignment

relative to the struts does have an impact on the static pressure distribution upstream of the

stators, and therefore will impact the noise signature.

8.4 Rig Modifications

Several modifications to the existing QHSF rig were required to accommodate the QHSF II

design. Key mechanical differences include a reduced hub/tip ratio, fan frame modifications to

accommodate additional instrumentation, and a stator actuation system that allows rotation of the

NASA/CR—2012-217451 163
entire stator set during rig operation. The actuation system facilitates aerodynamic stage

performance measurements, as well as the investigation of rotor-strut and rotor-stator interaction

noise mechanisms. External hardware for the actuation system can be removed, and the system

can be manually actuated, allowing the external nacelle to remain intact for far-field noise

measurements.

All aerodynamic, mechanical, and acoustic measurements on the QHSF II design will be

accomplished on the 22” rig, in the NASA-Glenn 9x15 Wind Tunnel facility, eliminating the

possibility for data discrepancies experienced between the 18” and 22” scale fans on the QHSF I

program. In order to accomplish all aerodynamic and mechanical testing at NASA, additional

hardware was provided for test purposes that included distortion screens and a screen rotator

mechanism for additional rotor mechanical and fan operability measurements. Additional

instrumentation (relative to the QHSF I 22” fan) included stage performance rakes and CAP

probes provided by Honeywell. NASA provided a rotor exit survey probe system.

NASA/CR—2012-217451 164
9. REFERENCES

1. Weir, D., “NAS3-27752, NASA AST – AOI 14, Design and Test of Fan and Nacelle

Models for Quiet High Speed Fan – Final Report,” CR-2003-212370, July 2003.

2. Lieber, L., Repp, R., and Weir, D., “Quiet High Speed Fan,” NASA CR-198518, November

1996.

3. Chen, J.P., and Whitfield, D.L., “Navier-Stokes Calculations for the Unsteady Flowfield of

Turbomachinery,” AIAA Paper 93-0676, Reno, NV, January 1993.

4. Chen, J.P., Celestina, M., and Adamczyk, J.J., “A New Procedure for Simulating Unsteady

Flows Through Turbomachinery Blade Passage,” ASME Paper 94-GT-151, The Hague, The

Netherlands, June 1994.

5. Bakhle, M.A., Srivastava, R., Stefko, G.L., and Janus, J.M., “Development of an Aeroelastic Code Based on an Euler / Navier-Stokes Aerodynamic Solver,” ASME Paper 96-GT311, 1996.

6. Bakhle, M.A., Srivastava, R., Keith, T.G., Jr., and Stefko, G.L., “Aeroelastic Calculations

Based on Three-Dimensional Euler Analysis,” AIAA Paper 98-3295, 1998.

7. ANSYS Analysis Guides, SAS IP Inc., 2001.

8. Adamczyk, J.J., “Model Equation for Simulating Flows in Multistage Turbomachinery,”

ASME Paper 85-GT-226, 1985.

9. Chen, J.P., and Barter, J., “Comparison of Time-Accurate Calculations for the Unsteady

Interaction in Turbomachinery Stage,” AIAA Paper AIAA-98-3292, 1998.

10. Sanders, A., “Stall Flutter Assessment of an Advanced Design Transonic Fan Using

TURBO-AE,” 6th National Turbine Engine High Cycle Fatigue Conference, Jacksonville,

FL, 2001.

11. Fite, E.B., private correspondence.

12. Dittmar, J.H., Elliott, D.M., and Fite, E.B., “The Noise of a Forward Swept Fan,” NASA

TM-2003-212208, November 2003.

13. Heidelberg, L.J., “Comparison of Tone Mode Measurements for a Forward Swept and

Baseline Rotor Fan,” 9th AIAA/CEAS Aeroacoustics Conference and Exhibit, Hilton Head,

SC, May 12–14, 2003.

NASA/CR—2012-217451 165
APPENDIX I

INSTRUMENTATION

(17 Pages)

NASA/CR—2012-217451 166
NASA/CR—2012-217451 167
NASA/CR—2012-217451 168
NASA/CR—2012-217451 169
NASA/CR—2012-217451 170
NASA/CR—2012-217451 171
NASA/CR—2012-217451 172
NASA/CR—2012-217451 173
NASA/CR—2012-217451 174
NASA/CR—2012-217451 175
NASA/CR—2012-217451 176
NASA/CR—2012-217451 177
NASA/CR—2012-217451 178
NASA/CR—2012-217451 179
Circumferential Placement of Plugs

 18 deg

90 deg 270 deg

Forward Looking Aft

Detailed Plug Configuations

Plug 1 Instrumentation Test Config Placement Responsibility

Inlet Boundary Layer/Inlet Distortion Aero/Dist 90 deg NASA

CAP probes (LE, Mid, TE) Aero 90 deg Honeywell

PS125O Static Pressures Aero 90 deg NASA

Rotor Exit Survey (Wedge Probe) Aero 90 deg NASA

Fan Rubstrip static pressure, Passage 1 Aero 90 deg Honeywell

Plug 2

Inlet Boundary Layer/Inlet Distortion Aero/Dist 18 deg NASA

Rotor Exit Survey (Combo PT/TT Probe) Aero 18 deg NASA

PS125O Static Pressures Aero 18 deg NASA

Plug 3

Inlet Boundary Layer/Inlet Distortion Aero/Dist 270 deg NASA

CAP probes (LE, Mid, TE) Aero 270 deg Honeywell

Optical light probes for flutter Aero 270 deg NASA

Fan Rubstrip static pressure, Passage 2 Aero 270 deg Honeywell

Plug 4

LDV Window Aero 90 deg NASA

Plug 5

Shock Position & Strut Potential kulites Aero 18 deg NASA

(15 total, 14+ one in fan case)

Plug 6

Blank (NASA)

Plug 7

Blank (NASA)

Plug 8

Blank (NASA)

NASA/CR—2012-217451 180
Fan Case Liner with Plugs

 A

B 

 plug (3)

 A 

Instrumentation to Modify Part Config Location Comment

A CAP probes (LE, Mid, TE) Aero 0 deg, 180 deg 2 in fan case, 2 in plugs for total of 4(x3)

B Rotor Exit Survey Ps, 10 total Aero in survey plane all 10 in case when no plugs

6 in case when plugs installed,

(2 in two plugs for a total of 4): 10 total

NASA/CR—2012-217451 181
18 deg

window

90 deg

window

Boundary layer rake 

(or distortion rake)

Combo (PT/TT) 

survey probe

PS125O(10)

0 deg cap probes 

(LE, MC,TE)

Rub strip Kulite (#11)

18 deg window plug

FLOW

Boundary layer rake 

(or distortion rake)

Wedge flow angle 

survey probe

PS125O(2)

90 deg cap probes 

(LE, MC,TE) 90 deg window plug

FLOW

Rub strip statics

PSVR(1 - 5)

PS125O(1)

PS125O(3)

NASA/CR—2012-217451 182
270 deg

window

18 deg

Kulite

window

270 deg window plug

Rub strip statics FLOW

PSVR(6 - 10)

Distortion rake

(when applicable)

Optical (Flutter) 

Light Probes

270 deg cap probes 

(LE, MC,TE)

18 deg Kulite window plug

FLOW

1

13

2

3

4

5

6

7

8

9

10

12

15

14

NASA/CR—2012-217451 183
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-08-2012 

2. REPORT TYPE

Final Contractor Report 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Quiet High Speed Fan II (QHSF II): Final Report 

5a. CONTRACT NUMBER

NAS3-01136 

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Kontos, Karen; Weir, Don; Ross, Dave 

5d. PROJECT NUMBER

5e. TASK NUMBER

2 

5f. WORK UNIT NUMBER

WBS 561581.02.08.03.45.04 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

Honeywell 

P.O. Box 52181 

Phoenix, Arizona 85072 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18181 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/CR-2012-217451 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 71 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

This report details the aerodynamic, mechanical, structural design and fabrication of a Honey Engines Quiet High Speed Fan II (lower 

hub/tip ratio and higher specific flow than the Baseline I fan). This fan/nacelle system incorporates features such as advanced forward sweep 

and an advanced integrated fan/fan exit guide vane design that provides for the following characteristics: (1) Reduced noise at supersonic tip 

speeds, in comparison to current state-of-the-art fan technology; (2) Improved aeroelastic stability within the anticipated operating envelope; 

and (3) Aerodynamic performance consistent with current state-of-the-art fan technology. This fan was fabricated by Honeywell and tested 

in the NASA Glenn 9- by 15-Ft Low Speed Wind Tunnel for aerodynamic, aeromechanical, and acoustic performance. 

15. SUBJECT TERMS

Low noise; Models; Scale models; Turbofan; Acoustic measurements; Aerodynamic characteristics 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

202 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18