Grant Gordon

Honeywell Aerospace, Phoenix, Arizona

Acoustic Database for Turbofan Engine Core-Noise 

Sources, Volume I—Final Report

NASA/CR—2015-218879/VOL1

September 2015

Honeywell Report No. 21–15066(10)
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Grant Gordon

Honeywell Aerospace, Phoenix, Arizona

Acoustic Database for Turbofan Engine Core-Noise 

Sources, Volume I—Final Report

NASA/CR—2015-218879/VOL1

September 2015

Honeywell Report No. 21–15066(10)

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC10BA07B, Task Order NNC11TA40T
Acknowledgments

This is the Final Report for the “Acoustic Database for Core Noise Sources Program,” which was sponsored by the

NASA Fundamental Aeronautics Program, Fixed-Wing Project, under Contract No. NNC10BA07B, Task Order 

NNC11TA40T. The NASA Technical Monitor was Dr. Lennart S. Hultgren, NASA Glenn Research Center.

The editor thanks the staff of the NASA Glenn Research Center for their support and technical insights.

A project of this magnitude and complexity is indebted to the contributions of many team members. In particular, the editor 

thanks the following Honeywell technical staff for their efforts:

Bill Schuster, Co-Principal Investigator Don Weir, Co-Principal Investigator

Ed Yaple, Development Shop Gary Meshew, Engine Testing

Harvey Niska, Lead Instrumentation Engineer Henry Gruenemeier, Engineering Test Services Hugh McDowell, Engineering 

Test Services Israel Picazo, Design

John Gerrard, Fabrication John Konves, Design Lead

Joshua Furner, Instrumentation Kaleb Lopez, Design

Marvin Springer, Engineering Test Services Ronald Goodwin, Lead Engine Test Engineer Ronald Swecker,

Engineering Test Services Taylor Marotta, Data Analysis

Thomas Gajdecki, Instrumentation Tom White, Engineering Test Services Travis Sorenson,

Control Engineer William Landram, Development Shop

Available from

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial 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.

NASA STI Program

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National Technical Information Service

5285 Port Royal Road

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This report is available in electronic form at http://www.sti.nasa.gov/ and http://ntrs.nasa.gov/
TABLE OF CONTENTS

1.0 EXECUTIVE SUMMARY ..................................................................................................... 1

2.0 INTRODUCTION ................................................................................................................. 2

2.1 Document Organization ........................................................................................... 2

2.2 Program Task Definition – As Defined in the Statement of Work ............................ 2

2.2.1 Task 1: Planning and Instrumentation ..................................................... 2

2.2.2 Task 2: TECH977 Engine Modifications .................................................. 3

2.2.3 Task 3: Data Analysis Methodology ......................................................... 3

2.2.4 Task 4: TECH977 Engine Test ................................................................ 4

2.2.5 Task 5: Data Processing and Distribution ................................................ 4

3.0 TASK 1: PLANNING AND INSTRUMENTATION .............................................................. 4

3.1 Identification of Candidate Techniques .................................................................... 4

3.2 Definition of Data Acquisition Systems .................................................................... 6

3.3 Development Probe Fabrication .............................................................................. 7

3.4 Development Probe Rig Testing .............................................................................. 8

3.5 Development Probe Bench Test .............................................................................. 9

3.6 Development Probe Oven Test .............................................................................. 11

3.7 Development Probe Burner Rig Test ..................................................................... 14

3.8 Development Probe Oven Calibration Test ............................................................ 16

3.9 Engine Probe Design ............................................................................................. 19

4.0 TASK 2: TECH977 ENGINE MODIFICATIONS ............................................................... 24

4.1 Engine Identification and Preparation .................................................................... 24

4.2 Preliminary Instrumentation Layout ....................................................................... 24

4.2.1 Engine Mechanical Design .................................................................... 25

4.3 Design Criteria ....................................................................................................... 25

4.4 Detailed Design ...................................................................................................... 28

5.0 TASK 3: DATA ANALYSIS METHODOLOGY ................................................................. 36

5.1 Review of Analysis Techniques ............................................................................. 36

5.2 Dynamic Gas Temperature Measurement System (DGTMS) Software ................ 37

5.3 Circumferential Mode Decomposition .................................................................... 40

5.4 Burner Rig Test Data Analysis ............................................................................... 42

6.0 TASK 4: TECH977 ENGINE TEST ................................................................................... 58

6.1 Engine Build with Modified Hardware .................................................................... 58

6.2 Engine Installation .................................................................................................. 64

6.3 Instrumentation Setup ............................................................................................ 68

6.4 Pretest Checkouts .................................................................................................. 71

6.5 Test Data Planned ................................................................................................. 72

6.6 Test Data Acquired ................................................................................................ 72

6.7 Test Data Export .................................................................................................... 78

NASA/CR—2015-218879/VOL1 iii
TABLE OF CONTENTS (CONT) 

7.0 TASK 5: DATA PROCESSING AND DISTRIBUTION ..................................................... 78

7.1 Task 5.1 Process Unsteady Temperature and Pressure Data ............................... 78

7.1.1 Unsteady Temperature Data .................................................................. 79

7.1.2 Unsteady Pressure Data Processing ................................................... 109

7.2 Task 5.2: Acoustic Modal Analysis....................................................................... 121

7.2.1 Combustor Circumferential Mode Decomposition ................................ 121

7.2.2 Trends in Spectra and Levels With Operating Condition ..................... 131

7.2.3 Time Delays for Matched Pairs of Kulites ............................................ 131

7.3 Task 5.3: Spectra and RMS Levels of Temperature and Pressure ...................... 135

7.3.1 Trend Evaluation of Combustor Temperature Data ............................. 135

7.3.2 Temperature Spectra at HPT and LPT Exits........................................ 140

7.3.3 Time Delays for Matched Pairs of Small Wire Thermocouples ............ 140

7.3.4 Temperature and Pressure Cross Correlations ................................... 144

7.4 Task 5.4: Distribution of Data and Analysis Results ............................................ 152

8.0 CONCLUSIONS .............................................................................................................. 153

8.1 Major Results ....................................................................................................... 153

8.2 Recommendations for Further Analysis ............................................................... 155

8.3 Recommendations for Future Testing .................................................................. 156

9.0 REFERENCES ................................................................................................................ 157

NASA/CR—2015-218879/VOL1 iv
LIST OF FIGURES

Figure 1. Preliminary Signal Conditioning and Data Acquisition Setup. ........................................................................ 7

Figure 2. Dual Thermocouple Probes Used in the Bench and Burner Rig Tests. ......................................................... 8

Figure 3. Voltage Output of Type B Thermocouples as a Function of Temperature. .................................................. 10

Figure 4. A Heat Gun Was Used to Generate Small Voltage From the Probes. ......................................................... 10

Figure 5. Final Signal Conditioning and Data Acquisition Setup. ................................................................................ 11

Figure 6. An Oven Test Was Conducted to Measure the DC Output of the Probes. ................................................... 12

Figure 7. Temperature Measured by the Thermocouples During an Oven Warmup. .................................................. 13

Figure 8. Temperature Differences Measured During the Oven Warmup. .................................................................. 13

Figure 9. “Large Temperature Fluctuations” Burner Rig Carousel Test. ..................................................................... 15

Figure 10. “Small Temperature Fluctuations” Burner Rig Test. .................................................................................... 16

Figure 11. An Isotech Pegasus Dry Block Calibrator Was Used to Evaluate the NASA Dual Thermocouple 

Probe Output Against Reference Standards. .......................................................................................... 17

Figure 12. Comparison of the Temperature Measured by Four Temperature Probes During a Warmup and 

Cooldown in the Isotech Pegasus Unit. .................................................................................................. 18

Figure 13. Comparison of the Probe Temperature Readings Made in the Dry Block Calibrator at 

Temperatures Near 1500°F [Bottom] and 2000°F [Top]. ........................................................................ 19

Figure 14. Preliminary Dual-Thermocouple Probe Design for the TECH977 Test. ..................................................... 21

Figure 15. A Drawing (70642386-1) Was Created for the Dual Thermocouple Probes for the Engine Test. .............. 22

Figure 16. The Thermocouple Wires Egress Through an Alumina Stick and Are Terminated in an Easily 

Accessible Connector. ............................................................................................................................ 23

Figure 17. Preliminary Instrumentation Layout for the TECH977 Engine Test. ........................................................... 25

Figure 18. Schematic of Temperature Probe Placement (at 273° CCW ALF). ............................................................ 26

Figure 19. View of the Modifications to the Combustor Case (70642392-1). .............................................................. 29

Figure 20. View of the Modifications to the Combustor (70642393-1). ........................................................................ 30

Figure 21. View of the Probe Integration and Actuation Setup for the Combustor. ..................................................... 30

Figure 22. View of the Modifications to the Turbine Case (70642394-1). .................................................................... 31

Figure 23. View of the Modifications to the LP Turbine Nozzle (70642395-1). ............................................................ 31

Figure 24. View of the Probe Integration and Actuation Setup for the LP Turbine. ..................................................... 32

Figure 25. View of the Modifications to the Mixer Nozzle (70642396-1). .................................................................... 32

Figure 26. View of the Probe Integration and Actuation Setup for the Mixer. .............................................................. 33

Figure 27. View of the Modifications to the C-Ducts (70642398-1). ............................................................................ 33

Figure 28. Bimba Flat-1 FOD-313-HV Actuators Were Used to Deploy the Probes. ................................................... 34

Figure 29. Actuation System Assembly for the Combustor and Turbine. .................................................................... 34

Figure 30. Actuation System Assembly for the Mixer Nozzle. ..................................................................................... 35

Figure 31. Details of the Actuation System for the Mixer Nozzle................................................................................. 35

Figure 32. Typical Data DTP Data Reduction Process (Figure 8 of CR-179513). ....................................................... 39

Figure 33. Aliasing of Circumferential Modes Is Predictable From Previous NASA EVNERT Array 

Measurements. ....................................................................................................................................... 40

Figure 34. Semi-Infinite Probes Have Been Successfully Used on Recent Engine Test Programs. ........................... 41

Figure 35. Example of a Kulite Tee Instrumented With a Thermocouple. ................................................................... 41

Figure 36. Representative Time History Voltage Output from the Burner Rig Test. .................................................... 43

Figure 37. An Example of the Splitting of a Time History Data Into “Mean” (DC) and “Dynamic” (AC) 

Components. .......................................................................................................................................... 44

Figure 38. Schematic of Ideal, As-Designed, and As-Manufactured Probes. .............................................................. 45

Figure 39. Sample Comparison From the Dimensional Sensitivity Analysis, for Honeywell Probe A in the 

Carousel Spinning at 1000 rpm. ............................................................................................................. 46

Figure 40. Sample Comparison From the Dimensional Sensitivity Analysis, for Stationary Honeywell Probe 

A at the Burner Flame Edge, Fuel Pressure of 155 psi. .......................................................................... 46

Figure 41. Results From the Carousel Testing of Honeywell Probe A at 500 rpm. ..................................................... 48

Figure 42. Results From the Carousel Testing of Honeywell Probe A at 1000 rpm. ................................................... 49

Figure 43. Results From the Carousel Testing of Honeywell Probe A at 2000 rpm. ................................................... 49

Figure 44. Results From the Carousel Testing of Honeywell Probe A at 2800 rpm. ................................................... 50

Figure 45. Results From the Carousel Testing of Honeywell Probe A at 0 rpm and With the Probe 

Positioned Directly Behind a Rod. .......................................................................................................... 51

Figure 46. Results From the Carousel Testing of Honeywell Probe A at 0 rpm and With the Probe Directly 

Exposed to the Flow. .............................................................................................................................. 52

Figure 47. Results From the Stationary Testing of Honeywell Probe A and the NASA Probe, Positioned at 

the Burner Flame Edge, at a Fuel Pressure of 94 psi. ............................................................................ 53

NASA/CR—2015-218879/VOL1 v
LIST OF FIGURES (CONT) 

Figure 48. Results From the Stationary Testing of Honeywell Probe A and the NASA Probe, Positioned at 

the Burner Flame Edge, at a Fuel Pressure of 155 psi. .......................................................................... 54

Figure 49. Results From the Stationary Testing of Honeywell Probe A and the NASA Probe, Positioned at 

the Burner Flame Core, at a Fuel Pressure of 94 psi. ............................................................................. 55

Figure 50. Results From the Stationary Testing of Honeywell Probe A and the NASA Probe, Positioned at 

the Burner Flame Core, at a Fuel Pressure of 155 psi. ........................................................................... 56

Figure 51. Example of Problematic Data From the Carousel Test of Honeywell Probe A at 2800 rpm (First 

Attempt). ................................................................................................................................................. 56

Figure 52. Example of Problematic Data From Honeywell Probe B From the Carousel Test of Probes A & B 

Together at 2000 rpm. ............................................................................................................................ 57

Figure 53. Example of Good Data From Honeywell Probe B From the Carousel Test of Probes B By Itself at 

2000 rpm. ............................................................................................................................................... 57

Figure 54. Photo of the TECH977 Engine Prior to Relocation for Disassembly. ......................................................... 58

Figure 55. Fabricated Combustor Case Wedges and Bosses. ................................................................................... 59

Figure 56. TECH977 Combustor Case Being Laser Cut. ............................................................................................ 59

Figure 57. Modified Combustor Case Showing Bosses and Wedges Welded in Place. ............................................ 60

Figure 58. Sensor End of the Three O’clock and Nine O’clock Orientations of the Dummy Probes Protruding 

Into the Combustor Case. ....................................................................................................................... 61

Figure 59. Assembled Actuators With Dummy Probes Installed. ................................................................................ 61

Figure 60. End View Showing Modification of the Drive Yoke Used to Actuate Inter Turbine Duct Thermal 

Probes at the Three O’clock Orientation. ................................................................................................ 62

Figure 61. End View Showing Modification of the Drive Yoke Used to Actuate Inter Turbine Duct Thermal 

Probes at the Nine O’clock Orientation. .................................................................................................. 62

Figure 62. Example Temperature Probe Originally Received From The Supplier. ...................................................... 63

Figure 63. Temperature Probes Fabricated for the TECH977 Engine Test. ............................................................... 63

Figure 64. Completed Temperature Probes. ............................................................................................................... 64

Figure 65. View of the TECH977 Installed at San Tan Cell 965 (Right Side, ALF). .................................................... 65

Figure 66. View of the TECH977 Installed at San Tan Cell 965 (Left Side, ALF). ....................................................... 65

Figure 67. View of the Nitrogen Purge Flow Supply System. ...................................................................................... 66

Figure 68. View of the Purge Flow Manifolds, Kulite Coils, and Kulites. ..................................................................... 66

Figure 69. View of the Installed Temperature Probe Hardware................................................................................... 67

Figure 70. Closeup of the Pressure Tubes and Temperature Probes. ........................................................................ 67

Figure 71. Location of the Two External Microphones. ............................................................................................... 68

Figure 72. View of the Cable Connections to the 64-Channel Dewetron Recorder. .................................................... 69

Figure 73. View of the Cable Connections to the 48-Channel Pulse Recorder. .......................................................... 69

Figure 74. View of the Front of the Sonoran Microsystems Signal Conditioning Rack Used With the Kulites. ............ 70

Figure 75. View of the Cable Connections on the Sonoran Microsystems Signal Conditioning Rack Used 

with the Kulites. ...................................................................................................................................... 71

Figure 76. Sample Readout from the Dewetron System at 48 Percent Speed (3/26). ................................................ 73

Figure 77 Instrumentation Clocking Locations, Aft Looking Forward. ......................................................................... 74

Figure 78. Fluctuating Temperatures of the Small and Large Wires at 48 Percent Speed. ......................................... 80

Figure 79. Fluctuating Temperatures of the Small and Large Wires at 54 Percent Speed. ......................................... 81

Figure 80. Fluctuating Temperatures of the Small and Large Wires at 60 Percent Speed. ......................................... 82

Figure 81. Fluctuating Temperatures of the Small and Large Wires at 65 Percent Speed. ......................................... 83

Figure 82. Representative Spectra of the Raw Temperature Signals at 48 Percent Speed. ....................................... 84

Figure 83. Representative Spectra of the Raw Temperature Signals at 60 Percent Speed. ....................................... 84

Figure 84. Example of Quantization of Raw Temperature Signals Recorded by the Pulse System. ........................... 85

Figure 85. Example of Quantization of Raw Temperature Signals Recorded By the Dewetron System. .................... 86

Figure 86. Dewetron System Electronic Noise Floor for Temperature Probes. ........................................................... 87

Figure 87. Nominal Dimensions of the Thermocouple Probe Elements. ..................................................................... 89

Figure 88. DC and AC Time Histories for the CF45 and CA45 Probes at the March 26 48 Percent 

Speed Point. ........................................................................................................................................... 93

Figure 89. DC and AC Time Histories for the CF12 and CA12 Probes at the March 26 48 Percent 

Speed Point. ........................................................................................................................................... 94

Figure 90. DC and AC Time Histories for the LF45 and LA45 Probes at the March 26 48 Percent 

Speed Point. ........................................................................................................................................... 95

Figure 91. DC and AC Time Histories for the LF12 and LA12 Probes at the March 26 48 Percent 

Speed Point. ........................................................................................................................................... 96

NASA/CR—2015-218879/VOL1 vi
LIST OF FIGURES (CONT) 

Figure 92. DC and AC Time Histories for the MX45 and MX12 Probes at the March 26 48 Percent 

Speed Point. ........................................................................................................................................... 97

Figure 93. An Example of the Coherence Between the Small Wire and Large Wire Thermocouple (T/C) Raw 

Time Histories. ........................................................................................................................................ 98

Figure 94. Small Wire Compensation Spectra for the Combustor Probes at 48 Percent. ........................................... 99

Figure 95. Small Wire Compensation Spectra for the Turbine Duct Probes at 48 Percent. ....................................... 99

Figure 96. Small Wire Compensation Spectra for the Mixer Probes at 48 Percent. .................................................. 100

Figure 97. Small and Large Wire Compensated Spectra for the CA45 Probe at 48 Percent. ................................... 100

Figure 98. Comparison of TECH977 and F100 Compensation Spectra. ................................................................... 101

Figure 99. Small and Large Wire Compensated Spectra for the Combustor Probes at 48 Percent. ......................... 102

Figure 100. Small and Large Wire Compensated Spectra for the Inter-Turbine Duct Probes at 48 Percent. ............ 103

Figure 101. Small and Large Wire Compensated Spectra for the Mixer Probes at 48 Percent. ................................ 104

Figure 102. Uncompensated DC (Top) and AC (Bottom) Time History Data for Combustor Probe CA12 

at 48 Percent. ....................................................................................................................................... 107

Figure 103. Small Wire and Large Wire Compensated Time Histories for Combustor Probe CA12 

at 48 Percent. ....................................................................................................................................... 108

Figure 104. Close-Up of Small and Large Wire Compensated Time History Data for Combustor Probe CA12 

at 48 Percent. ....................................................................................................................................... 108

Figure 105. A Comparison of the Noise Floor of the Dynamic Pressure Transducers, With and Without 

Purge Flow. .......................................................................................................................................... 110

Figure 106. Dynamic Pressure Spectra at 48 Percent Speed From Data Collected March 26. ................................ 112

Figure 107. Dynamic Pressure Spectra at 54 Percent Speed From Data Collected March 28. ................................ 113

Figure 108. Dynamic Pressure Spectra at 60 Percent Speed From Data Collected March 26. ................................ 114

Figure 109. Dynamic Pressure Spectra at 65 Percent Speed From Data Collected March 28. ................................ 115

Figure 110. Purge Flow Temperatures Measured During the March 26 Test. .......................................................... 117

Figure 111. Purge Flow Temperatures Measured During the March 28 Test. .......................................................... 118

Figure 112. Purge Flow Temperatures Measured During the March 28 Test (Continued). ....................................... 118

Figure 113. Cross-Spectral Phase Between Pressure Probes in the Combustor Sub-Array, at 60 Percent 

Speed, for Coherence Values above 0.05. ........................................................................................... 120

Figure 114. Coherence Between Pressure Probes in the Combustor Sub-Array at 60 Percent Speed. ................... 120

Figure 115. Cross-Spectral Phase for Combinations of Pressure Signals in the Combustor Sub-Array, at 60 

Percent Speed, from the EVNERT Run Without Fan Test. .................................................................. 121

Figure 116. Kulite Probe Locations for the Combustor Circumferential Mode Array. ................................................ 122

Figure 117. Comparison of Combustor Noise Spectra With Previous Measurements at 48 Percent Speed. ............ 122

Figure 118. Comparison of Combustor Noise Spectra With Previous Measurements at 60 Percent Speed. ............ 123

Figure 119. Mode Aliasing With a Reduced Number of Sensors at 48 Percent. ....................................................... 124

Figure 120. Mode Aliasing With a Reduced Number of Sensors at 54 Percent. ....................................................... 125

Figure 121. Mode Aliasing With a Reduced Number of Sensors at 60 Percent. ....................................................... 125

Figure 122. Comparison of Modal Decomposition Results at 48 Percent Speed. ..................................................... 126

Figure 123. Comparison of Modal Decomposition Results at 54 Percent Speed. ..................................................... 126

Figure 124. Comparison of Modal Decomposition Results at 60 Percent Speed. ..................................................... 127

Figure 125. Comparison of Modal Decomposition Results at 48 Percent Speed. ..................................................... 127

Figure 126. Comparison of Modal Decomposition Results at 54 Percent Speed. ..................................................... 128

Figure 127. Comparison of Modal Decomposition Results at 60 Percent Speed. ..................................................... 128

Figure 128. Comparison of DGTMS Modal Decomposition at 48 Percent Speed, With Temperature Probes 

Retracted and Inserted Into the Combustor. ......................................................................................... 129

Figure 129. Comparison of DGTMS and EVNERT Spectra at 48 Percent Speed..................................................... 129

Figure 130. Comparison of DGTMS and EVNERT Spectra at 54 Percent Speed..................................................... 130

Figure 131. Comparison of DGTMS and EVNERT Spectra at 60 Percent Speed..................................................... 130

Figure 132. OASPL Variation as a Function of Engine Speed. ................................................................................. 131

Figure 133. Cross-Spectral Phase Between the Upstream/Downstream Dynamic Pressure Probes in the 

Combustor and ITD, at 48 Percent (March 26). .................................................................................... 132

Figure 134. Cross-Spectral Phase Between the Upstream/Downstream Dynamic Pressure Probes in the 

Combustor and ITD, at 65 Percent (March 28). .................................................................................... 133

Figure 135. Cross-Spectral Phase Between the Combustor/Turbine and Turbine/Mixer Dynamic Pressure 

Probes, at 48 Percent (March 26). ........................................................................................................ 133

Figure 136. Cross-Spectral Phase Between the Combustor/Turbine and Turbine/Mixer Dynamic Pressure 

Probes, at 65 Percent (March 26). ........................................................................................................ 134

Figure 137. Mean Temperatures Measured in the Combustor. ................................................................................. 135

NASA/CR—2015-218879/VOL1 vii
LIST OF FIGURES (CONT) 

Figure 138. Mean Temperatures Measured in the ITD and Mixer. ............................................................................ 136

Figure 139. Speed Trend and Repeatability of the Compensated Small Wire Temperature Spectra at the 

Forward Combustor Location 12. ......................................................................................................... 137

Figure 140. Speed Trend and Repeatability of the Compensated Small Wire Temperature Spectra at the Aft 

Combustor Location 12. ........................................................................................................................ 138

Figure 141. Speed Trend and Repeatability of the Compensated Small Wire Temperature Spectra at the 

Forward ITD Location 12. ..................................................................................................................... 138

Figure 142. Speed Trend and Repeatability of the Compensated Small Wire Temperature Spectra at the Aft 

ITD Location 12. ................................................................................................................................... 139

Figure 143. Speed Trend and Repeatability of the Compensated Small Wire Temperature Spectra at the 

Mixer Location 12. ................................................................................................................................ 139

Figure 144. Coherence Between the Upstream and Downstream Small Wire Thermocouple Pairs at the 

March 26 48 Percent Speed Point. ....................................................................................................... 141

Figure 145. Cross-Spectral Phase Between the Upstream and Downstream Small Wire Thermocouple Pairs 

at the March 26 48 Percent Speed Point. ............................................................................................. 142

Figure 146. Example of Linear Fits of the Cross-Spectral Phase Between the Upstream and Downstream 

Small Wire T/C Pairs at the March 26 48 Percent Speed Point. .......................................................... 142

Figure 147. Measured Versus Predicted Convective Time Delays Between the Upstream and Downstream 

Probes in the ITD. ................................................................................................................................. 143

Figure 148. Coherence Between Dynamic Temperature and Dynamic Pressure at the Combustor Probe 

Locations, at 48 Percent Speed (March 26). ........................................................................................ 146

Figure 149. Cross-Spectral Phase Between Dynamic Temperature and Dynamic Pressure at the 

Combustor Probe Locations, at 48 Percent Speed (March 26). ........................................................... 146

Figure 150. Cross-Spectral Phase Between Dynamic Temperature and Phase-Corrected Pressure at the 

Combustor Probe Locations, at 48 Percent Speed (March 26). .......................................................... 147

Figure 151. Cross-Spectral Phase Between Dynamic Temperature and Phase-Corrected Pressure at the 

Combustor Probe Locations, at 54 Percent Speed (March 28). ........................................................... 147

Figure 152. Coherence Between Dynamic Temperature and Dynamic Pressure at the ITD Probe Locations, 

at 48 Percent Speed (March 26). .......................................................................................................... 148

Figure 153. Cross-Spectral Phase Between Dynamic Temperature and Dynamic Pressure at the ITD Probe 

Locations, at 48 Percent Speed (March 26). ........................................................................................ 148

Figure 154. Unwrapped Cross-Spectral Phase Between Dynamic Temperature and Pressure at the ITD 

Probe Locations, at 48 Percent Speed (March 26). .............................................................................. 149

Figure 155. Unwrapped Cross-Spectral Phase Between Dynamic Temperature and Pressure at the ITD 

Probe Locations, at 65 Percent Speed (March 28). .............................................................................. 149

Figure 156. Cross-Spectral Phase Between Dynamic Temperature and Phase-Corrected Pressure for ITD 

Probe Locations, at 48 Percent Speed (March 26). .............................................................................. 150

Figure 157. Cross-Spectral Phase Between Dynamic Temperature and Phase-Corrected Pressure for ITD 

Probe Locations, at 65 Percent Speed (March 26). .............................................................................. 151

Figure 158. Coherence Between Dynamic Temperature (Compensated) Upstream and Dynamic Pressure 

(Phase-Corrected) Downstream of the HP and LP Turbines. ............................................................... 152

Figure 159. Cross-Spectral Phase Between Dynamic Temperature (Compensated) Upstream and Pressure 

(Phase-Corrected) Downstream of the HP and LP Turbines. ............................................................... 152

LIST OF TABLES

Table 1. Sensor Evaluation Matrix. .............................................................................................. 5

Table 2. Summary of Test Points from the Burner Rig Test. ..................................................... 42

Table 3. Measured Probe Dimensions (cm). ............................................................................. 47

Table 4. Planned Operating Conditions for Data Collection. ..................................................... 72

Table 5. Matrix of Test Data Acquired. ...................................................................................... 76

Table 6. Operational Temperature Probes for Dewetron Data Files. ......................................... 77

Table 7. Location of Temperature Probes for the Testing on 3/26. ........................................... 77

Table 8. Location of Temperature Probes for the Testing on 3/28. ........................................... 77

Table 9. Temperature Increments Corresponding to 1.42 μV Quantization. ............................. 86

NASA/CR—2015-218879/VOL1 viii
LIST OF TABLES (CONT) 

Table 10. Thermocouple Probe Dimensions Used in the DGTMS Input File. ............................ 90

Table 11. Propagation Time Delays Assumed for Kulite Pressure Signals. ............................ 116

Table 12. Propagation Time Delays (in Milliseconds) Computed From Linear Fit of 

Measured Pressure Cross-Spectra. .............................................................................. 134

Table 13. Propagation Time Delays Computed for Small Wire T/C Pairs. .............................. 143

Table 14. Computed Time Delays Between Local T′ and P′ in the ITD. .................................. 150

ACRONYMS

APU Auxiliary Power Unit 

CAD Computer Aided Design 

CEL Cleveland Electric Labs 

CTA Constant Temperature Anemometry 

DGTMS Dynamic Gas Temperature Measurement System 

DTP Dual Thermocouple Probe 

EVNERT Engine Validation of Noise and Emission Reduction Technology 

FFT Fast Fourier Transform 

FPI Fabry Perot Interferometry

GRC Glenn Research Center 

HP High Pressure

HPT High Pressure Turbine 

ITD Inter-Turbine Duct

LP Low Pressure

LPT Low Pressure Turbine 

NIST National Institute of Standards and Technology 

OASPL Overall Sound Pressure Level 

PRT Platinum Resistance Thermometers 

PSDs Power Spectra Densities 

RMS Root Mean Square 

SOW Statement of Work 

SW Small Wire

T/C Thermocouple

TRL Technical Readiness Level 

UFF Universal File Format 

NASA/CR—2015-218879/VOL1 ix

1.0 EXECUTIVE SUMMARY 

In this program, a database of dynamic temperature and dynamic pressure measurements were 

acquired inside the core of a TECH977 turbofan engine to support investigations of indirect 

combustion noise. Dynamic temperature and pressure measurements were recorded for 

engine gas dynamics up to temperatures of 3100°F and transient responses as high as 1000 

Hz. These measurements were made at the entrance of the high pressure turbine (HPT) and at 

the entrance and exit of the low pressure turbine (LPT). Measurements were made at two 

circumferential clocking positions. In the combustor and inter-turbine duct (ITD), measurements 

were made at two axial locations to enable the exploration of time delays. The dynamic 

temperature measurements were made using dual thin-wire thermocouple probes. The 

dynamic pressure measurements were made using semi-infinite probes. 

Prior to the engine test, a series of bench, oven, and combustor rig tests were conducted to 

characterize the performance of the dual wire temperature probes and to define and 

characterize the data acquisition systems. A measurement solution for acquiring dynamic 

temperature and pressure data on the engine was defined. A suite of hardware modifications 

were designed to incorporate the dynamic temperature and pressure instrumentation into the 

TECH977 engine. In particular, a probe actuation system was developed to protect the delicate 

temperature probes during engine startup and transients in order to maximize sensor life. A set 

of temperature probes was procured and the TECH977 engine was assembled with the suite of 

new and modified hardware. The engine was tested at four steady state operating speeds, with 

repeats. Dynamic pressure and temperature data were acquired at each condition for at least 

one minute. At the two highest power settings, temperature data could not be obtained at the 

forward probe locations since the mean temperatures exceeded the capability of the probes. 

The temperature data were processed using software that accounts for the effects of convective 

and conductive heat transfer. The software was developed under previous NASA sponsored 

programs. Compensated temperature spectra and compensated time histories corresponding 

to the dynamic temperature of the gas stream were generated. Auto-spectral and crossspectral analyses of the data were performed to investigate spectral features, acoustic 

circumferential mode content, signal coherence, and time delays. The dynamic temperature 

data exhibit a wideband and fairly flat spectral content. The temperature spectra do not change 

substantially with operating speed. The pressure spectra in the combustor and ITD exhibit 

generally similar shapes and amplitudes, making it difficult to identify any features that suggest 

the presence of indirect combustion noise. 

Cross-spectral analysis reveal a strong correlation between pressure and temperature 

fluctuations in the ITD, but little correlation between temperature fluctuations at the entrance of 

the HPT and pressure fluctuations downstream of it. Temperature fluctuations at the entrance 

of the low pressure turbine were an order of magnitude smaller than those at the entrance to the 

high pressure turbine. Time delay analysis of the temperature fluctuations in the combustor was 

inconclusive, perhaps due to the substantial mixing that occurs between the upstream and 

downstream locations. Time delay analysis of the temperature fluctuations in the ITD indicate 

that they convect at the mean flow speed. Analysis of the data did not reveal any convincing 

indications of the presence of indirect combustion noise. However, this analysis has been 

preliminary and additional exploration of the data is recommended including the use of more 

sophisticated signal processing to explore subtle issues that have been revealed but which are 

not yet fully understood or explained. 

NASA/CR—2015-218879/VOL1 1
2.0 INTRODUCTION 

This report, prepared by Honeywell Aerospace, Phoenix, AZ, hereinafter referred to as 

Honeywell, presents the final report for the NASA sponsored program Acoustic Database for 

Core Noise Sources covering the period October 2011 through August 2014. This document 

covers the Phase I base effort, the Phase II engine test and posttest data reduction effort. 

2.1 Document Organization 

This document is organized according to the Statement of Work (SOW) tasks: 

 Section 2 summarizes the SOW elements

 Section 3 covers the Planning and Instrumentation task

 Section 4 covers the TECH977 Engine Modifications

 Section 5 covers the Data Analysis Methodology

 Section 6 covers the TECH977 Engine Test

 Section 7 covers the Data Processing and Distribution task

 Section 8 provides Conclusions and Recommendations for the overall effort

2.2 Program Task Definition – As Defined in the Statement of Work 

2.2.1 Task 1: Planning and Instrumentation 

Honeywell shall investigate the technology needed to acquire the dynamic gas temperature 

fluctuations with peak measurements potential of 2,000°F and frequency responses up to 1000 

Hz. To perform this task Honeywell shall: 

1. Configure data system to obtain the unsteady temperature fluctuations from sensor

signals.

2. Design probes and define installation requirements.

3. Provide thermocouple material for probe fabrication.

4. Identify, acquire, and test in a burner rig a limited number of probes.

5. Provide replacement hardware for the TECH977 engine as required for the

instrumentation for unsteady temperature.

6. Develop and deliver a detailed work plan for the base effort.

7. Develop and deliver a draft test plan containing a description of the test objectives and

success criteria, requirements to be satisfied, schedule, test description and test

conditions, number of tests and duration, and organizations involved with the respective

responsibilities identified for the Optional Effort (Phase II).

NASA/CR—2015-218879/VOL1 2
2.2.2 Task 2: TECH977 Engine Modifications 

Honeywell shall perform design modifications to the TECH977 engine for mounting the 

unsteady temperature and pressure sensors. This task also covers development of 

instrumentation to acquire the dynamic gas temperature fluctuations with peak measurements 

potential of 2000°F and frequency responses up to 1000 Hz. To perform this task Honeywell 

shall: 

1. Prepare an instrumentation layout; Honeywell shall seek NASA’s advice and

concurrence on needed measurements and probe locations.

2. Modify engine hardware and fabricate sensor mounts as required per design.

3. Configure measurement system.

4. Setup, checkout, and calibrate sensor system.

5. Provide estimates of unsteady temperature and pressure measurement uncertainties.

6. Coordinate test schedule and required data acquisition and test support equipment.

7. Refine, finalize and deliver a test plan containing a description of the test objectives and

success criteria, requirements to be satisfied, schedule, test description and test

conditions, number of tests and duration, and organizations involved with the respective

responsibilities identified for the optional Effort (Phase II).

2.2.3 Task 3: Data Analysis Methodology 

Honeywell shall identify all of the signal processing and analysis methods needed to process 

the data to characterize the noise. Techniques shall be configured to process unsteady 

temperature and pressure fluctuations time history data collected at multiple locations from 

within the engine. These techniques shall include consideration of circumferential acoustic 

modes and their axial propagation as well as the characterization of the levels of temperature 

and pressure fluctuations at the combustor exit plane. Honeywell shall seek NASA’s 

concurrence and advice on the selected data-analysis methods. To perform this task Honeywell 

shall: 

1. Identify existing methods needed for the analysis.

2. Configure and validate the analysis methods required by the instrumentation.

3. Document the selection and configuration of analysis methods.

NASA/CR—2015-218879/VOL1 3
2.2.4 Task 4: TECH977 Engine Test 

Honeywell shall conduct the TECH977 engine test to collect unsteady temperature and 

pressure fluctuations time history data at multiple locations from within the TECH977 engine. 

To perform this task Honeywell shall: 

1. Rebuild the engine with all modified hardware.

2. Perform an engine health check.

3. Install and terminate the instrumentation.

4. Collect unsteady time histories of temperature and pressure data in the area of the

combustor and turbine at various engine operating conditions; Honeywell shall seek

NASA’s advice and concurrence on measurement locations and engine operating

conditions.

5. Develop and deliver a test report providing a record of the testing performed, an

analysis and assessment of the test results, and the rationale for decisions.

2.2.5 Task 5: Data Processing and Distribution 

Honeywell shall process and distribute the data. This task includes the application of analytical 

techniques to extract the core-noise components from the unsteady temperature and pressure 

time history data collected at multiple engine locations during the test runs. To perform this task 

Honeywell shall: 

1. Process temperature and pressure data.

2. Perform acoustic modal analysis of the data.

3. Calculate spectra and root mean square (RMS) levels of the temperature and pressure

fluctuations at the combustor exit plane.

4. Distribute the data and analysis results; Honeywell shall provide a data package to

NASA with all resulting unsteady databases, including intermediate data sets, unsteady

sensor time histories, and documentation.

3.0 TASK 1: PLANNING AND INSTRUMENTATION 

3.1 Identification of Candidate Techniques 

At the beginning of this program, three technologies were identified as candidate sensors for the 

measurement of dynamic temperature: constant temperature anemometry (CTA), two wire 

thermocouple probes, and Fabry Perot Interferometry (FPI). These approaches were 

investigated to determine their ability to measure fluctuating temperatures at mean 

temperatures up to 2000°F. The candidates were ranked according to criteria that included 

frequency bandwidth, measurement accuracy, spatial resolution, sensor life, technical maturity, 

and technical readiness level (TRL). Table 1 summarizes the rankings of the candidate 

technologies. From among these alternatives it was determined that the two-wire thermocouple 

probe was the best approach by a significant margin when considering a variation of attributes. 

NASA/CR—2015-218879/VOL1 4
Table 1. Sensor Evaluation Matrix. 

A constant temperature anemometer measurement system can be purchased as a commercial 

off-the-shelf package from TSI Incorporated, but the fragility of the probe and the lack of high 

temperature maturity posed significant risk. The Fabry Perot Interferometer holds great promise 

for these types of measurements, but these systems are not of sufficient maturity to warrant 

their use on this program. On the other hand, the dual thermocouple approach has been well 

characterized and documented during previous research and development test programs. 

Several of these past investigations are documented in NASA technical memoranda and 

contactor reports (References 1, 2, 3, 4, 5, 6, and 7). This previous work provided an excellent 

background of the measurement technique and sensor capabilities. Furthermore, the data 

compensation software used to determine the fluctuating gas temperatures from the two-wire 

probe measurements had already been developed in previous NASA programs. Thus, the twowire approach had reached a high level of maturity, and was selected as the measurement 

technique for this program. In addition, NASA had available a sample temperature probe that 

could be used to validate the operation and robustness of the two-wire technique for the 

Honeywell engine application. This probe was loaned to Honeywell to develop experience and 

gain confidence with this measurement approach. 

For the measurement of dynamic pressure, the high temperatures in the combustor and turbine 

limit the type of instrumentation that can be employed. In the combustor, extreme temperatures 

preclude the direct mounting of transducers on the combustor wall. For example, water cooled 

piezoresistive pressure transducers are generally limited to mean temperatures below 1300°F. 

In such high temperature environments, the semi-infinite tube configuration is a mature and 

proven technique for measuring dynamic pressure. The sensor is mounted away from the 

hostile measurement zone. Acoustic waves are carried through a sensing tube past the sensor 

NASA/CR—2015-218879/VOL1 5
and then dissipated beyond the sensor by means of a long coil. Honeywell has gained 

experience with this method on several past engine test programs. To achieve higher accuracy 

for engine core measurements, where the mean pressures are very high, differential pressure 

sensors such as the Kulite XTE-190-10D model have been used. Semi-infinite probes using a 

minute nitrogen purge flow have been successfully used in previous turbine engine combustor 

tests to keep the transducers cool and protected from hot combustion gases. In addition to 

measuring dynamic pressure at a point, circular arrays of these semi-infinite tubes have been 

used to determine the circumferential mode content at a single axial station in the combustor of 

a Honeywell APU and a Honeywell propulsion engine (Reference 8). The amplitude and phase 

response of the transducers and the coils have been studied in these previous programs, and 

for the frequency ranges of interest in this program, excellent amplitude and phase response 

can be obtained. One alternative measurement option was considered. Kulite recently 

introduced the XTEL-SD-625 model transducer, which is a dual dynamic and static transducer 

with temperature capability to 900°F. The sensor incorporates an innovative mechanical 

filtering system, which takes the place of the semi-infinite tube, making it much more compact. 

However, this sensor was not used since the mean temperatures in the TECH977 combustor 

and turbine exceed its capability. 

3.2 Definition of Data Acquisition Systems 

NASA Glenn Research Center (GRC) loaned Honeywell a dual fine wire Type B thermocouple 

probe to assist with the design of the Honeywell probes and definition of the signal conditioning 

and data acquisition equipment. The probe was received from NASA GRC and used in a bench 

test to evaluate candidate arrangements. An initial instrumentation setup was proposed and 

evaluated. Two data acquisition systems readily available at Honeywell were identified as 

candidates for recording the temperature and pressure time history data. The first system was 

a Dewetron system, Model 801, configured with 64 channels of dynamic differential amplifiers, 

series MDAQ-SUB-ACC, and 24-bit delta-sigma analog-to-digital cards, model Orion 1624. The 

second system was Brüel and Kjær Pulse system, configured with 48 channels of 24-bit analogto-digital modules, using eight 6-channel LAN-XI Type 3050-A-060 cards, and driven with Pulse 

Labshop software. Both of these systems provide high-speed simultaneous acquisition of 

multiple data channels and have excellent dynamic range. The software driving the data 

systems also provides certain data analysis functionality. The Dewetron system has the 

capability to perform real time mathematical calculations on the data stream. Particularly useful 

in this test program was the ability to convert the instantaneous voltages from the temperature 

probes into physical units. The millivolt readings from the Type B thermocouples were 

converted to degrees Fahrenheit using National Institute of Standards and Technology (NIST) 

coefficients that were entered as calculation routines. This is quite helpful during engine testing, 

as it provides a check that the measured temperatures are in the expected ranges and provides 

feedback on the engine operation and probe deployment. The Pulse system does not have 

similar calculation capability, but it does provide the capability to perform real time auto and 

cross spectral analysis. The Dewetron system was selected as the primary data recorder, with 

the Pulse system functioning as a backup recorder. 

The primary challenge with the temperature measurement is that the voltages produced by 

Type B thermocouples are very small, typically on the order of a few millivolts. Furthermore, 

there is a need to resolve the small fluctuating voltages superposed on these mean levels, since 

this corresponds to the dynamic temperatures of interest. These voltages are on the low end of 

what the data acquisition systems are designed to record. This introduces two concerns. First, 

there is a concern about low signal-to-noise ratios due to the voltages being close to the 

electronic noise floor of the equipment. Second, there is a concern regarding signal 

NASA/CR—2015-218879/VOL1 6
quantization due to the analog-to-digital converters being optimized for a larger voltage range. 

An obvious solution to overcome these issues is to amplify the signals with a low noise amplifier 

prior to analog-to-digital conversion. 

Figure 1 shows the preliminary signal conditioning and data acquisition setup that was 

envisioned for the bench, rig, and engine testing. The Type B thermocouple wires are 

connected to a reference junction and a distribution amplifier, and the amplified signals are then 

routed into both the Dewetron and Pulse data acquisition systems. During bench testing of the 

temperature probes, it was determined that the signal-to-noise ratio of the amplified signals was 

actually degraded when readily available amplifiers were used (e.g. Vishay model 2310), due to 

the electronic noise associated with the signal amplifiers. It was determined that direct 

recording of the voltages from the temperature probes would provide a better overall signal-tonoise ratio and still provide acceptable signal quantization. As a result of the bench testing, 

direct recording of the thermocouple voltages was used in the rig and engine tests. 

Figure 1. Preliminary Signal Conditioning and Data Acquisition Setup. 

3.3 Development Probe Fabrication 

Honeywell needed to fabricate temperature probes for the bench and engine tests. Because 

the thermocouple wires used on these probes are quite small, special care is needed to 

assemble the thermocouple junctions and weld them to their supporting wires. In order to 

develop and gain experience with the probe fabrication process, two Type B dual-wire 

temperature probes were ordered from Cleveland Electric Labs (CEL) in Tempe, AZ. The 

probes feature two closely spaced beadless thermocouple junctions of different diameter. We 

NASA/CR—2015-218879/VOL1 7
had intended to manufacture the probes using the same diameter thermocouple elements that 

were used on the NASA probe, namely 0.003-inch and 0.010-inch. However, because the 

vendor could not locate a commercial source of 0.003-inch diameter Type B wire, the probes 

were built with 0.004-inch and 0.010-inch diameter wire. The thermocouple junctions were 

welded to support wires which are installed inside a ceramic stick. The development probes 

were designed with a thick Inconel housing so that they could be used in a rotating carousel 

test. These probes, shown in Figure 2, were used in the burner rig test to acquire dynamic 

temperature measurements and gain experience with the data acquisition systems and data 

reduction software. 

Figure 2. Dual Thermocouple Probes Used in the Bench and Burner Rig Tests. 

3.4 Development Probe Rig Testing 

Bench, oven, and burner rig tests were constructed to accomplish a number of objectives, 

including defining the optimal signal conditioning and data acquisition hardware, gaining 

experience in using the probes and data reduction software, and performing semi-quantitative 

quality checks on the data collected by the probes. For these tests, Honeywell procured two 

dual thermocouple probes (DTPs) from CEL. Special limits-of-error Inconel sheathed Type B 

thermocouples were also purchased for use as reference and control sensors for the oven test. 

For comparison purposes, NASA provided a DTP constructed with a 0.003-inch/0.010-inch wire 

diameter combination. NASA also provided the Fortran-based data reduction software routines. 

A test plan describing bench, oven, and burner rig testing of the Dynamic Gas Temperature 

Measurement System (DGTMS) was compiled and reviewed with NASA personnel. The 

DGTMS consists of (1) a DTP and (2) a collection of data processing software subroutines. The 

DTP consists of two closely spaced beadless Type B (platinum and 30 percent rhodium, 

platinum and 6 percent rhodium) thermocouple junctions installed in a ceramic stick, with one 

junction having a large diameter and the other having a small diameter. Each junction 

generates a small voltage that is related to the fluctuating temperature of the wire. The software 

processes these voltage signals and produces an estimate of the dynamic temperature of the 

gas convecting past the probe. The software uses the measured frequency response of the 

voltage signals produced by the two wire junction, along with the predicted frequency responses 

derived from an analytical model of the convective and conductive heat transfer in the probe, to 

compute a compensation spectrum that is applied to the measured data. 

NASA/CR—2015-218879/VOL1 8
3.5 Development Probe Bench Test 

The purpose of the bench test was to determine the signal conditioning and recording 

equipment that would be used to acquire data from the DGTMS probes. Some of the subtasks 

were to identify system hardware components and connectivity, evaluate the benefits and 

limitations of signal amplification, quantify the electronic background noise level and the impact 

of data system settings on the noise level, and identify and resolve any issues with cables and 

connectors. One of the challenges in acquiring data from the probes is that the voltage output 

from Type B thermocouples is very small (Figure 3), which places stringent demands on the 

signal conditioning and recording equipment. 

The bench testing was conducted by evaluating the voltage signal quality and electronic noise 

floor of various combinations of hardware. The capabilities of the data systems to measure very 

small voltages were examined by a couple of different methods. The first method was to 

directly inject low amplitude dynamic voltage signals into the data system inputs using a 

calibrated signal generator. The second method was to connect the probes to the data system 

inputs and use a heat gun to generate moderate temperature fluctuations at the probe ends, 

thus generating small fluctuating voltages (Figure 4). 

It was originally envisioned that signal amplification would be required, due to the very low 

voltage levels that are generated by the thermocouples and the desire to maximize the input 

range of the analog-to-digital converters on the data acquisition systems. However, it was 

found that use of the readily available signal amplifiers resulted in a decrease of the overall 

dynamic range due to increased electronic noise. Measurement of non amplified signals was 

determined to be a simpler and preferable solution, and provided a signal noise floor on the 

order of 20μV. Also, it was originally envisioned that a thermocouple reference junction would 

be used, but this was found to be unnecessary with the Type B thermocouples. The 

instrumentation setup that was defined as a result of the bench testing is indicated in Figure 5. 

This setup worked well in the subsequent oven and burner rig tests and was also used during 

the engine test. 

NASA/CR—2015-218879/VOL1 9
Figure 3. Voltage Output of Type B Thermocouples as a Function of 

Temperature. 

Figure 4. A Heat Gun Was Used to Generate Small Voltage From the Probes. 

Heat Gun   NASA Probe

  Honeywell Probes

NASA/CR—2015-218879/VOL1 10
Figure 5. Final Signal Conditioning and Data Acquisition Setup. 

3.6 Development Probe Oven Test 

The main purpose of the oven test was to verify that the Type B thermocouples used in the 

DTPs produce the same DC output as a reference Type B thermocouple when placed in a 

constant temperature environment. This would enable a comparison of the probes DC output 

against a known reference, ensure that the correct material was used in probe fabrication, and 

allow a comparison of the Honeywell and NASA probes. In addition, it was of interest to 

determine whether any additional noise is encountered in a laboratory test cell where there is 

more extraneous electrical noise, and when the probes are operating at higher mean 

temperatures. This test was accomplished by placing the NASA probe, the two Honeywell 

procured probes, and a commercially procured NIST traceable reference Type B thermocouple 

all together inside an oven. A Type K reference thermocouple was also used for an additional 

comparison. The probes were inserted through access holes in the back of the oven, and 

placed in close proximity to one another (Figure 6). The collection of probes were also 

surrounded by a small metal box, with an insulating material placed at the front and back, in 

order to minimize the impact of any convection currents generated inside the oven by the 

resistive heating elements in the oven walls. Steady state temperature tests were conducted 

with the oven set to provide a stabilized constant temperature in the range of 500°F to 2000°F, 

in steps of 150°F. Transient temperature tests were also conducted with the oven temperature 

slowly increasing or decreasing. The DC voltage readings from the large and small diameter 

thermocouple wires on the DTPs were compared to the voltage output from the reference Type 

B thermocouple. 

NASA/CR—2015-218879/VOL1 11
Figure 6. An Oven Test Was Conducted to Measure the DC Output of the Probes. 

At the steady state conditions, the voltage outputs of the Type B thermocouples were fairly 

similar, though not identical. At transient conditions, the voltage outputs from the Honeywell 

and NASA probes lagged behind the output of the Type B reference. Figure 7 shows a typical 

result during a continuous oven warm up from 1000°F to 1800°F, where the voltage outputs of 

the Type B and Type K thermocouples have been converted to temperature units using the 

appropriate NIST conversion coefficients. The temperature measured by the Type K beaded 

thermocouple was very similar to the temperature measured by the Type B beaded 

thermocouple. However, the temperatures measured by the small and large wires on the 

Honeywell and NASA DTPs lagged behind the Type B reference. This is illustrated more clearly 

in Figure 8, which plots the temperature difference from the Type B reference. The temperature 

differences were greater than what was expected. The rate of change of the oven temperature 

was sufficiently gradual (quasi steady state) that one would expect all thermocouples to read 

nearly the same temperature. A conclusive explanation for the difference was not determined. 

One possibility is that the mass of the ceramic stick and the cylindrical metal sheath on the 

Honeywell DTPs acts as a thermal source or sink during transients. However, this explanation 

is not entirely satisfactory. The response of Honeywell probe #1 is very similar to the NASA 

probe, which has a smaller diameter ceramic stick and no metal sheath near the tip. Another 

possibility is that the there might be some manufacturing differences in the thermocouple 

junctions. A second oven calibration test was made using a metrology thermocouple calibration 

device to further investigate the differences and is reported on in Section 3.8. 

HON Probe A HON Probe B

NASA

Probe

Type B, K

Reference

Oven

Box

NASA/CR—2015-218879/VOL1 12
Figure 7. Temperature Measured by the Thermocouples During an Oven Warmup. 

Figure 8. Temperature Differences Measured During the Oven Warmup. 

NASA/CR—2015-218879/VOL1 13
3.7 Development Probe Burner Rig Test 

The purpose of the burner rig test was to collect dynamic temperature data to support an 

evaluation of the data acquisition systems, the DTPs, and the DGTMS data processing 

software. Two types of tests were conducted. The first type was a test intended to collect 

“large-temperature-fluctuations” data by periodically exposing the Honeywell probes to a high a 

temperature burner rig gas stream, using a spinning carousel installed with rods. The second 

type was a test intended to collect “small-temperature-fluctuations” data by placing a one or 

more DGTMS probes in a fixed position in the burner gas stream and allowing the temperature 

fluctuations to arise naturally by the unsteady burning and turbulent mixing of the gas. 

Figure 9 shows a photo of a typical “large-temperature-fluctuations” test burner rig test setup. A 

stationary Honeywell DGTMS probe was positioned in the center of carousel that is installed 

with 12 rods. The carousel was positioned in front of the burner rig gas stream, and rotated at a 

fixed speed, thereby periodically exposing the probes to high temperatures. As the probes are 

periodically exposed to, and blocked from, the gas stream, large fluctuations in temperature of a 

semi-sinusoidal nature at a known frequency were created. The rotational speed of the 

carousel was varied to change the frequency of the main temperature oscillation. Data were 

collected at rotational speeds near 500, 1000, 1500, 2000, and 2800 rpm, corresponding to 

fundamental temperature fluctuations ranging from to 100 Hz to 560 Hz. Measurements were 

also made with the carousel stationary, and with the probe either directly behind a rod or in 

between two rods. Measurements were acquired for each of the Honeywell probes. A 

minimum of 60 seconds of data was recorded to provide sufficient averaging of the dynamic 

temperature spectra. 

As a side note, the original proposal for this test was to install the Honeywell DTP on the 

carousel and rotate the probe in and out of the burner flame, using slip ring to transfer the data 

from the rotating to stationary frames. This would provide a very large amplitude temperature 

fluctuation, with essentially two cycles per revolution. However, the maximum obtainable 

frequency of this large fluctuation would only be about 90 Hz, and the slip ring system would 

introduce additional noise. By holding the probe stationary in the center of a carousel that could 

accommodate 12 rods, temperature fluctuations at a much higher frequency (albeit with a lower 

amplitude) could be generated without any additional noise issues from the slip ring system. 

This latter approach was selected. 

NASA/CR—2015-218879/VOL1 14
Figure 9. “Large Temperature Fluctuations” Burner Rig Carousel Test. 

Figure 10 shows a photo of a typical “small-temperature-fluctuations” burner rig test setup. The 

NASA DTP, a Honeywell DTP, and a Kulite probe (semi-infinite coil arrangement) were placed 

at a fixed location in a hot gas stream. The probes were placed in the core of the gas stream 

with the intent of obtaining modest temperature fluctuations around a high mean temperature, 

with the fluctuations driven by turbulence and unsteady burning. The probes were also placed 

in the mixing region of the flow, with the intent of obtaining slightly larger temperature 

fluctuations around a lower mean temperature. The concept is that the observed temperature 

fluctuations arise naturally from the turbulent burning of the gas and from the turbulent mixing of 

the hot gas and surrounding air, with the expectation that this type of data might be 

representative of a measurement in the combustor or turbine in an engine test. 

Measurements were made for two fuel flow rates corresponding to mean gas flow temperatures 

near 1200°F and 1500°F, at two probe locations corresponding to the hot core and mixing 

region. The NASA probe and Honeywell probe #1 were installed simultaneously and placed in 

as close proximity as possible. One objective of this arrangement was to compare 

compensated frequency responses of the 0.003-inch/0.010-inch NASA probe and the 

0.004-inch /0.010-inch Honeywell probe, under the assumption that both probes experience 

comparable temperature fluctuations. A Kulite dynamic pressure probe was also installed next 

to the Honeywell DTP probe so that temperature-pressure cross spectra could be evaluated, 

after taking into account the time delay associated with propagation of pressure fluctuations in 

the Kulite probe tip. One observation was that the combustor flame front was very dynamic, so 

that it was not easy to identify the “edge” of the mixing region. The results of the burner rig test 

are discussed in Section 5.4, after a review of the data analysis methodology and software. 

HON Probe

Carousel

Rods

Burner

Carousel

HON Probe

NASA/CR—2015-218879/VOL1 15
Figure 10. “Small Temperature Fluctuations” Burner Rig Test. 

3.8 Development Probe Oven Calibration Test 

A follow-on DC temperature calibration test of the Honeywell and NASA probes was conducted 

using a metrology thermocouple calibration device to more carefully compare the mean outputs 

when the probes are placed in a uniform temperature environment. As described in Section 3.6, 

the initial oven test of the probes indicated that the voltage output (or temperature) of the 

Honeywell and NASA probes lagged substantially behind that of a reference Type B 

thermocouple during a slow oven warmup. The differences were larger than would be 

expected, especially since the small and large wires on the probes used the same Type B 

junction material as the reference thermocouple. The differences might have been due to 

convection currents or temperature stratification in the test chamber, due to the thermal mass of 

the ceramic stick and/or metal shroud of the probes acting as a source or sink of heat, or due to 

some other effect. Because the discrepancies were larger than expected, a second DC 

calibration test was proposed using a more controlled thermal environment. 

The Honeywell metrology department provided an Isotech Pegasus dry block thermocouple 

calibration device for this second evaluation. The Pegasus unit provides a much more 

controllable and uniform test environment than the oven. The device features a central well that 

is heated by a surrounding block to a precise, automatically controlled temperature. An 80mm 

deep metal insert sits inside the well, on top of which sits a 50mm ceramic insert that acts as a 

thermal insulator. The metal and ceramic inserts are drilled with holes to accommodate the 

installation of temperature sensors, whose sensing elements are positioned to an appropriate 

depth inside the metal insert. The available inserts had four holes with a diameter of 8mm 

(0.315-inch). The holes were large enough to accommodate the NASA probe (0.188-inch OD), 

but not the Honeywell burner rig “A” and “B” probes (0.5-inch OD). Although it was feasible to 

procure new inserts with larger holes that could accommodate the Honeywell development 

probes, this was not considered to be a value-added activity, since the probes considered for 

the engine test used a much different design, having a smaller diameter ceramic stick and no 

Kulite Probe

NASA Probe

HON

Probe

A

Burner Gas 

Stream

NASA/CR—2015-218879/VOL1 16
metal shroud near the sensing tip, similar to the existing NASA probe. The NASA dual 

thermocouple probe was tested in the Pegasus unit, along with three references, namely a Type 

B thermocouple (used in the oven test), a Type S thermocouple (used by metrology as a 

reference standard), and a platinum resistance thermometer (PRT, used per NASA request). 

The installation of the probes in the Pegasus unit is shown in Figure 11. 

Figure 11. An Isotech Pegasus Dry Block Calibrator Was Used to Evaluate the 

NASA Dual Thermocouple Probe Output Against Reference Standards. 

The four temperature sensors were placed in the calibrator, with each probe inserted into a hole 

in the ceramic and metal blocks, and with the sensing elements positioned in the metal block at 

approximately the same height above the bottom of the well. Ideally, when using a dry block 

calibrator, there should be a very close fit of the sensor within its hole, within a few thousandths 

of an inch if possible. In the present test, since standard size inserts were used, the fit was 

looser than ideal. The Type B reference had the loosest fit, with the Type S, NASA probe, and 

Platinum Resistance Thermometers (PRT) having tighter fits, in succession. The different gaps 

between the probes and their holes might have a small impact of the measurement (e.g. due to 

differences in conduction between the metal block and the probes, or differences in thermal 

gradients due to convective effects). 

A gradual heating and cooling cycle covering the range of 1000°F to 2000°F was performed, 

which lasted several hours. During the warmup, the unit was allowed to stabilize for 10 minutes 

at set point temperatures of 1000°F, 1500°F, and 2000°F. The voltage outputs of the NASA 

probe (small and large diameter thermocouple wires) and the Type B thermocouple were 

recorded on a Dewetron system and converted to temperature using the standard NIST 

coefficients. The temperature output of the Type S reference, which the metrology department 

normally uses as their standard reference for this unit, was recorded by a Fluke Black Stack 

thermometer readout. The temperature output of the PRT was recorded by a Hart Scientific 

(Fluke) Super Thermometer. The temperatures measured by the four probes during the 

warmup and cooldown are shown in Figure 12. The temperature measured by the NASA probe, 

Type S reference, and Type B reference all tracked closely together. The PRT, however, 

lagged behind significantly. Although PRTs are generally known for their accuracy, this 

particular PRT design is not very compatible with making measurements in the Pegasus device. 

The sensing element is not located at the very tip of the device, but an inch or so up the probe, 

so it is not exposed to the hottest part of the dry well. Because of this, the metrology 

department does not use the PRT as a reference standard with the Pegasus unit. Instead, the 

Type S thermocouple is used as a calibrated reference. 

Platinum resistance thermometer

NASA dual thermocouple probe

Type S thermocouple (metrology standard)

Type B thermocouple

50 mm ceramic insert (insulator) 

on top of 80mm metal insert

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Figure 12. Comparison of the Temperature Measured by Four Temperature 

Probes During a Warmup and Cooldown in the Isotech Pegasus Unit. 

Close-up comparisons of the readings from the probes at mean temperatures near 1500°F and 

2000°F are shown in Figure 13. The difference between the Type S and the Type B reference 

thermocouples is typically less than 5°F. The difference between the large and small 

thermocouples on the NASA probe is on the order of 5°F, and the difference between the Type 

S reference and NASA probe is typically less than 15°F. The results of the metrology test are 

very satisfactory. The magnitude of the observed differences in mean temperature will not have 

a significant impact on the computation of the compensated dynamic temperature spectra. The 

small differences between the NASA probe, the Type S reference, and the Type B reference 

could be caused by a number of factors, including: vertical or lateral variation of the temperature 

field inside the metal block; differences in pocket clearance; differences in data acquisition 

systems; and differences in the probes. A more refined test could identify what differences are 

due to the test setup as opposed to the actual probes. A more precise vertical alignment of the 

sensors might improve the agreement, along with longer stabilization times. Also, the 

agreement might be improved by fabricating a metal/ceramic inserts with closer tolerance holes 

that minimizes the gap between the sensors and the inserts. However, because the measured 

temperature differences would not have a significant impact on the compensated dynamic 

temperature spectra, it was not necessary to investigate these issues further. 

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Figure 13. Comparison of the Probe Temperature Readings Made in the Dry 

Block Calibrator at Temperatures Near 1500°F [Bottom] and 2000°F [Top]. 

3.9 Engine Probe Design 

The dual thermocouple probes were critical components of the engine test program and 

required special attention. The probes designed for the engine test were very similar to the 

NASA probe and the probes used in previous NASA/P&W programs (Reference 4, 6). The 

design criteria considered for the dual thermocouple temperature probes were much the same 

as those considered in previous reports (Reference 4). For this program, the design criteria 

were: 

1. Be able to operate in mean temperatures as high as 2000°F

2. Provide a frequency response up to 1000 Hz

3. Have a compact size (don’t want probe to affect temperature field)

4. Minimize adverse thermal conduction effects (no metal shroud near sensing tip)

5. Survive several hours of testing

6. Tolerate aerodynamic, thermal, and vibratory loads

7. Build on previous lessons learned

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The first criteria dictated the use of Type B sensing elements (at least for the combustor zone). 

The second criteria dictated the use of 0.003-inch/0.010-inch or 0.004-inch /0.010-inch wire 

combinations. The first, third, and fourth criteria dictated the use of a small ceramic stick to hold 

sensing elements. 

Figure 14 provides a schematic of the preliminary probe design. In addition to the metrology 

studies discussed in the previous section, the mechanical integrity of the probe during engine 

testing was a critical concern. Mechanical considerations for engine testing included 

(1) aerodynamic loading, (2) vibratory loads, and (3) thermal loads. Regarding aerodynamic 

loads, both the ceramic support element and the wires need to survive aerodynamic loads of the 

immersed flow field. Regarding vibratory loads, one needs to ensure that the ceramic sticks do 

not crack due to a resonant excitation from engine N1 or N2 shaft vibrations or from excitation 

induced by vortex shedding off of the probes. Regarding thermal loads, welds and wire 

arrangements must accommodate dramatic temperature changes without breaking. 

A mechanical analysis was performed to evaluate the aerodynamic loading of the probe at all 

measurement locations, since the probe protrudes into the operating engine gas stream like a 

cantilevered beam. A 3/16-inch diameter ceramic stick was found to be an acceptable size for 

the probe providing sufficient mechanical strength while minimizing the influence of the probe on 

the aerodynamic flow. Relative to the NASA probe design, two design changes were made 

based on recommendations in NASA CR-168267. The Honeywell probe (1) modifies the 

ceramic stick at the tip to allow more complete packing of ceramic cement into the holes, and 

(2) keeps the support wires straight to avoid introducing stress concentrations with bent wires 

which may lead to probe failure at high temperature. In addition, a reference Type B 

thermocouple was added to each probe in order to measure the mean temperature at the probe 

location with an independent thermocouple sensor. 

The typical operating range for Type B thermocouples is 400°F to 3100°F. The melting point of 

the platinum elements is near 3200°F. The ceramic stick, with 99.8 percent alumina content, 

has a maximum use, no load temperature of 3272°F. The Resbond ceramic adhesive has a 

maximum service temperature of 3000°F. The Inconel shroud supporting the ceramic stick, 

which remains outside of the combustor and turbine flow, has a melting point of about 2400°F. 

NASA/CR—2015-218879/VOL1 20
Figure 14. Preliminary Dual-Thermocouple Probe Design for the TECH977 Test. 

After the design criteria had been addressed, the design of the probe was finalized and an 

assembly drawing (70642386-1) was completed. A source of supply for 0.003-inch diameter 

Type B thermocouple wire was located, so the final design used 0.003′′ and 0.010′′ diameter 

wires for the small and large sensing elements, respectively. This combination of wire 

diameters enabled a good frequency response to be obtained for measurement of the dynamic 

temperature to a frequency of at least 1000 Hz. This combination was also used for most of the 

probes tested in the previous NASA/P&W program. Figure 15 shows the first page of the probe 

drawing, which defines the dimensions of the sensing elements and their 0.020-inch diameter 

support wires. The heights of the support wires and lengths of the sensing elements are similar 

to those used in previous test programs. The sensing elements are welded to the upstream 

side of the support wires. The support wires have been left straight to minimize the possibility of 

inducing local stress concentrations that could contribute to premature failure of the elements. 

An Inconel-sheathed Type B thermocouple was installed as a reference to cross-check the 

mean temperature measured by the fine wires. 

The thermocouple wires egress through a 15-inch long, six hole, 0.188-inch diameter alumina 

stick, and are terminated in a connector where the wiring transitions from platinum/rhodium to 

copper. The connector design, defined by the second page of the probe drawing (Figure 16), 

provides easy access for termination of the wires, and repair, if necessary. For structural 

support, the ceramic stick is housed in a 0.375-inch diameter Inconel tube, except for the last 

2.5-inch on the sensing end, which is inserted into the hot gas flow. 

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Figure 15. A Drawing (70642386-1) Was Created for the Dual Thermocouple Probes for the 

NASA/CR—2015-218879/VOL1 22
Figure 16. The Thermocouple Wires Egress Through an Alumina Stick and 

Are Terminated in an Easily Accessible Connector. 

NASA/CR—2015-218879/VOL1 23
4.0 TASK 2: TECH977 ENGINE MODIFICATIONS 

4.1 Engine Identification and Preparation 

The TECH977 engine S/N P119123 was located in engineering stores and was prepared for 

use in the engine test program. In Phase I, a draft engine test plan was prepared and design 

discussions were held to determine what modifications of the engine hardware would be 

required to mount the temperature and pressure instrumentation. 

4.2 Preliminary Instrumentation Layout 

A preliminary instrumentation layout was developed for the TECH977 engine test. The 

instrumentation consisted of 14 dynamic pressure Kulite probes, 10 dynamic temperature 

probes, and 2 external microphones, as shown in Figure 17. In the combustor, an array of 4 

Kulite probes at azimuthal positions of 0°, 90°, 180°, and 270° were planned to measure the 

circumferential mode content. The 4-probe array represents a subset of the 16-probe array that 

was previously used during the NASA EVNERT program (Reference 8). The purpose of this 

sub-array was to reaffirm the circumferential mode content that was previously measured 

(Reference 9). The modal content was reaffirmed and is discussed in Section 7.2.1. 

At the combustor exit, upstream of the high pressure turbine (HPT) a pair of dynamic 

temperature probes, spaced 2 inches apart axially (and offset azimuthally to avoid wake 

interference), would be located at 0°, with dynamic pressure probes located close to the 

temperature probes. Another pair of dynamic temperature and dynamic pressure probes, 

spaced 2.7 inches apart axially, would be located at 100° at the combustor exit. Similar pairs of 

probes would be located at the entrance to the low pressure turbine (LPT) at 0° and 100°. At 

the exit of the low pressure turbine, dynamic temperature and dynamic pressure probes would 

be located at 0° and 100°. Two external microphones would be located in the test cell. During 

the detailed design of the engine modifications, the instrumentation layout evolved and specific 

positions of the instrumentation were adjusted as manufacturing or installation constraints were 

identified. However, the general concept was maintained. 

NASA/CR—2015-218879/VOL1 24
Figure 17. Preliminary Instrumentation Layout for the TECH977 Engine Test. 

4.2.1 Engine Mechanical Design 

The overall goal of the engine mechanical design activity was to design a solution for 

incorporating the pressure and temperature instrumentation into the engine. This was a 

combined effort requiring input and coordination of multiple engineering disciplines, including 

instrumentation, computer aided design, engine project, manufacturing, mechanical analysis, 

acoustics, assembly and test. Numerous aspects were considered, including engine 

accessibility and sensor routing; availability and capabilities of actuator systems; impact of 

modifications on mechanical integrity of engine hardware; and modifications of engine assembly 

sequences to accommodate instrumentation. There were numerous challenges that had to be 

overcome, and numerous design iterations were required to meet the technical objectives. 

Rather than describe the full evolution and history of the design process, this section describes 

the final engine design solution, and provides salient comments on the design decisions and 

challenges encountered. 

4.3 Design Criteria 

The objective of the engine test was to collect unsteady time histories of temperature and 

pressure data in the area of the combustor and turbine at various engine operating conditions. 

The design criteria that went into the decision on the placement of the 10 temperature probes 

were: 

1. Measure at the combustor exit to acquire data relevant to HPT indirect noise generation

a. Two circumferential locations, to quantify azimuthal variation

b. Two axial positions, to quantify convection speeds

2. Measure downstream of HPT to acquire data relevant to LPT indirect noise generation

NASA/CR—2015-218879/VOL1 25
c. Two circumferential locations, to quantify azimuthal variation 

d. Two axial positions, to quantify convection speeds 

3. Measure downstream of LPT for residual temperature fluctuations 

e. Two circumferential locations, to quantify azimuthal variation 

The upstream and downstream probe pairs in the combustor exit and LPT entrance were 

azimuthally offset so that the downstream probe was not in the direct wake of the upstream 

probe, which could affect the fluctuating temperatures. The general schematic depicting the 

temperature probe placement relative to some engine hardware is shown in Figure 18. 

Figure 18. Schematic of Temperature Probe Placement (at 273° CCW ALF). 

Early in the program it was decided to use an actuation system to deploy the temperature 

probes. This decision was motivated in part by recommendations from previous testing of dual 

thermocouple probes on an F100 engine. The conclusions and recommendations from 

Reference 4 stated:

“A second area requiring further investigation concerns probe durability and 

obtaining increased probe life. The F100 engine test series demonstrated about one 

hour probe life for the present design. As described in the Task IV section of this 

report, the structural design of the probe was more than adequate for conditions in 

the engine. It is believed that the sensor element durability is extremely sensitive to 

particulate content in the airstream, especially during engine transients, and, at this 

time particulate content in the airstream is the hypothesis for failure. Increase in 

probe durability could be obtained by reducing sensor element and support wire 

lengths. Also, it is recommended that (1) the ceramic stick be modified at its tip to 

allow more complete packing of ceramic cement into the holes; and (2) the support 

NASA/CR—2015-218879/VOL1 26
wires be left straight rather than bent. Attempting to bend the support wires to mount 

the two elements in the same plane at probe centerline may have caused 

unacceptable stress concentrations and inadvertently aided probe failure. 

Fabrication and use of a probe retraction mechanism would also alleviate effects of 

particulates which occur during transients”. 

Thus, it was felt that a probe retraction system would be advantageous to extending probe life 

for the TECH977 engine testing. Recommendations regarding the probe design were also 

incorporated, as discussed in Section 3.9. This decision created a substantial design effort that 

included creation of computer aided design (CAD) models for relevant engine parts and 

externals, development of concepts for probe actuation, extension of the conceptual design to 

include a remote actuation and linkage method, and design of a method for sealing the probe 

when inserted into the combustor case. A number of design criteria were used to define the 

probe retraction solution: 

1. Deploy/retract probes gently after engine start/shutdown

2. Position probes at desired depth and orientation

 Ensure sensing wires are perpendicular to flow

3. Protect probe tip from flow when retracted

 Incorporate recessed cavity that shields probe tip

4. Prevent binding or breaking probes

 Address thermal growth differences between the combustor and combustor case

5. Maintain actuator temperatures below 400°F

6. Maintain temperatures at Type B/Cu-Cu junction below 400°F

7. Minimize air leakage around probe

 Small amount is acceptable, as influence on engine performance is not critical

 A 0.380-inch hole for a 0.375-inch probe produces about 0.02 lbm/sec/probe at

350 psi

A number of factors affected where the temperature probes could actually be placed, including 

accessibility around external plumbing on the engine cases, keep-away zones from flanges or 

critical areas, sufficient clearances from neighboring probes or other hardware, practical 

manufacturing constraints regarding reworked cases and fabricated bosses, and constraints on 

commercial available actuators. The design solution that emerged was to use commercially 

available actuators as follows: 

 Combustor exit: 4 temperature probes with 2 actuators, where each actuator drives an

upstream/downstream pair of probes via a yoke mechanism

 LPT entrance; 4 temperature probes with 2 actuators, where each actuator drives an

upstream/downstream pair of probes via a yoke mechanism

NASA/CR—2015-218879/VOL1 27
 LPT exit: 2 temperature probes with 2 actuators, where each actuator drives one probe

through the center of the actuator

4.4 Detailed Design

The instrumentation solution required the design and reworking of several existing engine 

components, the design and fabrication of several new parts, and the procurement of 

commercially available hardware. Detailed designs and formal drawings for the new and 

modified hardware were completed. New hardware included the following: 

 70642386-1 Thermocouple Probe Assembly

 70642387-1 Actuator Adapter (mounts between actuator and combustor/turbine case)

 70642388-1 Actuator Adapter (attaches mixer probe to actuator)

 70642389-1 Actuator Adapter (mounts between actuator and mixer nozzle)

 70642390-1/-2 Actuator Yoke (attaches combustor probes to actuators)

 70642391-1/-2 Actuator Yoke (attached turbine probes to actuators)

The reworked engine hardware included the following: 

 70642392-1 Combustor Case, Rework (from R3745568-1)

 70642393-1 Combustor, Rework (from R3745569-1)

 70642394-1 LP Turbine Case, Rework (from 3035657-1)

 70642395-1 LP Turbine Nozzle, Rework (from 3035658-2)

 70642396-1 Mixer Nozzle, Rework (from N10972-1)

 70642398-1 C-Duct Assembly, Rework (from P47G-09-0312-1)

In addition, a top level assembly drawing that defines the installation and integration of the dual 

thermocouple probes with the modified engine hardware was completed: 

 70642385-1 TC Probe Actuation System (Assembly)

CAD views of the modified combustor case and modified combustor are provided in Figure 19 

and Figure 20, respectively. Figure 21 shows the integration of the pressure and temperature 

probes in the combustor region. The two temperature probes are simultaneously deployed into 

(and retracted out of) the flow via a commercial pneumatic actuator via a single yoke. CAD 

views of the modified turbine case and modified LP turbine nozzle are shown in Figure 22 and 

Figure 23, respectively. Figure 24 shows the integration of the pressure and temperature 

probes in the inter-turbine duct (ITD). The two temperature probes are simultaneously deployed 

into (and retracted out of) the flow via a commercial pneumatic actuator via a single yoke. A 

CAD view of the modified mixer is shown in Figure 25. Figure 26 shows the integration of the 

pressure and temperature probes in the mixer. The temperature probe is actuated through the 

center of the actuator. Figure 27 shows modifications to the C-Ducts to accommodate passage 

of the probes through the bypass flowpath and outside the engine. Note that the actual 

hardware deviated slightly from these pictures due to manufacturing and assembly constraints 

encountered during the hardware fabrication and engine assembly process. 

NASA/CR—2015-218879/VOL1 28
Commercially available actuators, Bimba model Flat-1 FOD-313-HV, were chosen for this 

application (Figure 28). These are round cylinder, original style, double-acting, double-end rod, 

2-inch bore size, 3-inch stroke, hollow rod, high temperature actuators. The maximum 

operating air pressure is 200 psi, and the maximum operating temperature is 400°F. The range 

of travel – which controlled the location of the deployed and retracted probe element – was 

defined as follows. For the retraction stop, the probe was adjusted to the predefined retracted 

probe position, and the probe was locked to the fully retracted actuator via the yoke. The 

insertion depth of the probes was set by collars that were positioned around the Inconel shroud 

and locked in place with an adjustment screw, and which would hold the probe at the desired 

insertion depth when the actuators were activated. The probe translation rate was controlled by 

metering air flow to the actuator using a needle valve. Bench testing of the system using 

dummy rods was performed to fine tune the system and procedure. Figure 29 shows details of 

the actuation system assembly setup for the combustor and turbine. Figure 30 and Figure 31 

show details of the actuation system assembly setup for the mixer. 

Figure 19. View of the Modifications to the Combustor Case (70642392-1). 

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Figure 20. View of the Modifications to the Combustor (70642393-1). 

Figure 21. View of the Probe Integration and Actuation Setup for the Combustor. 

NASA/CR—2015-218879/VOL1 30
Figure 22. View of the Modifications to the Turbine Case (70642394-1). 

Figure 23. View of the Modifications to the LP Turbine Nozzle (70642395-1). 

NASA/CR—2015-218879/VOL1 31
Figure 24. View of the Probe Integration and Actuation Setup for the LP Turbine. 

Figure 25. View of the Modifications to the Mixer Nozzle (70642396-1). 

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Figure 26. View of the Probe Integration and Actuation Setup for the Mixer. 

Figure 27. View of the Modifications to the C-Ducts (70642398-1). 

NASA/CR—2015-218879/VOL1 33
Figure 28. Bimba Flat-1 FOD-313-HV Actuators Were Used to Deploy the Probes. 

Figure 29. Actuation System Assembly for the Combustor and Turbine. 

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Figure 30. Actuation System Assembly for the Mixer Nozzle. 

Figure 31. Details of the Actuation System for the Mixer Nozzle. 

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In February 2013, design reviews were held with Honeywell staff at Phoenix and NASA staff at 

GRC to review the approach to acquiring the dynamic temperature measurements on the 

TECH977 engine. During these reviews, the program objectives, prior experience, design 

criteria, measurement locations, probe design, engine/probe integration, engine modifications, 

engine operation, and risk and mitigation approaches were discussed. An agreement on the 

suitability of the measurement approach was obtained during these reviews. Several actions 

were taken, related to ensuring the structural integrity of critical hardware, validating the engine 

hardware and instrumentation assembly sequences, and conducting pretest instrumentation 

checkouts and engine check runs. The thermal differential growth of the combustor and 

combustor case was identified as a risk that could lead to binding and breakage of the 

temperature probes and potential subsequent damage to the engine. This risk was addressed 

during the hardware modification stage by enlarging the clearances in the combustor bosses. 

During engine assembly, the clearances were checked using dummy probes to ensure that the 

ceramic stick of the real probe would not contact the combustor boss (and exert a load on the 

stick that could cause it to snap) at the maximum predicted differential thermal growth. 

The draft engine test plan was updated to include the latest descriptions of the modified 

hardware and part numbers, a revised description of the probe design, and CAD views of the 

instrumentation. In addition, the pretest checkouts of the actuator and engine were updated, the 

engine test points were re-sequenced to avoid back-to-back repeats at the same condition, and 

the list of dynamic data channels was updated. A final test plan was released after all the 

drawings were completed, and provided details of the final engine and instrumentation 

configuration. The final plan included additional descriptions of the pneumatic actuation system, 

the dynamic pressure instrumentation, and the purge flow system, along with a discussion on 

measurement priorities and contingencies. 

5.0 TASK 3: DATA ANALYSIS METHODOLOGY 

5.1 Review of Analysis Techniques 

The data analysis methodology task focused on the signal processing and analysis methods 

needed to characterize the dynamic pressure and temperature fluctuations. This effort required 

consideration of the dynamic response of the sensors and sensor installations, and 

consideration of the subsequent processing of the dynamic signals from multiple sensors. The 

first question is concerned with measuring the dynamic properties of the gas at specific 

locations and involves identifying procedures to convert raw voltage signals into physical units 

that provide the best estimate of the “true” dynamic fluid properties of the system. The second 

question is concerned with gleaning additional information about the system, such as acoustic 

mode content or temperature fluctuation convection speeds, by examining relationships 

between signals at different locations. As with all dynamic measurements, it is important to 

consider the frequency responses of the transducers, the mounting configuration, the signal 

conditioning, the data acquisition system, and the subsequent signal processing. 

After down-selecting the dual thermocouple approach as the preferred technique for measuring 

dynamic temperature, previous NASA reports on the subject were reviewed to understand the 

technique and the processing methodology employed (References 1, 2, 3, 4, 5, 6, and 7). 

These reports document the development and validation of the technique, review 

measurements and analysis on combustor rig and engine tests, and describe the theoretical 

and computational procedures for compensating measured data to obtain the fluctuating 

temperature of the gas. The compensation procedure involves modeling both the convective 

heat transfer from the gas stream to the wires, and the heat conduction within the thermocouple 

NASA/CR—2015-218879/VOL1 36
wire and supports. The measured magnitude of the frequency response between large and 

small diameter wires, in conjunction with an analytical model of the frequency response ratio, 

enables the determination of the in-situ convective heat transfer coefficient, and subsequently a 

frequency response correction that is applied to the small diameter wire to obtain the local 

dynamic temperature. In this test program, the goal was to obtain compensated dynamic 

temperature measurements up to a maximum frequency range of 1 kHz. The software that 

performs this compensation was obtained from NASA and reviewed. 

Since both dynamic temperature and dynamic pressure fluctuations were being measured 

simultaneously in this program, an important consideration was to account for the phase delay 

in the pressure due to the propagation of acoustic waves in the semi-infinite tube from the probe 

opening to the transducer location, which requires knowledge of the temperature within the 

tube. This is of interest when computing cross-spectra between pressure measurements at 

different locations or between pressure and temperature signals. 

5.2 Dynamic Gas Temperature Measurement System (DGTMS) Software 

NASA provided a set of Fortran subroutines that comprise the DGTMS software for dual 

thermocouple probes. These routines were reviewed and compiled using the open source 

MinGW GNU Fortran compiler. Due to advances in data acquisition hardware and 

improvements in data processing methodologies, the software had evolved somewhat from its 

original implementation. This evolution is well documented in a series of published reports. 

However, the overall process flow remains largely the same. The original development and 

implementation of the DGTMS code is described in References 4 and 5. In that development 

program, the instrumentation approach and the data compensation methodology for measuring 

the fluctuating gas temperature using a DTP were defined and developed. The compensation 

procedure requires accounting for the dynamic conductive and convective heat transfer effects 

in the two thermocouples and their support wires. Laboratory burner and full-scale engine tests 

were conducted in which fluctuating temperature time history data were recorded on magnetic 

tape, and compensated spectra and time history data were computed using algorithms 

implemented on a HP5451C Fourier analyzer. In a follow-on program, documented in 

References 6 and 7, further development and testing of the DTP and DGTMS code were 

performed. The data processing algorithms were ported to Fortran IV and analog thermocouple 

time history data were acquired during various bench tests and processed on an IBM 

mainframe. More recently, as documented in Reference 3, enhancements were made to the 

DGTMS code to improve the speed of the fast Fourier transform (FFT) computations, implement 

an analytical solution for the probe frequency response in place of the original finite difference 

heat transfer model, port the software to a personal computer, and acquire the data digitally. 

This latter version of the software was used in this test effort. 

The overall process flow for the DGTMS is shown in Figure 32. The software reads a fixed 

format input file containing general information that includes: probe geometric parameters; 

operating point parameters such as fuel/air ratio, mean gas temperature, Mach number and 

mean gas pressure; analysis parameters; and plotting and output parameters. The software 

then reads a fixed format data file containing the digitized time history data from the DTP, 

specifically the AC voltage from the small diameter thermocouple, the AC voltage from the large 

diameter thermocouple, and the DC voltage from the large diameter thermocouple. The 

software then executes the data compensation procedure. Gas stream and thermocouple 

parameters are computed and an analytical estimate of the thermocouple-to-gas transfer 

function, as a function of an aerodynamic parameter gamma, is generated. The measured 

transfer function between the large and small thermocouples is computed, enabling a measured 

NASA/CR—2015-218879/VOL1 37
value of gamma to be interpolated from the analytical estimates. Subsequently, an appropriate 

transfer function can then be defined that relates the measured thermocouple temperature 

spectra to the gas temperature spectra. Application of this transfer function provides the 

compensated temperature spectra (i.e. gas temperature spectra), and Fourier inversion 

provides compensated time histories. 

The subroutines obtained from NASA were inspected to ensure that the methodology and 

procedures were consistent with the published documentation. The aforementioned reports 

describe a number of improvements and corrections that have been made to the software. It 

was verified that these modifications were in fact present in the version of the DGTMS software 

that was obtained from NASA. It was desired to check out and gain some experience with the 

software using laboratory or full-scale engine measurements acquired during the previous test 

programs. However, raw digital time history data and corresponding compensated spectra were 

not available from the previous test programs. A thorough test run of the software was 

conducted using data from the Honeywell burner rig testing. 

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Figure 32. Typical Data DTP Data Reduction Process (Figure 8 of CR-179513). 

NASA/CR—2015-218879/VOL1 39
5.3 Circumferential Mode Decomposition 

Options for the arrangement of dynamic pressure probe array and associated modal 

decomposition were investigated. The simplest arrangement for decomposing the 

circumferential mode content is an azimuthal array of equally spaced transducers. The 

circumferential mode content at the axial location of the array can be determined using a 

straightforward space-time FFT of the dynamic pressure time history data. In the NASA 

EVNERT program, a single circumferential array of 16 equally spaced semi-infinite Kulite 

probes was used to measure modes in the through-flow annular combustor of the TECH977 

engine (Reference 9). With this array, the circumferential mode content could be resolved to a 

frequency well beyond 3 kHz without any spatial aliasing, as shown in Figure 33. In order to 

allow more dynamic pressure probes to be used in other locations, a smaller array of 4 equally 

spaced probes was chosen for this program. The measurement serves as a validation that 

there are no dramatic changes in the circumferential mode content within the combustor. The 

reduced probe count results in the spatial aliasing of modes at a much lower frequency than 

occurred in the NASA EVNERT test. However, this effect is entirely predictable from the 

previous measurements. As shown in the right side of Figure 6, mode aliasing will occur above 

about 1000 Hz (e.g. there is an aliasing of the m = 2 and m = -2 modes near 1200 Hz). And as 

indicated by the arrows in the diagram, the aliased signals will cause the modes to appear 

within the measured modal analysis band. 

Figure 33. Aliasing of Circumferential Modes Is Predictable From Previous NASA 

EVNERT Array Measurements. 

Semi-infinite Kulite coils have been successfully used on several recent programs at Honeywell, 

including auxiliary power unit (APU) combustor noise measurements during the NASA RASER 

program and propulsion engine combustor noise measurements during the NASA EVNERT 

program. An example of the semi-infinite Kulite probe, and the purge flow system used to cool 

the probes, is shown in Figure 34. Recent work on semi-infinite coils was reviewed. Reference 

10 provides a useful comparison of flush mounted and semi-infinite probe measurements 

acquired using an impedance tube, a flow duct, and a micro-turbine engine, in the absence of 

any purge flow. Reference 11 provides an interesting computational assessment of the impact 

of temperature distributions on combustor probes that are cooled with nitrogen purge flow. The 

impact of using a cooled semi-infinite probe on the measured noise spectra, relative to a flush 

mounted microphone, is expected to be small, especially at the low frequencies that are of 

NASA/CR—2015-218879/VOL1 40
interest in this program. These references generally confirm this expectation, but the authors do 

provide a better quantification, or at least an approach to quantifying, the errors associated with 

the semi-infinite probe arrangement. 

One objective of the testing is to compute pressure-temperature cross-spectra at locations 

where both of these quantities are measured. The dynamic temperature measurements from 

the DTP are corrected using the DGTMS code to obtain the local gas temperature spectra 

and/or time history. The dynamic pressure data needs to be corrected to account for the 

frequency response of the probe system, and the most significant correction required is to 

account for the time delay in the signal as it propagates from the sensor tip to the Kulite 

location. A simple time delay can be computed assuming that the sound speed in the probe 

tube is known. This requires measurement of the air temperature in the tube. It was proposed 

to make measurements within the probe tube at one or more locations in order to provide an 

accurate time delay correction. Figure 35 shows an example of a Kulite tee that is instrumented 

with a thermocouple in order to measure the purge flow temperature. This type of tee was used 

to quantify the sound speed within the probe tube. 

Figure 34. Semi-Infinite Probes Have Been Successfully Used on Recent Engine 

Test Programs. 

Figure 35. Example of a Kulite Tee Instrumented With a Thermocouple. 

50 ft coil

combustor

Swagelok

fitting

transducer

tee

1 ft probe tube

0.125” diameter

combustor

case

case

boss

combustor

boss

reference line

XTE-190-10D

Kulite

plenum

reference 

pressure port

Nitrogen

supply

valve

lead

plenum

50 ft coil

combustor

Swagelok

fitting

transducer

tee

1 ft probe tube

0.125” diameter

combustor

case

case

boss

combustor

boss

reference line

XTE-190-10D

Kulite

plenum

reference 

pressure port

Nitrogen

supply

valve

lead

plenum

NASA/CR—2015-218879/VOL1 41
5.4 Burner Rig Test Data Analysis 

As described in Section 3.7, two general types of burner rig tests were conducted using the first 

two DTPs procured from CEL. The first was a rotating carousel test for use with the Honeywell 

probes which had the intent of generating “large temperature fluctuations” at a known fixed 

frequency (Figure 9). This test would explore the frequency response capability of the 

Honeywell probes. The second was a stationary probe test with the intent of creating “small 

temperature fluctuations” through the natural unsteady burning and turbulent mixing of the gas 

(Figure 10). This test would allow a comparison of the Honeywell and NASA probes when they 

were exposed to a comparable unsteady temperature environment. Time history data were 

acquired by both a Bruel & Kjaer Pulse system and a Dewetron system. The time history data 

from the B&K Pulse system were exported to Universal File Format (UFF) files for subsequent 

processing. The data were acquired at a fixed sample rate of 8192 S/s, and for a record length 

of 70 seconds. Table 2 summarizes the test points that were acquired in the burner rig test. 

Table 2. Summary of Test Points from the Burner Rig Test. 

An inspection of the time history data revealed that it was of good quality. Figure 36 shows a 

five second plot of the time history data from Honeywell probe #1 that was acquired by the 

Pulse system. The data is for the case of the probe installed in the center of the carousel 

rotating at 1000 rpm, at a mean gas temperature of 1580°F. The small diameter (0.004-inch) 

wire thermocouple clearly shows a higher response than the large diameter thermocouple wire 

(0.010-inch). The two wires appear to have approximately the same mean value, and the 

voltage fluctuations about the mean appear to be well synchronized. In this particular example, 

the small wire thermocouple has temperature fluctuations on the order of ±100°F (±3σ). If the 

data is examined closely, the quantization of the data can be observed. Although the input 

signal range of the Pulse A/D converter was fixed at ±10V, and no signal amplification was 

Probe 1 Probe 1 Probe 2 Probe 2

Carousel Expected Small T/C Large T/C Small T/C Large T/C

DGTMS Test Point Rotational Tone Mean Temp Mean Temp Mean Temp Mean Temp

Probe(s) Type Date & Time Speed Frequency Channel 1 Channel 2 Channel 3 Channel 4 Pulse UFF Filename

[rpm] [Hz] [°F] [°F] [°F] [°F]

HON B Carousel On 5/31/2012 11:38 1000 Test1_HonProbe1_1.uff

HON B Carousel On 5/31/2012 13:34 1000 200 1513 1508 n/a n/a 1500_1KRPM_HonProbe2_1.uff

HON B Carousel On 5/31/2012 13:41 2000 400 1568 1565 n/a n/a 1500_2KRPM_HonProbe2_1.uff

HON B Carousel On 5/31/2012 13:52 500 100 1526 1520 n/a n/a 1500_500RPM_HonProbe2_1.uff

HON B Carousel On 5/31/2012 14:00 1500 300 1523 1518 n/a n/a 1500_1500RPM_HonProbe2_1.uff

HON B Carousel On 5/31/2012 14:10 2000 400 1552 1545 n/a n/a 1500_2KRpm_HonProb2_B_1.uff

HON B Probe Behind Rod 5/31/2012 14:20 0 n/a 1536 1528 n/a n/a 1500_0RPM_HonProbe2_Blocked_1.uff

HON B Probe Between Rods 5/31/2012 14:26 0 n/a 1511 1503 n/a n/a 1500_0RPM_HonProbe2_NotBlocked_1.uff

HON B Ambient Noise, Burner Off 5/31/2012 14:32 0 n/a room temp room temp n/a n/a Ambient_1.uff

HON B Ambient Noise, Burner Off 5/31/2012 14:35 1000 n/a room temp room temp n/a n/a Amb_1KRpm_1.uff

HON B Carousel On 5/31/2012 15:16 2000 400 1464 1454 n/a n/a 1500_2KRpm_HonProb2_C_1.uff

HON B Carousel On 5/31/2012 15:22 2815 563 1508 1485 n/a n/a 1500_2KRpm_HonProb2_C1_1.uff

HON A Carousel On 6/1/2012 8:21 500 100 1425 1418 n/a n/a HP1_1500F_500Rpm_1.uff

HON A Carousel On 6/1/2012 8:33 500 100 1489 1480 n/a n/a HP1_1500F_500Rpm_A_1.uff

HON A Carousel On 6/1/2012 8:40 1000 200 1582 1567 n/a n/a HP1_1500F_1000Rpm_1.uff

HON A Carousel On 6/1/2012 8:44 1500 300 1539 1522 n/a n/a HP1_1500F_1500Rpm_1.uff

HON A Carousel On 6/1/2012 8:49 2000 400 1565 1547 n/a n/a HP1_1500F_2000Rpm_1.uff

HON A Carousel On 6/1/2012 8:53 2800 560 1465 1518 n/a n/a HP1_1500F_2800Rpm_1.uff

HON A Carousel On 6/1/2012 9:03 2800 560 1555 1539 n/a n/a HP1_1500F_2800Rpm_A_1.uff

HON A Probe Behind Rod 6/1/2012 9:11 0 n/a 1531 1513 n/a n/a HP1_1500F_0Rpm_Blocked_1.uff

HON A Removed Rod, Probe Exposed 6/1/2012 9:17 0 n/a 1009 1109 n/a n/a HP1_1500F_0Rpm_Not_Blocked_1.uff

HON A + HON B Carousel On 6/1/2012 10:00 500 100 1462 1415 1289 1527 HP12_1500F_500Rpm_1.uff

HON A + HON B Carousel On 6/1/2012 10:07 1000 200 1387 1342 1292 1569 HP12_1500F_1000Rpm_1.uff

HON A + HON B Carousel On 6/1/2012 10:12 2000 400 1404 1359 1193 1696 HP12_1500F_2000Rpm_1.uff

HON A + NASA Burner Edge, Fuel Press 94 6/1/2012 11:27 0 n/a 1109 1136 1144 1209 HP1NPKul_HotCore_1000_1.uff

HON A + NASA Burner Edge, Fuel Press 155 6/1/2012 11:39 0 n/a 1439 1465 1464 1545 HP1NPKul_HotCore_1500_1.uff

HON A + NASA Burner Core, Fuel Press 94 6/1/2012 12:16 0 n/a 1020 1036 1200 1247 HP1NPKul_Center_1000_1.uff

HON A + NASA Burner Core, Fuel Press 155 6/1/2012 12:20 0 n/a 1398 1415 1597 1657 HP1NPKul_Center_1500_1.uff

NASA/CR—2015-218879/VOL1 42
employed to maximize the A/D range, the quantization is seen to be small, since the 24-bit A/D 

converter has a very high dynamic range. The quantization corresponds to voltage level 

changes on the order of 1.4μV, which corresponds to about 0.3°F at a mean gas temperature of 

1500°F. 

Figure 36. Representative Time History Voltage Output from the Burner Rig Test. 

The test data was processed using the DGTMS software. The Fortran-based DGTMS software 

was obtained from NASA and compiled using the freely available GNU Fortran compiler. One 

code modification was made to the software. The subroutine TRFEM computes the analytical 

solution for the transfer function between the thermocouple wire and the gas stream (replacing 

the previous finite difference module). This subroutine was changed to increase the 

computational accuracy by using double precision in order to avoid numerical overflow. 

Although no test case was supplied with the code, processing of data from the burner rig test 

indicated that the software was operational and worked as described in the referenced reports. 

Input files were created containing the design specified probe dimensions, the approximate 

mean flow properties of the gas, and user-selectable program control options. The time history 

data files were converted into a format compatible with the DGTMS software. The DGTMS 

format consisted of a series of 280 blocks, with each block containing 2048 data points of the 

small thermocouple AC signal, 2048 data points of the large thermocouple AC signal, and 2048 

data points of the large thermocouple DC signal. In order to prepare these data files, the time 

history recordings from the large and small diameter thermocouples were each split into “mean” 

(DC) and “dynamic” (AC) components as follows. The DC component was generated by 

passing the data through a fourth order low pass Butterworth filter with a cutoff frequency of 1 

Hz. The AC component was generated by passing the data through a fourth order high pass 

Butterworth filter with a cutoff frequency of 1 Hz. The cut-off frequency is somewhat arbitrary, 

but a 1 Hz cutoff was chosen to be consistent with the analog filtering characteristics that were 

used on the previous NASA program (e.g. Figure 8 of Reference 4). The filters were designed 

in Matlab using the butter.m function and the filtfilt.m function in order to obtain zero phase 

distortion and to minimize the starting and ending transients. An example of the splitting of the 

time history files is shown in Figure 37. The phase preserving characteristics of the filtering 

process can be seen in the overlay of the raw and reconstructed time histories. 

NASA/CR—2015-218879/VOL1 43
Figure 37. An Example of the Splitting of a Time History Data Into “Mean” (DC) 

and “Dynamic” (AC) Components. 

The data analysis was performed using a frequency resolution of 4 Hz, which was chosen to 

have sufficient resolution and to allow convenient data manipulation. A total of 280 independent 

averages are available with the 70 seconds of data. A Matlab wrapper script was used to run 

the data files through the DGTMS software in a batch mode. The DGTMS software requires an 

input file that contains the probe dimensions, mean flow properties, and program control 

options. There can be some uncertainty in the probe dimensions and the mean flow properties. 

Regarding the probe dimensions, the DGTMS software employs a simplified physical model of 

the probe to compute the theoretical convective heat transfer between the gas stream and the 

probe wires, and the conductive heat transfer along the sensing and support wires. The actual 

physical construction is usually different from this idealized model. For example, there is often 

an overlap of the small sensing wires with their support wires to facilitate the welding of the two 

wires. In addition, there are often some differences between the design intent dimensions and 

the as-fabricated dimensions. 

Figure 38 shows a schematic of the ideal probe design (theoretical model), the Honeywell 

development probe design intent, the Honeywell development probe as-fabricated design, and 

the NASA probe configuration. To explore the impact of the different lengths of the 

thermocouple junction elements, selected data points were processed using input files with 

different dimensional inputs. An example of the impact of variation in the dimensional inputs is 

shown in Figure 39, for Honeywell Probe A in the carousel spinning at 1000 rpm. The 

difference in the compensated frequency spectra is on the order of a few tenths of a decibel 

absolute error over the frequency range of interest. Another example is shown in Figure 40, for 

a stationary Honeywell Probe A positioned at the edge of the burner rig flame, at a fuel pressure 

of 155 psi. The differences are more pronounced in this case. The addition of the measured 

small wire extensions to the DTGMS inputs increases the compensated spectrum amplitude by 

about 1.5 dB. Overall, the impact of using different dimensional assumptions for the lengths of 

the sensing elements did have a noticeable effect on the compensated spectra, but the 

magnitude of the differences were not substantial, and were dependent upon the operating point 

analyzed. Furthermore, the relative error between different points in the spectrum is small. 

Thus, the errors introduced due to these variations in as-fabricated probe dimension are not 

NASA/CR—2015-218879/VOL1 44
critical to the analysis. Careful attention to the fabrication of the probe sensing elements, and 

accurate measurement of the fabricated dimensions, help to minimize any errors associated 

with dimensional uncertainties. 

There is also some uncertainty in the Mach number, mean pressure and mean temperature of 

the flow in which the probes were immersed. A sensitivity study was also conducted to explore 

the impact of variations in assumed mean flow parameters on the compensated spectra. The 

compensated spectra were found to be insensitive to the mean flow inputs. These inputs are 

used to establish an initial theoretical estimate of an aerodynamic parameter Γ, which is then 

used to construct a family of theoretical frequency response function curves. The measured 

frequency response functions are interpolated using this family of curves to generate a 

measured in-situ value of Γ. Reasonable estimates of these parameters (within 25 percent) at 

the test operating conditions were found to be sufficient for the code to successfully compute 

this in-situ value of Γ. The resulting compensated temperature spectra are not affected by these 

mean flow inputs. 

Figure 38. Schematic of Ideal, As-Designed, and As-Manufactured Probes. 

NASA/CR—2015-218879/VOL1 45
Figure 39. Sample Comparison From the Dimensional Sensitivity Analysis, for 

Honeywell Probe A in the Carousel Spinning at 1000 rpm. 

Figure 40. Sample Comparison From the Dimensional Sensitivity Analysis, for 

Stationary Honeywell Probe A at the Burner Flame Edge, Fuel Pressure of 155 psi. 

NASA/CR—2015-218879/VOL1 46
Table 3. Measured Probe Dimensions (cm). 

Dimension NASA 

Probe 

Honeywell 

Probe A 

Honeywell 

Probe B 

Large Thermocouple 

Length of Support Wire 0.4064 0.4216 0.4864 

Length of the Small Wire 0.3277 0.2769 0.3480 

Diameter of Support Wire 0.0508 0.0508 0.0508 

Diameter of Small Wire 0.0254 0.0254 0.0254 

Small Thermocouple 

Length of Support Wire 0.2540 0.2921 0.3353 

Length of the Small Wire 0.2160 0.2768 0.3352 

Diameter of Support Wire 0.0457 0.0508 0.0508 

Diameter of Small Wire 0.00762 0.01016 0.01016 

The test data summarized in Table 2 were processed with the DGTMS software, using the 

measured probe dimensions in Table 3. Results from selected test points are presented here. 

First, a series of test points from the “large temperature fluctuation” carousel test of Honeywell 

Probe A are examined, for the carousel rotating at different speeds, and for the carousel 

stationary with the probe either blocked or not blocked by a rod (see Figure 9). Next, a series of 

test points from the “small temperature fluctuation” stationary test of Honeywell Probe A and the 

NASA probe (in close proximity) are examined, for the probes positioned at two locations in the 

burner flame, and for two fuel pressures (see Figure 10). 

Figure 41 shows the results of the carousel test of Honeywell probe A with the carousel rotating 

at 500 rpm. In this and subsequent figures, several items are shown. The DC and AC portions 

of the time histories for the small (blue) and large (red) thermocouples are shown in the top of 

the figure. Ideally, the DC signals of the large and small thermocouples would be identical, 

assuming that the two wires are located at essentially the same point in the flow and measuring 

the same mean gas temperature. In reality there is a small axial and vertical separation 

between the small and large thermocouple wires, so slight differences are expected. 

Furthermore, it can be expected that the AC signal from the small wire (0.004-inch) should look 

similar to the AC signal from the large wire (0.010-inch), but should have greater fluctuating 

amplitudes due to its greater responsiveness to temperature fluctuations. In the lower left of 

each figure, the power spectral density of the AC portion of the small and large wire 

thermocouple signals are plotted on a logarithmic scale, with a 0 dB reference of 1 K2

/Hz. 

These power spectra densities (PSDs) correspond to the spectra of the wire (not gas) 

temperatures. In the figures bottom left there is also an enlargement of a portion of the AC time 

history data. In the lower right of each figure, the compensated power spectral density is 

plotted, which is obtained by running the DGTMS software on the measured data. This 

compensated power spectral density represents the estimated temperature spectrum of the 

actual gas, and is determined by computing a small wire compensation spectrum using data 

from both thermocouples, and applying the compensation spectrum to the PSD from the small 

wire thermocouple. 

In Figure 41, the DC signals of the small and large wires are in good agreement, with some 

slight drift downwards with time due to a slight shift in the operating condition. The AC signals 

of the small and large wires oscillate in unison, with the small wire having higher 

responsiveness to temperature fluctuations as expected. The PSD of the small wire and large 

NASA/CR—2015-218879/VOL1 47
wires have similar shapes, with the amplitudes of the small wire being 5 to 10 dB above those of 

the large wire over most of frequency range. The compensated PSD reveals a peak at 100 Hz, 

which is equal to the rod passing frequency (500/60 rev/s*12 rods/rev), indicating that large 

amplitude temperature fluctuations are successfully being generated at the desired frequency. 

The broadband portion of the spectrum is quite flat with amplitude that is generally 10 dB below 

the 100 Hz peak. 

Figure 41. Results From the Carousel Testing of Honeywell Probe A at 500 rpm. 

Figure 42 shows the results of the carousel test of Honeywell probe A with the carousel rotating 

at 1000 rpm. The results are similar to the 500 rpm case, except for the expected shift in the 

rod passing peak to 200 Hz. Figure 43 shows the results for the carousel rotating at 2000 rpm. 

Again the results are similar, except for the shift in the rod passing frequency to 400 Hz. The 

broadband levels do not change significantly and remain flat over most of the frequency range. 

Figure 44 shows the results for the carousel rotating at its maximum speed of 2800 rpm. The 

compensated spectrum is more interesting in this case. In addition to the rod passing frequency 

at 560 Hz, there is a family of peaks at multiples of the carousel 1-per-rev frequency of 47 Hz, 

indicating that there are rod-to-rod temperature non-uniformities in the flow impinging on the 

probe. This could be due to the fact that there were physical differences between the rods (e.g. 

surface roughness and vertical alignment, see Figure 9) that are affecting the flow. The rotating 

carousel test confirmed the ability of the probes to measure temperature fluctuations to 

frequencies of at least 600 Hz, which encompasses the main expected combustion noise range. 

NASA/CR—2015-218879/VOL1 48
Figure 42. Results From the Carousel Testing of Honeywell Probe A at 1000 rpm. 

Figure 43. Results From the Carousel Testing of Honeywell Probe A at 2000 rpm. 

NASA/CR—2015-218879/VOL1 49
Figure 44. Results From the Carousel Testing of Honeywell Probe A at 2800 rpm. 

Figure 45 shows the results from the carousel testing of Honeywell Probe A at 0 rpm and with 

the probe positioned directly behind a rod. The DC signals from the small and large wires are in 

good agreement. The AC signals from the small and large wires, however, shows some 

anomalies starting at about 7 seconds into the time record and lasting for 38 seconds. As seen 

in the zoomed-in view, there appears to be some excess high frequency noise that appears 

suddenly. The PSDs of the AC signals are also suspicious, in that the spectra of the two wires 

quickly flatten out and merge together, calling into question the quality of the data. The source 

of the excess noise is unknown but it was noted that the potting material on the back of the 

probe was beginning to break off, perhaps allowing the wires to touch. The desired outcome for 

this test was to observe higher frequency temperature oscillation than could be obtained by the 

spinning carousel due to vortex shedding of the gas flow. It was expected that a half-inch rod 

immersed in the 1500°F flow with a typical Mach number of 0.3 to 0.4 would have a vortex 

shedding frequency well above 1000 Hz. Measurement of the vortex shedding phenomenon 

would provide confirmation of the upper frequency range capability of the probes. 

Unfortunately, due to the extraneous noise, the data above a few hundred Hertz is suspicious. 

The broadband levels of the compensated data are observed to increase strongly with 

frequency, which is questionable. Interestingly, there are some peaks at low frequencies, but 

their source is currently unknown. 

NASA/CR—2015-218879/VOL1 50
Figure 45. Results From the Carousel Testing of Honeywell Probe A at 0 rpm and 

With the Probe Positioned Directly Behind a Rod. 

Figure 46 shows the results from the carousel testing of Honeywell Probe A at 0 rpm and with 

the probe directly exposed to the flow. In this case a smooth, purely broadband spectrum was 

expected. The mean temperature was only around 1060°F, as opposed to about 1500°F on the 

previous test points. There is a larger discrepancy between the small and large DC signals, 

suggesting that the operating point or relative positioning of the probe in the flame changed. 

The broadband levels of the compensated spectrum are about the same amplitude as the 

rotating carousel test points. Interestingly, there is a tone near 700 Hz, and its harmonic at 

1400 Hz is also visible. This tone is also observable in many of the previous rotating carousel 

spectra. The origin of the tone is unknown, but it is hypothesized that it is related to the fuel 

delivery pump or system. 

NASA/CR—2015-218879/VOL1 51
Figure 46. Results From the Carousel Testing of Honeywell Probe A at 0 rpm and 

With the Probe Directly Exposed to the Flow. 

Of particular interest in this test series is how the Honeywell probe compares to the NASA probe 

when the two are exposed to the same fluctuating temperature environment (Figure 10). In 

particular, since the diameter of the small wire of the NASA probe (0.003-inch) is less than the 

diameter of the small wire on the initial prototype Honeywell probes (0.004-inch), it was of 

interest to examine any differences in the uncompensated and compensated small-wire spectra 

between the two probes. Note that the final design for the DTP used 0.003-inch and 0.010-inch 

diameter wires for the small and large sensing elements, respectively. Figure 47 shows a 

comparison of the data from the two probes, with the probes located near the burner flame 

edge, and with a fuel pressure of 94 psi. The mean temperature is about 1150°F. The DC 

signals from the four wires are in pretty good agreement and trend together, although the large 

wire thermocouple on the NASA probe appears to run a little high. The AC signals from the four 

wires also show the expected trends. The signals generally trend up and down together. The 

signals from the large wires on both probes match fairly well. And the signals from the small 

wires are similar, but with the small NASA wire having higher amplitude, as would be expected 

due to its smaller diameter. 

Comparing the uncompensated PSDs, the data is very encouraging. The spectra of the large 

thermocouples from the Honeywell and NASA probes (which both have wire diameters of 0.010-

inch) are almost identical, as would be expected if the probes are exposed to the same 

fluctuating gas temperature. The spectra of the small wire on the NASA probe has the same 

general shape, but higher amplitude, than the spectra from the small wire on the Honeywell 

probe, which is also expected due to the smaller diameter of the wire on the NASA probe. 

Comparing the compensated spectra, i.e. the estimated gas temperature spectra, one would 

ideally like to see the NASA and Honeywell probes to be in agreement if the probes are in the 

same point in the flow, or at least match well up to some high frequency limit, beyond which the 

NASA/CR—2015-218879/VOL1 52
two might be expected to deviate due to their different frequency response limits associated with 

the small wire thermocouple. The compensated spectra from the two probes are observed to 

be in fairly good agreement, with the Honeywell probe being about 2-3 dB higher than the NASA 

probe. It is difficult to conclude whether the small difference is a result of differences in the 

probes themselves or differences in the environment in which they are immersed. 

Figure 47. Results From the Stationary Testing of Honeywell Probe A and the 

NASA Probe, Positioned at the Burner Flame Edge, at a Fuel Pressure of 94 psi. 

Figure 48 shows a comparison of the data from the two probes, with the probes located near the 

burner flame edge, and with a fuel pressure of 155 psi. The mean temperature is close to 

1500°F. The conclusions are very similar to the previous comparison at 94 psi. The 

compensated spectra from the NASA and Honeywell probes are in even better agreement, with 

the Honeywell probe reading about 1-2 dB higher than the NASA probe. Also, the amplitudes of 

the compensated spectra are slightly higher than those at 94 psi, perhaps due to the higher 

mean operating temperature. 

NASA/CR—2015-218879/VOL1 53
Figure 48. Results From the Stationary Testing of Honeywell Probe A and the 

NASA Probe, Positioned at the Burner Flame Edge, at a Fuel Pressure of 155 psi. 

One might expect to see different levels of turbulent mixing, and thus different amplitudes of 

temperature fluctuations, when making measurements at different locations in the burner flame. 

Figure 49 shows a comparison of the data from the NASA and Honeywell probes, with the 

probes located near the burner core, and with a fuel pressure of 94 psi. The DC levels from the 

two probes show some significant differences, with the NASA probe reading higher (average 

temperature near 1200°F) than the Honeywell probe (average temperature near 1000°F), 

suggesting that there were significant differences in flow conditions for the two probes, despite 

their close physical proximity. But the AC signals for the two probes appear to oscillate in a 

similar fashion most of the time, indicating that there is still considerable similarity in the 

fluctuating temperatures that the probes are exposed to. Looking at the uncompensated 

frequency spectra, the large wire spectra for the two probes deviate slightly, whereas the small 

wire spectra are closer than was observed at the flame edge. This suggests that there is a 

discrepancy between the probes, or a difference in the environment in which they are placed, or 

both. The compensated spectra show large differences, on the order of 5 dB, with the 

Honeywell probe reading higher than the NASA probe. 

NASA/CR—2015-218879/VOL1 54
Figure 49. Results From the Stationary Testing of Honeywell Probe A and the 

NASA Probe, Positioned at the Burner Flame Core, at a Fuel Pressure of 94 psi. 

Figure 50 shows a comparison of the data from the NASA and Honeywell probes, with the 

probes located near the burner core, and with a fuel pressure of 155 psi. The results are very 

similar to the 94 psi core flame case. The higher fuel pressure translates to higher 

temperatures, with the NASA probe reading an average temperature near 1600°F compared to 

the Honeywell probe reading near 1400°F. 

NASA/CR—2015-218879/VOL1 55
Figure 50. Results From the Stationary Testing of Honeywell Probe A and the 

NASA Probe, Positioned at the Burner Flame Core, at a Fuel Pressure of 155 psi. 

Not all of the test data that was collected could be processed using the DGTMS software. 

There were some cases where there was a clear issue with the data itself. Figure 51 shows 

such an example, for the carousel test of Honeywell Probe A at the maximum rotational speed 

of 2800 rpm. The time history data for the small diameter wire was suspiciously noisy. The test 

setup was inspected, and it was found that the wires on the back side of the probe might have 

been touching due to the potting material falling off. The test point was repeated after the 

inspection of the setup, and good quality data was obtained that was successfully processed 

with the DGTMS software. 

Figure 51. Example of Problematic Data From the Carousel Test of Honeywell 

Probe A at 2800 rpm (First Attempt). 

As a second example, the data from Honeywell Probe B that was acquired when both Probe A 

and Probe B were in the carousel together, could not be processed, at either 500, 1000, or 2000 

rpm. Figure 52 shows the time history data for the 2000 rpm test point. The mean output from 

the small and large thermocouple wires on Probe B showed a very large split. The mean output 

from the small and large thermocouple wires on Probe A were in close agreement. This 

NASA/CR—2015-218879/VOL1 56
suggested that there was an issue with Probe B. An explanation for the problem is lacking. But 

data was acquired on the previous day for Probe B in isolation in the carousel, and the data did 

not show this behavior and was successfully processed. Figure 53 shows the time history data 

for Probe B in the carousel spinning at 2000 rpm, showing that the small and large wire DC data 

was in excellent agreement. 

Figure 52. Example of Problematic Data From Honeywell Probe B From the 

Carousel Test of Probes A & B Together at 2000 rpm. 

Figure 53. Example of Good Data From Honeywell Probe B From the Carousel 

Test of Probes B By Itself at 2000 rpm. 

Overall, the burner rig test provided a good exercise of the data acquisition systems and the 

DGTMS data reduction software. Data at a variety of test points were acquired, and the 

fluctuating voltages from the NASA and Honeywell probes were successfully processed with the 

DGTMS software. The software was robust and efficient, and well conditioned in that a failure 

to converge on a compensated spectrum often indicated that there was a fundamental problem 

with the data (e.g. noise on the signals, improper probe placement) or the input files (e.g. 

incorrect user inputs). The data showed the expected overall trends: the voltages of the small 

and large diameter thermocouple wires move in unison, and the small diameter wire is more 

responsive and has greater fluctuating amplitudes. Large fluctuating temperatures were 

produced at the expected frequencies in the carousel tests, both at the rod passing frequency, 

and to a lesser degree, at the one-per-rev frequency. Broadband temperature fluctuations were 

produced in both the carousel and stationary probe tests. 

In addition, the Honeywell and NASA probes, which had different designs and had different 

diameters for the small thermocouple wires, compared well. The spectra from the large 

diameter wires were generally in excellent agreement, as was expected since the wire diameter 

was the same for both, namely 0.010-inch. The spectra from the small diameter wire on the 

NASA probe (0.003-inch) was higher than that from the small diameter wire on the Honeywell 

probe (0.004-inch), which was also expected due to the higher responsiveness of the smaller 

NASA/CR—2015-218879/VOL1 57
diameter wire. The compensated spectra for the Honeywell and NASA probes were in close 

agreement, but not identical. The differences could be caused by different thermal 

environments to which the probes were exposed, despite being in close proximity. Overall, the 

results of the burner rig testing were very encouraging, providing a good check of the system 

capability, and providing experience and insight to ensure quality acquisition of fluctuating 

temperature data on an engine test. 

6.0 TASK 4: TECH977 ENGINE TEST 

6.1 Engine Build with Modified Hardware 

The TECH977 test engine, SN 119123, was moved from engineering stores to a development 

engine assembly area, and the engine was partially disassembled to gain access to the core 

engine hardware components requiring modification. Figure 54 shows a photo of the engine 

prior to disassembly. The C-ducts, which form the inner and outer flowpath of the bypass duct, 

are removed from the engine, so that the combustor and turbine cases are visible. This engine 

is the same serial number that was extensively tested during the NASA EVNERT program. 

Some of the pressure probe tubes (16 equally spaced Kulites) that were used to measure the 

dynamic pressure in the combustor are visible in the photo. 

Figure 54. Photo of the TECH977 Engine Prior to Relocation for Disassembly. 

Mixer 

Nozzle 

Combustor 

Case 

LP Turbine 

Case 

NASA/CR—2015-218879/VOL1 58
The combustor case was removed from the engine and modified to incorporate the guide 

bosses for the temperature and pressure probes in the combustor and turbine. The combustor 

development shop fabricated the adaptive hardware bosses (Figure 55) and modified the case 

by welding the adaptors and bosses in place. Figure 56 shows the combustor case being laser 

cut using an industrial Nd:YAG laser and the resulting access hole. Figure 57 shows the 

combustor case after the bosses had been welded on. 

Figure 55. Fabricated Combustor Case Wedges and Bosses. 

Figure 56. TECH977 Combustor Case Being Laser Cut. 

NASA/CR—2015-218879/VOL1 59
Figure 57. Modified Combustor Case Showing Bosses and 

Wedges Welded in Place. 

Significant amounts of rework were required to address issues discovered during fitting tests 

and during engine assembly. The modified combustor and combustor case were reworked in 

order to provide additional clearance for the thermal probes. Figure 58 shows how inaccurate 

engine modifications led to a misalignment of the thermal probes with the combustor case boss. 

This misalignment would have resulted in unacceptable interference between the ceramic shaft 

of the DTP and the boss clearance hole potentially resulting in probe fracture during use. 

Hence the bosses were reworked to extend the clearance hole. Figure 59 shows the 

relationship between the assembled and installed actuators and the extended bodies of the 

dummy probes. Notice that probes extending from the inter turbine duct assembly are not 

collinear due to the orientation of the underlying bosses that define the path of the probe. In 

order to avoid binding during actuation the drive yokes required modified to accommodate the 

differential travel between the two probe axes. Figure 60 and Figure 61 show how the yokes on 

either side of the engine were modified to accommodate this non collinear orientation of the 

probes. Slots were machined into the fixture brackets and locking knuckles were used on the 

probe bodies to carry the actuation load while accommodating motion in the plane of the yokes. 

NASA/CR—2015-218879/VOL1 60
Figure 58. Sensor End of the Three O’clock and Nine O’clock Orientations of the 

Dummy Probes Protruding Into the Combustor Case. 

Figure 59. Assembled Actuators With Dummy Probes Installed. 

NASA/CR—2015-218879/VOL1 61
Figure 60. End View Showing Modification of the Drive Yoke Used to Actuate 

Inter Turbine Duct Thermal Probes at the Three O’clock Orientation. 

Figure 61. End View Showing Modification of the Drive Yoke Used to Actuate 

Inter Turbine Duct Thermal Probes at the Nine O’clock Orientation. 

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In addition to the initial two DTP probes manufactured by CEL, a total of eleven probes were 

manufactured by Aerodyne Engineering of Indianapolis, IN. When initially received from the 

supplier these probes did not pass Honeywell’s quality standards. Figure 62 shows a typical 

example of their original as-received condition. Evidence can clearly be seen of welding blobs, 

uneven posts and wire lengths, excess ceramic coverage, thermocouple necking and overall 

poor quality fabrication. As a result, the probes were returned to the supplier for rework and 

remanufacture. Figure 63 shows that the subsequently delivered remanufactured probes were 

significantly improved. The complete collection of the assembled DGTMS probes ready for 

assembly into the modified engine is presented in Figure 64. 

Figure 62. Example Temperature Probe Originally Received From The Supplier. 

Figure 63. Temperature Probes Fabricated for the TECH977 Engine Test. 

NASA/CR—2015-218879/VOL1 63
Figure 64. Completed Temperature Probes. 

6.2 Engine Installation 

We had originally planned to run the engine in an indoor test cell, but when an outdoor cell 

became available, i.e. Honeywell’s San Tan Cell 965, we elected to use this cell since it 

provided better access around the engine. In addition this outdoor cell did not suffer from the 

reverberant environment effects of an indoor cell. The one change that was made to the 

acoustic measurement setup was to locate the two external microphones on inverted ground 

plane stands, rather than on poles. Figure 65 and Figure 66 show photos of the overall engine 

installation at Cell 965. The engine was configured with a laboratory bellmouth and screen, and 

a laboratory exhaust nozzle. 

NASA/CR—2015-218879/VOL1 64
Figure 65. View of the TECH977 Installed at San Tan Cell 965 (Right Side, ALF). 

Figure 66. View of the TECH977 Installed at San Tan Cell 965 (Left Side, ALF). 

NASA/CR—2015-218879/VOL1 65
The nitrogen purge flow systems for the combustor and the turbine Kulite instrumentation are 

shown in Figure 67 and Figure 68. The systems were installed according the purge flow 

schematic in the engine test plan. Figure 69 shows the installation of the temperature probes in 

the combustor, ITD, and mixer. Figure 70 shows a close-up of the installed combustor and 

turbine probes at the right hand side (ALF). Figure 71 shows the locations of the two external 

microphones. 

Figure 67. View of the Nitrogen Purge Flow Supply System. 

Figure 68. View of the Purge Flow Manifolds, Kulite Coils, and Kulites. 

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Figure 69. View of the Installed Temperature Probe Hardware. 

Figure 70. Closeup of the Pressure Tubes and Temperature Probes. 

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Figure 71. Location of the Two External Microphones. 

6.3 Instrumentation Setup 

As described in Section 3.2, a Dewetron system, Model 801, configured with 64 channels of 

dynamic differential amplifiers, series MDAQ-SUB-ACC, and 24-bit delta-sigma analog-to-digital 

cards, model Orion 1624, was used as the primary data acquisition system. A Brüel & Kjær 

Pulse system, configured with 48 channels of 24-bit analog-to-digital modules, using eight 6-

channel LAN-XI Type 3050-A-060 cards, and driven with Pulse Labshop software (Version 

18.0.1.1) was used as a backup system. The voltages from the temperature probes were 

recorded directly without amplification and without the use of a reference junction, using the 

system connectivity described in Section 3.7, Figure 5. In order to minimize signal losses and 

electrical interference, the cable lengths for the temperature probes were kept as short as 

possible. Each of the ten temperature probes were connected to the Dewetron data acquisition 

system using a 10.5-foot long cable. Each cable contained three wire pairs: one pair for the 

small wire thermocouple, one pair for the large wire thermocouple, and one pair for the 

reference thermocouple. The wires were shielded to minimize electrical interference. A set of 

4-foot long patch cables of similar construction were used to tee the voltages from the Dewetron 

system to the Pulse system. Figure 72 shows the connection of the temperature probe cables 

to the Dewetron system, and the cables that tee the signals over to the Pulse system. Figure 73 

shows the connection of the tee cables to the Pulse system. 

NASA/CR—2015-218879/VOL1 68
Figure 72. View of the Cable Connections to the 64-Channel Dewetron Recorder. 

Figure 73. View of the Cable Connections to the 48-Channel Pulse Recorder. 

The signals from the thirteen Kulites and the two external microphones were recorded 

simultaneously with the temperature signals. A rack of Sonoran Microsystems amplifiers was 

used to provide the signal conditioning (voltage supply, filtering, and amplification) for the 

Kulites, as shown in Figure 74. The amplified signal outputs were connected to both the 

Dewetron and Pulse data recorders. Figure 75 shows the cable connections on the Sonoran 

NASA/CR—2015-218879/VOL1 69
Microsystems unit; the gray cables provide the power supply to and the signal response from 

the Kulites, and the white cables are for the amplified output signals. The Kulite transducers 

were calibrated as follows. On the Dewetron system, each Kulite channel was calibrated by 

entering the factory calibrated sensitivity from the vendor supplied data sheets. The factory 

sensitivities were very close to the nominal transducer sensitivity of 10.0 mV/psi for the XTE190-

10D model Kulite. On the Pulse system, each Kulite channel was calibrated by applying a 

pistonphone signal with the use of a special adapter. The resulting calibration was nearly the 

same as the nominal sensitivity, so there was very little difference between the two systems. 

Due to the short cable lengths of the temperature probes, the Dewetron and Pulse data 

acquisition front ends needed to be placed close to the engine. Each data system was placed 

on a cart located beneath the exhaust nozzle. Each data system was operated remotely from 

the console room using a dedicated laptop connected to the front end via an Ethernet cable. 

Pretest checkouts of the data systems were conducted after the probes were installed to verify 

that the data systems were operating properly and that there were no sources of electrical noise 

from the test facility that would contaminate the data. 

The Kulite measurements were made using semi-infinite probes. The length of the sensing end 

of the tube was selected to position all of the Kulites outside of the engine bypass duct. This 

was done to enable access to the Kulite sensors should any fail during the test and need to be 

replaced. This also ensured access to the Kulite tees and fittings in case any leaks were 

discovered during the checkout of the purge flow systems. The overall length of the sensing 

tube, from the sensing end, located flush with the outer wall of the core flowpath, to the Kulite 

location in the tee, was 35.875-inch for all of the Kulite sensors. The sensing line was 

constructed by brazing a one foot section and a two foot section together using an outer sleeve 

(to avoid internal discontinuities in the flowpath), to produce a 36-inch line. A 0.5-inch length of 

the sensing line was cut off during installation of the Kulite tees (the ends of the tubes had been 

previously brazed shut during the checkout of the purge flow system). When the sensing line is 

installed onto the tee, the distance from the end of the sensing line to the center of the tee 

where the Kulite sensor sits is 3/8-inch. Thus, the overall length from the sensing end of the 

tube to the sensor face is 35.875-inch. This length was made identical for all 13 Kulites to 

simplify the subsequent data corrections. 

Figure 74. View of the Front of the Sonoran Microsystems Signal Conditioning 

Rack Used With the Kulites. 

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Figure 75. View of the Cable Connections on the Sonoran Microsystems Signal 

Conditioning Rack Used with the Kulites. 

6.4 Pretest Checkouts 

Pretest checks of the data acquisition system, the purge flow system, and the probe actuation 

system were conducted prior to the actual test to ensure that all systems were functioning as 

desired. A checkout of the data acquisition system was conducted to ensure that there were no 

new sources of electrical interference introduced at the test cell, relative to the testing that was 

performed in the laboratory. The background electronic noise floor associated with the 

temperature probes and cabling was consistent with levels measured in the bench testing. A 

checkout of the purge flow system was performed prior to installation of the Kulites to ensure 

that the purge flow supply pressure tracked the reference pressure in the combustor and turbine 

sections of the engine. The delta-p pressure regulators were adjusted to provide an initial 

pressure (prior to engine start) of a few psi above the (ambient) reference pressure. During 

engine operation, the regulators adjust the supply pressure in unison with changes in the 

reference pressure. However, there was a tendency of the delta-p to drop slightly as the engine 

speed (and core pressure) increased. The delta-p pressure regulators were adjusted to both (1) 

maintain a positive delta-p to ensure that purge flow is supplied through the Kulite tubes at all 

speeds and (2) minimize the delta-p so that the purge flow does not create excessive 

hydrodynamic noise over the Kulites or disrupt the temperature measurements. 

A checkout of the probe actuation system was performed to ensure that the system was 

operational and to make final adjustments to the amount of manifold pressure. The desire was 

to provide enough pressure to ensure full deployment, but low enough to provide a gentle 

insertion so as not to jar the delicate thermocouple elements. The initial checkout was 

accomplished using a set of dummy probes of the same dimensions as the actual probes. The 

dummy and actual probes were scribed to ensure that the correct insertion depths were 

obtained at the deployed condition. The checkout indicated that at some locations, a chamfer 

was required on the Inconel housing to enable the probe to be installed into the guide boss. 

After the initial checkouts with the dummy probes, the sheaths of the temperature probes were 

coated with a nickel-based anti-seize compound and the real probes were installed. The probes 

were installed with the sensing elements aligned normal to the flow (engine axis). A final test of 

NASA/CR—2015-218879/VOL1 71
the actuation system was performed prior to starting the engine using the actual probes. Video 

cameras were placed to observe the actuation of the probes during engine operation. 

6.5 Test Data Planned 

The test plan called for collection of data at the operating conditions defined in Table 4 where 

NFC is the corrected fan speed and N1C is the corrected low pressure spool speed. The 

operating conditions consisted of ambient noise with and without purge flow, ground idle, and 

four low power operating speeds. These engine speeds cover the range of interest for an 

aircraft approach condition. Since jet noise is relatively low on approach, this is the regime 

where combustion noise would be of practical importance to aircraft noise. The four speeds 

were to be repeated to provide an indication of data repeatability. The objective was to 

systematically step through the test matrix and record 70 seconds of data at each condition. 

The data would be acquired by both the Dewetron and Pulse systems at roughly the same time, 

although the two data systems were not precisely synchronized in time with one another. The 

data would be acquired at a sample rate of 10 kHz on the Dewetron system and at 65.536 kHz 

on the Pulse system. The temperature probes were to be deployed and retracted before and 

after each engine speed point to preserve probe life while the next operating point was 

established. Engine operational and performance data would also be collected during testing at 

a sample rate of 10 Hz. 

Table 4. Planned Operating Conditions for Data Collection. 

Test Code NFC [N1C] Percent Kulite Purge Flow Temperature Probe 

977CNAMNP Engine Off OFF Retracted

977CNAMWP Engine Off ON Retracted

977CNAMGI Ground Idle ON Deployed

977CNAM4801 48.0 [47.0] ON Deployed

977CNAM5401 54.0 [52.9] ON Deployed

977CNAM6001 60.0 [58.8] ON Deployed

977CNAM6501 65.0 [63.7] ON Deployed

977CNAM4802 48.0 [47.0] ON Deployed

977CNAM5402 54.0 [52.9] ON Deployed

977CNAM6002 60.0 [58.8] ON Deployed

977CNAM6502 65.0 [63.7] ON Deployed

6.6 Test Data Acquired 

The desired test matrix was completed, but the test execution differed from the plan due to 

several challenges that emerged during testing. The initial part of the test worked exactly as 

desired. On 3/26/2014, after performing the ambient noise measurements, the engine was 

started, warmed up at idle, and then set to the first operating condition of 48 percent corrected 

fan speed. The data systems were turned on (both the high sample rate Dewetron and Pulse 

systems, and the low sample rate engine data recorder). The probe actuators were enabled 

and the probes were successfully deployed into the flowpath. The successful deployment and 

operation of the probes indicated that the thermal growth of the engine was properly accounted 

for in the design of the actuation system. All of the temperature probes responded to the high 

core temperatures, as witnessed by the calculated temperatures on the Dewetron system, and 

data was collected for 70 seconds. Figure 76 shows a sample readout from the Dewetron 

system. The ability to have real time computation of the wire temperatures with this system was 

extremely helpful. 

NASA/CR—2015-218879/VOL1 72
Figure 76. Sample Readout from the Dewetron System at 48 Percent Speed 

(3/26). 

All of the probes in the combustor region survived this initial test point, which was a 

considerable achievement as there was significant uncertainty regarding the peak operational 

temperatures at the combustor probe locations. The mean temperatures measured in the 

combustor, particularly at the forward locations (2900°F), were substantially higher than 

expected based on engine cycle estimates at the combustor exit location (1800°F), but not out 

of line with what might be expected in the flame zone. After collecting the data, the probes were 

retracted and the data systems were stopped. It was noticed that the retraction of the probes 

was much slower than what was observed during the checkout, possibly indicating that the 

probes were starting to stick, which raised a concern about whether the repeated deployment 

and retraction of subsequent test points would be successful. The test team had a brief 

discussion on whether to proceed with the next planned test point at 54 percent, or skip to a 

higher power point first, in case any probes or actuators started to give problems. It was 

estimated that the 65 percent test point was likely too hot for the forward probes to survive, 

based on temperatures observed at 48 percent and the expected temperature increase with 

speed. So the 60 percent speed point was attempted. The probes were inserted and appeared 

to deploy, but the Dewetron data system locked up. The probes were retracted and the engine 

was brought to idle to reset the data system. A second attempt at the 60 percent speed point 

was then attempted. However, the forward and aft combustor probes at clocking location 12 

(see Figure 77 for instrument clocking locations) did not appear to fully insert, and the small wire 

signal on the aft combustor probe at location 12 was lost (presumably due to a failure in the 

connector). Also, the small wire thermocouple signal on the combustor aft probe at location 4.5 

NASA/CR—2015-218879/VOL1 73
was lost shortly after insertion, even though the temperatures at this location were modest. The 

probes were retracted to diagnose why the combustor probes at location 12 did not insert fully. 

During retraction, the aft combustor probe at location 4.5 would not retract fully. The engine 

was brought to idle to try to free the stuck probes, and an attempt at the 54 percent speed point 

was made. The probes were deployed and data was collected, but the combustor probes on 

both sides did not insert fully. The engine was shutdown to diagnose the probe actuation 

system lockups and the data system lockups. 

Figure 77 Instrumentation Clocking Locations, Aft Looking Forward. 

On 3/27/2014, repairs were made to the installation. The aft probe at combustor location 4.5 

was completely frozen and could not be removed from the engine, even with significant force. It 

was disconnected from the actuator. The combustor probes at location 12 were not operational 

after shutdown. The spare probe was installed in the aft combustor position, and the mixer 

probe at location 4.5 was moved to the forward combustor position. This was done to provide a 

full complement of operational probes on the left side of the engine (ALF). Anti-seize lubricant 

was applied to all of the probes. Checkouts of the data systems were also performed but did 

not reveal any anomalies that would explain the lockups. The Dewetron data file at 60 percent 

speed was found to be corrupted. The sample rate on the Dewetron recorder was increased 

from 10 kHz to 20 kHz, since fewer channels were being recorded and the data system could 

process the remaining channels at a higher rate. 

Data collection continued on 3/28/2014. Because of the difficulties with the actuation system, it 

was decided to actuate the probes after reaching the first operating point, and leave them 

deployed until the last test point. Data was collected at 48 percent and 54 percent, and these 

points were repeated. During these test points, the Dewetron and Pulse systems again 

experienced intermittent lockups. On the Dewetron system, the displays showing the real time 

data would briefly freeze up. Subsequent inspection of the data file showed warning flags 

NASA/CR—2015-218879/VOL1 74
stating that data records were dropped, but this appeared to be a rare event. On the Pulse 

system, the software would also occasionally lock up. Data was next collected at 60 percent 

speed. The forward combustor probe at location 4.5 was lost due to excessive temperatures 

(>3100°F). After data collection, the probes were retracted and the engine was shut down to 

again diagnose the data system lockups. 

Subsequent troubleshooting identified that a potential cause of the data system lockups was 

high sound levels near the engine might have induced excessive vibration of the hard drives, 

causing problems with streaming the data to disk. The acquisition systems could not be 

relocated due to the short cable lengths used with the temperature probes, so instead they were 

covered with heavy Kaowool acoustic blankets to shield them from the engine noise. After this 

was done, no further problems were encountered. 

The probes were recoated with anti-seize compound and the engine was started and stabilized 

at the 54 percent speed point. The probes were deployed, and data was collected. Data was 

then collected at 60 percent and 65 percent corrected fan speed (NFC). At the 65 percent 

speed point, the forward combustor probe at location 12 was lost due to excessive temperatures 

(>3100°F). After collecting repeat data at 65 percent, the probes were retracted and the engine 

was shut down. This completed the test matrix. No problems were ever encountered with the 

turbine or mixer temperature probes, and no problems were ever encountered with the Kulite 

probes or microphones. 

Because of the issues encountered with the combustor probe actuator lockups and the data 

system lockups, the collection of the test data was not in the order or the format originally 

planned. On the Pulse system, individual recordings were made for each of the test point, and 

binary universal file format (UFF) files were created from each of these separate recordings. On 

the Dewetron system, the recordings generally contained multiple test points, and so the data 

for a particular test point needed to be exported to a Matlab file in a post-processing step. Table 

5 summarizes the test data that was acquired on the Dewetron and Pulse systems. 

NASA/CR—2015-218879/VOL1 75
Table 5. Matrix of Test Data Acquired. 

The temperature probes that were operational during the Dewetron recordings are provided in 

Table 6. It is easy to tell by inspection of the data files which temperature probes are 

operational. Table 7 and Table 8 provide the serial numbers and location of the temperature 

probes used for the testing on March 26 and 28, respectively. 

Test Date Test Code NFC 

Percent 

Kulite Purge 

Flow Status 

Temperature 

Probes 

Dewetron 

Data 

Pulse 

Data 

03/26/2014 977CNAMNP Engine Off OFF Retracted Y Y 

03/26/2014 977CNAMWP Engine Off ON Retracted Y Y 

03/26/2014 977CN4801 48.6 ON Deployed Y Y 

03/26/2014 977CN6001 60.1 ON Deployed N Y 

03/26/2014 977CN6002 60.0 ON Deployed Y Y 

03/26/2014 977CN5401 54.0 ON Deployed Y Y 

03/28/2014 977CNNP Engine Off OFF Retracted Y Y 

03/28/2014 977CNWP Engine Off ON Retracted Y Y 

03/28/2014 977CN48RT 48.0 ON Retracted Y Y 

03/28/2014 977CN4801 48.0 ON Deployed Y Y 

03/28/2014 977CN5401 54.0 ON Deployed N Y 

03/28/2014 977CN5402 54.0 ON Deployed Y Y 

03/28/2014 977CNGI01 23.2 ON Deployed Y N 

03/28/2014 977CN4802 48.4 ON Deployed Y Y 

03/28/2014 977CN5403 53.9 ON Deployed Y Y 

03/28/2014 977CN6001 60.2 ON Deployed Y Y 

03/28/2014 977CNGI02 23.1 ON Deployed Y N 

03/28/2014 977CNGI03 23.0 ON Deployed Y N 

03/28/2014 977CN5404 54.1 ON Deployed Y Y 

03/28/2014 977CN6002 60.1 ON Deployed Y Y 

03/28/2014 977CN6501 64.9 ON Deployed Y Y 

03/28/2014 977CNGI04 23.1 ON Deployed Y N 

03/28/2014 977CN6502 65.1 ON Deployed Y Y 

03/28/2014 977CNGI06 23.1 ON Deployed Y N 

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Table 6. Operational Temperature Probes for Dewetron Data Files. 

Table 7. Location of Temperature Probes for the Testing on 3/26. 

Left Hand Side (ALF) Right Hand Side (ALF) 

Location Probe 

Number 

 Location Probe 

Number 

Combustor FWD 8 Combustor FWD 6 

Combustor AFT 3 Combustor AFT 4 

LP Turbine FWD 7 LP Turbine FWD 10 

LP Turbine AFT 1 LP Turbine AFT 9 

Mixer 11 Mixer 5 

Table 8. Location of Temperature Probes for the Testing on 3/28. 

Left Hand Side (ALF) Right Hand Side (ALF) 

Location Probe 

Number 

 Location Probe 

Number 

Combustor FWD 5 Combustor FWD 6 

Combustor AFT 2 Combustor AFT 4 (dead) 

LP Turbine FWD 7 LP Turbine FWD 10 

LP Turbine AFT 1 LP Turbine AFT 9 

Mixer 11 Mixer Dummy 

Date % NFC REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW REF SW LW

3/26/2014 48

3/26/2014 60 1a 0 0 0 0 1b 0

3/26/2014 54 0 0 0 0 0 0 0 0 0 0

3/28/2014 48 2 2 2 3 3 3 4 4 4

3/28/2014 54 2 2 2 3 3 3 4 4 4

3/28/2014 48 2 2 2 3 3 3 4 4 4

3/28/2014 54 2 2 2 3 3 3 4 4 4

3/28/2014 60 5 5 5 2 2 2 3 3 3 4 4 4

3/28/2014 23 2 2 2 3 3 3 4 4 4

3/28/2014 23 2 2 2 3 3 3 4 4 4

3/28/2014 54 2 2 2 3 3 3 4 4 4

3/28/2014 60 2 2 2 3 3 3 4 4 4

3/28/2014 65 2 2 2 5 5 3 3 3 4 4 4

3/28/2014 65 2 2 2 5 3 3 3 4 4 4

Probe Operating, Data Presumed to be Good

Probe Operating, Data is Questionable

Probe Not Operating

Probe Not In Use

0 Data unusable, probe was not properly inserted

1 (a) element died during insertion, temp < 2300F, (b) element dead before probe was inserted

2 This probe was stuck in an uninserted position and could not be removed/replaced; Removed from data collection

3 Spare probe was used in this location

4 Dummy probe inserted, removed from data collection; The mixer probe from this location was moved to the CF12 location

5 Probe elements destroyed by excessive temperatures, > 3100F

CB FWD 4.5 CB AFT 4.5 CB FWD 12 CB AFT 12 IT FWD 4.5 IT AFT 4.5 IT FWD 12 IT AFT 12 Mixer 4.5 Mixer 12

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6.7 Test Data Export 

On the Dewetron system, recordings were made throughout several test points due to issues 

encountered with software lockups. After the test, separate data files were exported from the 

Dewetron system to Matlab binary format using the DEWESoft X program. The software is 

freely available for posttest analysis and can be downloaded from www.dewetron.com. The 

software can be used in the “Analysis” mode without a license to manipulate existing data files. 

The data file of interest was opened up by navigating to the file using the folder browser and 

double clicking the file. Data was exported to MATLAB format by zooming in on the region of 

interest in the time history, selecting the “export” tab, selecting “absolute time” for the time units, 

selecting the channels of interest, clicking “File export” and selecting *.mat format with the 

“MATLAB 7.3 MAT file (> 2 GB)” option, and clicking “Export.” The test data from the Pulse 

system was exported to a binary UFF for ease of subsequent analysis. This format is readable 

into Matlab using a special-purpose script. 

7.0 TASK 5: DATA PROCESSING AND DISTRIBUTION 

7.1 Task 5.1 Process Unsteady Temperature and Pressure Data 

The purpose of this subtask was to perform initial data processing on the temperature and 

pressure data. The statement of work for this subtask consisted of the following elements: 

1. Process the unsteady temperature data (10 probes) for all test points

 Perform initial data review and quality assessment

 Define DGTMS software process for compensated spectra

 Create compensated time histories of the fluctuating temperature at each probe

location for each operating condition

2. Process the unsteady pressure data (13 probes) for all test points

 Perform initial data review and quality assessment

 Process power spectral density spectra

 Create phase-corrected pressure time histories that correspond to the dynamic

pressure at the tip of the pressure probe

3. Process engine external acoustic data (2 microphones) for all test points

 Perform initial data review and quality assessment

 Process power spectral density spectra

The results from these subtasks are discussed in the following subsections. 

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7.1.1 Unsteady Temperature Data 

7.1.1.1 Data Review and Quality Assessment 

There are several aspects of data quality that are of interest to the temperature data. These 

include an initial inspection of the data, an assessment of data quantization, an evaluation of the 

electronic noise floor, and statistical characterization of the time history data. The electronic 

noise floor of the data systems is of critical interest. As previously mentioned, the voltage 

signals from the data systems were recorded directly without amplification. 

An initial inspection of the data is provided in Figure 78 through Figure 81 for the test points 48, 

54 and 60 percent respectively. In these figures, dynamic portion of the small wire and large 

wire thermocouple time histories are examined for the four tested engine speeds. For these 

plots, the time history data was converted from millivolts to degrees Fahrenheit using the NIST 

coefficients for Type B thermocouples. The signal was then split into DC and AC components 

using a cut-off frequency of 1 Hz, using a process that is described in the next section. Figure 

78 shows the AC components of the small and large wire thermocouples for the five probes on 

the left hand side of the engine (aft looking forward) at the 48 percent speed point. Visual 

inspection of these plots reveals several features. First, it provides an indication of the 

magnitude of temperature fluctuations of the wire elements (note: this is not the fluctuating gas 

temperature). In the combustor, the small wire temperature typically fluctuates between -200°F 

and 200°F, and the large wire temperature typically fluctuates between -50°F and 50°F. In the 

ITD, the fluctuating wire temperatures are one order of magnitude smaller than those observed 

in the combustor. And in the mixer, they are about 50 percent smaller than those observed in 

the turbine. And the fluctuations of the small and large wires in the mixer region are roughly in 

unison. There appears to be a slight degree of coherence between the upstream and 

downstream probes in the combustor. There appears to be a very high degree of coherence 

between the upstream and downstream probes in the ITD. These characteristics are similar at 

the higher engine speeds. 

A first look at the spectral content on these signals is shown in Figure 82 and Figure 83, where 

the power spectral density of the small wire and large wire temperature signals is presented for 

the 48 percent and 60 percent speeds. The spectra show a smooth, wideband content, with 

amplitude decreasing with frequency. The overall shape is similar to that observed in the burner 

rig test and previous contractor reports. The spectral levels in the combustor are much higher 

than those in the ITD or mixer. The amplitudes of the inter-turbine and mixer probe elements 

are seen to flatten out at high frequency, which is a result of reaching the electronic noise floor, 

as discussed later. The changes in the spectral content due to changes in the engine speed is 

not dramatic.

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Figure 78. Fluctuating Temperatures of the Small and Large Wires at 48 Percent S

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Figure 79. Fluctuating Temperatures of the Small and Large Wires at 54 Percent S

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Figure 80. Fluctuating Temperatures of the Small and Large Wires at 60 Percent S

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Figure 81. Fluctuating Temperatures of the Small and Large Wires at 65 Percent S

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Figure 82. Representative Spectra of the Raw Temperature Signals 

at 48 Percent Speed. 

Figure 83. Representative Spectra of the Raw Temperature Signals 

at 60 Percent Speed. 

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Data quantization is another important consideration. Since the thermocouple voltage signals 

were recorded without amplification, the signals could not be arbitrarily gained to optimize the 

fixed input range of the analog-to-digital converters on the data systems. The nominal input 

range of the DEWETRON ORION 1624 series analog-to-digital cards is ±10 V. This is also the 

nominal input range of the Brüel & Kjær LAN-XI 3050-A-060 analog-to-digital modules. Both 

systems provide 24-bit digital conversion, so the dynamic range is large enough to accurately 

measure very small voltages. There is the added consideration of the data type and precision 

used in outputting the measured data (e.g. whether raw A/D counts are considered, or floating 

point values). In order to get a sense for the practical impact of quantization on the measured 

data that is subsequently analyzed with these data systems measured data were inspected. 

The small wire thermocouple voltages from the forward combustor probe at the 4.5 location 

were examined. 

Figure 84 shows a snippet of data from the Pulse system, taken directly from the binary UFF file 

(sample rate of 65536 Hz). The oscillation in the voltage caused by the temperature fluctuations 

is evident in the top part of the figure. The bottom part of the figure shows the difference in 

voltage between neighboring data points, and is useful for spotting quantization effects. There 

is a clear horizontal banding of these voltage increments, indicating quantization of the data. 

The increment between bands is 1.417 μV, which is an extremely small voltage change. The 

temperature increments corresponding to this voltage quantization, for the range of 

temperatures of interest in this test, are shown in Table 9, where the conversion from voltage to 

temperature was performed using the NIST curves for Type B thermocouples. The temperature 

increments range from 0.7°F at mean temperatures near 700°F (e.g. the mixer location) to 0.2°F 

at mean temperatures near 3000°F (e.g. at the combustor forward location). From the 

standpoint of indirect noise generation, being able to resolve the temperature fluctuations to this 

level of resolution (note that these are uncompensated) is considered adequate. 

Figure 84. Example of Quantization of Raw Temperature Signals Recorded by the 

Pulse System. 

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Table 9. Temperature Increments Corresponding to 1.42 μV Quantization. 

mV T (°F) mV T (°F) mV T (°F) mV T (°F) 

11.845 3000.0 5.7115 2000.0 1.4300 1000.0 0.6812 700.0 

11.8464 3000.2 5.7129 2000.3 1.4314 1000.5 0.6826 700.7 

Figure 85 shows a snippet of data from the Dewetron system, taken from the Matlab data file 

that was exported from the Dewetron system (sample rate of 10000 Hz). The time corresponds 

to the same snippet of data from the Pulse system in Figure 84. The oscillation in the voltage 

caused by the temperature fluctuations is evident in the top part of the figure. The bottom part 

of the figure shows the difference in voltage between subsequent data points, and is useful for 

spotting quantization effects. There is no clear banding of these voltage increments, suggesting 

that the quantization of the data is substantially smaller than is observed with the Pulse system. 

The banding can only be observed by zooming in much more closely. The increment between 

bands appears to be about 0.03 μV. The temperature increments corresponding to this voltage 

quantization are more than an order of magnitude smaller than the Pulse system, or less than 

0.1°F at mean temperatures of 700°F and above. Thus, quantization of the temperature data 

acquired by the Dewetron system (the primary data system) is likely of no consequence. There 

is one caveat, in that the compensation gain can be quite large at high frequencies (20 to 30 dB 

at 500 Hz) so evaluating the impact of quantization more carefully needs to consider the 

frequency content of the measured signals and the amplitudes of those frequency components 

at which indirect combustion noise becomes important. 

Figure 85. Example of Quantization of Raw Temperature Signals Recorded 

By the Dewetron System. 

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The background electronic noise floor is also an important consideration. The extremely small 

voltage fluctuations of interest are riding on top of a small DC voltage level corresponding to the 

mean temperature of the environment in which the thermocouples are immersed. If the 

background electronic noise of the recording system is too high, the mean voltage and/or the 

fluctuating component could be lost in the noise. Figure 86 shows the power spectral density of 

the electronic noise floor of the small wire (SW in the figure key) and large wire (LW in the figure 

key) thermocouples measured prior to engine start, compared to the levels measured at the 48 

percent speed test point. The measurements made prior to engine start are shown as dotted 

lines in the figure (‘AWB Purge On’ in the figure key) and the measurement made at 48 percent 

are shown with solid lines (‘0326 4801’ in the figure key). The background electrical noise of all 

of the thermocouples is essentially the same. The noise floor ranges from -20 dB at 20 Hz 

down to about -30 dB at 1000 Hz, with 0 dB corresponding to 1 μV. This data was processed 

with a Hanning window, 50 percent overlap, and a frequency resolution of 4 Hz. The noise floor 

corresponds to extremely low temperature fluctuations. For example, a 500 Hz sinusoidal 

signal with a root-mean-square amplitude of 0.1 μV would just barely protrude above the noise 

floor. This corresponds to measured temperature fluctuations below 0.1°F. 

Figure 86. Dewetron System Electronic Noise Floor for Temperature Probes. 

The electronic noise floor is clearly acceptable for raw measurement of temperature using an 

ideal temperature probe that requires no compensation. However, the physical size and limited 

frequency response of the probes used here have compensation factors on the order of 20 to 30 

dB, which dictates that the measurement system should be capable of resolving much smaller 

fluctuations. For example, if it is deemed necessary to resolve fluctuating temperatures on the 

order of 10°F at 1000 Hz, then the measurement system must have a noise floor that is one or 

two orders of magnitude smaller than those corresponding voltages to accommodate the 

compensation process. A fundamental question is, “what amplitude of fluctuating temperatures 

need to be resolved, and over what frequencies?” To answer this, one needs to know how the 

magnitude of these temperature fluctuations relates to the magnitude of the pressure 

NASA/CR—2015-218879/VOL1 87
fluctuations that are generated upon convection of these entropy waves through the turbine, and 

how these pressure fluctuations compare to the other noise sources of the engine. At the 

present time, the answer to this question is not known. It is not currently known even what 

order of magnitude of temperature fluctuations is important. In addition, the conversion from 

temperature to pressure might be highly dependent on the specifics of the engine. 

Figure 86 indicates that the voltage signals from the combustor probes i.e. CF45, CA45, CF12 

and CA12 in the figure, are well above the noise floor to at least 1000 Hz, so that no limitation is 

imposed on the compensation process. The ITD probes i.e. LF45, LA45, LF12 and LA12 in the 

figure, should only be compensated to about 500 Hz if the small wire is used for compensation, 

or 250 Hz if the large wire is used for compensation. Applying compensation beyond these 

frequencies will only result in the amplification of signal noise. The mixer probes i.e., MX45 and 

Mx12 in the figure should only be compensated to about 250 Hz if the small wire is used for 

compensation. Considering that the noise floor roughly corresponds to fluctuations on the order 

of 0.1°F, fluctuations on the order of 1°F or less cannot be resolved at the higher frequencies. 

From the standpoint of the indirect noise generation, this currently does not appear to impose 

any limitation. The fluctuations in the ITD and the mixer beyond these frequency limits are, from 

a practical perspective, probably too small to be of any interest. Thus the electronic noise floor 

of the Dewetron system appears to be acceptable for practical purposes. 

7.1.1.2 DGTMS Software Process for Compensated Spectra 

The DGTMS software was used to determine the compensation spectra and compute the 

compensated temperature spectra. No major modifications were made to the program, and the 

code worked well in the form received from NASA. The various subroutines of the code are well 

documented in previous contractor reports (References 1, 6, 7). The process for using the 

DGTMS software to compensate the data is similar to that used during the burner rig test. 

Because of the large number of probe/test point combinations, a semi-automated process was 

used. A Matlab script was written to perform the following steps: 

1. Create the DGTMS input file. This file contains the probe dimensions, mean flow

properties, and various program control options.

2. Create the DGTMS time history data file.

 Resample the data from its acquired sample rate to 8192 Hz.

 Split the time histories into AC and DC components using fourth order Butterworth

filters with a cut-off frequency of 4 Hz.

 Export the data to a fixed format text file.

3. Run the DGTMS program.

4. Read the compensation spectra and the compensated temperature spectra from the

DGTMS output files and export these to Excel.

5. Create compensated time histories using an interpolated compensation spectrum and

save these out as Matlab binary files for additional post-processing.

Pertinent details of these steps are now described in more detail, with reference to the data set 

acquired at 48 percent corrected fan speed collected on March 26. To run the DGTMS 

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software, the user must supply an input file and a data file. Both files are in ASCII text format. 

The input file contains information regarding the thermocouple probe geometry and the mean 

flow properties, along with parameters that specify various aspects of the data processing, 

program control, and output. The input file is described in Table 6 of Reference 7. 

The input file requires the thermocouple probe dimensions to use in the convective and 

conductive heat transfer model. The dimensions of the individual probe elements were not 

measured, so the nominal dimensions from the engineering drawing, shown in Figure 87, were 

used with one modification. The diameter of the support wires is nominally 0.020-inch. 

However, there is an overlap region where the smaller diameter sensing wires are welded on 

top of the support wires. As a result, the actual construction deviates from the ideal theoretical 

model that is used in the DGTMS code. To approximately account for this difference, an 

effective diameter was defined, given by the diameter of a cylindrical wire that has the same 

combined volume (mass) as the 0.020-inch support wire and the overlapped portion of the 0.10-

inch or 0.003-inch sensing wire. Table 10 provides the nominal values of the probe elements 

along with the computed “effective diameter” of the support wires. These values were used in 

the DGTMS input file. 

Figure 87. Nominal Dimensions of the Thermocouple Probe Elements. 

 

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Table 10. Thermocouple Probe Dimensions Used in the DGTMS Input File. 

Feature Thermocouple Value Adjusted

Length of support wire Large Diameter 0.105-inch (0.267 cm) - 

Half of total length of the smaller wire Large Diameter 0.073-inch (0.184 cm) - 

Diameter of support wire Large Diameter 0.020-inch (0.0508 cm) 0.0557 cm 

Diameter of smaller wire Large Diameter 0.010-inch (0.0254 cm) - 

Length of support wire Small Diameter 0.075-inch (0.191 cm) - 

Half of total length of the smaller wire Small Diameter 0.035-inch (0.0889 cm) - 

Diameter of support wire Small Diameter 0.020-inch (0.0508 cm) 0.0512 cm 

Diameter of smaller wire Small Diameter 0.003-inch (0.00762 cm) - 

Next, the DGTMS time history data file was created. The first step is to resample the data from 

the acquired sample rate (namely 10 kHz or 20 kHz for the Dewetron data acquired on March 

26 or 28, respectively) to a more modest sample rate. A sample rate of 8192 Hz was chosen for 

this purpose, which significantly over samples the frequency range of interest. The data was 

resampled in Matlab using the resample function. The initial and final 1/8th second of the time 

history was clipped off to remove portions of the signal that might have been distorted by 

starting and ending transient artifacts. The time histories for the large and small wire 

thermocouples were then split into AC and DC components. Low-pass and high-pass fourth 

order Butterwoth filters with a cutoff frequency of 4 Hz were defined using the Matlab butter

function and applied to the data using the filtfilt function so that the filtered time history has zero 

phase distortion. The cut-off frequency is somewhat arbitrary, and a value between 1 Hz and 

20 Hz would probably be appropriate. The 4 Hz cut-off was chosen to coincide with the 4 Hz 

analysis bandwidth used later in the processing. 

Figure 88 through Figure 92 plot the time history data for the 10 temperature probes after 

resampling and splitting into DC and AC components. Inspection of the DC signals for the small 

wire and large wire thermocouples shows that DC amplitudes are very similar (ideally, they 

would be identical) and that the low frequency undulations are the same for both sensing 

elements. This provides confidence that both sensing elements are operating correctly. The 

reason for the small difference in DC levels in some cases is unknown. The small and large 

thermocouple junctions are not at the exact same point in space (the two junctions are 

displaced 0.030-inch in the radial direction and 0.056-inch in the axial direction), so some of the 

difference might be due to actual differences in flow temperatures, but further investigation is 

required. Inspection of the AC signals shows that the small wire thermocouple has much 

greater amplitude than the large wire thermocouple, as expected, and that the fluctuations are in 

unison. Inspection of the full time history plots for the various test conditions is helpful, as there 

are occasionally anomalies in the data that are readily apparent. For example, the small wire 

AC time history for the CA12 probe in Figure 89 indicates an unexplained anomalous behavior 

near 30 seconds. The data in this region is suspect. 

The data file for the DGTMS code is a formatted file containing the small wire thermocouple AC 

data, the large wire thermocouple AC data, and the DC data from either the small or large wire 

(which is usually very similar). For this report, the DC data from the large wire was used. The 

data is arranged in blocks, as described on page 16 of Reference 7, in units of millivolts. Once 

the input file and the data file have been created, the DGTMS software can be executed. 

In order for the DGTMS code to be able to successfully compute the in-situ heat transfer 

coefficient for a particular probe at a given operating point, there must be sufficient coherence 

(greater than 0.8) between the small and large thermocouple wires at low frequencies (say 

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below 50 Hz). Thus it is of interest to examine the coherence between the small and large 

thermocouple wires. An example is presented in Figure 93, which plots the coherence for the 

ten probes at the March 26 48 percent speed point. In the combustor, there is substantial 

coherence over a wide range of frequencies. The aforementioned anomaly in the small wire 

time history data for the CA12 probe results in a loss of coherence relative to the other 

combustor probes. This probe was replaced for the second day of testing, and the coherence 

improved in the repeat of the test point. The coherence in the ITD and mixer are sufficiently 

high at low frequencies, but drops off very rapidly with frequency. This is a result of the 

temperature fluctuations at frequency becoming extremely small and the corresponding voltage 

fluctuations reaching the electronic noise floor of the data acquisition system. The Matlab 

function mscohere was used to compute coherence in these analyses. 

The DGTMS software was run for all functioning temperature probes at all operating conditions. 

For each probe, at each condition, an input file and a time history data file was created. The 

DGTMS software was executed from the Matlab script. Program control options in the input file 

were set to compute the compensation spectrum and the compensated temperature spectrum 

for the small wire thermocouple (IFLAGS(3)=1). For comparison purposes, a second input file 

was created with the program control options set to compute the compensation spectrum and 

the compensated temperature spectrum for the large wire thermocouple (IFLAGS(3)=2). The 

computed small wire compensation spectra for the combustor, ITD, and mixer probes at the 

March 26 48 percent speed operating point are shown in Figure 94, Figure 95, and Figure 96, 

respectively. The amplitude and phase of the compensation spectra are smooth, well behaved, 

and have magnitudes and shapes that are similar to previous results. 

Within the DGTMS code, the compensation spectrum for the small (or large) wire thermocouple 

is applied to the measured spectrum of the small (or large) wire thermocouple to produce the 

compensated temperature spectrum, which is the estimated spectrum of the fluctuating 

temperature of the gas stream. Figure 97 shows a representative result for the CA45 probe at 

48 percent speed. This location is of particular interest, since the temperature fluctuations at 

the aft end of the combustor convect into the high pressure (HP) turbine and are hypothesized 

to create pressure fluctuations via the indirect noise generating mechanism. The figure shows 

the results for both the small wire and large wire compensation. In this case, the results are 

similar regardless of which wire is chosen for the compensation process, lending confidence 

regarding the consistency of the compensation process. The compensated temperature spectra 

are smooth and wideband. On a linear frequency scale, there are no significant peaks in the 

spectrum. The PSD amplitude (in dB) decreases with frequency in a nearly linear fashion. 

As one point of comparison, in Figure 98, the compensation spectra for the combustor probes at 

48 percent speed are compared to the compensation spectra that were computed during the 

P&W F100 testing (Reference 4). Caution needs to be exercised here, because the probe 

design, probe placement, engine characteristics, and operating condition are different between 

the two test programs. However, there are enough similarities to make the comparison 

worthwhile. The combustor compensation spectra in the TECH977 test are seen to be quite 

similar to the turbine nozzle compensation spectrum in the F100 test, lending confidence that 

the DGTMS measurement and software analysis processes are producing results that are 

consistent with previous testing. 

The compensated spectra for the four combustor probes at 48 percent speed are shown in 

Figure 99. There is very little difference in the compensated spectra for the two aft probes, 

indicating that the temperature fluctuations on either side of the engine have similar magnitude 

near the exit of the combustor. This is an encouraging result, since the aft combustor probe at 

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location 4.5 became inoperative early in the test series. There is also not a dramatic difference 

in the compensated amplitude between the forward probes and aft probes. For the CF45 probe, 

there is a difference of about 5 dB between the small wire and large wire compensated spectra. 

The reason for the discrepancy is not currently known. 

The compensated spectra for the four ITD probes at 48 percent speed are shown in Figure 100. 

Here, because of the electronic noise floor, the compensated spectra are not reliable above a 

certain frequency. For the large wire thermocouple, the compensated spectra are only reliable 

below about 200 Hz. For the small wire thermocouple, the compensated spectra are only 

reliable below about 500 Hz. There is no dramatic difference in compensated amplitude 

between the upstream and downstream probes at the location 4.5, or between the upstream 

and downstream probes at location 12, indicating that the temperature fluctuations are not 

changing with axial distance. There is slight difference between the probes on the two sides of 

the engine, with the probes on location 4.5 reading about 3 dB higher than those on the 

opposite side. At 200 Hz, where both the large and small compensated spectra should be valid, 

there is a 5 dB difference between the compensated spectra. Presumably, the small wire 

compensated spectra would be more reliable due to the higher signal-to-noise ratio. Although 

the small wire compensated spectra can only be computed up to 500 Hz, this does not appear 

to impose any practical limitation, because the temperature fluctuations, even at low 

frequencies, are so small that convection through the LP turbine nozzle is unlikely to create any 

significant indirect noise. However, this presumption requires validation, which is outside the 

scope of this work effort. 

The compensated spectra for the two mixer probes at 48 percent speed are shown in Figure 

101. Because of the electronic noise floor, the compensated spectra are not reliable above a 

certain frequency. For the large wire thermocouple, the compensated spectra are only reliable 

below about 50 Hz. For the small wire thermocouple, the compensated spectra are only reliable 

below about 250 Hz. There is no dramatic difference in compensated amplitude between the 

probes on the two sides of the engine. Although the small wire compensated spectra can only 

be computed up to 250 Hz, this does not appear to impose any practical limitation, because the 

temperature fluctuations are so small that convection through the remainder of the nozzle is 

unlikely to create any significant indirect noise. 

NASA/CR—2015-218879/VOL1 92
Figure 88. DC and AC Time Histories for the CF45 and CA45 Probes at the March 26 48 Perce

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Figure 89. DC and AC Time Histories for the CF12 and CA12 Probes at the March 26 48 Perce

Data Anomaly 

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Figure 90. DC and AC Time Histories for the LF45 and LA45 Probes at the March 26 48 Perce

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Figure 91. DC and AC Time Histories for the LF12 and LA12 Probes at the March 26 48 Perce

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Figure 92. DC and AC Time Histories for the MX45 and MX12 Probes at the March 26 48 Perc

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Figure 93. An Example of the Coherence Between the Small Wire and Large Wire Thermo

Raw Time Histories. 

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Figure 94. Small Wire Compensation Spectra for the Combustor 

Probes at 48 Percent. 

Figure 95. Small Wire Compensation Spectra for the Turbine Duct 

Probes at 48 Percent. 

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Figure 96. Small Wire Compensation Spectra for the Mixer Probes at 48 Percent. 

Figure 97. Small and Large Wire Compensated Spectra for the 

CA45 Probe at 48 Percent. 

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Figure 98. Comparison of TECH977 and F100 Compensation Spectra.

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Figure 99. Small and Large Wire Compensated Spectra for the Combustor Probes at 4

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Figure 100. Small and Large Wire Compensated Spectra for the Inter-Turbine Duct Probes 

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Figure 101. Small and Large Wire Compensated Spectra for the Mixer Probes at 48 P

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7.1.1.3 Compensated Time Histories 

Compensated time histories – providing the estimate of the fluctuating temperature of the gas 

stream (up to the frequency range limits) – can be generated by applying the compensation 

spectra to the measured time history data in the frequency domain via complex division. The 

DGTMS code has some built-in functionality to create compensated time histories. Because of 

computer hardware limitations at the time the code was written, the compensated time histories 

are only generated one FFT block at a time. A full compensated time history could be 

generated by connecting the individual blocks, but it is time consuming and prone to 

discontinuities at the block boundaries. 

An alternative approach is to compute a compensated time history in one shot by interpolating 

the compensation spectrum. Such a scheme works well because the compensation spectra are 

smooth and well behaved functions. The steps of the process are as follows: 

1. Compute the FFT of the entire SW time history (one single block). 

2. Interpolate the compensation spectrum to the same number of points in the FFT. 

3. Apply the compensation spectrum via complex division in the frequency domain. 

4. Take the inverse FFT to get back to the time domain. 

Unlike the block-by-block process used in the DGTMS software, this approach does not apply a 

Hanning window to the data prior to taking the Fourier transform. Because the data sequence is 

very long, typically 0.5 million data points, the errors introduced by the discontinuity between the 

beginning and ending sample is negligible. The FFT assumes the sequence to be periodic over 

that time interval, and the Fourier components of the step discontinuity are smeared into the 

spectrum. However, with a long data sequence, the amplitude of these contributions is small 

compared to the frequency content in the (random) signal. 

An important consideration when generating the time histories is that the compensation should 

not be applied to portions of the raw time history that are dominated by electrical noise. In the 

DGTMS code, the ITHRSH parameter is a user-selected relative threshold level that specifies 

the compensation limit. If there are frequency components with amplitudes below the threshold 

level, then the compensation is not applied and the amplitude of these frequency components is 

zeroed out. This parameter can be determined by comparing the uncompensated spectrum 

with the noise floor. An alternative approach is to only compensate the time history up to a 

certain frequency, so that the result becomes a band-limited compensated time history. 

Because it is pretty clear at what frequency the electronic noise floor is reached, this alternative 

approach was used. 

The compensation process was applied in Matlab. The interpolation of the compensation 

spectrum was performed using the Matlab interp1 function, and the forward and inverse discrete 

Fourier transforms were performed using the Matlab fft and ifft functions. As an example of the 

generation of the compensated time history, consider the time history data from the aft 

combustor probe at location 12 (CA12) for the 48 percent speed point, using the March 28 4802 

test point data. The raw time histories of the probe are shown in Figure 102 after splitting into 

DC and AC components. The time history data appear to be of good quality (compare this to 

Figure 89, which had a data anomaly). Figure 103 compares the entire compensated time 

histories generated in two ways, (1) by applying the interpolated small wire compensation 

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spectrum to the small wire time history data, and (2) by applying the interpolated large wire 

compensation spectrum to the large wire time history data. It is apparent that the temperature 

fluctuations of the compensated time histories, which represent the fluctuating gas temperature, 

are substantially larger than the measured fluctuating wire temperatures. Fluctuations on the 

order of ±750°F are common. Furthermore, the small wire and large wire compensated time 

histories are very similar. Figure 104 shows this more clearly by plotting a close-up of the time 

histories over a short time span of 0.1 second.

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Figure 102. Uncompensated DC (Top) and AC (Bottom) Time History Data fo

Combustor Probe CA12 at 48 Percent. 

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Figure 103. Small Wire and Large Wire Compensated Time Histories for Combustor Probe CA

Figure 104. Close-Up of Small and Large Wire Compensated Time History Dat

for Combustor Probe CA12 at 48 Percent.

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7.1.2 Unsteady Pressure Data Processing 

7.1.2.1 Data Review and Quality Assessment 

The internal dynamic pressure data was acquired using Kulite model XTE190-10D transducers. 

These are 10 psi differential piezoresistive transducers with a nominal sensitivity of 10.0 mV/psi. 

They have been extensively used at Honeywell for measuring internal engine dynamic 

pressures. The differential style provides excellent signal-to-noise characteristics, but requires 

that the sensors be supplied with a mean reference pressure that is close to the pressure in the 

environment being measured. As previously mentioned, the transducers were installed into 

tees and used with semi-infinite coils due to the high temperatures in the measurement 

environment. A very small nitrogen purge flow was also blown through the coils to keep the 

transducers cool, even though they are located outside of the engine compartments in this 

particular test. Signal conditioning was provided by a rack of Sonoran Microsystems amplifier 

modules. 

Two important data quality assessments are signal-to-noise ratio and the impact of purge flow. 

The plots within the figure are labeled using a naming convention where P is pressure, E is 

EVNERT, C is combustor, A is aft, F is forward, L is inter-turbine duct, MX is mixer and the 

numbers refer to the clocking pattern defined through Figure 77. Figure 105 shows a 

comparison of the electronic noise floor with the engine off, the noise floor when the engine is 

off and the purge flow is on, and the noise levels measured at the 48 percent speed point, for 

the 13 Kulites and 2 external microphones. The electronic noise floor is at least 20 dB below 

the measured noise levels when the engine is running. When the purge flow is turned on, there 

is an increase in ambient noise, particularly for the Kulites in the combustor (which may have 

had a higher velocity purge flow), but the ambient noise is still at least 20 dB below the 

measured noise levels when the engine is running. Thus, there is no concern regarding either 

electronic noise or noise associated with the purge flow. It should be mentioned that the purge 

flow does change somewhat over the range of tested engine speeds, but the delta-pressure 

between the supply manifold and the reference manifold tends to decrease with speed, so that 

the noise associated with the purge flow system should also decrease slightly with increasing 

engine speed. 

NASA/CR—2015-218879/VOL1 109
Figure 105. A Comparison of the Noise Floor of the Dynamic Pressure Transduc

With and Without Purge Flow. 

NASA/CR—2015-218879/VOL1 110
7.1.2.2 Power Spectral Density Spectra 

Power spectral density spectra were created from the dynamic pressure time histories for all of 

the 13 Kulite sensors and the two external microphones, for all engine operating conditions. 

PSD spectra for the 48 percent, 54 percent, 60 percent, and 65 percent operating speeds are 

shown in Figure 106, Figure 107, Figure 108, and Figure 109, respectively. The spectra for 

PCE03, PCE07, PCE11, and PCE15 are tightly grouped and nearly overlay one another. This 

is not surprising, as these sensors, which comprise the four-sensor sub-array from the 2007 

NASA EVNERT test, are all located at the same axial position on the combustor. In the 

EVNERT test, the dynamic pressure spectra in the combustor array were very similar. 

The spectrum for PCF45 is very similar to the four sub-array sensors. This is also not 

surprising, as this sensor is located nearly at the same axial location (recall the PCF45 

temperature probe – and it accompanying pressure probe – is the most forward temperature 

probe in the combustor). The PCA45 and PCA12 pressure spectra have similar spectral 

shapes, and the amplitudes are fairly close, but they don’t overlay each other as well as one 

might expect for sensors that are at the same axial station. The PCA12 pressure spectrum 

often has the lowest amplitude of all the combustor sensors at high frequency (1000 Hz). The 

PCF12 pressure spectrum is most similar to the PCA45 and PCA12 spectra. 

The PLA12 and PLA45 pressure probes have fairly similar spectra, and have the highest 

amplitudes at high frequency. Interestingly, the PLF12 pressure probe has a similar spectral 

shape but is a good 5 dB lower in amplitude than the PLA12 and PLA45 probes. The reason for 

this difference in amplitude is not known. There is no probe in a corresponding forward location 

at the opposite side, since the PLF45 pressure line was used as a reference pressure tap for 

the turbine purge flow system (a separate reference pressure tap had not been incorporated on 

the turbine case). Furthermore, the PLF12 spectrum often has a lower amplitude at 400 Hz 

than any of the combustor probes. 

The spectra of the two mixer probes, PMX45 and PMX12, are very similar to one another but 

have substantially lower amplitudes than the combustor and turbine sensors. Comparing the 

spectra across different engine speeds, these conclusions generally persist, with some 

exceptions. The most notable change with higher engine speed is a gradual increase in 

spectral amplitudes for all of the sensors. 

Comparing the combustor and turbine spectra, it is difficult to make an argument for the 

presence of indirect noise. For example, there is not a frequency range where the turbine 

pressure spectra all show a large increase over the combustor pressure spectra, as one might 

expect if substantial indirect noise was generated downstream of the high pressure turbine. 

However, this viewpoint might be overly simplistic. A portion of the indirect combustion noise 

could propagate forward into the combustor, confusing the picture. A solid argument for the 

presence or absence of indirect combustion noise will need to rely on more sophisticated 

analyses. 

NASA/CR—2015-218879/VOL1 111
Figure 106. Dynamic Pressure Spectra at 48 Percent Speed 

From Data Collected March 26. 

NASA/CR—2015-218879/VOL1 112
Figure 107. Dynamic Pressure Spectra at 54 Percent Speed 

From Data Collected March 28. 

NASA/CR—2015-218879/VOL1 113
Figure 108. Dynamic Pressure Spectra at 60 Percent Speed 

From Data Collected March 26. 

NASA/CR—2015-218879/VOL1 114
Figure 109. Dynamic Pressure Spectra at 65 Percent Speed 

From Data Collected March 28. 

7.1.2.3 Phase-Corrected Pressure Time Histories 

Phase-corrected pressure time histories were generated which correspond to the dynamic 

pressure that would have been measured at the tip of the sensing tube. That is, when using 

semi-infinite probes, there is a propagation delay due to the time that it takes the pressure 

signal to propagate up the length of tubing, from the engine duct wall to the Kulite location in the 

tee. To compute the time delay, the length of the sensing line tube and the sound speed in the 

tube must be known. The length of the sensing line was a constant 35.875-inch for all of the 

Kulite probes (the semi-infinite coils where the sound is attenuated were 40 feet long). Nitrogen 

purge flow was used in the combustor and turbine pressure probes to prevent hot gases from 

being compressed into the tubes during engine operation. As a result, the temperature and thus 

sound speed in those sensing lines is expected to be quite uniform. Purge flow was not 

required for the mixer probes, since the mean pressures and temperatures were much lower. 

Half of the Kulite tees were instrumented with thermocouples to measure the temperature of the 

gas in sensing tubes. The thermocouples were positioned just slightly beyond the Kulite tap 

location, on the semi-infinite coil side of the tee. 

Figure 110 shows the Kulite tee temperatures measured during the testing on March 26. Most 

of the thermocouples registered nearly the same temperature. Furthermore, the temperatures 

NASA/CR—2015-218879/VOL1 115
were not very sensitive to changes in engine operating condition. On this particular day, the 

temperatures were typically 76°F. The temperatures were only measured at a single point, so 

the temperature distribution throughout the three foot length of the sensing tubes as they 

snaked around the engine case and into the engine core is not known. For this analysis, it was 

assumed that the temperature in the tube was uniform over its entire length. At a temperature 

of 76°F, the sound speed in the sensing line is 1134 ft/s and there will be a time delay of 2.64 

ms for the acoustic wave to propagate from the end of the sensing line (tapped flush to the outer 

wall of the engine core duct) to the Kulite location. Figure 111 and Figure 112 show the Kulite 

tee temperatures measured during the testing on March 28. Aside from the CE15 probe, all of 

the tee temperatures were nearly the same. The temperature was around 62°F for the first 

engine run and 72°F for the second engine run, corresponding to time delays of approximately 

2.67 ms and 2.65 ms, respectively. These propagation time delays are summarized in Table 

11. 

Table 11. Propagation Time Delays Assumed for Kulite Pressure Signals. 

Test Date Assumed Time Delay, ms 

03/26/2014 2.64

03/28/2014 First Engine Run 2.67 

03/28/2014 Second Engine Run 2.65 

The purge flow temperature measured in the CE15 tee (labeled TKCE15 in the figures) appears 

to be an outlier. It is not known why the thermocouple in the CE15 tee was reading higher than 

the others, but this tee was close to the engine, near the cutout in the bypass duct, where warm 

secondary flow (used for cooling the engine core) was escaping from the seals around the 

cutout and impinging on the tee. It is possible that the tee housing was simply heated by this 

impinging air, and that the thermocouple was in turn heated via conduction, leading to an 

erroneous reading. On the other hand, it is possible that the purge flow in the tube was also 

heated and that the reading is accurate. In order to decide whether special treatment is needed 

for this sensing line, the following analysis procedure was considered. 

Since the CE15 pressure probe is part of the sub-array in the forward part of the combustor, it is 

possible to inspect the cross-spectral phase between the probes in this sub-array and infer 

whether there are substantial differences in the temperatures (sound speeds) of the purge flow 

inside the sensing tubes. The sensing tubes in this array are equally spaced and have equal 

length. As will be shown in the next section, certain acoustic modes are dominant in the 

combustor over certain frequency ranges, and as a result there is a well defined phase 

relationship between the signals measured at different locations. For example, the plane wave 

is dominant below 500 Hz. So if the purge flow temperatures (sound speeds) inside the sensing 

lines are all the same, then the signals measured by the Kulites will all be in phase and the 

cross-spectral phase will be zero for all pairs of Kulites. On the other hand, if the temperatures 

inside these tubes are substantially different, then the cross spectral phase will be non-zero, 

with a difference that increases linearly with frequency. 

Consider the data from March 26 at the 60 percent speed point, where the CE15 Kulite tee 

temperature read 120°F, corresponding to a time delay of 2.53 ms, and the CE11 Kulite tee 

temperature read 78°F, corresponding to a time delay of 2.63 ms. If the probes are sensing the 

same plane wave pressure field (dominant up to 500 Hz), and if the temperature readings are 

accurate, then there should be a time delay difference of 0.1 ms between the CE11 and CE15 

signals, corresponding to a phase difference of 18 degrees at 500 Hz (the phase difference 

being a linear function of frequency). 

NASA/CR—2015-218879/VOL1 116
Figure 110. Purge Flow Temperatures Measured During the March 26 Test. 

NASA/CR—2015-218879/VOL1 117
Figure 111. Purge Flow Temperatures Measured During the March 28 Test. 

Figure 112. Purge Flow Temperatures Measured During 

the March 28 Test (Continued). 

NASA/CR—2015-218879/VOL1 118
Figure 113 shows the phase of the cross spectrum for different combinations of probes in the 

combustor sub-array. The phase has been plotted only for those frequencies where the 

coherence between the same two channels, shown in Figure 114, is greater than 0.05. There 

does appear to be a very slight phase shift associated with the CE15 signal, as indicated by the 

slight negative slope of the phase involving the CE15 signals. However, the slope of 

approximately -0.02°/Hz is about half of that associated with a 0.1 ms time delay (-0.04°/Hz). 

Therefore, it is concluded that there may have been some heating of the CE15 purge flow, but 

not as much as indicated by the measured temperature. The CE15 signal appears to have, at 

worst, a 0.05 ms shorter time delay than the other probes in the sub-array. For the analyses 

considered here, and for the frequency range of interest (< 1000 Hz), there appears to be little 

error in assuming a constant time delay for all of the probes in the combustor, and further, all of 

the pressure probes in the engine. Therefore, a common time delay correction was used for all 

of the Kulite probes, differing slightly by test time, as summarized in Table 11. 

As a point of comparison, Figure 115 shows the cross-spectral phase for the same 

combinations of sensors from the EVNERT program, References 8 and 9, Run Without Fan 

test, for coherence values above 0.05. It appears that in the EVNERT test there may have been 

slightly larger tube-to-tube variations in purge flow temperature, even though all of the sensing 

tube lengths in that combustor array were only 12 inches long. 

There are additional effects related to the propagation of acoustic waves in the tube that have 

not been accounted for. There is a frequency dependent attenuation that occurs in the sensing 

line due to viscous dissipation. The reduction in amplitude has not been quantified, but is 

considered to be small over the frequency range of interest in this program (<1000 Hz). Also, in 

order to route the sensing tubes outside the engine, there are numerous bends in the tubes. 

These bends were made to be as gradual as possible so as not to create an impedance 

discontinuity. There impact of these bends is considered to be small over the frequency range 

of interest, but a precise quantification of the effects has not been made. 

Phase-corrected pressure time histories were generated to be consistent with the compensated 

temperature time histories. These pressure time histories were created as follows. First, the 

time delays were determined using the measured tee temperatures and known length of tubing, 

as summarized in Table 11. Next, a sequence of “delayed” times was constructed at the same 

sample rate as the original pressure data. The pressure data was then interpolated to these 

new delay times using the Matlab interp1 function. The first sample of interpolated pressure 

data corresponds to the same physical time as the first sample of temperature data, as if the 

pressure measurement had been made directly at the wall. The interpolated pressure signals 

were then resampled at a frequency of 8192 Hz using the Matlab resample function, and then 

high pass filtered using the Matlab filtfilt function, with the first and last 1/8th second clipped off 

to be consistent with the processing of the dynamic temperature data. The net result of the 

correction is that the pressure time histories which are resampled to 8192 Hz are shifted by 

about 22 samples in order to align the pressure data with the temperature data. For some 

analyses, this time delay does not have a substantial impact. For example, the time delay 

results in a loss of coherence between pressure and temperature, but with a 4 Hz analysis 

bandwidth, the reduction in coherence is quite small. 

NASA/CR—2015-218879/VOL1 119
Figure 113. Cross-Spectral Phase Between Pressure Probes in the Combustor 

Sub-Array, at 60 Percent Speed, for Coherence Values above 0.05. 

Figure 114. Coherence Between Pressure Probes in the Combustor 

Sub-Array at 60 Percent Speed. 

NASA/CR—2015-218879/VOL1 120
Figure 115. Cross-Spectral Phase for Combinations of Pressure Signals in the 

Combustor Sub-Array, at 60 Percent Speed, from the 

EVNERT Run Without Fan Test. 

7.2 Task 5.2: Acoustic Modal Analysis 

7.2.1 Combustor Circumferential Mode Decomposition 

A circumferential mode decomposition of the upstream 4-microphone combustor array was 

performed. As a background, during the Run Without Fan work element of the 2007 NASA 

EVNERT program, the TECH977 was instrumented with an array of 16 equally-spaced dynamic 

pressure sensors, in the form a semi-infinite coils using a 12-inch sensing line. The ends of the 

pressure sensing lines were flush with the outer wall of the combustor, and were oriented 

azimuthally at the same angular positions as the 16 fuel nozzles. Details of the testing and 

analysis results can be found in References 8 and 9. For the DGMTS test, measurements were 

made using a subset of this array to provide a comparative reference to the previous dataset. 

An array of four equally spaced sensors was used. Figure 116 shows a comparison between 

the locations of the probes that were used in the 2007 EVNERT test and the 2014 DGTMS test. 

The biggest difference in the instrumentation setup between the two tests was that the sensing 

lines used in the DGTMS test were three times longer than the EVNERT test. Figure 117 

compares the noise spectra at the four common locations for the two tests, at the 48 percent 

speed condition. Figure 118 compares the noise spectra at the four common locations at 60 

percent speed. Overall, there is a good match in the shapes and amplitudes of the noise 

spectra measured in the two tests. 

NASA/CR—2015-218879/VOL1 121
Figure 116. Kulite Probe Locations for the Combustor 

Circumferential Mode Array. 

Figure 117. Comparison of Combustor Noise Spectra With Previous 

Measurements at 48 Percent Speed. 

NASA/CR—2015-218879/VOL1 122
Figure 118. Comparison of Combustor Noise Spectra With Previous 

Measurements at 60 Percent Speed. 

The modal structure of the sound field in the combustor was explored in References 8 and 9. 

When using a finite number of transducers, there is a limit to the number of modes that can be 

uniquely identified. Mode aliasing will occur if the number of sensors is too small. Figure 119 

shows what happens to the mode decomposition when the 4-sensor array is used instead of the 

full 16-sensor array, using the NASA EVNERT measurements at 48 percent speed. With an 

array of 16 sensors, the modes can be fully distinguished to a frequency well beyond 3 kHz. 

However, when only four sensors are used, mode aliasing occurs. Above 1000 Hz, the 16-

sensor analysis shows the m=-2 and m=2 modes are present with roughly equal amplitude. 

With a 4-sensor array, these two modes are aliased to the same mode analysis bin, and cannot 

be distinguished. Above 2000 Hz, the 16-sensor analysis shows that the m=-3 and m=3 modes 

are present with roughly equal amplitude. In the 4-sensor analysis, the m=-3 mode is aliased to 

the m=+1 bin, and the m=+3 mode is aliased to the m=-1 bin. Although the two modes are 

distinguished from one another, the mode orders have been aliased. Figure 120 and Figure 

121 shows the same comparison at 54 percent speed and 60 percent speed (the 65 percent 

speed point was not tested in the EVNERT Run Without Fan test). The modal plots and 

conclusions are nearly the same. One can conclude that below 1000 Hz, the m=0 mode, m=-1, 

and m=1 modes can be resolved without aliasing. 

Next, a comparison is made between the EVNERT test data and the DGTMS test data. A sideby-side comparison of the 4-sensor modal decomposition results at 48 percent speed is 

presented in Figure 122. A similar comparison at 54 percent and 60 percent speed are 

presented in Figure 123 and Figure 124, respectively. Visually, the results for the two tests are 

extremely similar, both in terms of the modal content and the amplitudes of the modes. This is 

to be expected, since the same engine core was used in both tests (although the rest of the 

engine setups were quite different). 

NASA/CR—2015-218879/VOL1 123
A closer inspection of the comparison at these three speeds is presented in Figure 125 through 

Figure 127, where the modal decomposition results have been rotated at an angle to highlight 

subtle differences. Comparing the two, the DGTMS modal amplitudes appear to be nearly the 

same as the EVNERT modal amplitudes, but with an additional amplitude modulation. The 

frequency of this amplitude modulation appears to be near 200 Hz. The reason for this 

difference is not known, but the combustor in the DGTMS test has additional holes to 

accommodate the extra temperature and pressure instrumentation. It can be determined that 

the undulation is not due to the presence of the temperature probes in the combustor. Figure 

128 shows a comparison at 48 percent speed with the temperature probes retracted and 

deployed, and no difference is seen in the modal decomposition. 

The individual pressure spectra for the four probes for 48 percent, 54 percent, and 60 percent 

speed are compared in Figure 129 through Figure 131. There is very good agreement in the 

individual spectral amplitudes between the two tests, with the exception that the DGTMS data 

shows additional undulations. Also noteworthy is that the Kulite tees for the PCE11 and PCE15 

sensors were instrumented with thermocouples (which protrude slightly into the flowpath), 

whereas the PCE03 and PCE07 sensors did not. There is no noticeable qualitative difference in 

the spectra between tubes that have thermocouples and the tubes that don’t. This suggests 

that the thermocouples installed into the Kulite tees have a negligible influence on the noise 

propagation in the tube. 

Figure 119. Mode Aliasing With a Reduced Number of Sensors at 48 Percent. 

NASA/CR—2015-218879/VOL1 124
Figure 120. Mode Aliasing With a Reduced Number of Sensors at 54 Percent. 

Figure 121. Mode Aliasing With a Reduced Number of Sensors at 60 Percent. 

NASA/CR—2015-218879/VOL1 125
Figure 122. Comparison of Modal Decomposition Results at 48 Percent Speed. 

Figure 123. Comparison of Modal Decomposition Results at 54 Percent Speed. 

NASA/CR—2015-218879/VOL1 126
Figure 124. Comparison of Modal Decomposition Results at 60 Percent Speed. 

Figure 125. Comparison of Modal Decomposition Results at 48 Percent Speed. 

NASA/CR—2015-218879/VOL1 127
Figure 126. Comparison of Modal Decomposition Results at 54 Percent Speed. 

Figure 127. Comparison of Modal Decomposition Results at 60 Percent Speed. 

NASA/CR—2015-218879/VOL1 128
Figure 128. Comparison of DGTMS Modal Decomposition at 48 Percent Speed, 

With Temperature Probes Retracted and Inserted Into the Combustor. 

Figure 129. Comparison of DGTMS and EVNERT Spectra at 48 Percent Speed. 

NASA/CR—2015-218879/VOL1 129
Figure 130. Comparison of DGTMS and EVNERT Spectra at 54 Percent Speed. 

Figure 131. Comparison of DGTMS and EVNERT Spectra at 60 Percent Speed. 

NASA/CR—2015-218879/VOL1 130
7.2.2 Trends in Spectra and Levels With Operating Condition 

Section 7.1.2.2 provided a discussion on the trends in the dynamic pressure spectra as a 

function of operating condition. Figure 132 presents a comparison of the overall sound pressure 

level, computed over the frequency range from 50 Hz to 1000 Hz, for the internal and external 

pressure sensors, as a function of engine speed. There is a general increase in overall sound 

pressure level (OASPL) as the engine speed increases, on the order of 6-7 dB from the 48 

percent speed point to the 65 percent speed point. All of the sensors exhibit the same general 

trend. 

Figure 132. OASPL Variation as a Function of Engine Speed. 

7.2.3 Time Delays for Matched Pairs of Kulites 

Cross-spectra between pairs of dynamic pressure signals were computed to examine time 

delays. The transducer pairs examined were the two upstream/downstream probe pairs in the 

combustor, PCA45xPCF45 and PCA12xPCF12; the two aft probe combustor-to-turbine pairs on 

the same side of the engine, PLA45xPCA45 and PLA12xPCA12; the upstream/downstream pair 

in the turbine, PLA12xPLF12, and the two turbine-to-mixer pairs on the same side of the engine, 

PMX45xPLA45 and PMX12xPLA12. The cross-spectra were computed for all operating 

conditions and repeat points, in order to identify trends in the results. 

NASA/CR—2015-218879/VOL1 131
Figure 133 presents the cross spectral phase between the upstream and downstream pressure 

probes in the combustor and ITD, at 48 percent speed (March 26). The cross-spectral phase at 

higher operating speeds is very similar. Figure 134 shows the cross-spectral phase at 65 

percent speed (March 28). There appears to be a fairly well defined phase relationship at low 

frequencies (the phase is not random), but the behavior is not entirely simple. Between the 

CF45 and CA45 probes, the phase appears to change linearly from 0 to 100 Hz, remains fairly 

constant from 100 Hz to 400 Hz, and changes in a somewhat linear fashion from 400 Hz to 800 

Hz. Between the CF12 and CA12 probes, there is a nonlinear relationship between 0 and 200 

Hz, after which the phase remains relatively constant at 0 degrees up to 800 Hz. Between the 

LF12 and LA12 probes, the phase is near zero up to about 300 Hz. Overall, the phase change 

is not a simple linear function of frequency, and thus the interpretation of the results does not 

lend itself to a simple time delay analysis as one would have for a simple propagation problem, 

where disturbances would be expected to propagate at either at the sound speed or the flow 

speed. 

There are several issues that contribute to this complexity. The geometry where the pressure 

measurements are being made is semi-enclosed, leading to the possibility of reflections of 

acoustic waves and acoustic resonances. Furthermore, there are changes in the flowpath 

geometry between the upstream and downstream probes, leading to the possibility of mean flow 

variation and acoustic mode evolution. In any case, one might still attempt to fit a linear curve 

through the low frequency portion of the data. This was attempted over a range of 50 Hz to 800 

Hz for the pairs of pressure probes in the combustor, and between 20 Hz and 200 Hz for the 

pair of pressure probes in the ITD. The result of the fit is presented in Table 12. In some cases 

there is a reasonable close correspondence with the time delays for convective propagation. 

For example, the average fit-derived time delay of 0.13 ms between PLF12 and PLA12 

corresponds fairly closely with the expected convective time delay of 0.1 ms, discussed later in 

Section 7.3.3. These results must be interpreted with caution however. 

Figure 133. Cross-Spectral Phase Between the Upstream/Downstream Dynamic 

Pressure Probes in the Combustor and ITD, at 48 Percent (March 26). 

NASA/CR—2015-218879/VOL1 132
Figure 134. Cross-Spectral Phase Between the Upstream/Downstream Dynamic 

Pressure Probes in the Combustor and ITD, at 65 Percent (March 28). 

Figure 135 presents the cross spectral phase between the aft combustor and aft ITD pressure 

probes (using sensors on the same sides of the engine) and between the aft ITD and mixer 

probes (using sensors on the same sides of the engine), at 48 percent speed (March 26). The 

cross-spectral phase at higher operating speeds is very similar. Figure 136 shows the crossspectral phase at 65 percent speed (March 28). There is substantial scatter in the data, but 

there appears to be a region below about 200 Hz where the phase change exhibits a roughly 

linear change with frequency. A linear fit of the data was attempted over a frequency range 

from 20 to 200 Hz. The result of the fit is presented in Table 12. 

Figure 135. Cross-Spectral Phase Between the Combustor/Turbine and 

Turbine/Mixer Dynamic Pressure Probes, at 48 Percent (March 26). 

NASA/CR—2015-218879/VOL1 133
Figure 136. Cross-Spectral Phase Between the Combustor/Turbine and 

Turbine/Mixer Dynamic Pressure Probes, at 65 Percent (March 26). 

Table 12. Propagation Time Delays (in Milliseconds) Computed From Linear Fit of 

Measured Pressure Cross-Spectra. 

Test Date and 

Operating Point 

PCA45

x

PCF45

PCA12

X

PCF12

PLA45

X

PCA45

PLA12

X

PLF12

PLA12

X

PCA12

PMX45

X

PLA45

PMX12

X

PLA12

20140326_977CNAM4801_1 0.27 0.08 5.09 0.08 4.22 1.25  ‐0.43

20140326_977CNAM5401_1 0.22 0.10 4.84 0.10 4.01 1.12 0.66

20140326_977CNAM6001_1 0.15 0.11 4.44 0.08 3.53 1.20 0.93

20140326_977CNAM6002_1 0.14 0.12 3.89 0.09 4.07 1.24 0.81

20140328_977CNAM4801_1 0.29 0.06 5.03 0.18 4.35 1.18 0.35

20140328_977CNAM4802_1 0.29 0.05 5.04 0.15 3.84 0.56 0.26

20140328_977CNAM5401_1 0.26 0.05 5.22 0.15 3.21 0.86 0.70

20140328_977CNAM5402_1 0.26 0.07 4.76 0.07 4.10 1.56 0.75

20140328_977CNAM5403_1 0.29 0.07 5.09 0.13 3.38 0.70 0.63

20140328_977CNAM5404_1 0.25 0.07 4.96 0.24 4.19 1.12 0.50

20140328_977CNAM6001_1 0.22 0.10 4.01 0.13 3.79 1.20 0.91

20140328_977CNAM6002_1 0.20 0.12 3.87 0.03 4.03 0.92 1.19

20140328_977CNAM6501_1 0.16 0.11 2.96 0.18 2.95 1.16 0.62

20140328_977CNAM6502_1 0.16 0.13 2.99 0.17 2.51 1.11 0.64

AVERAGE 0.23 0.09 4.44 0.13 3.73 1.08 0.61

NASA/CR—2015-218879/VOL1 134
7.3 Task 5.3: Spectra and RMS Levels of Temperature and Pressure 

7.3.1 Trend Evaluation of Combustor Temperature Data 

It is of interest to examine trends in both the mean and fluctuating temperatures measured by 

the probes in the combustor, the ITD, and the mixer. Figure 144 displays the mean 

temperatures measured in the combustor as a function of engine speed. All of the probes show 

a nearly linear trend of the mean temperature increasing with engine speed. At an engine 

speed between 60 percent and 65 percent, the critical temperature of 3100°F is reached at the 

forward probe locations, at which point the platinum/rhodium thermocouple wires melt. In the 

test, the CF12 probe reached this critical temperature at a slightly higher speed than the CF12 

probe, perhaps because the CF12 probe is not as far forward or because it is in a less active 

part of the flame zone (being located behind a fuel nozzle, and not between fuel nozzles like the 

CF45 probe). The mean temperatures at the aft combustor probes CA45 and CA12 are very 

similar and much smaller than the forward probe mean temperatures. There is evidently 

substantial mixing that occurs between the forward and aft locations, a conclusion that becomes 

important when considering cross-correlations between the upstream and downstream probes. 

There is not a significant difference between the mean temperatures of the two aft probes which 

are at the same axial station. The measured mean temperatures at the aft probe locations were 

about 300°F more than expected based on engine cycle predictions. There might be additional 

flow mixing that occurs downstream of the aft probes. Figure 145 displays the mean 

temperatures measured in the ITD and the mixer as a function of engine speed. Again, the 

mean temperatures increase almost linearly with engine speed. The ITD probes at the 4.5 

location appear to run slightly hotter (about 80°F) than the ITD probes at the 12 location. The 

measured temperatures in the ITD and mixer are in good agreement with the temperatures 

expected from engine cycle predictions. 

Figure 137. Mean Temperatures Measured in the Combustor. 

NASA/CR—2015-218879/VOL1 135
Figure 138. Mean Temperatures Measured in the ITD and Mixer. 

It is of particular interest to know how the compensated temperature spectra change with 

operating speed, and how repeatable the measurements are. Figure 139 presents the small 

wire compensated temperature spectra for the forward combustor probe CF12 at speeds of 48 

percent, 54 percent, and 60 percent, along with the repeat points. The repeatability of the 

measurements is excellent, and there is very little difference in the amplitude of the 

compensated spectra as the engine speed changes. Figure 140 presents the small wire 

compensated temperature spectra for the aft combustor probe CA12 at speeds of 48 percent, 

54 percent, 60 percent, and 65 percent, along with the repeat points. The spectral amplitudes 

at the aft combustor probe location are about 5 dB less than those at the forward combustor 

probe location. The repeatability of the measurements is excellent, and again there is very little 

difference in the amplitude of the compensated spectra as the engine speed changes. 

Figure 141 presents the small wire compensated temperature spectra for the forward ITD probe 

LF12 at speeds of 48 percent, 54 percent, 60 percent, and 65 percent along with the repeat 

points. Here, the frequency range has been truncated at 500 Hz, since the data becomes 

contaminated by the electronic noise floor above this frequency. The repeatability of the 

measurements is excellent, and there is very little difference in the amplitude of the 

compensated spectra as the engine speed changes. The compensated temperature spectra 

are 20 dB below the compensated spectra at the aft combustor location, and thus the 

temperature fluctuations in the ITD are an order of magnitude smaller than they are in the 

combustor. Figure 142 presents the small wire compensated temperature spectra for the aft 

ITD probe LA12 at speeds of 48 percent, 54 percent, 60 percent, and 65 percent along with the 

repeat points. The results are very similar to the forward probe LF12, which might be expected 

if the residual temperature fluctuations are simply convected by the mean flow. 

NASA/CR—2015-218879/VOL1 136
Figure 143 presents the small wire compensated temperature spectra for the mixer probe MX12 

at speeds of 48 percent, 54 percent, 60 percent, and 65 percent along with the repeat points. 

Here the repeatability is not as good. Here, the frequency range has been truncated at 250 Hz, 

since the data becomes contaminated by the electronic noise floor above this frequency. Even 

below 250 Hz, the voltages are very low and the coherence between the large and small wires 

is not high, and for some cases the compensation process becomes questionable. The spectra 

with the lower amplitudes are believed to be more correct, since these test points show higher 

coherence and the compensated spectral amplitudes trend downwards with increasing 

frequency. More careful processing of the mixer probe data is warranted, and might improve 

these results. This was not pursued here since these temperature fluctuations, being of such 

low amplitude and at the exit of the engine, are probably not of practical importance. 

Figure 139. Speed Trend and Repeatability of the Compensated Small Wire 

Temperature Spectra at the Forward Combustor Location 12. 

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Figure 140. Speed Trend and Repeatability of the Compensated Small Wire 

Temperature Spectra at the Aft Combustor Location 12. 

Figure 141. Speed Trend and Repeatability of the Compensated Small Wire 

Temperature Spectra at the Forward ITD Location 12. 

NASA/CR—2015-218879/VOL1 138
Figure 142. Speed Trend and Repeatability of the Compensated Small Wire 

Temperature Spectra at the Aft ITD Location 12. 

Figure 143. Speed Trend and Repeatability of the Compensated Small Wire 

Temperature Spectra at the Mixer Location 12. 

NASA/CR—2015-218879/VOL1 139
7.3.2 Temperature Spectra at HPT and LPT Exits 

Section 7.1.1.2 presented compensated temperature spectra at the combustor exit (HPT 

entrance), the LPT entrance, and the LPT exit (mixer). Comparing the compensated spectra, 

the temperature fluctuations at the LPT entrance are an order of magnitude smaller than those 

at the HPT entrance. If the indirect noise generation process through the HP and LP turbines 

are comparable (i.e. if a given fluctuating temperature produces a fluctuating pressure of the 

same order of magnitude through either turbine), then the indirect noise generated in the LP 

turbine would be an order of magnitude smaller in amplitude than the indirect noise generated in 

the HP turbine. The compensated temperature spectra in the LPT exit (mixer) have about half 

the amplitude as the compensated temperature spectra in the LPT entrance (ITD). The residual 

temperature fluctuations at the LPT exit convect through the mixer nozzle and into the exhaust 

nozzle. Since there are no rotating blade rows, the mean flow gradients are not as pronounced 

as those in the LPT, and one might surmise that any indirect noise generation would be weaker 

than the indirect noise generated in the LPT. Thus, one concludes that any further study of 

indirect noise generation in the TECH977 engine should focus first on the combustor 

temperature fluctuations convecting through the HPT. The indirect noise generated in the HPT 

should dominate the indirect noise generated elsewhere. 

7.3.3 Time Delays for Matched Pairs of Small Wire Thermocouples 

The computation of phase delays between matched pairs (upstream and downstream) of small 

wire thermocouples is of interest because this can shed light on the convection speed of the 

temperature fluctuations. It is first useful to examine the coherence between the two time 

history signals. The coherence between the pairs of small wire thermocouples at the 3/26 48 

percent speed point is shown in Figure 144, over the frequency range from 0 to 500 Hz. For 

this analysis, the compensated time histories were used. 

The coherence in the combustor is quite low. There appears to be some slight coherence at 

combustor location 12 at low frequency, but the level is only 0.1 at 50 Hz. The coherence levels 

at combustor location 4.5 are less than 0.05. The higher coherence at location 12 may be due 

to the smaller probe axial spacing of 1.5-inch at that location, versus the 2.0-inch spacing at 

location 4.5. Recall also that the upstream and downstream probes were intentionally offset in 

the azimuthal direction, so that the downstream probe would not be in the wake of the upstream 

probe. This offset may contribute to some loss of coherence, as the probes would be sensing 

slightly different temperature fields. However, it appears that continued mixing between the 

upstream and downstream probes is the main reason for the reduced coherence in the 

combustor. As shown in Section 7.3.1, there is a dramatic difference in the mean temperature 

of the upstream combustor probes versus the downstream combustor probes, indicating that 

substantial mixing of the hot and cold gases, and potentially continued burning of the fuel, is still 

occurring. 

On the other hand, the coherence between the upstream and downstream small wire pairs in 

the ITD is quite high. This is to be expected, since the flow downstream of the HPT is well 

mixed with no residual burning. The remaining temperature fluctuations should be largely 

convected down the duct by the mean flow. Taylor’s hypothesis, where the turbulent eddies 

(and temperature fluctuations) are considered to be frozen in the flow, is probably a good 

approximation here. The coherence levels at combustor location 4.5 are slightly higher than 

those at location 12. This is a bit unexpected, since the axial spacing of 1.2-inch at location 4.5 

is slightly larger than axial spacing of 0.9-inch at location 12. However, the coherence at both 

locations is quite high regardless, being above 0.8 at low frequencies. 

NASA/CR—2015-218879/VOL1 140
The cross-spectral phase between the upstream and downstream small wire thermocouples is 

presented in Figure 145. As expected from the low coherence, there is not an obvious trend in 

the combustor data. However, for the CF12 and CA12 pair, there does appear to be a slight 

indication of a time delay in the cross-spectral phase below 100 Hz, as suggested by a roughly 

constant slope in the phase. In the turbine, the results are much more dramatic, as might be 

expected from the high coherence levels. There is a clearly identifiable constant slope region in 

the cross-spectral phase from 0 to 500 Hz, suggestive of a convective time delay. The slope at 

location 4.5 is larger than the slope at location 12, which is consistent with a larger spacing 

between the probes. 

To obtain an estimate of the convection speeds, a linear fit of the cross-spectral phase data is 

applied. For the combustor locations, a reliable linear fit could not be obtained due to the high 

amount of scatter in the cross-spectral phase. It appears that the residual mixing and burning is 

too large to derive a reliable estimate of the eddy convection speed. For the ITD locations, a 

linear fit is applied to the phase data between 100 and 400 Hz. This fit is expected to be very 

good due to the high coherence and strong linear relationship in the cross-spectral phase. An 

example of the linear fit through the data is shown in Figure 146. The time delays are then 

calculated from the measured slopes as ߬	= (dߠ/d݂ሻ/360°. Time delays for the 

upstream/downstream ITD pairs were calculated for all of the test points. The computed time 

delays are shown in Table 13. The average computed time delay is 0.15 ms for location 4.5 

with the 1.24-inch probe spacing, and 0.09 ms for location 12 with the 0.93-inch probe spacing. 

This is very close to the expected values. The expected time delays in the ITD, computed using 

the engine cycle estimated flow velocity (computed from the Mach number and temperature or 

sound speed) and assumed sensor spacing, is 0.13 ms for location 4.5 with the 1.24-inch probe 

spacing and 0.1 ms for location 12 with the 0.93-inch probe spacing. The expected time delay 

is roughly constant over the tested engine speeds. A comparison between the measured and 

cycle-predicted time delays is provided in Figure 147. Considering the various uncertainties in 

the measurements and cycle predictions (probe spacing, mean flow direction, cycle accuracy, 

etc.) the agreement is quite good. This provides a convincing confirmation that the temperature 

fluctuations in the ITD region are simply convected by the mean flow. 

Figure 144. Coherence Between the Upstream and Downstream Small Wire 

Thermocouple Pairs at the March 26 48 Percent Speed Point. 

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Figure 145. Cross-Spectral Phase Between the Upstream and Downstream Small 

Wire Thermocouple Pairs at the March 26 48 Percent Speed Point. 

Figure 146. Example of Linear Fits of the Cross-Spectral Phase Between the 

Upstream and Downstream Small Wire T/C Pairs at the 

March 26 48 Percent Speed Point. 

 

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Table 13. Propagation Time Delays Computed for Small Wire T/C Pairs. 

Test Date and 

Operating Point 

Computed Time Delay 

Between LF45 and LA45, ms 

Computed Time Delay 

Between LF12 and LA12, ms 

03/26/2014 4801 0.156 0.077

03/26/2014 6001 0.148 0.096

03/26/2014 5401 0.130 0.076

03/28/2014 4801 0.144 0.064

03/28/2014 5402 0.153 0.085

03/28/2014 4802 0.159 0.061

03/28/2014 5403 0.153 0.085

03/28/2014 6001 0.146 0.105

03/28/2014 5404 0.140 0.095

03/28/2014 6002 0.144 0.123

03/28/2014 6501 0.155 0.127

03/28/2014 6502 0.149 0.128

AVERAGE 0.15 0.09

Figure 147. Measured Versus Predicted Convective Time Delays Between the 

Upstream and Downstream Probes in the ITD. 

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7.3.4 Temperature and Pressure Cross Correlations 

Of particular interest to the indirect noise problem is whether there is an identifiable correlation 

between the dynamic temperature fluctuations and the dynamic pressure fluctuations inside the 

engine core. For example, a convincing argument for the existence of indirect combustion noise 

could be made if there is a significant correlation between the temperature fluctuations that are 

convected through the HPT and the pressure fluctuations that are propagating downstream of 

the turbine. In this section, correlations between temperature and pressure at various locations 

in the TECH977 core engine are examined. Reference 1 provides an early experimental 

investigation of coherence and cross-spectral measurements between pressure and 

temperature fluctuations inside a long duct downstream of a combustor. The experiment 

identified non-random phase relationships between pressure and temperature at low 

frequencies. The authors explored the coherence and cross-spectra between pressure and 

temperature fluctuations both at a single point and at spatially separated points. Similar 

explorations are considered here. 

First, consider the correlation between the dynamic temperature and the dynamic pressure at 

the same (or nearly so) location in space. For each of the pressure/temperature probes in the 

TECH977 test, the temperature measurement is made in the flow, roughly near the centerline of 

the flowpath, whereas the pressure measurement is made at the outer wall of the duct. There 

may be some loss in correlation because the temperature and pressure measurements are not 

made at the exact same location in space. For example, the length scale of the temperature 

fluctuations might be smaller than the width of the annular duct, so there is no guarantee that 

the dynamic temperature near the duct wall will be highly coherent with the dynamic 

temperature in the flow. Similarly, there is no guarantee that the dynamic pressure measured at 

the wall would be highly coherent with the dynamic pressure measured at the flowpath 

centerline1

. 

Figure 148 shows the coherence between the dynamic compensated temperature fluctuations 

measured by the small wire thermocouple in the centerline of the duct and the dynamic 

pressure measured at the duct wall (close to where the temperature probe enters the duct), for 

the four probes in the combustor, at 48 percent speed. The coherence at all four locations is 

very small, below 0.05 at all frequencies. Figure 149 plots the cross-spectral phase between 

the compensated dynamic temperature and the dynamic pressure (not adjusted for propagation 

time delay in the sensing tube) at the four locations, at 48 percent speed. Although the phase 

does not appear to be completely random, there is little indication of any strongly dominant 

pattern. Figure 151 plots the cross-spectral phase between the compensated dynamic 

temperature and the phase-corrected dynamic pressure (after adjustment for propagation time 

delay in the sensing tube) at 48 percent speed. There is perhaps a slight linear trend at the 4.5 

location, but there is significant scatter in the data. Figure 151 plots the cross-spectral phase 

between the compensated dynamic temperature and the phase-corrected dynamic pressure at 

54 percent speed. The results are similar to those at 48 percent speed. There does not appear 

to be a simple phase relationship between the dynamic pressure and dynamic temperature 

measured at a single point in the combustor. 

1

 A higher degree of coherence might be expected between dynamic pressure measured at the wall and 

at the centerline, since the geometry and mean flow conditions are such that higher order radial modes 

do not propagate in the core ducts at low frequencies. For example, the modal measurements in the 

combustor show that the plane wave is dominant at frequencies below 500 Hz and the (m=±1, n=0) 

modes are dominant at frequencies between 500 Hz and 1000 Hz). 

NASA/CR—2015-218879/VOL1 144
Figure 152 shows the coherence between the dynamic compensated temperature fluctuations 

measured by the small wire thermocouple on the temperature probe and the dynamic pressure 

measured at the duct wall, for the three locations in the inter-turbine duct. Note that the pressure 

tap at location LF45 was used to provide the purge flow reference pressure, so there is no 

dynamic pressure measurement available at that point. The frequency range is restricted to 500 

Hz, because the compensated temperature is unreliable above that frequency due to the 

electronic noise floor. There is much greater coherence in the inter-turbine duct than in the 

combustor, although the coherence levels are still modest, being generally below 0.3. The 

cross-spectral phase between the compensated dynamic temperature and the dynamic 

pressure (not adjusted for propagation time delay in the sensing tube) is shown in Figure 153. 

Here, the results are dramatically different than those from the combustor, as is more easily 

seen by plots of the unwrapped cross-spectral phase shown in Figure 154. There is a very 

definitive phase relationship between pressure and temperature fluctuations. 

At the LF12 and LA12 locations, the phase appears to be a linear function of frequency. At the 

LA45 location, the behavior is more complex; over certain frequency ranges, the phase changes 

linearly with frequency, but there are unusual step changes in the phase, e.g. at 175 Hz. The 

behavior of the phase is similar at other speeds, and becomes somewhat more consistent at the 

higher power settings. The unwrapped cross-spectral phase between the compensated dynamic 

temperature and the dynamic pressure (not adjusted for propagation time delay in the sensing 

tube) at 65 percent speed is shown in Figure 155. A linear fit of the phase data for locations 

LF12 and LA12 is applied over a frequency range from 100 to 350 Hz. The result of the fit is 

shown in Table 14. The average time delay is 2.7 ms, which corresponds very nicely with the 

expected time delay associated with propagation of the pressure signal from the wall location 

through the 3 foot long pressure sensing line to the Kulite location, summarized in Table 11. 

As an additional confirmation, the cross-spectral phase between the compensated dynamic 

temperature and the phase-corrected dynamic pressure is shown in Figure 156 and Figure 157 

for 48 percent and 65 percent speed, respectively. The phase is predominantly constant with 

frequency. At location LF12, the temperature and pressure are out of phase by 60 degrees. At 

location LA12, the temperature and pressure are out of phase by 120 degrees at 48 percent 

speed and by 60 degrees at 65 percent speed. Location LA45 exhibits a more complex 

behavior, and reveals the existence of an additional linear phase shift. An explanation for this 

behavior is currently lacking. 

The high degree of correlation between the local pressure and temperature at the inter-turbine 

duct probe locations is unexpected and intriguing. Are the pressure fluctuations in the interturbine duct hydrodynamic or acoustic in nature? Recall that the cross-spectral measurements 

between upstream and downstream small wire thermocouple pairs indicated that the 

temperature fluctuations in the inter-turbine duct are convected with the mean flow. Since the 

temperature and pressure are well correlated, does this then imply that the pressure fluctuations 

are convected with the mean flow? Does this then imply that the pressure fluctuations are 

dominated by hydrodynamic disturbances? Or is it possible that indirect noise is being 

generated in the HP turbine, and that the residual temperature fluctuations are highly correlated 

with the pressure fluctuations? This is an area that requires further exploration. 

NASA/CR—2015-218879/VOL1 145
Figure 148. Coherence Between Dynamic Temperature and Dynamic Pressure at 

the Combustor Probe Locations, at 48 Percent Speed (March 26). 

Figure 149. Cross-Spectral Phase Between Dynamic Temperature and Dynamic 

Pressure at the Combustor Probe Locations, at 48 Percent Speed (March 26). 

NASA/CR—2015-218879/VOL1 146
Figure 150. Cross-Spectral Phase Between Dynamic Temperature and PhaseCorrected Pressure at the Combustor Probe Locations, 

at 48 Percent Speed (March 26). 

Figure 151. Cross-Spectral Phase Between Dynamic Temperature and 

Phase-Corrected Pressure at the Combustor Probe Locations, 

at 54 Percent Speed (March 28). 

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Figure 152. Coherence Between Dynamic Temperature and Dynamic Pressure at 

the ITD Probe Locations, at 48 Percent Speed (March 26). 

Figure 153. Cross-Spectral Phase Between Dynamic Temperature and Dynamic 

Pressure at the ITD Probe Locations, at 48 Percent Speed (March 26). 

NASA/CR—2015-218879/VOL1 148
Figure 154. Unwrapped Cross-Spectral Phase Between Dynamic Temperature 

and Pressure at the ITD Probe Locations, at 48 Percent Speed (March 26). 

Figure 155. Unwrapped Cross-Spectral Phase Between Dynamic Temperature 

and Pressure at the ITD Probe Locations, at 65 Percent Speed (March 28). 

NASA/CR—2015-218879/VOL1 149
Table 14. Computed Time Delays Between Local T′ and P′ in the ITD. 

Test Date and 

Operating Point 

Computed Time Delay at 

Location LF12, ms 

Computed Time Delay at 

Location LA12, ms 

03/26/2014 4801 2.75 2.80 

03/26/2014 6001 2.72 2.58 

03/26/2014 5401 2.70 2.76 

03/28/2014 4801 2.77 2.95

03/28/2014 5402 2.71 2.77

03/28/2014 4802 2.77 2.86

03/28/2014 5403 2.71 2.77

03/28/2014 6001 2.66 2.59

03/28/2014 5404 2.73 2.71

03/28/2014 6002 2.73 2.71

03/28/2014 6501 2.64 2.43

03/28/2014 6502 2.52 2.44

AVERAGE 2.70 2.70

Figure 156. Cross-Spectral Phase Between Dynamic Temperature and PhaseCorrected Pressure for ITD Probe Locations, at 48 Percent Speed (March 26). 

NASA/CR—2015-218879/VOL1 150
Figure 157. Cross-Spectral Phase Between Dynamic Temperature and PhaseCorrected Pressure for ITD Probe Locations, at 65 Percent Speed (March 26). 

Next, the coherence and cross-spectra between pressure and temperature fluctuations at 

spatially separated points is examined. Perhaps the most interesting comparison is between 

the temperature fluctuations upstream of either the HPT or LPT, and the pressure fluctuations 

directly downstream of the corresponding turbine. If there is a high degree of correlation, this 

might indicate the presence of indirect combustion noise. However, one could expect that there 

will be many localized incoherent temperature fluctuations convecting through the turbine at the 

same time, and thus multiple incoherent sources generating noise simultaneously. Pressure 

measurements downstream of these sources might contain contributions from multiple 

incoherent sources, and thus the coherence between temperature upstream and pressure 

downstream of the turbine might be substantially reduced. Figure 158 presents the coherence 

between the compensated dynamic temperature upstream of the turbine and the phasecorrected dynamic pressure at the nearest location directly downstream of the turbine. The 

coherence levels are quite low in general. Figure 159 presents the corresponding crossspectral phase between the same temperature and pressure measurements. Surprisingly, there 

is a fairly well defined phase relationship. Across the HPT, the relationship is not very clear, but 

the phase certainly does not appear to be random. There is a hint of a linear relationship 

between the dynamic temperature from the aft combustor probe at location 12 and the dynamic 

pressure from the forward pressure probe at location 12. The phase relationship between the 

temperature probes upstream of the LPT and the pressure probes downstream is even more 

definitive, with the phase dominated by a linear behavior. Further analysis is required to 

understand the reasons for this behavior and whether the associated time delay matches with a 

convective time scale. 

NASA/CR—2015-218879/VOL1 151
Figure 158. Coherence Between Dynamic Temperature (Compensated) Upstream 

and Dynamic Pressure (Phase-Corrected) Downstream of the HP and LP Turbines. 

Figure 159. Cross-Spectral Phase Between Dynamic Temperature (Compensated) 

Upstream and Pressure (Phase-Corrected) Downstream of the HP and LP Turbines. 

7.4 Task 5.4: Distribution of Data and Analysis Results 

Honeywell provided NASA with an initial data package containing all of the raw dynamic 

pressure and dynamic temperature time history data that was acquired during the test. This 

data package contained the following data: 

1. Dewetron Data Files from March 26, 2014 (Native Dewetron format and Matlab format

exported using DEWESoft X)

2. Dewetron Data Files from March 28, 2014 (Native Dewetron format and Matlab format

exported using DEWESoft X)

NASA/CR—2015-218879/VOL1 152
3. Excel spreadsheet describing details of the Matlab exported files (engine operating

point, beginning and ending times)

4. B&K Pulse Data Files from March 26, 2014 (Binary Universal File Format)

5. B&K Pulse Data Files from March 28, 2014 (Binary Universal File Format)

A second package of information and analysis results was prepared containing the following 

data: 

1. Probe dimensional characteristics

2. Relevant engine operating condition data

3. Input and output files for the DGTMS software

4. Narrowband compensated temperature and pressure spectra

5. Analysis results

6. Conclusions and recommendations for further analysis/testing (See Section 8.0)

7. Honeywell’s updated version (source code) of the NASA DGTMS software

8.0 CONCLUSIONS 

8.1 Major Results 

A comprehensive set of dynamic temperature and dynamic pressure measurements have been 

acquired on the TECH977 engine. Dynamic temperature measurements were successfully 

acquired in the combustor, the ITD, and the mixer, using dual wire thermocouple probes. The 

temperature probes were constructed using 0.003-inch and 0.010-inch diameter Type B 

thermocouple sensing elements mounted in a ceramic stick. A pneumatic actuation system was 

developed to deploy the temperature probes into the middle of the annular flowpath. Dynamic 

pressure measurements were simultaneously acquired next to each temperature probe, and at 

four additional locations in the combustor, using Kulite differential pressure transducers 

mounted in semi-infinite tubes. The probes in the combustor and ITD were cooled using a small 

amount of nitrogen purge flow. Dynamic temperature and pressure time history measurements 

were made at four steady-state operating speeds covering the low speed regime where 

combustion noise is of practical interest. At each speed, the time history data were acquired for 

at least 70 seconds using both a Dewetron data acquisition system (primary) and a Pulse data 

acquisition system (backup). After the test, the NASA DGTMS software was used to process 

the raw thermocouple voltages and generate compensated temperature spectra corresponding 

to the fluctuating temperature of the gas stream. Compensated temperature spectra were 

generated up to a frequency range of 1000 Hz in the combustor, 500 Hz in the ITD, and 250 Hz 

in the mixer. Compensated time histories were also generated, representative of the fluctuating 

temperature of the gas stream. The compensated time history spectra at a particular 

measurement location were found to be fairly constant with engine speed, over the range of 

speeds tested. 

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In the combustor, fluctuating temperature data were acquired at mean temperatures up to the 

limits of probe survivability (3100°F). At the two higher power settings, mean temperatures at 

the forward combustor probe locations exceeded 3100°F, so data could not be obtained at 

those speeds. Mean temperatures at the aft combustor locations were between 2000°F and 

2400°F and did not impose any limitations on testing. Difficulties with the actuation of the aft 

combustor probe on the right side of the engine limited the amount of temperature data acquired 

at that location. The compensated temperature spectra in the combustor have a wideband 

spectral content. On a narrowband basis, the spectra are relatively flat, with a slight decrease in 

amplitude with frequency. The temperature measurements in the combustor indicate that 

significant burning of fuel and/or mixing of the hot and cold gas occurs between upstream and 

downstream probe locations. As a result, correlations between upstream and downstream 

temperature probes, such as the determination of convective time delays, were not easily 

established. Pressure measurements made using a circumferential array of four equally spaced 

probes in the forward end of the combustor confirmed the modal content that was previously 

measured during the NASA EVNERT program, although some small differences were noted. 

Cross-spectra between upstream and downstream pressure probes revealed a non-random 

phase relationship that is more complicated than a simple propagation delay. Cross-spectra 

between temperature and pressure fluctuations measured at a common location did not reveal 

any simple phase relationship between the two. 

In the ITD (downstream of the HPT) and the mixer (downstream of the LPT), the measured 

temperature fluctuations were an order of magnitude smaller than those measured in the 

combustor. Mean gas temperatures were well below the survivability limits of the probes, and 

did not impose any limitations on testing. A high coherence between upstream and downstream 

ITD probes was observed, and inspection of the cross spectral phase data showed a linear 

relationship with frequency indicating that the temperature fluctuations are convected by the 

mean flow. The measured time delays are consistent with expectations. There is very little 

coherence between the temperature fluctuations measured on opposite sides of the engine, 

suggesting that the length scales of the temperature fluctuations are much smaller than the 

circumference of the annulus. Interestingly, there is a strong coherence between the pressure 

and temperature fluctuations measured at the same location in the ITD. Inspection of the crossspectral phases indicates that, after accounting for the propagation time delay along the sensing 

tube of the semi-infinite coil, the pressure and temperature fluctuations have a predominately 

constant phase offset. In addition, cross-spectra between the compensated temperature 

fluctuations measured upstream of a turbine and the phase-corrected pressure fluctuations 

measured directly downstream of the turbine indicated a non-random phase relationship 

between the two variables. An explanation for this behavior is currently lacking. 

A fundamental question of interest is whether the temperature and pressure measurements 

acquired on the TECH977 engine provide a clear indication of the presence of indirect 

combustion noise. For example, can it be shown that indirect combustion noise dominates over 

direct combustion noise across a particular frequency range? This is a challenging question to 

answer, even with temperature and pressure measurements in hand. In order to answer it, one 

must first define what constitutes “convincing evidence” of indirect combustion noise. At the 

simplest of approaches, one can compare pressure spectra measured upstream and 

downstream of a turbine stage and look for dramatic changes in the spectra. If there is a region 

where the downstream spectra is higher than the upstream spectra (e.g. a broadband peak), 

one might consider this to be an indication of indirect noise generated by the convection of 

temperature fluctuations through the flow gradients in the turbine. In the current test, no clear 

and convincing spectral amplification was observed. However, such a comparison is 

complicated by the possibility of indirect noise propagating upstream of the turbine (thus 

NASA/CR—2015-218879/VOL1 154
confounding the measurements upstream) and by flow and geometry changes that alter the 

acoustic resonances (and thus the broadband peaks) in the vicinity where the pressure 

measurements are made. An alternative approach is to look at correlations between variables, 

such as between temperature fluctuations upstream of a turbine and pressure fluctuations 

downstream, and consider high coherence over a particular frequency range as positive 

indication of indirect combustion noise. A high coherence between temperature fluctuations in 

the combustor and pressure fluctuations downstream of the HPT was not observed in the 

current test. However, this criteria also oversimplifies the problem for reasons discussed 

subsequently. Alternatively, one could explore more sophisticated metrics and signal 

processing techniques. One example is provided by the entropy wave generator [Reference 

16], where indirect noise has been studied in a controlled laboratory setting with a simplified 

nozzle geometry and where fluctuating temperature sources could be turned on or off. The 

increase in noise power when the entropy sources were turned on was found to trend with 

nozzle Mach number according to theoretical expectations, providing convincing evidence of the 

presence of indirect noise. In an engine test, where the direct and indirect sources coexist and 

cannot be independently varied, where there are dramatic changes in the flowpath geometry, 

and where multiple turbine rows are present, the task of establishing “convincing evidence” is 

much more daunting. With the simplified approaches considered here, no convincing 

identification of indirect combustion noise could be made. More sophisticated analysis will have 

to be carried out to identify the indirect noise contribution. 

8.2 Recommendations for Further Analysis 

Much of the analysis presented in this report is exploratory in nature. Some fundamental 

conclusions have been presented, but more sophisticated signal processing techniques will be 

required to yield additional insights and conclusions. Although the main objective of acquiring 

dynamic temperature and pressure data has been achieved, the answer to the underlying 

question of whether indirect combustion noise can be detected remains elusive. The spectra 

and correlations between the dynamic pressure and dynamic temperature measurements are 

often non-trivial, and a more careful dissection of the data is believed to be necessary. In 

particular, the analyses such as those explored in References 12–15, which have proven 

insightful on previous TECH977 testing, are recommended. 

One aspect of the indirect noise problem that probably needs thoughtful consideration is the 

impact of multiple noise sources. The length scales of the temperature fluctuations in the 

combustor are likely to be small (there are multiple fuel nozzles, each creating burning and 

mixing zones that are uncorrelated with the others). When multiple uncorrelated temperature 

fluctuations convect through the turbine and create pressure fluctuations, one would expect that 

the pressure fluctuations generated by each source would be uncorrelated. Furthermore, the 

pressure fluctuations measured at locations downstream of the HPT are likely to be a 

combination of multiple uncorrelated sources (not only from indirect combustion noise, but also 

direct combustion noise and possibly other sources). The presence of multiple uncorrelated 

sources will result in diminished coherence between pressure and temperature fluctuations at 

various points. The cross-spectral phase, time delay, or other computed quantities will also be 

affected, and the interpretation of the results should consider the impact of multiple sources. 

How the sources couple to the acoustic modes in the duct is also likely to be a relevant concern, 

since only low order modes are able to propagate in the engine core. 

It is recommended that computational simulations of indirect combustion noise generation be 

explored to gain additional insight into the measurements. In particular, it would be helpful to 

better understand the magnitude of pressure fluctuations generated by the indirect mechanism 

NASA/CR—2015-218879/VOL1 155
as a function of incident temperature fluctuations. Also, it would be helpful to explore the 

coherence and cross-spectra between the dynamic temperature and dynamic pressure at 

different locations. Although the simulations often incorporate major simplifications regarding 

the geometry, the mean flow, the incident disturbance field, and most notably the turbine 

aerodynamics, and although the simulations come with their own set of complications, there are 

notable advantages. The simulation datasets can be explored much more thoroughly; the 

indirect noise mechanism can be observed more directly in the absence of competing sources; 

and the limitations and uncertainties associated with physical instrumentation are not present. 

Furthermore, the strength, length scales, and number of sources of the incident temperature 

disturbances can be varied to explore how this affects the correlations between pressure and 

temperature. These investigations might also provide recommendations for instrumentation 

types and locations for future test programs, along with recommendations for data acquisition 

and signal processing procedures. 

8.3 Recommendations for Future Testing 

Based on a recommendation from previous test efforts, a probe actuation system was 

developed for this test effort to deploy and retract the temperature probes into the flow, with the 

purpose of enhancing probe survivability. Despite significant planning and risk mitigation 

efforts, difficulties were still encountered with some of the probe actuators. The system 

achieved the objective of removing the probes from the flowpath during engine startup and rapid 

transients, and may have contributed to the survivability of the probes. However, significant 

effort and expense was involved in designing, procuring, assembling, and testing the actuation 

system. In future testing on other engines, it may be more cost effective to use fixed probes 

and manufacture more spare probes to replace any that fail. Although the failure rate of probes 

during an engine start has not been assessed in this program, the use of a fixed probe system 

might be an acceptable risk. 

In a realistic engine such as the TECH977 where space is a premium, significant mixing and 

burning of the gas occurs throughout the combustor. There is not a substantial axial distance 

where the flow has been fully mixed and simply convects downstream into the HPT. 

Additionally, hardware constraints such as the presence of flanges limit how closely probes can 

be placed to the turbine nozzle. As a result, it is difficult to obtain spatially separated combustor 

temperature measurements in a purely convective flow regime. If a rigorous determination of 

the convection speeds of the temperature fluctuations at the exit of the combustor is a priority, it 

might be better to perform a rig test where a long duct could be installed downstream of the 

combustor, to allow for additional stabilization of the flow. 

The fluctuating temperature measurements in the ITD and the mixer had a limited frequency 

range due to the noise floor of the data acquisition system. This was not considered to be a 

limitation in the TECH977 test, since the magnitude of the fluctuating temperatures in these 

areas is very small, and any fluctuating pressures generated by the indirect noise process is 

likely small in comparison to direct and indirect noise generated by the combustor and HPT. 

However, if accurate measurement of low amplitude temperature fluctuations are needed at 

higher frequencies, (e.g. for use in a validation study of a numerical simulation) then low noise 

signal conditioning amplifiers that are tailored to amplifying very low voltages will need to be 

identified and employed. 

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9.0 REFERENCES 

1. Miles, J.H., Wasserbauer, C.A., and Krejsa, E.A., “Cross Spectra Between

Temperature and Pressure in a Constant Area Duct Downstream of a Combustor,”

NASA TM 83351, 1983.

2. Miles, J.H., Wasserbauer, C.A., and Krejsa, E.A., “Cross Spectra Between

Temperature and Pressure in a Constant Area Duct Downstream of a Hydrogen

Fueled Combustor,” NASA TM 83463, 1983.

3. Fralick, G.C., Oberle, L.G., and Greer, L.C. III, “Dynamic Gas Temperature

Measurements Using a Personal Computer for Data Acquisition and Reduction,”

NASA TM106119, 1993.

4. Elmore, D.L., Robinson, W.W, and Watkins, W.B., “Dynamic Gas Temperature

Measurement System, Volume I – Technical Efforts,” United Technologies Pratt &

Whitney Division, NASA CR-168267, 1983.

5. Purpura, P. T., “Dynamic Gas Temperature Measurement System, Volume II –

Operation and Program Manual,” United Technologies Pratt & Whitney Division,

NASA CR-168267, 1983.

6. Elmore, D.L., Robinson, W.W, and Watkins, W.B., “Further Development of the

Dynamic Gas Temperature Measurement System, Volume I – Technical Efforts,”

United Technologies Pratt & Whitney Division, NASA CR-179513, 1986.

7. Stocks, D.R. and Elmore, D.L., “Further Development of the Dynamic Gas

Temperature Measurement System, Volume II – Computer Program User’s Manual,”

United Technologies Pratt & Whitney Division, NASA CR-179513, 1986.

8. Weir, D., “Engine Validation of Noise and Emission Reduction Technology Phase I,”

NASA/CR—2008-215225, 2008.

9. Royalty, C. M. and Schuster, B., “Noise from a Turbofan Engine Without a Fan from

the Engine Validation of Noise and Emission Reduction Technology (EVNERT)

Program,” AIAA-2008-2810, 14th AIAA/CEAS Aeroacoustics Conference,

Vancouver, British Columbia Canada, 5-7 May 2008.

10. Wegner, M.A., Nance, D., and Ahuja, K. K., “Characterization of Short and Infiniteline Pressure Probes for In-duct Acoustic Measurements under Hostile

Environment,” AIAA 2007-3443, 13th AIAA/CEAS Aeroacoustics Conference, Rome,

Italy, 21-23 May 2007.

11. Lourier, J.M., Reichling, G., Stöhr, M., Domenico, M.D., Noll, B.E., and Aigner, M.,

“Numerical Analysis of Probe Microphones Used for Thermoacoustic

Measurements”, AIAA 2012-0545, Proc. 50th AIAA Aerospace Sciences Meeting,

Nashville TN, 2012.

12. Miles, J. H., “Time Delay Analysis of Turbofan Engine Direct and Indirect

Combustion Noise Sources,” J. of Propulsion and Power, V. 25, N1, Jan.-Feb. 2009.

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13. Hultgren, L. S. and Miles, J. H., “Noise-Source Separation Using Internal and FarField Sensors for a Full-Scale Turbofan Engine,” NASA Tm 2009-215834, 2009.

14. Miles, J. H., “Separating Direct and Indirect Turbofan Engine Combustion Noise

While Estimating Post-Combustion (Post-Flame) Residence Time Using the

Correlation Function,” NASA TM 2011-216248, 2011.

15. Miles, J. H., “Spatial Correlation in the Ambient Core Noise Field of a Turbofan

Engine,” J. Acoustical Society of America, V131, N6, June 2012.

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