E. Allen Arrington and Christine M. Pastor

Sierra Lobo, Inc., Cleveland, Ohio

Jose C. Gonsalez and Monroe R. Curry III

QSS Group, Inc., Cleveland, Ohio 

Aero-Thermal Calibration of the NASA Glenn Icing 

Research Tunnel (2004 and 2005 Tests)

NASA/CR—2010-216733

July 2010
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question via the Internet to help@

sti.nasa.gov

• Fax your question to the NASA STI Help Desk 

at 443–757–5803

• Telephone the NASA STI Help Desk at

 443–757–5802

• Write to:

 NASA Center for AeroSpace Information (CASI)

 7115 Standard Drive

 Hanover, MD 21076–1320
E. Allen Arrington and Christine M. Pastor

Sierra Lobo, Inc., Cleveland, Ohio

Jose C. Gonsalez and Monroe R. Curry III

QSS Group, Inc., Cleveland, Ohio 

Aero-Thermal Calibration of the NASA Glenn Icing 

Research Tunnel (2004 and 2005 Tests)

NASA/CR—2010-216733

July 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contracts NAS3–00145 and NNC05CA95C
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://gltrs.grc.nasa.gov

Level of Review: This material has been technically reviewed by expert reviewer(s).
NASA/CR—2010-216733 iii 

Contents 

Summary ....................................................................................................................................................... 1

Introduction ................................................................................................................................................... 1

Symbols ........................................................................................................................................................ 2

Subscripts ............................................................................................................................................. 3

Description of Facility .................................................................................................................................. 4

Test Hardware, Instrumentation and Data Systems ...................................................................................... 5

9-ft Survey Rake ................................................................................................................................... 5

2–D RTD Array .................................................................................................................................... 8

Facility Instrumentation ..................................................................................................................... 10

Steady-State Data Acquisition System ............................................................................................... 11

Thermal Anemometry Instrumentation .............................................................................................. 11

Test Procedures and Execution ................................................................................................................... 12

Test Matrix ......................................................................................................................................... 12

Test Procedures .................................................................................................................................. 13

Data Reduction and Analysis ...................................................................................................................... 15

Facility Instrumentation ..................................................................................................................... 15

9-ft Survey Rake Calculations ............................................................................................................ 16

Basic Flow Parameters ............................................................................................................. 16

Flow Angularity ....................................................................................................................... 17

Turbulence Intensity ................................................................................................................. 18

2–D RTD Array Calculations ............................................................................................................. 20

Measurement and Data Uncertainty ................................................................................................... 20

Data Analysis ..................................................................................................................................... 20

Discussion of Results .................................................................................................................................. 21

Aero-Thermal Calibration Curves ...................................................................................................... 21

Updates Based on 2004 Calibration Test ................................................................................. 21

Inputs to IRTAT ....................................................................................................................... 22

Test Section Conditions; Calibrated Values ............................................................................. 23

Test Section Conditions; Calculated Values ............................................................................ 24

Updates Based on 2005 Interim Calibration ............................................................................ 24

Aero-Thermal Flow Quality Data ...................................................................................................... 24

Total Pressure ........................................................................................................................... 25

Mach Number........................................................................................................................... 32

Total Temperature .................................................................................................................... 32

Flow Angularity ....................................................................................................................... 55

Turbulence ............................................................................................................................... 55

Summary of Test Results ............................................................................................................................ 70

Recommendations ....................................................................................................................................... 70

Appendix A.—Uncertainty Estimates ......................................................................................................... 71

A.1 General Uncertainty Analysis ................................................................................................... 71

A.2 Hardware RTD Temperature Uncertainty ................................................................................. 71

A.3 Hardware Pressure Uncertainty ................................................................................................ 72

A.3.1 Total Pressure .............................................................................................................. 72

A.3.2 Dynamic Pressure ....................................................................................................... 73

A.3.3 Static Pressure ............................................................................................................. 74

A.4 Airspeed Calculations ............................................................................................................... 74

A.5 Mach Number Calculations ......................................................................................................76

A.6 Flow Angle Calculations ........................................................................................................... 76
NASA/CR—2010-216733 iv 

Appendix B.—RTD Probe Flow Recovery Coefficients ............................................................................ 79

Appendix C.—Calculated Test Section Conditions .................................................................................... 81

Appendix D.—2005 Interim Aero-Thermal Calibration of the IRT ........................................................... 83

D.1 Total Temperature Calibration .................................................................................................. 84

D.2 Flow Quality Comparison ......................................................................................................... 85

D.3 Calibration Curve Comparison ................................................................................................. 87

References ................................................................................................................................................. 115

List of Tables 

Table 1.—Test matrix for the IRT aero-thermal calibration using the 9-ft survey rake (2004 test) .......... 12

Table 2.—Temperature calibration test matrix using the 2–D RTD array (2004 test). .............................. 13

Table 3.—Typical test conditions in the IRT during the 9-ft survey rake testing as part of the 

2004 aero-thermal calibration ............................................................................................................ 14

Table 4.—Aero-thermal flow quality goals for an icing tunnel test section ............................................... 25

Table A.1.—RTD total temperature probe system component uncertainties. ............................................ 72

Table A.2.—Total Pressure Component Bias uncertainties ........................................................................72

Table A.3.—Total Pressure Component Precision uncertainties ................................................................ 73

Table A.4.—Facility Uncertainties ............................................................................................................. 77

Table B.1.—Flow recovery correction coefficients for the RTD probes mounted on the 2–D RTD 

Arra..................................................................................................................................................... 79

Table C.1.—Summary of the parameters used to determine the test section aero-thermal 

conditions in the IRT .......................................................................................................................... 82

Table D.1.—Test matrix for the IRT aero-thermal calibration using the 9-ft survey rake (2005 

interim calibration). ............................................................................................................................ 83

Table D.2.—Temperature calibration test matrix using the 2–D RTD array (2005 interim 

calibration). ........................................................................................................................................ 84

Table D.3.—General equation and curve fit coefficients used in the calculation of the calibrated 

test section total temperature. ............................................................................................................. 85

List of Figures 

Figure 1.—Layout of the NASA Glenn Icing Research Tunnel. .................................................................. 4

Figure 2.—Installation photo of the replacement inlet guide vanes (IGV) upstream of the IRT 

drive fan. ............................................................................................................................................... 5

Figure 3.—Photograph of the 9-ft horizontal survey rake installed in the IRT test section. ........................ 6

Figure 4.—Plan view of the IRT test section showing the installed location of the 9-ft survey 

rake. ...................................................................................................................................................... 6

Figure 5.—Plan and elevation views of the 9-ft survey rake. ....................................................................... 7

Figure 6.—Cross section of the 9-ft survey rake showing flow probes details. ........................................... 7

Figure 7.—Details of RTD probes used on the 2–D RTD array (dimensions in inches). ............................. 8

Figure 8.—7x7 grid arrangement of the RTD probes in the IRT test section. .............................................. 9

Figure 9.—RTD Array installed in the IRT test section. (a) View looking downstream. (b) View 

looking upstream. ................................................................................................................................. 9

Figure 10.—RTD array vertical member cross sectional details. ............................................................... 10

Figure 11.—Typical total temperature flow recovery curve for the 2–D RTD array probes. ..................... 10

Figure 12.—Total temperature flow recovery data for RTD probes 1 through 23 with average 

curve fit. ............................................................................................................................................. 10

Figure 13.—Layout of the RTD probes mounted on the D corner turning vanes. ...................................... 11

Figure 14.—Schematic depicting terminology on one wire of X probe. .................................................... 19
NASA/CR—2010-216733 v 

Figure 15.—Aero-thermal calibration curves based on the 2004 full calibration test ................................ 22

Figure 16.—Total pressure distribution data from the IRT test section. Data collected using the 9-

ft survey rake mounted at 11 vertical positions over several tunnel runs. The test section 

total pressure data is normalized using the bellmouth total pressure measurement. Data from 

the 66-in. vertical position has been omitted due to boundary layer influence. Each y-axis 

grid (0.0002 total pressure ratio) represents approximately 0.003 psia. (a) Uts=50 mph (43 

knots), Pair=0 psig. (b) Uts=50 mph (43 knots), Pair=70 psig. (c) Uts=100 mph (87 knots), 

Pair=0 psig. (d) Uts=100 mph (87 knots), Pair=70 psig. ...................................................................... 26

Figure 16.—Continued. (e) Uts= 115 mph (100 knots), Pair=0 psig. (f) Uts=115 mph (100 knots), 

Pair=70 psig. (g) Uts=150 mph (130 knots), Pair=0 psig. (h) Uts=150 mph (130 knots), Pair=70 

psig. .................................................................................................................................................... 27

Figure 16.—Continued. (i) Uts= 175 mph (152 knots), Pair=0 psig. (j) Uts=175 mph (152 knots), 

Pair=5 psig. (k) Uts=175 mph (152 knots), Pair=20 psig. (l) Uts=175 mph (152 knots), Pair=70 

psig. .................................................................................................................................................... 28

Figure 16.—Continued. (m) Uts= 200 mph (174 knots), Pair=0 psig. (n) Uts=200 mph (174 knots), 

Pair=70 psig. (o) Uts=250 mph (217 knots), Pair=0 psig. (p) Uts=250 mph (217 knots), Pair=70 

psig. .................................................................................................................................................... 29

Figure 16.—Continued. (q) Uts= 225 mph (195 knots), Pair=0 psig. (r) Uts=225 mph (195 knots), 

Pair=20 psig (no z=36-in. vertical position data). (s) Uts=225 mph (195 knots), Pair=70 psig. ........... 30

Figure 16.—Concluded. (t) Uts= 300 mph (261 knots), Pair=0 psig. (u) Uts=300 mph (261 knots), 

Pair=70 psig. (v) Uts=350 mph (304 knots), Pair=0 psig. (w) Uts=350 mph (304 knots), 

Pair=70 psig. ........................................................................................................................................ 31

Figure 17.—Mach number distribution data from the IRT test section. Data collected using the 9-

ft survey rake mounted at 11 vertical positions over several tunnel runs. The test section 

Mach number data was normalized using the bellmouth Mach number measurement. Data 

from the 66-in. vertical position was been omitted due to boundary layer influence. 

Approximate Mach number delta is indicated for each test section setting. ( a) Uts=50 mph, 

Mts=0.066, Pair=0 psig. (b) Uts=50 mph, Mts=0.066, Pair=70 psig. (c) Uts=100 mph, 

Mts=0.133, Pair=0 psig. (d) Uts=100 mph, Mts=0.133, Pair=70 psig. .................................................. 33

Figure 17.—Continued. (e) Uts=115 mph, Mts=0.152, Pair=0 psig. (f) Uts=115 mph, Mts=0.152, 

Pair=70 psig. (g) Uts=150 mph, Mts=0.199, Pair=0 psig. (h) Uts=150 mph, Mts=0.199, Pair=70 

psig. .................................................................................................................................................... 34

Figure 17.—Continued. (i) Uts=175 mph, Mts=0.232, Pair=0 psig. (j) Uts=175 mph, Mts=0.232, 

Pair=5 psig. (k) Uts=175 mph, Mts=0.232, Pair=20 psig. (l) Uts=175 mph, Mts=0.232, Pair=70 

psig. .................................................................................................................................................... 35

Figure 17.—Continued. (m) Uts=200 mph, Mts=0.265, Pair=0 psig. (n) Uts=200 mph, Mts=0.265, 

Pair=70 psig. (o) Uts=250 mph, Mts=0.332, Pair=0 psig. (p) Uts=250 mph, Mts=0.332, Pair=70 

psig. .................................................................................................................................................... 36

Figure 17.—Continued. (q) Uts=225 mph, Mts=0.305, Pair=0 psig. (r) Uts=225 mph, Mts=0.265, 

Pair=20 psig. (s) Uts=225 mph, Mts=0.305, Pair=70 psi. ...................................................................... 37

Figure 17.—Concluded. (t) Uts=300 mph, Mts=0.399, Pair=0 psig. (u) Uts=300 mph, Mts=0.399, 

Pair=70 psig. (v) Uts=350 mph, Mts=0.464, Pair=0 psig. (w) Uts=350 mph, Mts=0.464, Pair=70 

psig. .................................................................................................................................................... 38

Figure 18.—Total temperature distribution along the D-corner turning vanes as measured by the 

D-corner RTD probes. See Figure 13 for location of probes relative to the heat exchanger. 

(a) TT,davg>4 °C. (b) –0.5>TT,davg>–3 °C. (c) –3>TT,davg> –9 °C. .......................................................... 40

Figure 18.—Concluded. (d) –11>TT,davg>–22 °C. (e) –22>TT,davg>–31 °C. ................................................. 41

Figure 19.—IRT test section total temperature distribution data collected using the 2–D RTD 

Array (a) Uts=50.1 mph (43.5 knots), TT,davg=4.3 °C, Pair=0 psig. (b) Uts=50.4 mph (43.8 

knots), TT,davg=5.2 °C, Pair=70 psig. (c) Uts=50.6 mph (44.0 knots), TT,davg=0.2 °C, Pair=0 

psig. (d) Uts=50.5 mph (43.9 knots), TT,davg=0.2 °C, Pair=70 psig. ..................................................... 42
NASA/CR—2010-216733 vi 

Figure 19.—Continued. (e) Uts=50.4 mph (43.8 knots), TT,davg=–5.4 °C, Pair=0 psig. (f) Uts=49.9 

mph (43.3 knots), TT,davg=–5.3 °C, Pair=70 psig. (g) Uts=49.9 mph (43.4 knots), TT,davg=–17.6 

°C, Pair=0 psig. (h) Uts=49.9 mph (43.4 knots), TT,davg=–17.6 °C, Pair=70 psig. ................................. 43

Figure 19.—Continued. (i) Uts=50.0 mph (43.5 knots), TT,davg=–29.7 °C, Pair=0 psig. (j) Uts=50.2 

mph (43.7 knots), TT,davg=–29.6 °C, Pair=70 psig. (k) Uts=115.6 mph (100.5 knots), 

TT,davg=4.8 °C, Pair=0 psig. (l) Uts=116.4 mph (101.2 knots), TT,davg=5.3 °C, Pair=70 psig. ................ 44

Figure 19.—Continued. (m) Uts=116.8 mph (101.5 knots), TT,davg=1.3 °C, Pair=0 psig. (n) 

Uts=116.5 mph (101.2 knots), TT,davg=1.3 °C, Pair=70 psig. (o) Uts=116.4 mph (101.1 knots), 

TT,davg=–4.2 °C, Pair=0 psig. (p) Uts=117.0 mph (101.7 knots), TT,davg=–4.2 °C, Pair=70 psig. ........... 45

Figure 19.—Continued. (q) Uts=116.8 mph (101.5 knots), TT,davg=–16.4 °C, Pair=0 psig. (r) 

Uts=116.7 mph (101.4 knots), TT,davg=–16.4 °C, Pair=70 psig. (s) Uts=116.6 mph (101.3 

knots), TT,davg=–28.4 °C, Pair=0 psig. (t) Uts=116.4 mph (101.2 knots), TT,davg=–28.6 °C, 

Pair=70 psig. ........................................................................................................................................ 46

Figure 19.—Continued. (u) Uts=151.5 mph (131.7 knots), TT,davg=2.1 °C, Pair=0 psig. (v) 

Uts=151.8 mph (131.9 knots), TT,davg=2.3 °C, Pair=70 psig. (w) Uts=152.3 mph (132.4 knots), 

TT,davg=–3.5 °C, Pair=0 psig. (x) Uts=151.8 mph (131.9 knots), TT,davg=–3.3 °C, Pair=70 psig. ........... 47

Figure 19.—Continued. (y) Uts=152.2 mph (132.3 knots), TT,davg=–15.4 °C, Pair=0 psig. (z) 

Uts=151.6 mph (131.6 knots), TT,davg=–15.6 °C, Pair=70 psig. (aa) Uts=151.8 mph (131.9 

knots), TT,davg=–27.9 °C, Pair=0 psig. (bb) Uts=152.5 mph (132.5 knots), TT,davg=–27.7 °C, 

Pair=70 psig. ........................................................................................................................................ 48

Figure 19.—Continued. (cc) Uts=233.1 mph (202.5 knots), TT,davg=5.1 °C, Pair=0 psig. (dd) 

Uts=234.8 mph (204.0 knots), TT,davg=5.5 °C, Pair=70 psig. (ee) Uts=233.9 mph (203.3 knots), 

TT,davg=–0.4 °C, Pair=0 psig. (ff) Uts=234.0 mph (203.4 knots), TT,davg=–0.2 °C, Pair=70 psig. ........... 49

Figure 19.—Continued. (gg) Uts=232.8 mph (202.3 knots), TT,davg=–12.5 °C, Pair=0 psig. (hh) 

Uts=232.7 mph (202.2 knots), TT,davg=–12.5 °C, Pair=70 psig. (ii) Uts=233.6 mph (203.0 

knots), TT,davg=–24.7 °C, Pair=0 psig. (jj) Uts=232.5 mph (202.0 knots), TT,davg=–24.7 °C, 

Pair=70 psig. ........................................................................................................................................ 50

Figure 19.—Continued. (kk) Uts=254.7 mph (221.3 knots), TT,davg=4.9 °C, Pair=0 psig. (ll) 

Uts=254.5 mph (221.2 knots), TT,davg=4.8 °C, Pair=70 psig. (mm) Uts=253.9 mph (220.6 

knots), TT,davg=0.6 °C, Pair=0 psig. (nn) Uts=253.7 mph (220.5 knots), TT,davg=0.7 °C, Pair=70 

psig. .................................................................................................................................................... 51

Figure 19.—Continued. (oo) Uts=253.3 mph (220.1 knots), TT,davg=–11.4 °C, Pair=0 psig. (pp) 

Uts=253.3 mph (219.9 knots), TT,davg=–11.3 °C, Pair=70 psig. (qq) Uts=254.1 mph (220.8 

knots), TT,davg=–23.7 °C, Pair=0 psig. (rr) Uts=252.9 mph (219.7 knots), TT,davg=–23.7 °C, 

Pair=70 psig. ........................................................................................................................................ 52

Figure 19.—Continued. (ss) Uts=304.7 mph (264.8 knots), TT,davg=5.3 °C, Pair=0 psig. (tt) 

Uts=304.6 mph (264.7 knots), TT,davg=5.0 °C, Pair=70 psig. (uu) Uts=303.3 mph (264.6 

knots), TT,davg=–8.8 °C, Pair=0 psig. (vv) Uts=303.8 mph (264.0 knots), TT,davg=–8.9 °C, 

Pair=70 psig. ........................................................................................................................................ 53

Figure 19.—Concluded. (ww) Uts=302.9 mph (263.2 knots), TT,davg=–20.3 °C, Pair=0 psig. (xx) 

Uts=303.8 mph (264.0 knots), TT,davg=–20.2 °C, Pair=70 psig. ............................................................ 54

Figure 20.—Flow angle data from the IRT test section collected using the 9-ft survey rake 

mounted at 11 vertical positions over several tunnel runs. Nominal conditions are used to 

denote each data set. (a) Uts=50 mph (43 knots), Pair=0 psig. (b) Uts=50 mph (43 knots), 

Pair=70 psig. (c) Uts=100 mph (87 knots), Pair=0 psig. (d) Uts=100 mph (87 knots), Pair=70 

psig. .................................................................................................................................................... 56

Figure 20.—Continued. (e) Uts=115 mph (100 knots), Pair=0 psig. (f) Uts=115 mph (100 knots), 

Pair=70 psig. (g) Uts=150 mph (130 knots), Pair=0 psig. (h) Uts=150 mph (130 knots), 

Pair=70 psig. ........................................................................................................................................ 57
NASA/CR—2010-216733 vii 

Figure 20.—Continued. (i) Uts=175 mph (152 knots), Pair=0 psig. (j) Uts=175 mph (152 knots), 

Pair=5 psig. (k) Uts=175 mph (152 knots), Pair=20 psig. (l) Uts=175 mph (152 knots), Pair=70 

psig. .................................................................................................................................................... 58

Figure 20.—Continued. (m) Uts=200 mph (174 knots), Pair=0 psig. (n) Uts=200 mph (174 knots), 

Pair=70 psig. (o) Uts=250 mph (217 knots), Pair=0 psig. (p) Uts=250 mph (217 knots), Pair=70 

psig. .................................................................................................................................................... 59

Figure 20.—Continued. (q) Uts=225 mph (195 knots), Pair=0 psig. (r) Uts=225 mph (195 knots), 

Pair=20 psig. (s) Uts=225 mph (195 knots), Pair=70 psig. .................................................................. 60

Figure 20.—Concluded. (t) Uts=300 mph (261 knots), Pair=0 psig. (u) Uts=300 mph (261 knots), 

Pair=70 psig. (v) Uts=350 mph (304 knots), Pair=0 psig. (w) Uts=350 mph (304 knots), 

Pair=70 psig. ........................................................................................................................................ 61

Figure 21.—Average flow angularity vector magnitude in the core (center 4.5- by 3- ft area) of 

the IRT test section at various Pair settings. ........................................................................................ 62

Figure 22.—Summary of the axial and transverse turbulence levels in the IRT test section as 

measured by probes mounted on the 9-ft survey rake. The probes were arrayed laterally 

across the span of the rake as follows: Probe 1 was 31.8 in. from the inside wall, probe 2 

was 53.7 in. and probe 3 was 76.2 in. (a) Test section airspeed of 50 mph ....................................... 63

Figure 22.—Continued. (b) 100 mph .......................................................................................................... 64

Figure 22.—Continued. (c) 115 mph .......................................................................................................... 65

Figure 22.—Continued. (d) 150 mph .......................................................................................................... 66

Figure 22.—Continued. (e) 175 mph, 0 and 70 psi. .................................................................................... 67

Figure 22.—Continued. (f) 175 mph, 5 and 20 psi. .................................................................................... 68

Figure 22.—Concluded. (g) 200 mph. ........................................................................................................ 69

Figure D.1.—Total temperature calibration curves for the IRT based on the 2005 aero-thermal 

interim calibration data. ...................................................................................................................... 88

Figure D.2.—Total temperature calibration curve coefficients for Uts <50 knots. This information 

is used to interpolate slope and intercept values for spray bar air pressure settings between 0 

and 40 and between 40 and 70 psig. ................................................................................................... 89

Figure D.3.—Comparison of Mach number ratio distribution data across the tunnel at test section 

centerline. (a) Mnom= 0.066, Unom= 43 knots. (b) Mnom= 0.100, Unom= 65 knots. (c) Mnom= 

0.133, Unom= 87 knots. (d) Mnom= 0.152, Unom= 100 knots. ................................................................ 90

Figure D.3.—Continued. (e) Mnom= 0.200, Unom= 130 knots. (f) Mnom= 0.234, Unom= 150 knots. (g) 

Mnom= 0.266, Unom= 175 knots. (h) Mnom= 0.306, Unom= 200 knots. ................................................... 91

Figure D.3.—Continued. (i) Mnom= 0.336, Unom= 220 knots. (j) Mnom= 0.370, Unom= 240 knots. (k) 

Mnom= 0.402, Unom= 260 knots. (l) Mnom= 0.435, Unom= 280 knots. .................................................... 92

Figure D.3.—Concluded. (m) Mnom= 0.470, Unom= 300 knots. ................................................................... 93

Figure D.4.—IRT test section total temperature distribution data collected using the 2–D RTD 

Array. (a) Uts=49.02 mph (42.6 knots), TT,davg=–29.8 °C, Pair=0 psig. (b) Uts=55.2 mph (48.0 

knots), TT,davg=–29.8 °C, Pair=70 psig. (c) Uts=49.14 mph (42.7 knots), TT,davg=–17.6 °C, 

Pair=0 psig. (d) Uts=54.89 mph (47.7 knots), TT,davg=–17.5 °C, Pair=70 psig. ..................................... 94

Figure D.4.—Continued. (e) Uts=49.71 mph (43.2 knots), TT,davg=0.2 °C, Pair=0 psig. (f) Uts=49.94 

mph (43.4 knots), TT,davg=0.3 °C, Pair=40 psig. (g) Uts=49.59 mph (43.1 knots), TT,davg=0.2 

°C, Pair=70 psig. ................................................................................................................................. 95

Figure D.4.—Continued. (h) Uts=116.46 mph (101.2 knots), TT,davg=–28.4 °C, Pair=0 psig. (i) 

Uts=118.76 mph (103.2 knots), TT,davg=–28.4 °C, Pair=70 psig. (j) Uts=116.46 mph (101.2 

knots), TT,davg=–16.3 °C, Pair=0 psig. (k) Uts=118.30 mph (102.8 knots), TT,davg=–16.3 °C, 

Pair=70 psig. ........................................................................................................................................ 96

Figure D.4.—Continued. (l) Uts=116.57 mph (101.3 knots), TT,davg=–4.3 °C, Pair=0 psig. (m) 

Uts=116.80 mph (101.5 knots), TT,davg=–4.2 °C, Pair=70 psig. (n) Uts=116.34 mph (101.1 

knots), TT,davg=1.3 °C, Pair=0 psig. (o) Uts=116.34 mph (101.1 knots), TT,davg=1.3 °C, Pair=70 

psig. .................................................................................................................................................... 97
NASA/CR—2010-216733 viii 

Figure D.4.—Continued. (p) Uts=151.91 mph (132.01 knots), TT,davg=–15.4 °C, Pair=0 psig. (q) 

Uts=174.53 mph (151.6 knots), TT,davg=0.06 °C, Pair=0 psig. .............................................................. 98

Figure D.4.—Continued. (r) Uts=233.38 mph (202.8 knots), TT,davg=–24.4 °C, Pair=0 psig. (s) 

Uts=233.61 mph (203.0 knots), TT,davg=–24.7 °C, Pair=70 psig. (t) Uts=235.10 mph (204.3 

knots), TT,davg=–12.3 °C, Pair=0 psig. (u) Uts=234.87 mph (204.1 knots), TT,davg=–12.4 °C, 

Pair=70 psig. ........................................................................................................................................ 99

Figure D.4.—Continued. (v) Uts=230.5 mph (200.3 knots), TT,davg=0.0 °C, Pair=0 psig. (w) 

Uts=231.19 mph (200.9 knots), TT,davg=0.0 °C, Pair=70 psig. ........................................................... 100

Figure D.4.—Continued. (x) Uts=253.11 mph (219.95 knots), TT,davg=0.67 °C, Pair=0 psig. (y) 

Uts=292.07 mph (253.8 knots), TT,davg=0.8 °C, Pair=0 psig. (z) Uts=291.61 mph (253.4 

knots), TT,davg=0.7 °C, Pair=70 psig. .................................................................................................. 101

Figure D.4.—Concluded. (aa) Uts=338.33 mph (294.0 knots), TT,davg=–17.4 °C, Pair=0 psig. (bb) 

Uts=338.1 mph (293.8 knots), TT,davg=–17.7 °C, Pair=70 psig. (cc) Uts=338.9 mph (294.5 

knots), TT,davg=–5.5 °C, Pair=0 psig. (dd) Uts=339.6 mph (295.1 knots), TT,davg=–5.6 °C, 

Pair=70 psig. ...................................................................................................................................... 102

Figure D.5.—IRT test section total temperature standard deviation data collected using the 2–D 

RTD Array at the center line of the RTD Array. .............................................................................. 103

Figure D.5.—Concluded. IRT test section total temperature standard deviation data collected 

using the 2–D RTD Array for the full RTD Array. .......................................................................... 104

Figure D.6.—Comparison of test section centerline flow angularity vectors from the 2004 full 

calibration and 2005 interim calibration at various Pair settings. (a) Mnom = 0.066, Vnom= 43 

knots. (b) Mnom= 0.100, Vnom= 65 knots. (c) Mnom= 0.133, Vnom= 87 knots. (d) Mnom = 0.152, 

Vnom = 100 knots. ............................................................................................................................... 105

Figure D.6.—Continued. (e) Mnom = 0.200, Vnom= 130 knots. (f) Mnom= 0.234, Vnom= 150 knots. (g) 

Mnom= 0.266, Vnom= 178 knots. (h) Mnom= 0.306, Vnom= 200 knots. .................................................. 106

Figure D.6.—Continued. (i) Mnom = 0.336, Vnom= 220 knots. (j) Mnom= 0.370, Vnom= 240 knots. (k) 

Mnom= 0.402, Vnom= 260 knots. (l) Mnom= 0.435, Vnom= 280 knots. .................................................. 107

Figure D.6.—Concluded. (m) Mnom = 0.470, Vnom= 300 knots. ................................................................. 108

Figure D.7.—Comparison of average data from flow angularity measurements made in the IRT 

test section (horizontal survey at centerline) during the 2004 and 2005 test entries. (a) Fullspan data (all eleven probes included). (b) Core data (center seven probes only). .......................... 109

Figure D.8.—Comparison of standard deviation data from flow angularity measurements made in 

the IRT test section (horizontal survey at centerline) during the 2004 and 2005 test entries. 

(a) Full-span data (all eleven probes included). (b) Core data (center seven probes only). ............. 110

Figure D.9.—Comparison of 2004 and 2005 turbulence data (horizontal survey at centerline). ............. 111

Figure D.10.—Comparison of the aero-thermal calibration curves created based on the 2004 and 

2005 test results. (a): Mach number calibration curve comparison with 2004 and 2005 test 

data. .................................................................................................................................................. 112

Figure D.10.—Continued. (b) Static Pressure calibration curves for the IRT based on the 2004 

and 2005 calibration data. ................................................................................................................ 113

Figure D.10.—Concluded. (c) Total temperature calibration curves for the IRT based on the 2004 

and 2005 calibration data. ................................................................................................................ 114
Aero-Thermal Calibration of the NASA Glenn Icing Research 

Tunnel (2004 and 2005 Tests) 

E. Allen Arrington and Christine M. Pastor 

Sierra Lobo, Inc. 

Cleveland, Ohio 44135 

Jose C. Gonsalez and Monroe R. Curry III 

QSS Group, Inc. 

Cleveland, Ohio 44135 

Summary 

A full aero-thermal calibration of the NASA Glenn Icing Research Tunnel was completed in 2004 

following the replacement of the inlet guide vanes upstream of the tunnel drive system and improvement 

to the facility total temperature instrumentation. This calibration test provided data used to fully document 

the aero-thermal flow quality in the IRT test section and to construct calibration curves for the operation 

of the IRT. The 2004 test was also the first to use the 2–D RTD array, an improved total temperature 

calibration measurement platform. 

Introduction 

Updating and maintaining the icing cloud and aero-thermal calibration and flow quality data are 

critical tasks in the NASA Glenn Icing Research Tunnel (IRT). Establishing and maintaining the 

calibrations is the only way to insure that correct test section operating conditions are being set for a given 

set of facility control inputs. Without knowledge of the calibration and flow quality data, it is impossible 

to provide high-quality test data for customers and researchers. To insure that the data quality from the 

IRT is always at the highest standards, IRT staff and wind tunnel calibration personnel have developed 

and continue to refine the calibration methodologies used. 

The aero-thermal calibration test program provides the calibration relationships needed to accurately 

determine the test section operating conditions for total pressure, airspeed (comparable Mach number) 

and total temperature. Once these calibrated conditions are determined, all other aero-thermal operating 

conditions can be calculated. In addition to the calibration information, detailed flow quality data (spatial 

uniformity) is also collected. The collected calibration and flow quality data provide a complete 

description of the aero-thermal operation of the IRT. 

The current aero-thermal calibration methodology and instrumentation has evolved based on lessons 

learned by the calibration team over the past 15 years, input from IRT customers, in-house research staff, 

IRT facility personnel, and recommendations from References 1 and 2. Reference 1 defines a time table 

and level of calibration that should be conducted following modifications to an icing tunnel. 

Correspondingly, Reference 1 recommends that short check calibrations should be conducted multiple 

times each year to insure repeatable operation of the wind tunnel. Based on these recommendations, the 

following aero-thermal calibration tests have been defined: 

• Full calibration: Includes collection of aero-thermal calibration data over the entire operating 

range of the IRT as well as a complete mapping of the IRT flow quality at one cross section of the 

test section, typically at the centerline of the turntable. A full calibration should be conducted 

every 5 years or following a major facility modification. The full calibration provides all of the 

information needed to develop the aero-thermal calibration relationships, or calibration curves, 

for the facility and provides extensive detailed information on the aero-thermal flow quality 

characteristics of the test section. 

NASA/CR—2010-216733 1
• Interim calibration: Includes collection of aero-thermal calibration data over the entire operating 

range of the IRT, however, data is only required from the test section centerline. An interim 

calibration is conducted one year following a full calibration. The interim calibration is used to 

validate the data collected during the full calibration, therefore, if the full calibration and interim 

calibration data match then the facility is operating properly, however if not, then an investigation 

must commence immediately to rectify the discrepancies. 

• Check calibration: Ideally a one-day test using a very simple rake. Check calibration testing 

covers the entire operating range, but the test matrix is not as extensive as a full or interim 

calibration. This calibration should be conducted 2 to 3 times per year. The check calibration data 

is used in a statistical process control analysis to monitor the stability and repeatability of the 

facility operation. 

This paper provides a detailed description of the full aero-thermal calibration conducted in the IRT in 

2004. This calibration was required due to replacement of the drive system inlet guide vanes, which is 

detailed in the subsequent Description of Facility section. The 2004 full calibration was also needed to 

update the total temperature calibration based on improvements to facility instrumentation (Ref. 3). 

Previous to the 2004 calibration, the most recent full calibration was completed in 2000 following the 

replacement of the facility heat exchanger (Ref. 4). 

The goals for the 2004 aero-thermal full calibration are listed below: 

• Qualify for operation the 2–D RTD array as the new measurement platform for total temperature 

calibration and flow quality data in the IRT test section. 

• Complete data collection and construct aero-thermal calibration curves for the IRT. The three 

calibration relationships used are total pressure, Mach number and total temperature. 

• Complete detailed flow quality surveys using the 9-ft survey rake and 2–D RTD array. 

While the bulk of this report focuses on the full aero-thermal calibration conducted in 2004, a 

discussion of the 2005 interim calibration is included in an appendix. The 2005 interim calibration 

provided additional information on the total temperature calibration of the IRT and allowed an 

opportunity to verify additional planned improvements to the total temperature calibration.1

Symbols 

a speed of sound, ft/sec

A0, A1, A2 Hot wire curve fit coefficients 

C0, C1, C2 RTD probe total temperature flow recovery curve fit coefficients 

Cα Pitch angle pressure coefficient 

Cβ Yaw angle pressure coefficient 

CTT,0, CTT,1 Test section total temperature calibration curve fit coefficients 

Co Total pressure coefficient (9-ft rake pressure probes) 

Cq Static pressure coefficient (9-ft rake pressure probes) 

Ct Total temperature recovery coefficient (9-ft rake total temperature probes) 

E Hot wire anemometer output voltage, volts 

K0 to K2 Flow angle prediction coefficients, degrees 

l characteristic length, ft 

M Mach number 

N Number of data points 

P Pressure, psia 

 

1

Check calibration data are collected as part of both full and interim calibration test entries, as well as during 

dedicated check calibration test entries. However, the check calibration tests are not covered in this report. 

NASA/CR—2010-216733 2
P1 to P9 Flow angle probe pressures, psia 

Pair Spray bar air pressure, psig 

Pavg Average of P1, P2, P3, and P4, psia 

Prat Ratio of static to total pressure 

PS Static pressure, psia (psf) 

PS,rake Average of P6, P7, P8, and P9 (probe static pressure), psia 

PT Total pressure, psia (psf) 

PT,rake Equal to P5 (probe total pressure), psia 

q Dynamic pressure, psi 

R Specific gas constant for air, 1716 ft2

/(sec2 °R) 

Re Reynolds number 

RSS Root-sum-squared uncertainty estimate 

T Temperature, °C (°F) 

TIu Axial turbulence intensity, percent 

TIv Transverse turbulence intensity, percent 

TS Static temperature, °C (°F) 

TT Total temperature, °C (°F) 

TT,davg Average of the 24 D-corner total temperature measurements, °C (°F) 

TT,dc D-corner total temperature measurement, °C (°F) 

U Velocity, knots (ft/sec, mph) 

u Hot wire axial velocity, ft/sec 

v Hot wire transverse velocity, ft/sec 

Vm Flow angle resultant vector magnitude, degrees 

Vφ Flow angle resultant vector orientation, degrees 

X Axial coordinate with axis origin at bellmouth/test section weld seam, in. 

Y Spanwise coordinate with axis origin at the test section inner wall, in. 

Z Vertical coordinate with axis origin at the test section floor, in. 

α Pitch flow angle, degrees 

β Yaw flow angle, degrees 

γ Ratio of specific heats, 1.4 

Δ Measured pitch or yaw angle offset, degrees 

ΔTDC D-Corner average temperature, °C (°F) 

σ Standard deviation 

σTT,DC Standard deviation of the D-corner total temperature, °C (°F) 

θ Hot wire sensor inclination angle, degrees 

φ Parameter used to calculate flow angle resultant vector orientation, degrees 

Subscripts 

arm Pertaining to the 6 degree-of-freedom digitizing arm 

array Pertaining to the 2–D RTD array 

avg Average 

bm Bellmouth conditions 

bm,north Measured by the north bellmouth rake 

bm,south Measured by the south bellmouth rake 

eff Pertaining to hot wire effective velocities 

HW Pertaining to a hot wire probe 

i Data point index or hot wire sensor index 

level Pertaining to a digital level 

NASA/CR—2010-216733 3
local Pertaining to parameters in the test section after all probe calibration coefficients have been 

taken into account 

nom Nominal test section condition or setting 

rake Pertaining to the 9-ft survey rake 

ruler Pertaining to a tape measure or ruler 

ts Test section conditions (based on calibration data) 

Description of Facility 

The NASA Glenn Icing Research Tunnel (IRT) is a closed-loop atmospheric tunnel illustrated in 

Figure 1. The airflow is driven by a 25-ft diameter twelve blade fan that is powered by a 5000-hp electric 

motor. The test section, shown at the bottom center of Figure 1, is 6-ft high, 9-ft wide, and 20-ft long, 

with no divergence along any of the test section surfaces. The velocity in the empty test section can be 

varied between 25 and 340 knots at –17.8 °C (37 to 575 ft/sec at 0 °F). A pair of heat exchangers located 

in the C-D leg of the tunnel loop, seen at the left of Figure 1, allow for total temperature settings as low as 

–30 °C. A set of ten horizontally oriented spray bars, located in the settling chamber at the bellmouth inlet 

prior to the test section, inject atomized water into the airflow to create icing conditions. The IRT has 

been in service since 1944 and has undergone several major upgrades and modifications over the years to 

insure that it remains the lead ground test facility for icing research (Refs. 5 and 6). In 2004, the inlet 

guide vanes (IGVs) upstream of the drive fan were replaced, primarily for structural reasons, but also to 

improve aerodynamic performance. Figure 2 shows the final configuration of the replacement IGVs 

upstream of the IRT drive fan prior to being painted. The view in Figure 2 looks downstream from Corner 

B showing the inlet and outlet guide vanes, fan blades, and floor drain. During the downtime for the IGV 

replacement, the new facility instrumentation for total temperature measurement was installed in Dcorner, consisting of 24 RTD probes. More information about the IRT is located in Reference 7. 

Figure 1.—Layout of the NASA Glenn Icing Research Tunnel. 

NASA/CR—2010-216733 4
Figure 2.—Installation photo of the replacement inlet guide vanes (IGV) 

upstream of the IRT drive fan. 

Test Hardware, Instrumentation and Data Systems 

The 9-ft survey rake and the 2–D RTD array are two pieces of test hardware used to complete the 

aero-thermal calibration of the IRT. The 9-ft survey rake is used to collect pressure, flow angle, 

turbulence calibration, and flow quality data. The 2–D RTD array is used to collect total temperature 

calibration and flow quality data2

. Detailed descriptions of the hardware, instrumentation and data 

systems are included in this section. 

9-ft Survey Rake 

Figure 3 shows the 9-ft horizontal survey rake installed in the IRT test section at the vertical 

centerline, a height of Z=36-in. The flow angle pressure probes, total temperature probes, and hot wire 

probes mounted on the 9-ft survey rake in Figure 3, will be discussed in more detail later in this section. 

The reference axis is also defined in Figure 3 depicting the vertical direction as the Z-axis, the horizontal 

direction as the Y-direction, and the axial direction as the X-axis. The origin of the left-handed X-Y-Z

coordinate system is as follows: the X axis origin is at the bellmouth/test section weld seam as seen in 

Figure 4, the Y axis origin is at the inner wall as seen in Figure 4, and the Z axis origin is at the floor as 

seen in Figure 3. This rake is used to collect calibration and flow quality data for pressure, airspeed, flow 

angle and turbulence in the IRT test section. 

Figure 4 illustrates the plan view of the IRT test section with the survey rake installed. The rake was 

installed at an axial position of X=179.3-in. This was also the axial location of the cross-sectional plane 

surveyed during the 2004 test program as well as the test program in 1997. Figure 5 shows the plan and 

front views of the survey rake. The following illustration, Figure 6 details a cross-section of the survey 

rake with exploded views of the flow angle pressure probe ports. 

 

2

Prior to the 2004 calibration, the 9-ft survey rake was used to collect all calibration data, including total temperature 

data. Starting in 2004, an improved total temperature measurement platform, the 2–D RTD Array, was introduced 

for calibration and flow quality data collection. 

NASA/CR—2010-216733 5
Figure 3.—Photograph of the 9-ft horizontal survey rake installed in the IRT test section. 

Figure 4.—Plan view of the IRT test section showing the installed location of the 9-ft survey rake. 

NASA/CR—2010-216733 6
Figure 5.—Plan and elevation views of the 9-ft survey rake. 

Figure 6.—Cross section of the 9-ft survey rake showing flow probes details. 

The instrumentation mounted on the 9-ft survey rake includes eleven flow angle pressure probes, 

eleven total temperature thermocouple probes, and three thermal anemometry probes. The spanwise, Y

positions of all probes with respect to the test section inner wall are defined in Figure 5. The survey rake 

was supported at both ends by bolting the rake to wall mounted support plates as seen in Figure 3. These 

support plates are 6-ft tall and have a bolt pattern that allows the rake to be positioned vertically every 

NASA/CR—2010-216733 7
6-in. above or below the test section centerline. A vertical strut, shown in Figure 3, was also used to 

support the rake in the center. As seen in Figure 6, the main body of the rake is formed by two aluminum 

I-beams and thin aluminum plates riveted together. Thin sheets of aluminum were bent to form the 

leading and tailing edge of the rake body. 

The flow angle pressure probes are bolted to the rake body I-beams as illustrated in Figure 6. These 

probes are of hemi-spherical head design, bottom left Figure 6, and have five total pressure ports in the 

head. The bottom center Figure 6 shows the center port in the hemi-spherical head design which measures 

total pressure, and the four circumferential total pressure ports which measure pitch and yaw angle. Four 

static pressure ports are located downstream of the head depicted bottom right Figure 6. Top Figure 6, 

shows the heads of the probes located 20.75-in. upstream of the rake leading edge and the static pressure 

taps on each probe located 5.75-in. downstream of the head. The probes were calibrated for Mach 

numbers between 0.1 and 0.6 in the NASA Glenn 3.5-in. diameter free jet calibration facility. Details of 

the probe calibration can be found in Reference 8. 

Aspirated total temperature type T thermocouple probes with copper/constantan wires were used to 

monitor test section temperature. These probes are also depicted in Figure 3, Figure 5, and Figure 6. The 

probes are mounted to the bottom surface of the survey rake with the tips of the probes approximately 

0.5-in. upstream of the rake leading edge and 2.0-in. below the rake centerline. The total temperature 

probes were calibrated for total temperature recovery in the NASA Glenn 3.5-in. diameter free jet 

calibration facility. 

Three single sensor and dual sensor (cross-wire) hot-wire probes with 0.00015-in. diameter tungsten 

wires, were used to measure the turbulence levels in the test section. The probes were mounted to the 

upper surface of the rake body as shown in Figure 3, Figure 5, and Figure 6. The top of Figure 6 more 

clearly illustrates the probe tips were located 9.75 in. upstream of the rake body leading edge and 3.0-in. 

above the rake centerline. 

2–D RTD Array 

The 2–D RTD Array was designed and built to provide a more accurate total temperature calibration 

and temperature mapping of the test section, with the benefit of reducing overall test time and cost. The 

total temperature probes used on the array are 4-wire resistive temperature device (RTD) with a ceramic 

capsule sensor. Figure 7 shows the details of the RTD probe. See Reference 3 for a discussion of the 

thought process used to arrive at the design of the array and selection of the probes. 

Reference 1 was utilized to determine the number and arrangement of probes on the array. Figure 8 

shows the resulting layout is a 7-by-7 grid with the probes spaced at 0, ±25, ±50 and ±75 percent of the 

test section dimension from the horizontal and vertical centerlines. The intersections of the dotted lines in 

this figure indicate the probe positions. The 7-by-7 grid arrangement provided the test section coverage 

required without incurring an unreasonable cost for the RTD probes. 

Figure 7.—Details of RTD probes used on the 2–D RTD array (dimensions in inches). 

NASA/CR—2010-216733 8
Figure 9 shows the 2–D RTD array installed in the IRT test section (views looking downstream and 

upstream). The structure of the 2–D RTD array includes a set of seven vertical members and horizontal 

members. Each vertical member houses seven RTD probes and instrumentation wiring, while the four 

horizontal members provide structural rigidity. The vertical members seen in Figure 9 are 0.75 in. thick 

by 4 in. wide and have a round leading edge. Figure 10 illustrates the cross section of an array vertical 

member assembly. Also, Figure 10 shows how the tips of the RTDs extend 7 in. forward of the leading 

edge, and the 0.75 in. deep channel that was attached to the trailing edge to house the instrumentation 

lines. The horizontal members seen in Figure 9 are 0.5 in. thick by 3 in. wide and also have a round 

leading edge. The array was secured to the test section walls, floor and ceiling, with allowances made to 

account for thermal expansion of the array. The tips of the RTD probes were positioned 194.5 in. 

downstream of the tunnel datum line or 121.5 in. from the inlet of the test section (the thermocouples on 

the 9-ft survey rake were 199.5 in. downstream of the datum line). 

Figure 8.—7x7 grid arrangement of the RTD probes in the IRT 

test section. 

Figure 9.—RTD Array installed in the IRT test section. (a) View looking downstream. (b) View looking upstream. 

(a) (b)

Vertical 

member

Horizontal 

member

RTD Probes

NASA/CR—2010-216733 9
Figure 10.—RTD array vertical member cross sectional details. 

Figure 11.—Typical total temperature flow recovery curve 

for the 2–D RTD array probes.

Figure 12.—Total temperature flow recovery data for 

RTD probes 1 through 23 with average curve fit.

A portion of the RTD probes (serial numbers 1 through 23) were calibrated for temperature flow 

recovery in the 3.5-in. Diameter Free Jet Facility. The recovery testing was conducted to determine how 

well the probes sense the total temperature over a range of airflow conditions. During the testing, the free 

jet air temperature was set to a nominal condition of about 520 °R (60.33 °F) and held constant while the 

free jet Mach number was varied from 0 to 0.5. Figure 11 shows the total temperature recovery results for 

a typical probe (serial number 15) with its individual flow recovery curve. The recovery curves were 

subsequently used to correct data collected in the IRT. The 20 RTD probes that were individually 

calibrated for flow recovery were mounted in the center of the 2–D RTD array. The probes not measured 

directly for their recovery characteristics were mounted around the periphery of the array and used the 

average recovery from all the calibrated probes, seen in Figure 12. The probes have an absolute 

measurement uncertainty of ±0.10 °C (±0.18 °F) for a probe that was individually calibrated for total 

temperature flow recovery or ±0.16 °C (±0.29 °F) for a probe that was not specifically calibrated for 

recovery (uses the average recovery values). The details of this uncertainty analysis are found in 

Appendix A. A summary of the flow recovery correction coefficients is provided in Appendix B. 

Facility Instrumentation 

The following standard facility instrumentation was used during this test program. 

1) Bellmouth pitot-static probes: A probe is mounted on the south and north wall near the exit of the 

bellmouth. Transducers associated with these probes measure the total and differential pressures. 

These measurements are used to arrive at PT,hm and PS,bm. The location of these probes is shown in 

Figure 4, as the outer (north) probe and the inner (south) probe. 

NASA/CR—2010-216733 10
2) Total temperature probes in D-corner: Twenty-four RTD probes are arrayed on the leading edges 

of the turning vanes in D-corner to measure the temperature profile exiting the facility cooler. The 

distribution of the measured temperatures and the average of these twenty-four probes, TT,davg, is 

used to monitor the operation of the heat exchanger. The locations of these probes are shown in 

Figure 13. 

Steady-State Data Acquisition System 

Real-time steady-state data acquisition and data display are provided by the NASA Glenn Escort 

Alpha system. This is the standard data acquisition and data display system used in the large test facilities 

at NASA Glenn. The system accommodates inputs from the Electronically Scanned Pressure System 

(ESP), the facility distributed process control system, and any analog devices such as thermocouples, 

RTDs and pressure transducers. This system records all steady-state pressures and temperatures from the 

9-ft survey rake, 2–D RTD array, the tunnel bellmouth pitot-static probes, and the total temperature 

probes in the D-corner. It also records important facility parameters such as fan rotational speed and spray 

bar air pressure. The Escort program, designated D046, used for data collection, online data analysis and 

display was developed specifically for the aero-thermal calibration testing. 

The Electronically Scanned Pressure system used during this test program utilized plug-in modules 

that each contained 32 individual transducers with individual ports. Each transducer/port can be addressed 

separately and scanned at a rate of 10,000 ports per second. Calibration of all ESP transducers was 

performed automatically by the system at least every 2 hr. For this test program, ±5 psid modules were 

used. The accuracy in the pressure measurements made with the ESP system is 0.1 percent of full-scale of 

the module, or 0.005 psia for the ±5 psid modules used. 

Thermal Anemometry Instrumentation 

Commercially available constant temperature hot wire anemometers were used for the turbulence 

intensity measurements. Three anemometers were required for the three single sensor probes used. Each 

anemometer was equipped with its own signal conditioner for low-pass filtering, DC offsetting, and 

amplifying. A commercial off the shelf (COTS) 3-channel 12-bit analog to digital converter with an input 

range of ±5 V was used to digitize the signal conditioned anemometer bridge voltages. A COTS personal 

computer with COTS software was used to control the data acquisition process and to store the digitized 

data on computer hard disk. 

Figure 13.—Layout of the RTD probes mounted on the D corner turning vanes.

NASA/CR—2010-216733 11
The hot wire probes were calibrated in place using the velocities measured by the flow angle pressure 

probes. Data was acquired at small enough velocity increments to generate calibration curves for the hot 

wires. The probes were generally used from 50 mph to 175 mph, i.e., within the incompressible range. 

Over this speed range, the bridge output voltage varied between 1.3 and 1.8 V. 

Test Procedures and Execution 

This section and the subsequent subsections describe aspects of the test planning, setup and execution 

that were critical to the success of the aero-thermal calibration in the IRT. As indicated in Reference 2, it 

is important that tunnel calibration tests are given proper attention to detail and with respect to project 

management. Also, customer requirements must be considered when developing the tunnel calibration 

test plan, test matrix and hardware. This specific test considered needs from both internal and external 

customers requiring improved total temperature measurement accuracy, correspondingly improving total 

temperature calibration. These requirements had a direct impact on the development of the new test 

hardware as well as the test planning and execution. 

Test Matrix 

The recommendations made following the 2000 aero-thermal calibration (Ref. 4) were implemented 

for the 2004 full aero-thermal calibration. The “aero” and “thermal” data collection were separated; the 

9-ft survey rake was used to collect the “aero” data at ambient temperature conditions and the 2–D RTD 

array was used to collect all temperature calibration data. The “aero” test matrix, seen in Table 1 was 

completed in 3 test days. Only 1 day was needed to complete the RTD array test matrix, seen in Table 2. 

Table 2 also shows the update in how the temperature calibration test matrix was constructed. In 

previous tests, the total temperature calibration test matrix was set up to cover the operating range of the 

IRT by selecting test points at minimum and maximum conditions and at regular intervals in between 

based on the D-corner total temperature. Starting in 2004, the temperature calibration test matrix was 

based on the test section static temperature as outlined in the guidelines provided in Reference 1. The 

actual test points are still the test section airspeed and D-corner total temperature, but the total 

temperature setting is determined from the airspeed and desired test section static temperature. The order 

of test points were arranged from highest to lowest in order to more efficiently use the available test time. 

TABLE 1.—TEST MATRIX FOR THE IRT AERO-THERMAL 

CALIBRATION USING THE 9-ft SURVEY RAKE (2004 TEST)

Rake height, 

Y, 

in. 

Total temperature, 

TT,davg, 

°C 

Test section airspeed, 

Uts, 

knots 

Spraybar air 

pressure, 

Pair, 

psig 

36 Ambient 

(nominally 5 °C) 

0, 43, 65, 87, 100, 130, 152, 173, 200, 217, 240, 260, 282, 304, 339 0, 70 

6, 12, 18, 24, 30, 

42, 48, 54, 60, 66 

ambient 0, 43, 87, 100, 130, 173, 200, 217, 260, 304 0 

NASA/CR—2010-216733 12
TABLE 2.—TEMPERATURE CALIBRATION TEST MATRIX USING THE 2–D RTD ARRAY (2004 TEST). 

Test section 

static 

temperature, 

TS,ts, 

°C 

Test 

section 

airspeed, 

Uts, 

knots 

D-corner total 

temperature, 

TT,davg, 

°C 

Spraybar 

air 

pressure, 

Pair, 

psig 

Test section 

static 

temperature, 

TS,ts, 

°C 

Test 

section 

airspeed, 

Uts, 

knots 

D-corner total 

temperature, 

TT,davg, 

°C 

Spraybar 

air 

pressure, 

Pair, 

psig 

–30 43 –29.8 0, 70 –5.6 43 –5.3 0, 70 

–30 100 –28.7 0, 70 –5.6 100 –4.2 0, 70 

–30 130 –27.8 0, 70 –5.6 130 –3.3 0, 70 

–30 200 –24.5 0, 70 –5.6 200 –0.3 0, 70 

–30 217 –23.8 0, 70 –5.6 217 0.7 0, 70 

–30 304 –17.8 0, 70 0 43 0.2 0, 70 

–17.8 43 –17.6 0, 70 0 100 1.3 0, 70 

–17.8 100 –16.4 0, 70 0 130 2.2 0, 70 

–17.8 130 –15.6 0, 70 –0.6 200 4.7 0, 70 

–17.8 200 –12.5 0, 70 –1.7 217 4.6 0, 70 

–17.8 217 –11.6 0, 70 –7.2 304 4.9 0, 70 

–17.8 304 –5.6 0, 70 4.4 43 4.7 0, 70 

 3.3 100 4.7 0, 70 

One area where the test matrices deviate from the recommendations in Reference 1 was at the higher 

airspeed and temperature settings because it was not possible to set static temperatures of 4.4 °C at 174 

and 260 knots. The total temperatures were 8.3 and 13.3 °C, respectively, since temperature control is not 

available at total temperature settings above about 4 °C). Therefore, the total temperature (D-corner 

temperature) was set at 5 °C with the lower related static temperatures in the test section as listed in 

Table 2. 

Test points were added to the 2–D RTD Array test matrix to cover static temperature conditions 

around the freezing point, 0 °C, since this is an important operating range for the icing community. Test 

points were also added to the test matrices corresponding to airspeed settings of 100, 200 and 300 knots 

so that flow quality information would be available at these specific settings. 

Test Procedures 

For the 2004 test entry, testing using the 9-ft survey rake was completed in May. Testing using the 

2–D RTD array was delayed until July due to additional time allowed for icing cloud calibration. 

The first test configuration was with the 9-ft survey rake positioned at the test section centerline. Prior 

to the first tunnel run, the flow angle probe, pitch and yaw offset angles were measured with respect to the 

tunnel centerline using a six-degree of freedom digitizing arm. Providing a check, in addition to these 

measurements pitch and yaw offset angles were verified with a digital level, tape measure or ruler. The 

combination of these measurements formed the baseline pitch and yaw offset angles. For each change in 

Z, vertical rake height change, the flow angle probe pitch and yaw offset angles were measured again 

with the digital level and tape measure and compared to the baseline values. All changes from the 

baseline were recorded and were later used to correct the pitch and yaw angles generated by the data 

reduction process. Table 3 shows the typical test conditions encountered during the 9-ft survey rake 

testing. The values listed in Table 3 are the average of all readings collected for each nominal setting 

across all test runs. It is important to note that all testing using the 9-ft survey rake was conducted at 

temperatures of approximately 4 °C (TT,bm ≥ 4 °C). This is done to avoid having the tips of the pressure 

probes freeze during tunnel operation and to decrease the possibility of getting water in the 

instrumentation lines. 

NASA/CR—2010-216733 13
TABLE 3.—TYPICAL TEST CONDITIONS IN THE IRT DURING THE 9-ft SURVEY 

RAKE TESTING AS PART OF THE 2004 AERO-THERMAL CALIBRATION 

Test 

section 

airspeed, 

Uts

(knots) 

Spray bar 

air 

pressure, 

Pair

(psig) 

Bellmouth 

Mach 

number, 

Mbm 

Bellmouth 

total 

pressure, 

PT,bm

(psia) 

Bellmouth 

static 

pressure, 

PS,bm

(psia) 

D-corner average total 

temperature, 

TT,davg 

 (°R) (°C) 

Drive 

system 

fan speed 

(rpm) 

43.5 0 0.067 14.308 14.264 510.3 5.4 76 

51.3 70 0.079 14.315 14.252 502.7 6.1 76 

64.5 0 0.099 14.310 14.213 505.7 7.8 109 

70.1 70 0.107 14.314 14.200 507.1 8.6 109 

87.3 0 0.134 14.308 14.131 507.9 9.0 143 

90.4 70 0.139 14.311 14.120 504.1 6.9 143 

99.4 0 0.153 14.305 14.074 504.7 7.2 163 

101.6 70 0.156 14.310 14.070 505.7 7.8 163 

131.0 0 0.201 14.299 13.902 507.5 8.8 212 

131.5 70 0.202 14.304 13.905 508.0 9.0 212 

150.9 0 0.232 14.296 13.770 505.8 7.9 242 

150.6 5 0.232 14.300 13.775 506.0 8.0 242 

151.1 20 0.233 14.301 13.771 503.9 6.8 242 

152.1 70 0.234 14.301 13.768 507.3 8.7 242 

172.2 0 0.265 14.299 13.618 507.2 8.6 275 

173.4 70 0.268 14.306 13.611 503.9 6.8 275 

198.8 0 0.306 14.290 13.394 510.5 10.4 315 

200.1 70 0.309 14.295 13.381 507.5 8.8 315 

216.4 0 0.334 14.290 13.230 509.1 9.7 342 

218.6 70 0.337 14.293 13.212 509.3 9.8 342 

237.3 0 0.365 14.291 13.037 515.1 13.0 373 

237.9 70 0.366 14.294 13.032 514.4 12.6 373 

257.6 0 0.397 14.284 12.817 515.4 13.2 404 

259.0 70 0.399 14.287 12.803 515.0 13.0 404 

280.3 0 0.430 14.280 12.573 521.1 16.4 439 

281.8 70 0.434 14.283 12.552 519.0 15.2 439 

300.9 0 0.466 14.270 12.296 515.2 13.1 473 

302.9 70 0.468 14.271 12.283 518.0 14.6 472 

The Escort system was used to collect data for both the 9-ft survey rake and the 2–D RTD array. For 

each test condition, two to three readings were taken at a scan rate of one per second, for 20 sec, and 

averaged to generate the final reading. 

All three probe types were recorded simultaneously by individual analog-to-digital converter data 

acquisition boards integrated into a personal computer. The data acquisition boards were configured to 

sample three or six channels simultaneously, depending on whether a single sensor or dual sensor X probe 

was used. The sensors were set to acquire data at a rate of 10,000 samples per second and 100,000 data 

points per channel. This is the equivalence of 10 sec of data for each channel. The signal conditioners for 

each hot wire probe were configured to a low pass filter at 4000 Hz, DC offset of –1 V, and a gain of 5. 

The addition of the 2–D RTD array increased efficiency of gathering total temperature calibration and 

flow quality data. The array is mounted at the test section test plane and requires minimal documentation 

of axial station and probe measurement prior to operation. Unlike the 9 ft survey rake no detailed 

alignment procedure of the RTD probes is required when using the 2–D RTD array. Checks of the RTD 

NASA/CR—2010-216733 14
array instrumentation connections and structure are completed prior to the tunnel run. The most efficient 

way to complete the RTD array test matrix is to start at the coldest bellmouth total temperature condition 

and work through the matrix toward the higher conditions. This is done because it is faster and easier to 

change airspeed settings at a given total temperature, rather than hold the airspeed constant and change 

temperature. The order of test points for the RTD array test matrix, seen in Table 2, is listed by increasing 

total temperature. 

Data Reduction and Analysis 

This section provides a detailed description of the steps taken in the reduction and analysis of all data 

collected using the 9-ft survey rake and the 2–D RTD array. Included in this section are instrumentation 

calibrations, data corrections and conversion equations. The conversion equations were used to convert 

the measured pressures and temperatures into Mach number, flow angle and turbulence data. For all of 

the equations listed, the specific heat ratio, γ, is 1.4 and the specific gas constant for air, R, is equal to 

1716 ft2

 /(sec2 °R). 

U.S. customary units of measurement are used in the following data reduction. However, the icing 

community prefers a mixed set of units including airspeed in knots and temperature in degrees Celsius 

(°C). Therefore, the details on the data reduction are presented as they are carried out in English units. 

The final results have been converted to and are presented in the units used by the icing community. 

Facility Instrumentation 

The measured and calculated facility parameters that are of greatest importance to the IRT aerothermal calibration are the bellmouth total pressure, bellmouth static pressure, and the D-corner total 

temperature. The bellmouth total pressure is the average of the two measured total pressures from the 

north wall and south wall 

, = ( , + PPP ,southTnorthTbmT ) 2

In the IRT, the total pressure and differential pressures are measured. The static pressure for each 

bellmouth probe is calculated from their difference. The north and south probe are each calculated using 

the equation 

= TS − ΔPPP

The bellmouth static pressure is again the average of the two probe calculations 

, ( ,northSbmS += PPP ,southS ) 2

The facility total temperature, TT,davg is the average of the 24 RTDs mounted on the leading edge of the Dcorner turning vanes 

∑=

=

24

1

, ,, 24

1

i

T davgT T idT

In addition to the basic parameters, some calculated parameters were also needed based on the bellmouth 

total and static pressure measurements, PT,bm and PS,bm, respectively. The calculated bellmouth Mach 

number is 

NASA/CR—2010-216733 15
⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ = γ

−γ

1

1

2

1

,

,

bmS

bmT

bm P

P

M

The dynamic pressure based on bellmouth conditions is 

2 , 2 bm bmS MPq bm ⋅⋅ γ =

9-ft Survey Rake Calculations 

Data collected by the 9-ft survey rake is used to construct calibration curves for total pressure and 

Mach number. Additionally, the collected data provides flow quality information on total pressure, static 

pressure, Mach number, flow angularity and turbulence. For the following calculations, subscript “bm” 

refers to measured or calculated parameters associated with the two bellmouth pitot-static probes located 

on the inner and outer tunnel walls at the exit of the bellmouth, as shown in Figure 4. The subscript 

“rake” refers to measured or calculated parameters associated with any of the probes on the 9-ft survey 

rake. Additionally, the subscript “local” refers to the measured or calculated corrected “rake” parameters, 

which have been corrected by individual probe calibration coefficients and represent the true “local” 

properties in the test section. Note that all “local” total pressures, static pressures, and Mach numbers are 

normalized by bellmouth (“bm”) parameters to arrive at recovery ratios. All of the compressible flow 

equations used in the data reduction can be found in Reference 9. 

Basic Flow Parameters 

The average static pressure sensed by each pressure probe was calculated by averaging the four static 

pressure taps on the probe body 

( )

4

9876 ,

PPPP P rakeS+++ =

The probe total pressure is simply the pressure measured by port P5 (PT,rake = P5). Mach number is 

calculated using the ratio of the static to total pressure measured by each probe 

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ = γ

−γ

1

1

2

1

,

,

rakeS

rakeT

rake P

P

M

The correction equations for total and static pressure are based on the rake Mach number. Co and Cq are 

functions of Mrake and are experimentally determined during calibration of the pressure probe flow angle. 

Values for these coefficients can be found in Reference 8. 

,localT , −= ( ) [ , −⋅ PPMCPP ,rakeSrakeTrakeorakeT ]

( )

( ) rakeq

rakeSrakeT

localS localT MC

PP

PP , ,

, ,

− −=

The local test section Mach number for each probe is determined using the PT,local and PS,local calculated 

above 

NASA/CR—2010-216733 16
⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ = γ

−γ

1

1

2

1

,

,

localS

localT

local P

P

M

Local dynamic pressure in the test section is then determined 

2 , 2 local localS MPq local ⋅⋅ γ =

The thermocouples on the 9-ft survey rake are only used to monitor the temperature during the aero 

portion of the calibration. Total temperature calibration and flow quality data are collected using the 2–D 

RTD array. The following equations are used to reduce the collected temperature data. The equations 

containing “local” and “rake” parameters apply to all eleven individual total temperature probes mounted 

on the 9-ft survey rake. The corrected local total temperature, TT,local, is calculated as a function of TT,rake

and Mlocal 

, TMCT ,rakeTlocaltlocalT = )( ⋅

Ct is the recovery coefficient for each individual total temperature probe and is a function of Mlocal. These 

coefficients were determined experimentally when the probes were calibrated in the NASA Glenn 3.5-in. 

diameter free jet calibration facility. The total temperature probes gradually predict temperatures lower 

than the actual temperature as the local Mach number increases. For example, at a local Mach number of 

0.1, typical values of Ct are 1.001 and at a local Mach number of 0.6, the values of Ct are generally 

around 1.005. The values of Ct are listed in Reference 10. The local static temperature was calculated to 

monitor data and conditions during the test 

1

2 , , 2

1 1

−

⎟

⎠

⎞ ⎜

⎝

⎛ −γ +⋅= localS localT Mlocal TT

The local airspeed is based on a matched set of one pressure probe and one thermocouple from the 9-

ft survey rake. Each pressure probe was matched to the thermocouple mounted directly under it on the 

rake. The local airspeed for each set of probes was calculated using the following equation 

local MU local γ⋅= RT ,localS

Flow Angularity 

The equations in this section were used to generate pitch and yaw flow angle data from the raw 

pressure data obtained from the flow angle pressure probes (recall Figure 6). The average of the four flow 

angle pressure measurements is used as part of the calculation for both pitch and yaw flow angles 

4

PPPP 4321 Pavg+++ =

Pressure coefficients are calculated for both the pitch angle and yaw angle based on the five pressure 

ports on the hemispherical head of each probe. The pitch angle is related to the pressure difference 

between the ports in the pitch plane, ports P1 and P3, while the yaw angle is related to the difference 

between the yaw plane ports, P2 and P4, as shown in the following equations 

NASA/CR—2010-216733 17
 ( )

( ) PP avg

PP C −− α = 5

13 and ( )

( ) PP avg

PP C −− β = 5

24

The pitch angle, α, and yaw angle, β, are calculated as a function of Mrake, Cα and Cβ. 

,0 rake ,1 rake CMKMK ,arm ,level =α α + α )()( αα Δ+Δ+⋅ α

,0 rake ,2 rake CMKMK ,arm ,ruler )()( β β + Δββ + Δβ =β + ⋅

Pressures P1 through P5 are measured directly from the flow angle pressure probes. The coefficients 

K0,α, K1,α , K0,β , and K2,β were experimentally determined when the flow angle pressure probes were 

calibrated in the NASA Glenn 3.5-in. diameter free jet calibration facility and are all functions of Mrake. 

Values for these coefficients are documented in Reference 8. The Δα,arm and Δβ,arm values were determined 

by measuring the pitch and yaw offset angles for the probes with a 6 degree of freedom digitizing arm and 

were constant for every Z elevation. The Δα,level and Δβ,ruler values were equal to 0° for the Z=36-in. 

elevation since this elevation is where the baseline measurements were taken and were different for every 

other Z elevation. These values were determined by taking measurements with a digital level and a ruler 

and comparing these measurements to the baseline values. Deviations from the baseline numbers were 

entered as Δα,level and Δβ,ruler. 

Turbulence Intensity 

As was previously indicated, the hot wire probes were calibrated in situ in the IRT. This is because it 

is difficult to match the temperature and static pressure conditions of the IRT with those in the probe 

calibration facility. Data were acquired so that calibration curves could be constructed after acquisition 

was completed. Time averaged bridge voltages were matched to time averaged airspeeds from the flow 

angle pressure probes from the 9-ft survey rake for the same condition. Calibration curves, exponential 

functions with effective air velocity as a function of hot wire bridge voltage, were generated for each hot 

wire probe sensor. The equations used in this curve-fit process are from Reference 113

 and are given 

below, 

U ,, ieffHW Ulocal ⋅= cos θ ,iHW and EA iHWi U iiHWiieffHW eAEA ,,2 ,1,,0,, ⋅ ⋅+⋅=

where i = 1 for a single, straight-wire sensor and i = 1 to 2 for a dual cross-wire sensor (X probe). θHW,,i is 

the angle of the hot wire sensor relative to the freestream velocity vector minus 90°. For a single sensor 

probe, the sensor is normal to the velocity vector and θHWr,1 is (90° – 90° =) 0° for this case. For a dual 

sensor X probe at 0° angle of attack, one sensor is at 45° and the other is at 135° relative to the velocity 

vector. For these two sensors, θHWr,1 is (45° – 90° =) –45° and θHW,2 is (135° – 90° =) +45°. Ulocal is the 

freestream velocity, and UHW,eff,i is the velocity component normal to the hot-wire sensor. EHW,i is the hotwire voltage measured by the anemometer. Once the values of UHW,eff,i are computed, these values along 

with the corresponding values of hot wire anemometer bridge voltage, EHW,i can be used to arrive at 

values of A0,i, A1,i, and A2,i via non-linear least squares curve-fitting using the exponential equation given 

above. 

 

3

The procedure was changed in subsequent calibrations to use the King’s Law equation and single-wire probes. 

NASA/CR—2010-216733 18
Figure 14.—Schematic depicting terminology on one wire of X probe. 

The reader may be wondering how accurate flow angles and transverse velocities can be computed by 

the dual sensor X probe if no angular calibration data is used. To answer this question, two key 

assumptions are called upon. The first key assumption is that the X probes will only be used to assess the 

unsteadiness in flow angle and will not be used to quantify the absolute flow angle. The resulting flow 

angles and transverse velocities will all be centered around zero and have averages essentially at zero. 

This should work fine in the IRT test section since we know that the flow angles will be small, typically 

±2° or less, and will be centered about zero. The second key assumption is that the X probes used in this 

test behave like other X probes previously calibrated through various flow angles. The results for an X 

probe calibrated through various flow angles show that the calibration map will collapse down to one 

curve when effective velocity is plotted versus hot wire anemometer bridge voltage. This typically holds 

true for angles between ±20°. 

Given the exponential curve fit equation used to relate hot wire effective velocity to hot wire 

anemometer bridge voltage, data reduction can continue with the equations given below for a dual sensor 

hot wire X probe. 

 Test section airspeed: 2

2,,

2

HW UU effHW 1,, U effHW = +

Pitch flow angle: ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−°=α −

1,,

2,, 1 tan45

effHW

effHW

HW U

U

Axial velocity component: HW =Uu HW ⋅ cos(αHW )

Transverse velocity component: HW = Uv HW ⋅sin(αHW )

For a single sensor hot wire probe, the following equalities are used instead of the four equations 

above. 

HW HW == UUu effHW 1,,

Once uHW is available for a single sensor probe, and uHW and vHW are available for dual sensor X 

probe, statistics necessary for computing turbulence intensity can be calculated using the equations given 

below. The value of N was 100,000. The average axial and transverse air velocity components are 

determined using the following equations. 

∑=

=

N

i

avgHW u iHW N

u

1

, ,

1 and ∑=

=

N

i

avgHW v iHW N

v

1

, ,

1

The magnitude of the velocity vector, also known as airspeed, is determined by combining the axial and 

transverse velocity components. 

Ulocal

UHW,eff,1

θHW,1 

NASA/CR—2010-216733 19
2 , 2 U ,avgHW = ,avgHW + vu avgHW

The standard deviation of the axial and transverse velocity components are calculated using the following 

standard statistical equations. 

∑( ) =

− − =σ

N

i

u uu avgHWiHW N HW

1

2

,, 1

1 and ∑( ) =

− − =σ

N

i

v vv avgHWiHW N HW

1

2

,, 1

1

Finally, the turbulence intensity for the axial and transverse velocity components are defined as 

 %100 ,

×

σ = avgHW

u

u U

TI HW and %100 ,

×

σ = avgHW

v

v U

TI HW

Since in-situ calibration curves were generated, there was no need to correct for changes in 

temperature or static pressure. The calibration curves took care of these inherent variations. In addition, 

new calibration curves were generated for every velocity sweep. 

2–D RTD Array Calculations 

The total temperature measurements made by the 49 RTD probes on the 2–D RTD array only 

required a correction of flow recovery. The determination of the total temperature flow recovery 

correction was described in a previous section of this report. The resulting correction equation is 

localT ( ts ) TMCMCCT .arrayTts

2 , 10 2 = + ⋅ + ⋅ ⋅

Where TT,array is the total temperature measured by each probe and TT,local is the corrected total temperature. 

The correction coefficients are listed in Appendix B. Only 20 of the 49 probes on the array were directly 

calibrated for flow recovery and therefore have specific coefficients. These 20 probes were mounted in 

the interior portion of the array. The remaining 29 RTD probes used the average of the correction 

coefficients as listed in Appendix B. 

Measurement and Data Uncertainty 

Details of the data uncertainty analysis are contained in Appendix A. The uncertainty analysis follows 

the standard procedure set in Reference 12. 

Data Analysis 

Standard test procedure was to collect three data readings for each test condition listed in the test 

matrices for the 9-ft survey rake and the 2–D RTD array. Each data reading was the average of 20 scans 

of data, at a data acquisition rate of 1 scan per second. The data readings were used in two combinations 

for development of calibration curves versus assessment of test section flow quality. 

• For the construction of the calibration curves, the readings were used as they were collected 

(average of 20 scans). So, each test condition would contribute three individual readings to the 

development of the calibration curves. 

• Flow quality distribution data plots were created by averaging the three data readings collected at 

each test condition. For data from the 9-ft survey rake, the data represented at each rake position, 

or height, is the average of three separate data readings collected at one test condition over the 

course of 11 tunnel runs (one run for each of the 11 vertical rake positions). For the 2–D RTD 

NASA/CR—2010-216733 20
array, the data is again plotted by height of the RTD probes above the test section floor and each 

data set of the average of three readings collected at a given test setting. However, unlike the 9-ft 

rake, for the array, only one tunnel run was needed to collect all the data for a given test setting. 

• To combine data collected using the 9-ft survey rake over several test runs to show the composite 

flow quality in the test section, the pressure, Mach number and airspeed data were normalized 

using data collected simultaneously by the facility bellmouth rakes. 

• The corrected pitch and yaw flow angle data were converted into flow angle vectors to more 

easily visualize the flow direction. The length of the vector represents the resultant flow angle 

magnitude and the orientation of the vector shows the resultant flow direction. The magnitude 

and orientation were calculated as shown below. 

 Vector magnitude, Vm 22 Vm β+α=

 Vector orientation, Vφ 2

2

1 tan

β

α =φ − and 

⎪

⎪

⎩

⎪

⎪

⎨

⎧

≥β<αφ

<β<α+φ

>β≥αφ

≥β≥αφ

φ =

0 and 0 if ,-360

0 and 0 if 180,

0 and 0 if ,-180

0 and 0 if ,

V

Discussion of Results 

There are three sets of results described in this section: (1) the aero-thermal calibration curves for 

determining the operating conditions in the IRT, (2) detailed aero-thermal flow quality mapping data for 

the IRT test section and (3) uncertainty estimates for the measured and calculated data. 

Aero-Thermal Calibration Curves 

As indicated, one of the primary goals of this test entry was to update the aero-thermal calibration 

curves used to set operating conditions in the IRT test section. A computer program subroutine was 

created following the 2000 calibration. This subroutine contained the calibration relationships and 

additional calculations to fully describe the test section aero-thermal operating point. This subroutine is 

the IRT Aero-Thermal, and abbreviated as IRTAT. The data collected during the 2004 calibration test 

were used to update the IRTAT. 

Updates Based on 2004 Calibration Test 

The existing calibration curves, based on 2000 calibration data, were compared to calibration curves 

developed from the 2004 data. It was not expected that the 2000 and 2004 curves would match due to the 

changes made to the facility, but large variations were also not anticipated. Figure 15 shows the four 

calibration relationships that were created based on the 2004 data, total pressure, static pressure, Mach 

number and total temperature. The primary calibration relationships were changed after review of the 

2004 calibration curves and comparison to the existing year 2000 calibration curves. This change was 

completed by replacing the total pressure curve with the Mach number calibration curve, due to a reduced 

sensitivity on spraybar air injection effects. The 2000 calibration curves do not include spraybar air 

injection. Spraybar air injection and high temperature test points, above 37 °F or 3 °C, were not included 

in the total temperature calibration curve. The computing requirements for IRTAT were updated per the 

new calibration curves (Revision 4, dated July 29, 2004).4

 The calculations used in IRTAT (Revision 4) 

are described below. 

 

4

Additional corrections were made to IRTAT so the final version based on the 2004 data was Revision 5, dated 

August 13, 2004. 

NASA/CR—2010-216733 21
Figure 15.—Aero-thermal calibration curves based on the 2004 full calibration test 

Inputs to IRTAT 

The measured inputs for determining the calibrated test section aero-thermal conditions are acquired 

from two areas. The first area is the north and south bellmouth, or test section inlet, pitot-static probes, 

and the second area is the 24 D-corner RTD probes. Details and definitions on these inputs and their use 

for determining the calibrated parameters are listed below, with the appropriate units listed in parenthesis. 

Bellmouth total pressure (psia), PT,bm, is the average of the two measured pressures, one each from the 

north and south bellmouth rakes 

( ) ,bmT PPP ,,,, sbmTnbmT 2

1 = +

NASA/CR—2010-216733 22
Likewise, the bellmouth delta-pressure (psid), ΔPbm, is the average of the two delta-pressure 

transducers 

( ) bm , PPP ,sbmnbm 2

1 Δ+Δ=Δ

The bellmouth static pressure (psia), Ps,bm, is calculated from the measured total and differential pressures. 

First the static pressure for each bellmouth rake is determined 

,,,, Δ−= PPP ,nbmnbmTnbmS and ,,,, Δ−= PPP ,sbmsbmTsbmS

The bellmouth static pressure is then the average of the static pressure calculated from each bellmouth 

probe 

( ) ,bmS PPP ,,,, sbmSnbmS 2

1 = +

The facility total temperature is the average of the 24 D-corner total temperature RTD measurements 

(°C), TT,davg 

∑=

=

24

1

, ,,

i

T davgT T iDT

The Mach number based on bellmouth conditions, Mbm, is calculated from the ratio of the bellmouth 

static to bellmouth total pressure using the standard isentropic flow equation (Eq. 44 from Ref. 9 with 

γ=1.4) 

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ =

−

5 1

2857.0

,

,

bmT

bmS

bm P

P

M

Test Section Conditions; Calibrated Values 

The test section calibrated values, Mach number, static pressure and total temperature, are determined 

from calibration curves based on bellmouth pitot-static probes and D-corner RTD measurements. Figure 

15 shows the above mentioned calibration curves. The relationships for the test section conditions are 

listed in the subsequent section. The calibration relationships were based on data collected using the 9-ft 

survey rake and the RTD Array. The 9-ft survey rake measured total and static pressure which is used to 

compute Mach number and the RTD Array measured the total temperature. The calibrated test section 

values are based on conditions measured at the test section turntable centerline and represent average flow 

field values at the core of the test section (the center 25 measurement locations from both the 

aerodynamic and thermal surveys). The pressure and Mach number calibration curves include data for all 

airspeed and spraybar air pressure settings, however, the total temperature calibration curve based on the 

2004 calibration excludes data at the following conditions: TT,davg>3 °C, airspeed< 50 knots, Pair> 0 psig. 

The calibrated test section static pressure (psia), Ps,ts, is determined from the static pressure calibration 

curve, seen in 

Figure 15(b), based on the least squares curve fit to the data 

P ,tsS 37706681.0 02636653.1 P ,bmS = − + ⋅

NASA/CR—2010-216733 23
The calibrated test section Mach number, Mts, is likewise calculated from the Mach number calibration 

curve, shown in 

Figure 15(c), and best fit line 

Mts = 00082798.0 + 01631617.1 ⋅ Mbm

Test section total temperature (°C), TT,ts, is determined from the calibration curve shown in 

Figure 15(d) and calculated from the following equation 

T ,tsT 998554.0246981.0 T ,davgT = + ⋅

It is important to note that the units for the calibrated test section conditions are not in a consistent set 

of engineering units. For subsequent calculations, the total temperature was converted from metric to 

English 

() () 32

5

9 , F , C ⎟ +

⎠

⎞ ⎜

⎝

⎛ T tsT T tsT ⋅°=° and T ,tsT (°R) = T ,tsT (° ) + 67.459F

Test Section Conditions; Calculated Values 

The remaining test section aero-thermal operating conditions are calculated based on the values of the 

calibrated parameters (PS,ts, Mts and TT,ts). The calculated parameters include the total pressure, static 

temperature, airspeed and Reynolds number. The details of these calculations are included in Appendix C. 

Updates Based on 2005 Interim Calibration 

As described in Appendix D of the IRT and Reference 3, the total temperature calibration was 

improved in 2005. The improved total temperature calibration was incorporated into IRTAT as 

Revision 6 (dated April 11, 2005). These updates are documented in Appendix D. 

Aero-Thermal Flow Quality Data 

In addition to creating the aero-thermal calibration relationships, the data collected using the 9-ft 

survey rake and the 2–D RTD array are also used to construct detailed descriptions of the test section flow 

field. The flow quality data for total and static pressure, total temperature, airspeed and Mach number are 

presented in x-y plots (rather than contour plots) to provide as much detail as possible about the flow 

field. Pitch and yaw flow angle data were combined to create flow vectors that provide a combined flow 

angle magnitude and direction. Note that for most of the data presented, the data collected with the 9-ft 

survey rake at 66-in. above the test section floor (maximum height) was omitted as this data was affected 

by boundary layer (either the rake was in the boundary layer or tripped the boundary layer). Also, in some 

cases, data near the test section walls has been truncated due to the choice of y-axis scale; this was done 

in order to show the details of the flow field in the test section core. The total pressure measurement on 

probe number two (18 in. from the inside test section wall) was inoperative during testing (plugged tube). 

These data were omitted from the flow quality plots and account for the gap in the total pressure and 

Mach number distribution data at the position 18 in. from the wall. As noted earlier, 9-ft survey rake data 

were collected at bellmouth total temperature settings greater than 4 °C (about 40 °F). 

Aero-thermal flow quality goals for an icing tunnel are defined in Reference 1 and are summarized 

here in Table 4. In general, the IRT flow quality is very good in terms of the pressure, temperature and 

Mach number (airspeed) distributions and meets or exceeds the goals listed in Table 4 (as will be 

discussed in subsequent sections). While the IRT does not have any conventional flow field manipulators 

(i.e. turbulence reduction screens or flow straightening honeycomb), the heat exchanger does act a flow 

NASA/CR—2010-216733 24
manipulator and provides a significant pressure drop that evens the flow field in the tunnel. The flow 

angularity in the IRT test section is rather high compared with an aerodynamic wind tunnel of similar 

operating range. This is due to the lack of flow straightening and also the presence of the spraybars in the 

settling chamber upstream of the test section. The IRT is a classic example of the trade-offs made in 

ground testing; in order to create icing conditions, test section aerodynamic flow quality is somewhat 

sacrificed by the omission of normal flow field manipulators since those devices would collect ice and 

effect operation and icing performance of the tunnel. 

TABLE 4.—AERO-THERMAL FLOW QUALITY GOALS FOR AN ICING TUNNEL TEST SECTION 

Parameter Measurement 

Uncertainty 

Test section spatial 

uniformity 

Tunnel centerline 

temporal stability 

Airspeed ±1% ±2% ±2% 

Static air temperature between –30 and +5 °C ±2 °C ±1 °C ±0.5 °C 

Flow angularity ±0.25° ±2° N/A 

Turbulence (Pair=0 psig) ±0.25% <2% ±2% 

Turbulence (Pair=70 psig) ±0.25% <2% ±2% 

Total Pressure 

Figure 16 shows the total pressure flow quality data over the operating range of the IRT. As described 

above, the total pressure data from the 9-ft survey rake were divided by the corresponding bellmouth total 

pressure to remove the run-to-run variations. The data collected from various rake heights for like 

operating conditions were then combined to provide a representation of the test section flow quality in a 

plane located axially at the centerline of the test section turntable. 

With the spraybar air off (Pair = 0 psig), the total pressure data show a very uniform distribution 

across the test section with very little variation in the core of the test section. At the lower airspeed 

settings (Uts≤150 mph or 130 knots) the total pressure variation throughout the entire surveyed area of the 

test section is 0.003 psia or less. At about 200 mph (174 knots) a very slight gradient trend begins to form 

in the test section, with higher total pressures recorded near the inside test section wall. Worst case is at 

the 350 mph (304 knots) setting, where the total pressure variation is about 0.006 psia. 

Adding spraybar air into the tunnel airflow has a measurable effect on the total pressure distribution 

over the entire airspeed range of the facility. In order to gage the worst case effect, data were collected 

with the spraybar air at 70 psig, although there was some data collected at lower air pressure settings. The 

effects of the spraybar air injection are larger at lower test section airspeeds. The largest effects are also 

seen near the test section boundaries, so the spraybar air injection appears to have an effect on the test 

section boundary layer. The flow in the core of the test section is affected minimally. For example, at 100 

mph (87 knots) the total pressure variation in the central 5-ft core area is about 0.003 psia for both the 0 

and 70 psig spraybar air injection cases. The difference is seen in the 2-ft area near each test section wall 

where larger variations were measured (about 0.010 psia near the outside tunnel wall). Similar results are 

seen at higher test section airspeeds. The spraybar air injection also increases the total pressure gradient 

recorded in the test section at airspeeds above 200 mph. At the 300 mph (261 knots), the gradient 

increases from about 0.003 psia at 0 psig spraybar air to about 0.006 psia at 70 psig spraybar air. 

Overall, the IRT test section flow quality is excellent for the total pressure distributions. The 

magnitude of the variations discussed above is quite small and should not have any significant impact of 

the testing conducted in the IRT. 

NASA/CR—2010-216733 25
Figure 16.—Total pressure distribution data from the IRT test section. Data collected using the 9-ft survey rake mounted at 11 vertical positions over several tunnel 

runs. The test section total pressure data is normalized using the bellmouth total pressure measurement. Data from the 66-in. vertical position has been omitted due to boundary layer influence. Each y-axis grid (0.0002 total pressure ratio) represents approximately 0.003 psia. (a) Uts=50 mph (43 knots), Pair=0 psig. (b) 

Uts=50 mph (43 knots), Pair=70 psig. (c) Uts=100 mph (87 knots), Pair=0 psig. (d) Uts=100 mph (87 knots), Pair=70 psig.

NASA/CR—2010-216733 26
Figure 16.—Continued. (e) Uts= 115 mph (100 knots), Pair=0 psig. (f) Uts=115 mph (100 knots), Pair=70 psig. (g) Uts=150 mph (130 knots), Pair=0 

psig. (h) Uts=150 mph (130 knots), Pair=70 psig. 

NASA/CR—2010-216733 27
Figure 16.—Continued. (i) Uts= 175 mph (152 knots), Pair=0 psig. (j) Uts=175 mph (152 knots), Pair=5 psig. (k) Uts=175 mph (152 knots), Pair=20 

psig. (l) Uts=175 mph (152 knots), Pair=70 psig.

NASA/CR—2010-216733 28
Figure 16.—Continued. (m) Uts= 200 mph (174 knots), Pair=0 psig. (n) Uts=200 mph (174 knots), Pair=70 psig. (o) Uts=250 mph (217 knots), Pair=0 psig. 

(p) Uts=250 mph (217 knots), Pair=70 psig.

NASA/CR—2010-216733 29
Figure 16.—Continued. (q) Uts= 225 mph (195 knots), Pair=0 psig. (r) Uts=225 mph (195 knots), Pair=20 psig (no z=36-in. vertical position data). 

(s) Uts=225 mph (195 knots), Pair=70 psig.

NASA/CR—2010-216733 30
Figure 16.—Concluded. (t) Uts= 300 mph (261 knots), Pair=0 psig. (u) Uts=300 mph (261 knots), Pair=70 psig. (v) Uts=350 mph (304 knots), Pair=0 psig. 

(w) Uts=350 mph (304 knots), Pair=70 psig.

NASA/CR—2010-216733 31
Mach Number 

The Mach number ratio distribution data are shown in Figure 17. As was the case with the total 

pressure data, the Mach number measurements from the 9-ft survey rake mounted in the test section were 

normalized by the bellmouth Mach number so that the data could be combined to show the overall flow 

quality for a given test section condition. For each set of Mach number ratio data plots, an approximate 

delta Mach number (∆M) was added to provide a gage of the absolute variation in the Mach number 

within the survey plane. Note that the y-axis scale is expanded to show flow field detail. 

The overall Mach number variation for conditions with no spraybar air injection (Pair = 0 psig) was 

0.002 to 0.004, neglecting points near the test section surfaces that may have been affected by the 

boundary layer. The shape of the lateral distribution were consistent at each rake vertical position with 

slightly higher Mach numbers recorded nearer the outside test section wall and the variation was greater 

at the higher Mach number settings (note that the difference in Mach number from inside wall to outside 

wall was only 0.001 to 0.003). 

Spraybar air injection did not have a significant effect on the Mach number distributions, particularly 

in the core of the test section. At the lowest airspeed settings, the spraybar air injection appeared to 

increase the boundary layer thickness near the test section walls (as was also seen in the total pressure 

data) resulting in lower Mach number readings at some rake positions for probes near the walls. For 

nominal airspeed settings of 200 mph (174 knots) and higher however, 70 psig spraybar air injection had 

little or no effect on the Mach number distribution data. 

The Mach number distributions at the test section survey plane are very good over the entire 

operating range of the IRT, including those conditions with 70 psig spraybar air injection. Considering 

that there are no flow quality manipulators (i.e. screens or honeycomb) installed in the settling chamber 

upstream of the test section, the Mach number flow quality is remarkably good (the flat heat exchangers 

and D-corner turning vanes do have a positive effect on the test section flow quality, as described in 

Ref. 4). 

Total Temperature 

For the icing tests that are conducted in the IRT, temperature is the most important of the aerothermal measurements. Both the accuracy of the total temperature measurement and the uniformity of 

temperature in the test section are critical. Because of the critical nature of the temperature measurements, 

the IRT staff is always looking at ways to improve the quality of the total temperature data collected in 

the IRT and the overall flow quality in terms of the total temperature distributions in the test section. As 

mentioned above and in Reference 3, both the accuracy of the total temperature measurements and the 

test hardware used to survey the IRT test section were improved prior to the 2004 aero-thermal 

calibration. These improvements include upgrading the facility and calibration total temperature 

instrumentation from thermocouples to RTDs and using a dedicated total temperature survey tool, the 2–

D RTD array, for test section measurements. 

The total temperature setting in the IRT is controlled by the two flat heat exchangers located in the CD leg of the tunnel loop (Figure 1). The 24 RTDs mounted on the leading edge of the D-corner turning 

vanes (Figure 14) are used to monitor the temperature uniformity of the air exiting the heat exchangers. 

The control system is used to vary the amount of coolant to each of the 8 heat exchanger zones in order to 

produce a uniform total temperature profile in D-corner. The goal is to control the temperature 

distribution to within ±0.5 °C within the D-corner survey plane (this exceeds the goal set forth in 

Ref. 1 and listed in Table 4). Temperature conditions can be accurately set and controlled from about 4 to 

–30 °C. The critical operating point for the IRT is Uts = 230 mph (200 knots) and TT,davg = 0 °C. Note that 

there were some points called out in the test matrix (Table 4) that required total temperature settings 

greater than 4 °C to achieve the desired static temperature. These conditions are beyond the effective 

controllable range of the IRT. 

NASA/CR—2010-216733 32
Figure 17.—Mach number distribution data from the IRT test section. Data collected using the 9-ft survey rake mounted at 11 vertical positions over 

several tunnel runs. The test section Mach number data was normalized using the bellmouth Mach number measurement. Data from the 66-in. vertical position was been omitted due to boundary layer influence. Approximate Mach number delta is indicated for each test section setting. ( 

a) Uts=50 mph, Mts=0.066, Pair=0 psig. (b) Uts=50 mph, Mts=0.066, Pair=70 psig. (c) Uts=100 mph, Mts=0.133, Pair=0 psig. (d) Uts=100 mph, 

Mts=0.133, Pair=70 psig.

NASA/CR—2010-216733 33
Figure 17.—Continued. (e) Uts=115 mph, Mts=0.152, Pair=0 psig. (f) Uts=115 mph, Mts=0.152, Pair=70 psig. (g) Uts=150 mph, Mts=0.199, Pair=0 psig. (h) 

Uts=150 mph, Mts=0.199, Pair=70 psig.

NASA/CR—2010-216733 34
Figure 17.—Continued. (i) Uts=175 mph, Mts=0.232, Pair=0 psig. (j) Uts=175 mph, Mts=0.232, Pair=5 psig. (k) Uts=175 mph, Mts=0.232, Pair=20 psig. (l) 

Uts=175 mph, Mts=0.232, Pair=70 psig.

NASA/CR—2010-216733 35
Figure 17.—Continued. (m) Uts=200 mph, Mts=0.265, Pair=0 psig. (n) Uts=200 mph, Mts=0.265, Pair=70 psig. (o) Uts=250 mph, Mts=0.332, 

Pair=0 psig. (p) Uts=250 mph, Mts=0.332, Pair=70 psig. 

NASA/CR—2010-216733 36
Figure 17.—Continued. (q) Uts=225 mph, Mts=0.305, Pair=0 psig. (r) Uts=225 mph, Mts=0.265, Pair=20 psig. (s) Uts=225 mph, Mts=0.305, Pair=70 psi. 

NASA/CR—2010-216733 37
Figure 17.—Concluded. (t) Uts=300 mph, Mts=0.399, Pair=0 psig. (u) Uts=300 mph, Mts=0.399, Pair=70 psig. (v) Uts=350 mph, Mts=0.464, Pair=0 psig. 

(w) Uts=350 mph, Mts=0.464, Pair=70 psig. 

NASA/CR—2010-216733 38
Figure 18 shows the resulting uniformity of the D-corner total temperature for the settings used 

during the aero-thermal calibration. Note that for most all conditions below 3 °C the variation is D-corner 

total temperature readings are within or nearly within the goal. Exceptions are at very low temperatures 

and very low speeds where variations greater than ±1 °C where recorded. The critical operating regime 

for the icing community is airspeed from 115 to 300 mph (100 to 260 knots) and temperatures from 

freezing to –20 °C; for those settings the D-corner total temperature distribution was consistently within 

the ±0.5 °C goal. 

The test section total temperature data were collected using the 2–D RTD array. The total temperature 

distributions for each test condition are shown in Figure 19. The data in Figure 19 were sorted first by 

airspeed (lowest to highest) then by total temperature (highest to lowest) for each airspeed setting group. 

Each data plot in Figure 19 shows the test section total temperature distribution for a given airspeed, total 

temperature and spraybar air setting; that data are grouped by vertical height within each plot. Note that 

probe number 3 (top row of probes on the array, third from the inside test section wall) was inoperative 

for most of the test. 

The following are general observations of the test section total temperature uniformity: 

• D-corner total temperature settings greater than 3 °C (TT,davg>3 °C).—These settings are warmer 

than the controlled operating range of the IRT heat exchange and therefore produce a much 

higher variation in the test section and D-corner total temperature distributions as compared to 

settings less than 3 °C (Figure 18 and Figure 19). The D-corner and test section total temperature 

gradient both show higher temperatures near the tunnel ceiling with a gradual decrease in 

temperature toward the floor for all conditions where TT,davg>3 °C. As seen by the subsequent data 

plots, these “high temperature” settings are not indicative of the IRT total temperature flow 

quality at settings used for icing conditions. 

• D-corner total temperature settings less than 3 °C (TT,davg<3 °C).—For total temperature settings 

in the effective operating range of the heat exchangers (3 to –30 °C), the total temperature 

variation in the core of the test section is reduced to 1 °C or less (the core of the test section is 

defined as the center 5- by 5-ft area). Many conditions show variation on the order of 0.5 °C 

(peak-to-peak). 

• Nominal test section airspeed of 50 mph (Uts=50 mph =43 knots).—At the very low speed 

condition, the spraybar air injection has a significant effect on the test section total temperature 

distribution at all temperature settings. At Pair = 0 psig, the distributions are uniform and show 

little variation over test section cross-section (horizontal and vertical distributions), generally on 

the order of 1 °C or less. Spraybar air is heated to nominally 80 °C at Pair = 70 psig there is a 

noticeable horizontal gradient of about 2 °C across the test section (higher temperatures recorded 

near the outside wall) as well as vertical variation of about 2 °C. 

• Temperature spikes near test section boundaries.—There are several instances of temperature 

spikes recorded near the test section floor or ceiling (probe vertical position of 9 and 63 in. above 

the floor, respectively). For example, there is a spike seen near the ceiling, 54 in. from the inside 

wall, at the Uts=50 mph (43 knots); worst case is a 2 °C spike at TT,davg=–5.4 °C. Also, there is an 

occasional temperature spike at the floor, generally 72 in. from the inside wall at Uts of 115 mph 

(100 knots) and higher). These spikes are near enough to the test section boundaries that they 

should not have any impact on the core flow temperature distributions. 

• Primary operating range (115<Uts<300 mph, 3<TT,davg<–20 °C).—For all of the critical operating 

conditions (airspeed and total temperature combinations), the total temperature profiles in the 

core of the test section show very uniform conditions that meet the ±0.5 °C goal. In fact, for many 

conditions, the total temperature variation in the core of the test section is 0.5 °C peak-to-peak for 

both spraybar air (Pair) off and 70 psig conditions. 

• Critical (freezing) conditions (TT,davg=0 °C).—Temperature conditions at or about the freezing 

point of water are of critical importance in the aircraft icing community for all airspeed settings in 

NASA/CR—2010-216733 39
the IRT. These total temperature distributions in the IRT test section at freezing conditions are 

excellent; at all airspeed settings the variation in the core of the test section is less than the 

recommended goal of ±1 °C stipulated in Reference 1 and for most cases meets the ±0.5 °C goal 

used in the IRT. 

Figure 18.—Total temperature distribution along the D-corner turning vanes as measured by the D-corner RTD 

probes. See Figure 13 for location of probes relative to the heat exchanger. (a) TT,davg>4 °C. (b) –0.5>TT,davg>–3 °C. 

(c) –3>TT,davg> –9 °C. 

NASA/CR—2010-216733 40
Figure 18.—Concluded. (d) –11>TT,davg>–22 °C. (e) –22>TT,davg>–31 °C. 

NASA/CR—2010-216733 41
Figure 19.—IRT test section total temperature distribution data collected using the 2–D RTD Array (a) Uts=50.1 mph (43.5 knots), TT,davg=4.3 °C, 

Pair=0 psig. (b) Uts=50.4 mph (43.8 knots), TT,davg=5.2 °C, Pair=70 psig. (c) Uts=50.6 mph (44.0 knots), TT,davg=0.2 °C, Pair=0 psig. (d) Uts=50.5 mph (43.9 knots), TT,davg=0.2 °C, Pair=70 psig.

NASA/CR—2010-216733 42
Figure 19.—Continued. (e) Uts=50.4 mph (43.8 knots), TT,davg=–5.4 °C, Pair=0 psig. (f) Uts=49.9 mph (43.3 knots), TT,davg=–5.3 °C, Pair=70 psig. 

(g) Uts=49.9 mph (43.4 knots), TT,davg=–17.6 °C, Pair=0 psig. (h) Uts=49.9 mph (43.4 knots), TT,davg=–17.6 °C, Pair=70 psig. 

NASA/CR—2010-216733 43
Figure 19.—Continued. (i) Uts=50.0 mph (43.5 knots), TT,davg=–29.7 °C, Pair=0 psig. (j) Uts=50.2 mph (43.7 knots), TT,davg=–29.6 °C, Pair=70 psig. 

(k) Uts=115.6 mph (100.5 knots), TT,davg=4.8 °C, Pair=0 psig. (l) Uts=116.4 mph (101.2 knots), TT,davg=5.3 °C, Pair=70 psig. 

NASA/CR—2010-216733 44
Figure 19.—Continued. (m) Uts=116.8 mph (101.5 knots), TT,davg=1.3 °C, Pair=0 psig. (n) Uts=116.5 mph (101.2 knots), TT,davg=1.3 °C, Pair=70 psig. 

(o) Uts=116.4 mph (101.1 knots), TT,davg=–4.2 °C, Pair=0 psig. (p) Uts=117.0 mph (101.7 knots), TT,davg=–4.2 °C, Pair=70 psig. 

NASA/CR—2010-216733 45
Figure 19.—Continued. (q) Uts=116.8 mph (101.5 knots), TT,davg=–16.4 °C, Pair=0 psig. (r) Uts=116.7 mph (101.4 knots), TT,davg=–16.4 °C, 

Pair=70 psig. (s) Uts=116.6 mph (101.3 knots), TT,davg=–28.4 °C, Pair=0 psig. (t) Uts=116.4 mph (101.2 knots), TT,davg=–28.6 °C, Pair=70 psig. 

NASA/CR—2010-216733 46
Figure 19.—Continued. (u) Uts=151.5 mph (131.7 knots), TT,davg=2.1 °C, Pair=0 psig. (v) Uts=151.8 mph (131.9 knots), TT,davg=2.3 °C, Pair=70 psig. 

(w) Uts=152.3 mph (132.4 knots), TT,davg=–3.5 °C, Pair=0 psig. (x) Uts=151.8 mph (131.9 knots), TT,davg=–3.3 °C, Pair=70 psig. 

NASA/CR—2010-216733 47
Figure 19.—Continued. (y) Uts=152.2 mph (132.3 knots), TT,davg=–15.4 °C, Pair=0 psig. (z) Uts=151.6 mph (131.6 knots), TT,davg=–15.6 °C, Pair=70 

psig. (aa) Uts=151.8 mph (131.9 knots), TT,davg=–27.9 °C, Pair=0 psig. (bb) Uts=152.5 mph (132.5 knots), TT,davg=–27.7 °C, Pair=70 psig. 

NASA/CR—2010-216733 48
Figure 19.—Continued. (cc) Uts=233.1 mph (202.5 knots), TT,davg=5.1 °C, Pair=0 psig. (dd) Uts=234.8 mph (204.0 knots), TT,davg=5.5 °C, Pair=70 psig. 

(ee) Uts=233.9 mph (203.3 knots), TT,davg=–0.4 °C, Pair=0 psig. (ff) Uts=234.0 mph (203.4 knots), TT,davg=–0.2 °C, Pair=70 psig. 

NASA/CR—2010-216733 49
Figure 19.—Continued. (gg) Uts=232.8 mph (202.3 knots), TT,davg=–12.5 °C, Pair=0 psig. (hh) Uts=232.7 mph (202.2 knots), TT,davg=–12.5 °C, 

Pair=70 psig. (ii) Uts=233.6 mph (203.0 knots), TT,davg=–24.7 °C, Pair=0 psig. (jj) Uts=232.5 mph (202.0 knots), TT,davg=–24.7 °C, Pair=70 psig. 

NASA/CR—2010-216733 50
Figure 19.—Continued. (kk) Uts=254.7 mph (221.3 knots), TT,davg=4.9 °C, Pair=0 psig. (ll) Uts=254.5 mph (221.2 knots), TT,davg=4.8 °C, Pair=70 psig. 

(mm) Uts=253.9 mph (220.6 knots), TT,davg=0.6 °C, Pair=0 psig. (nn) Uts=253.7 mph (220.5 knots), TT,davg=0.7 °C, Pair=70 psig. 

NASA/CR—2010-216733 51
Figure 19.—Continued. (oo) Uts=253.3 mph (220.1 knots), TT,davg=–11.4 °C, Pair=0 psig. (pp) Uts=253.3 mph (219.9 knots), TT,davg=–11.3 °C, 

Pair=70 psig. (qq) Uts=254.1 mph (220.8 knots), TT,davg=–23.7 °C, Pair=0 psig. (rr) Uts=252.9 mph (219.7 knots), TT,davg=–23.7 °C, Pair=70 psig. 

NASA/CR—2010-216733 52
Figure 19.—Continued. (ss) Uts=304.7 mph (264.8 knots), TT,davg=5.3 °C, Pair=0 psig. (tt) Uts=304.6 mph (264.7 knots), TT,davg=5.0 °C, 

Pair=70 psig. (uu) Uts=303.3 mph (264.6 knots), TT,davg=–8.8 °C, Pair=0 psig. (vv) Uts=303.8 mph (264.0 knots), TT,davg=–8.9 °C, Pair=70 psig. 

NASA/CR—2010-216733 53
Figure 19.—Concluded. (ww) Uts=302.9 mph (263.2 knots), TT,davg=–20.3 °C, Pair=0 psig. (xx) Uts=303.8 mph (264.0 knots), TT,davg=–20.2 °C, Pair=70 psig. 

NASA/CR—2010-216733 54
Flow Angularity 

The test section flow angularity as measured by the 5-hole probes mounted on the 9-ft survey rake is 

shown in Figure 20. As described above, the 5-hole probes measure the pitch and yaw components of the 

flow angle and are then combined to produce the resultant flow vectors. The vectors shown in Figure 20 

only reflect the magnitude and direction of the local air flow, but not airspeed. For the low speed test 

points (Uts=50 mph) the data from rake position 66-in. above the test section floor was deleted due to the 

low speeds measured at that location (the airspeeds were much lower than the calibrated range of the flow 

angle probes). Larger flow angles were consistently recorded near the test section ceiling which may be 

due to the suspected boundary layer interaction noted in Mach number and total pressure distribution data. 

The flow angularity in the IRT test section generally meets the ±2° goal stated in Reference 1 

(Table 4), particularly in the core of the test section (roughly the center 4.5 by 3 ft area). In this central 

core area, the average flow angle vector magnitude is less than 2° for all conditions (including maximum 

spraybar air pressure) as shown in Figure 21. The average flow angle vector magnitude hovers between 

0.7° and 0.9° over the entire airspeed range with no spraybar air injection (Pair=0 psig) This data also 

indicates that there is a significant effect on the average flow angle vector magnitude at maximum 

spraybar air pressure (Pair=70 psig) up to about Uts=150 mph. Above Uts=150 mph (130 knots), spraybar 

air pressure effect on the average flow angle vector magnitude is less than 0.1 degree between Pair=70 and 

Pair=0 psig. The limited data collected at lower spraybar air pressure settings (Pair=5 and 20 psig) show no 

significant difference in average flow angle vector magnitude compared to spraybar air off (Pair=0 psig). 

Overall, the flow angularity in the IRT test section is acceptable for an icing facility and actually is 

not bad for a subsonic wind tunnel that does not have any flow straightening honeycomb in the settling 

chamber. Other than the vortices in the test section corners at low airspeed and high spraybar air settings, 

there are no obvious flow patterns shown in the vector data. Most of the variation in the flow angle 

orientation is probably due to disturbances generated by the spraybars (wakes). 

Turbulence 

Thermal anemometry was used to measure the turbulence levels in the IRT test section. Three hotwire probes were mounted on the 9-ft survey rake. The turbulence data is shown in Figure 22. The IRT 

will not have the low-turbulence levels of a wind tunnel used for aerodynamics testing due to the lack of 

flow manipulators like screens and honeycomb and due to the influence of the spraybars. In general, the 

axial turbulence varies between 0.2 and 1.0-percent with no air injection from the spraybar (Pair = 0). The 

traverse turbulence is a little higher, ranging from 0.5 to 2.0-percent. Note also that the turbulence levels 

are higher near the test section centerline due to the wake from the spraybar centerline support. Higher 

turbulence levels were also recorded near the test section floor and ceiling, possible due to boundary layer 

effects. The injection of spraybar air has a significant effect on the test section turbulence, particularly at 

the lower airspeed settings. 

NASA/CR—2010-216733 55
Figure 20.—Flow angle data from the IRT test section collected using the 9-ft survey rake mounted at 11 vertical positions over several tunnel runs. 

Nominal conditions are used to denote each data set. (a) Uts=50 mph (43 knots), Pair=0 psig. (b) Uts=50 mph (43 knots), Pair=70 psig. (c) Uts=100 mph 

(87 knots), Pair=0 psig. (d) Uts=100 mph (87 knots), Pair=70 psig. 

NASA/CR—2010-216733 56
Figure 20.—Continued. (e) Uts=115 mph (100 knots), Pair=0 psig. (f) Uts=115 mph (100 knots), Pair=70 psig. (g) Uts=150 mph (130 knots), 

Pair=0 psig. (h) Uts=150 mph (130 knots), Pair=70 psig. 

NASA/CR—2010-216733 57
Figure 20.—Continued. (i) Uts=175 mph (152 knots), Pair=0 psig. (j) Uts=175 mph (152 knots), Pair=5 psig. (k) Uts=175 mph (152 knots), 

Pair=20 psig. (l) Uts=175 mph (152 knots), Pair=70 psig. 

NASA/CR—2010-216733 58
Figure 20.—Continued. (m) Uts=200 mph (174 knots), Pair=0 psig. (n) Uts=200 mph (174 knots), Pair=70 psig. (o) Uts=250 mph (217 knots), 

Pair=0 psig. (p) Uts=250 mph (217 knots), Pair=70 psig. 

NASA/CR—2010-216733 59
Figure 20.—Continued. (q) Uts=225 mph (195 knots), Pair=0 psig. (r) Uts=225 mph (195 knots), Pair=20 psig. (s) Uts=225 mph 

(195 knots), Pair=70 psig. 

NASA/CR—2010-216733 60
Figure 20.—Concluded. (t) Uts=300 mph (261 knots), Pair=0 psig. (u) Uts=300 mph (261 knots), Pair=70 psig. (v) Uts=350 mph 

(304 knots), Pair=0 psig. (w) Uts=350 mph (304 knots), Pair=70 psig. 

NASA/CR—2010-216733 61
Figure 21.—Average flow angularity vector magnitude in the core (center 4.5- by 3- ft area) of 

the IRT test section at various 

Pair settings. 

NASA/CR—2010-216733 62
Figure 22.—Summary of the axial and transverse turbulence levels in the IRT test section as measured by probes mounted on 

the 9-ft survey rake. The probes were arrayed laterally across the span of the rake as follows: Probe 1 was 31.8 in. from the inside wall, probe 2 was 53.7 in. and probe 3 was 76.2 in. (a) Test section airspeed of 50 mph 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1

2

3

Anemometer Position

Probe 1

2

3

NASA/CR—2010-216733 63
Figure 22.—Continued. (b) 100 mph 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1 2 3

Anemometer Position

Probe 1 2 3

NASA/CR—2010-216733 64
Figure 22.—Continued. (c) 115 mph 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1 2 3

Anemometer Position

Probe 1 2 3

NASA/CR—2010-216733 65
Figure 22.—Continued. (d) 150 mph 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1 2 3

Anemometer Position

Probe 1 2 3

NASA/CR—2010-216733 66
Figure 22.—Continued. (e) 175 mph, 0 and 70 psi. 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1

2

3

Anemometer Position

Probe 1

2

3

NASA/CR—2010-216733 67
Figure 22.—Continued. (f) 175 mph, 5 and 20 psi. 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1 2 3

Anemometer Position

Probe 1 2 3

NASA/CR—2010-216733 68
Figure 22.—Concluded. (g) 200 mph. 

Axial turbulence intensity, TIu Transverse turbulence intensity, TIv

Probe 1 2 3

Anemometer Position

Probe 1 2 3

NASA/CR—2010-216733 69
Summary of Test Results 

Aero-thermal calibration testing was completed following facility modifications to the NASA Glenn 

Icing Research Tunnel. The original inlet guide vanes upstream of the IRT drive fan were replaced due to 

the age and structural concerns. The new inlet guide vanes were designed to provide more efficient 

operation of the IRT. During January 2004 a full aero-thermal calibration was completed to fully assess 

the impact of the new inlet guide vanes on the IRT test section calibration and flow quality. An interim 

calibration was completed in January 2005. The following were the primary accomplishments and results 

from the 2004 to 2005 aero-thermal calibration test entries in the IRT. 

From the January 2004 full aero-thermal calibration: 

• Completed the planned test matrix that covered the entire operating of the IRT (airspeed, 

temperature and spraybar air injection). 

• From the collected data set, a new set of aero-thermal calibration curves were constructed. These 

calibration curves provided an accurate characterization of the IRT aero-thermal test section 

conditions over the operation range of the facility. 

• The data set also provided a detailed flow field map of the aero-thermal flow quality in the IRT 

test section. 

In addition to providing a general check of the IRT aero-thermal operation and calibration, the data 

collected during the January 2005 interim aero-thermal calibration provided an improved total 

temperature calibration model. 

Recommendations 

The following recommendations were developed following each calibration test entry. Some of the 

recommendations from the 2004 test were implemented during the 2005 test. 

• Following the 2004 test, it was noted that the aero-thermal calibration curves should be refined to 

better model the affects of the spraybar air. This was implemented in the calibration curves 

developed following the 2005 interim calibration. 

• During the 2004 calibration, there some test conditions that required additional stabilization time, 

particularly in terms of total temperature. It was recommended that a means of monitoring the 

stability of the total temperature setting was needed to ensure steady test conditions. The facility 

electrical engineers developed a graphical representation that monitored the real-time, running 

average and standard deviation in the D-corner temperature measurements to monitor data 

stability. This was implemented for the 2005 testing. 

• The data collected in 2004 and 2005 showed that spraybar air injected into the airstream has a 

very noticeable effect on the test section operating conditions as the low speed settings (less than 

100 knots). Additional testing is needed to determine (1) the test section airspeed at which 

spraybar air pressure no longer has a significant effect and (2) the minimum spraybar air pressure 

that causes a noticeable effect at the low speed settings. 

NASA/CR—2010-216733 70
Appendix A.—Uncertainty Estimates 

This analysis was conducted based on the AIAA Standard “Assessment of Experimental Uncertainty 

with Application to Wind Tunnel Testing” and its companion Guide, References 12 and 13. 

A.1 General Uncertainty Analysis 

The accuracy of a measurement is the true value in relation to the measured value. In 

experimentation, the true value is unknown. Therefore, the uncertainty of the measured value is 

determined using the bias and precision errors. The bias error is the fixed or constant error that occurs 

with each measurement, while the precision error is the random or repeatability error. 

General uncertainty analysis (Ref. 14) uses the experimental result, r, with variables up to XJ as seen 

in Equation (A1) 

),....,( = 21 XXXrr J (A1) 

The general uncertainty equation takes the form seen in Equation (A2) with the expanded form shown as 

2

1

1

22

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

∑θ=

=

J

i

r i Xi U U (A2) 

2

1

2 2

2

2

1

... 1 2

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

∂

∂

++ ⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂

+ ⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂ = XJ

J

r X X U

X

r U

X

r U

X

r U (A3) 

In this analysis, the uncertainties of the velocity, Mach number, and flow angularity were determined with 

respect to total temperature, total pressure and static pressure. Additionally, the pressure measurement 

uncertainty was determined with respect to the Escort D+ data acquisition system for the IRT facility, 

including the pressure transducer and amplifier. 

A.2 Hardware RTD Temperature Uncertainty 

The individual component uncertainties were reviewed for the RTD total temperature probe system. 

Uncertainty values are given in Table A.1. The component uncertainties were calculated using a rootsum-squared (RSS) technique which is a method of determining errors. 

222 ++= cbaRSS (A4)

The component uncertainties calculated using the RSS technique yielded a system uncertainty of 

±0.10 °C for a probe that had been individually calibrated for flow recovery and ±0.16 °C for a probe that 

had not been individually calibrated for flow recovery. The uncertainties associated with the RTD probes 

are more than an order of magnitude improvement over the uncertainty estimate for the thermocouples on 

the 9-ft survey rake. The system uncertainty associated with the 9-ft survey rake is ±1.1 °C.4,5 The RSS 

computations for the RTD probes are show below: 

For probes that were tested for total temperature flow recovery: 

C10.0)030.0040.0026.0006.0080.0030.0(RSS 2 2 2 2 2 2 D = ±=+++++

For probes using average flow recovery coefficients: 

NASA/CR—2010-216733 71
C16.0)123.0040.0026.0006.0080.0030.0(RSS 2 2 2 2 2 2 D = ±=+++++

TABLE A.1.—RTD TOTAL TEMPERATURE PROBE SYSTEM COMPONENT UNCERTAINTIES. 

Component 

uncertainty, 

°C 

Component description 

±0.030 RTD capsule, 0.01% DIN curve 

±0.080 RTD temperature transmitter 

±0.006 Resistor to convert transmitter 4 to 20 mA output to DC V (USF 240 250 Ω 0.01% 2 ppm/°C) 

±0.026 Escort data system analog to digital conversion error 

±0.040 Flow recovery reference temperature measurement error (jet plenum reference) 

±0.030 Flow recovery curve uncertainty (individual probe cal) 

±0.123 Flow recovery curve uncertainty (probe not calibrated for flow recovery, average coefficients) 

A.3 Hardware Pressure Uncertainty 

A.3.1 Total Pressure 

Two pitot-static probes, one mounted on the IRT test section vertical north wall and the other on the 

vertical south wall, are used to measure total and differential pressure. The total pressure is measured 

using a 0- to 20-psia Setra Model 270 transducer (Setra Systems, Inc., Boxborough, MA), and the 

differential pressure is measured using a 0- to 3-psid Setra Model 239 transducer. 

The IRT facility data acquisition system utilized the ESCORT D+ with 470 Neff amplifier system 

and the Setra Model 270 pressure transducer for the total pressure measurement. 

The calculations shown below detail the bias error of the system. 

222222222 ++++++++±= BBBBBBBBBB 987654321

2 2 2 2 2 2 2 B ±= + +++ + + )001.0()004.0()010.0()001.0()001.0()010.0()002.0(

B = ± psia0149.0

TABLE A.2.—TOTAL PRESSURE COMPONENT BIAS UNCERTAINTIES 

Component uncertainty, 

psia 

Component description 

±0.002 Cal Lab calibration error 

±0.010 Non-linearity 

±0.001 Thermal Zero shift 

±0.001 Thermal span shift 

±0.010 Neff gain accuracy 

±0.004 Neff non-linearity 

±0.000 Channel-to-channel offset 

±0.000 Common mode voltage 

±0.001 Digitizing Error 

Precision error 

222222 +++++±= SSSSSSS 654321

2 2 2 2 2 2 S ±= + ++ + + )002.0()003.0()004.0()001.0()002.0()002.0(

S = ± psia0117.0

NASA/CR—2010-216733 72
TABLE A.3.—TOTAL PRESSURE COMPONENT PRECISION UNCERTAINTIES 

Component uncertainty, 

psia 

Component description 

±0.002 Repeatability 

±0.002 Hysteresis 

±0.001 Excitation voltage effect 

±0.004 Neff Gain Stability 

±0.003 Zero Stability 

±0.002 Noise 

Total pressure measurement uncertainty 

( )

2

2 2

⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛ +±=

N

S URSS B (A5) 

UU psia RSS PT ±== 027741.0

= TRSS = T PUPUU TP = = %13.0001387.0)()(%

A.3.2 Dynamic Pressure 

Equation (A6) was found using the Ames Tables (4) and assuming an isentropically perfect gas. 

1 2 2

2

1 1

2

−γ

γ−

⎟

⎠

⎞ ⎜

⎝

⎛ −γ + γ = T MPq M (A6) 

1

2

2

1

2

1 2 −γ

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

γ = ∂

∂

M

M

P

q

T

 (A7) 

( )

1

1

2

32

1

2 1

2

1 1

2

1

1

2

1 + −γ

γ

−γ

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ −γ

⎟

⎠

⎞ ⎜

⎝

⎛ −γ γ

−

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

γ = ∂

∂

M

PM

M

MP

M

q

T

T (A8) 

2

1

2 2

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

⎟

⎠

⎞ ⎜

⎝

⎛

∂

∂

+ ⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂ = P M

T

q U

M

q U

P

q U T (A9) 

( )

2

1

2

2

1

1

2

32

1

2

1

2

2

42 2

1

2

1 1

2

1

1

2

1 1

2

1 4

⎥

⎥

⎥

⎥

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎢

⎢

⎢

⎢

⎣

⎡

⎟

⎟

⎟

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎜

⎜

⎜

⎝

⎛

⎟

⎟

⎟

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎜

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ −γ

⎟

⎠

⎞ ⎜

⎝

⎛ −γ γ

−

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

γ +

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

γ = + −γ

γ

−γ

γ

−γ

γ M

T

T

q U

M

PM

M

MP

M

UM

U PT (A10) 

NASA/CR—2010-216733 73
A.3.3 Static Pressure 

Using the Ames Tables (4) and assuming an isentropically perfect gas. 

1 2

2

1 1 −γ

γ−

⎟

⎠

⎞ ⎜

⎝

⎛ −γ PP TS += M (A11) 

1

2 1

2

1

1

−γ

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ − ⎠

⎞ ⎜

⎝

⎛ −γ

= ∂

∂

M

P

P

T

S (A12) 

( )

1

1

2 1

2

1 1

2

1 2

+ −γ

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ − ⎠

⎞ ⎜

⎝

⎛ −γ −γ

⎟

⎠

⎞ ⎜

⎝

⎛ −γ γ

= ∂

∂

M

PM

M

P T

S (A13) 

2

1

2 2

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

⎟

⎠

⎞ ⎜

⎝

⎛

∂

∂

+ ⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂ = M

S

P

T

S

P U

M

P U

P

P U S T (A14) 

( )

2

1

2

2

1

2

2

2

2222

1

2

2

2

1 1

2

1

2

1 4

1

2

1

⎥

⎥

⎥

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎢

⎢

⎢

⎣

⎡

−γ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

⎟

⎠

⎞ ⎜

⎝

⎛ −γ γ

+

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ⎟ + ⎠

⎞ ⎜

⎝

⎛ −γ

= + −γ

γ

−γ

γ

M

UPM

M

U

U

MT

P

P T

S (A15) 

A.4 Airspeed Calculations 

Using the velocity equation, Equation (A16), the speed of sound equation, Equation (A18), and the 

isentropic identity shown, Equation (A19), the velocity was solved for 

= ⋅ aMV (A16)

2

1

1

1

1

2

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ =

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

−γ

S

T

P

P M (A17) 

γ= RTa (A18) 

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

−γ

⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛ =

1

T

S

P

P

T

T (A19) 

NASA/CR—2010-216733 74
2

1

1

1

1

2

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ =

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

−γ

S

T

T P

P V RT (A20) 

The general expression for uncertainty is given by Equation (A21). UTT , is the uncertainty in total 

temperature, U PS , is the uncertainty in static pressure, and UPT , is the uncertainty in total pressure. 

2

1

2 2 2

⎥

⎥

⎦

⎤

⎢

⎢

⎣

⎡

⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂

+ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

∂

∂

+ ⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

∂

∂ = T S PT

T

P

S

T

T

V U

P

V U

P

V U

T

V U (A21) 

To produce algebraic simplification, the uncertainty expression is divided by V2

 and shown below. 

2 2 2 2

1 1 1 ⎥

⎦

⎤ ⎢

⎣

⎡

∂

∂

+ ⎥

⎦

⎤ ⎢

⎣

⎡

∂

∂

+ ⎥

⎦

⎤ ⎢

⎣

⎡

∂

∂ = ⎥

⎦

⎤ ⎢

⎣

⎡ T S PT

T

P

S

T

T

V U

P

V

V

U

P

V

V

U

T

V

VV

U (A22) 

Taking the partial derivative of the velocity equation with respect to total temperature, TT, static 

pressure, PS, and total pressure, PT, the uncertainty of velocity is seen in Equation (A26). 

2

1

2

1

1

2

1

)(

1

1

2

2

1

T

S

T

T RT

P

P R

T

V

γ

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ

γ

= ∂

∂

γ

−γ

 (A23) 

( )

2

1

1

2

1 1

2

1

1

1

1

2 5.0

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

−γ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ −= ∂

∂

γ

−γ

− γ

−γ

S

T

S

S

T

T

S

P

P P

P

P P

P

V (A24) 

( )( )

2

1

1

2

1

1 1

2

1

1

1

1

2 5.0

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

γ−γ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ = ∂

∂

γ

−γ

− γ

−γ

S

T

S

T

S

T

T

P

P P

RT

P

P

P

V (A25) 

NASA/CR—2010-216733 75
( )

( )

( )( )

( )

( )( )

2

1

1

2 4

2

2 12

22

1

2 2

2

2 12

2

1

22 2

1

1

1

2 25.0

1

1

1

2 25.0

1

1

2 25.0

⎟

⎟

⎟

⎟

⎟

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎜

⎜

⎜

⎜

⎜

⎝

⎛

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

γ−γ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ +

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

γ−γ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ + γ

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−γ

γ

=

γ

−γ

− γ

−γ

γ

−γ

− γ

−γ γ

−γ

S

T

S

T

S

T

T P

S

T

S

T

S

T

P

T

S

T

T

V

P

P P

RT

P

P UP

P

P P

RT

P

P U

RT

P

P UR

U

T S (A26) 

A.5 Mach Number Calculations 

Using the MuPAD feature in the symbolic toolbox in Matlab the partial derivatives for Mach number 

were solved for and are shown in Equation (A27) and (A28). 

2

1

1

1 1

1

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

= ∂

∂

γ

−γ

− γ

−γ

S

T

S

S

T

T

P

P P

P

P

P

M (A27) 

2

1

1

2

1 1

1

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

−= ∂

∂

γ

−γ

− γ

−γ

S

T

S

S

T

T

S

P

P P

P

P P

P

M (A28) 

The partial derivatives were then plugged back into Equation (A3) yielding the following for the 

uncertainty for Mach. 

2

1

1

4

2 )1(2

22

1

2

2 )1(2

2

1 1

⎥

⎥

⎥

⎥

⎥

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎢

⎢

⎢

⎢

⎢

⎣

⎡

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

+

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

γ

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

=

γ

−γ

− γ

−γ

γ

−γ

− γ

−γ

S

T

S

S

T

T P

S

T

S

S

T

P

M

P

P P

P

P UP

P

P P

P

P U

U T S

 (A29) 

A.6 Flow Angle Calculations 

The flow angle calculations are completed utilizing the 9-ft rake flow angularity probes. Figure 6 in 

the main body of the document illustrates the locations and descriptions of each of the pressure orifices. 

Equations (A31) and (A32) are the equations used to calculate pitch and yaw utilizing the flow angularity 

probes. 

4

4321

,41

PPPP P avg+++ − = (A30) 

NASA/CR—2010-216733 76
( )

( ) POFF DPOFF

PP

PP AKAK

avg

++ −

− +=α − ,415

13 10 (A31) 

( )

( ) YOFF DYOFF

PP

PP BKBK

avg

++ −

− +=β − ,415

24 20 (A32) 

The partials of the pitch and yaw equations were taken with respect to P1 through P5 and plugged 

into the uncertainty equations. Equation (A33) is the uncertainty equation for pitch, and Equation (A34) is 

the uncertainty equation for yaw. 

( )

( )

( )

( )

( )

( )

( )

( )

( )

( )

5.0

4

5 4444

2

31

2

5

4

5 4444

2

31

2

4

2

4

5 4444

2

31

2

2

2

2

2

5 4444

31

5 4444

2

3

2

2

5 4444

31

5 4444

2

1

21 3 4 21 3 4 21 3 4

21 3 4 21 3 4

21 3 4 21 3 4

16

1

16

1

16

1

4

1 1

4

1 1

⎟

⎟

⎟

⎠

⎞

−+++

− +

−+++

− +

−+++

− +

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

−+++

− − −+++

+

⎜

⎜

⎜

⎝

⎛

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

−+++

− − −+++ =

P

PPAUK

P

PPUAK

P

PPUAK

P

PPAK

P

AK U

P

PPAK

P

AK UU

PP P P

P

PP P P

P

PP P P

P

PP P P P PP P P

PP P P

Pitch P PP P P

 (A33) 

( )

( )

( )

( )

( )

( )

( )

( )

( )

( )

5.0

4

5 4444

2

42

2

5

4

5 4444

2

42

2

3

2

4

5 4444

2

42

2

1

2

2

2

5 4444

42

5 4444

2

4

2

2

5 4444

42

5 4444

2

2

21 3 4 21 3 4 21 3 4

21 3 4 21 3 4

21 3 4 21 3 4

16

2

16

2

16

2

4

2 2

4

2 1

⎟

⎟

⎟

⎠

⎞

−+++

− +

−+++

− +

−+++

− +

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

−+++

− − −+++

+

⎜

⎜

⎜

⎝

⎛

⎟

⎟

⎟

⎠

⎞

⎜

⎜

⎜

⎝

⎛

−+++

− − −+++ =

P

PPBUK

P

PPUBK

P

PPUBK

P

PPBK

P

BK U

P

PPAK

P

BK UU

PP P P

P

PP P P

P

PP P P

P

PP P P P PP P P

PP P P

Yaw P PP P P

 (A34)

Utilizing the following inputs from the 2004 data set RDG 1530, the following uncertainties were 

determined. The inputs and the uncertainties calculated for the single data point are shown in Table A.4. 

TABLE A.4.—FACILITY UNCERTAINTIES 

Input Uncertainties 

Total Pressure (psi) 14.315 Total Pressure (psia) 0.027741 

Static Pressure (psi) 12.258 Total Temperature (°C/R) 0.16/0.08889 

K0A –0.1682 Static Pressure (psia) 0.0411 

K1A 12.1366 Mach Number 0.0061 

K0B 0.5052 Dynamic Pressure 0.0582 

K2B 11.9173 Velocity (ft/sec) 6.7359 

 Pitch (degrees) 0.0435 

 Yaw (degrees) 0.0427 

NASA/CR—2010-216733 77

Appendix B.—RTD Probe Flow Recovery Coefficients 

Table B.1 contains the total temperature flow recovery curve fit coefficients for each of the RTD 

probes used on the 2–D RTD Array. The corrected total temperature measurement for a given probe is 

determined from the following equation 

localT ( ts ) TMCMCCT .arrayTts

2 , 10 2 = + ⋅ + ⋅ ⋅

where TT,array is the total temperature measured by the RTD probe and TT,local is the corrected local total 

temperature. 

TABLE B.1.—FLOW RECOVERY CORRECTION COEFFICIENTS FOR 

THE RTD PROBES MOUNTED ON THE 2–D RTD ARRA. 

Probe Serial 

Number 

C2 C1 C0

1 5.681198×10–3 8.859325×10–4 1.0 

2 7.608867×10–3 5.487745×10–6 1.0 

3 6.518485×10–3 2.305113×10–4 1.0 

4 6.562552×10–3 9.399041×10–4 1.0 

5 5.788899×10–3 6.373217×10–4 1.0 

6 6.482487×10–3 1.749681×10–4 1.0 

7 1.011491×10–2 –1.646186×10–3 1.0 

8 5.661808×10–3 5.089893×10–4 1.0 

9 6.506136×10–3 4.493981×10–4 1.0 

10 5.600892×10–3 5.937824×10–4 1.0 

11 6.159188×10–3 7.799326×10–4 1.0 

12 4.848719×10–3 1.197796×10–3 1.0 

13 5.634439×10–3 1.523486×10–4 1.0 

14 5.412994×10–3 1.068442×10–3 1.0 

15 6.099340×10–3 6.670401×10–4 1.0 

16 5.399611×10–3 6.660572×10–4 1.0 

17 6.631578×10–3 6.137670×10–4 1.0 

18 6.718590×10–3 6.487126×10–4 1.0 

19 4.910175×10–3 8.683076×10–4 1.0 

20 6.384046×10–3 6.479808×10–4 1.0 

21 6.823889×10–3 3.565925×10–4 1.0 

22 5.007236×10–3 1.218866×10–3 1.0 

23 6.436245×10–3 5.701839×10–4 1.0 

Average 6.881316×10–3 1.490507×10–4 1.0 

NASA/CR—2010-216733 79

Appendix C.—Calculated Test Section Conditions 

The following set of equations is used to compute the test section flow conditions based on the 

calibrated test section conditions. The calibrated test section conditions are static pressure, PS,ts, Mach 

number, Mts and total temperature, TT,ts. For all calculations, γ=1.4 and R=1716 ft-lb/(slug⋅°R). English 

units are used for all calculations; however some units conversions are required to provide operating 

conditions in the units desired by the aircraft icing community. 

The first parameter calculated is the test section total pressure, PT,ts. The ratio of static to total 

pressure is determined from calibrated Mach number found in Reference 9 as Equation 44. 

2/7 2

,

,

, 5

1

−

⎟

⎟

⎠

⎞ ⎜

⎜

⎝

⎛

+== ts

tsT

tsS

tsrat

M

P

P

P

The total pressure is then calculated using the pressure ratio and the calibrated static pressure. Note that 

the units used in the calculation are psf. 

tsrat

tsS

tsS

tsS

tsT

tsT P

P

P

P

P

P ,

,

,

,

,

, =⋅ ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

=

The test section static temperature, TS,ts, is calculated, in °R, as shown below using Reference 9, 

Equation 43 

2

,

, 2.01 ts

tsT

tsS M

T

T

⋅+ =

Test section airspeed Uts in ft/sec, is determined from the definition of Mach number, M=V/a. Solving 

for V and substituting the definition for the speed of sound produces the following equation 

tstststs TRMaMU ,tsS ⋅⋅γ=⋅=

Many of the IRT test customers are used to working with airspeed in knots, so the following unit 

conversion is needed 

( ) 592484.0knots (ft/sec) Uts Uts = ⋅

The test section Reynolds number is calculated beginning with determining the air density 

tsS

tsS

ts TR

P

,

,

⋅ =ρ

The test section air viscosity, μts in units of slugs/(ft⋅sec), is determined from the following equation 

8

,

5.1

, 10

6.198

270.2 − ⋅ +=μ

tsS

tsS

ts T

T

These parameters are used in the Reynolds number, (106

/ft), Rets, equation seen below 

NASA/CR—2010-216733 81
ts

tsts ts

lU

μ

ρ Re =

In this equation, the characteristic length, l, is equal to 1 ft. The test section Reynolds number can also be 

calculated based on test article dimension (i.e., chord length) 

cxtsc = ReRe ⋅

Table C.1 summarizes the inputs, calibrated test section conditions and calculated test section 

conditions as in this report. The program name column lists the variable names for each parameter as they 

appear in the computer subroutine IRTAT. At present, the parameter name and program names do not 

exactly match because the parameter naming nomenclature has been updated to be consistent with naming 

recommendations, detailed in Reference 2. The program names will eventually be updated to be 

consistent with the naming recommendations. Also, note that Pair was added as an input to IRTAT as part 

of revision 6. 

TABLE C.1.—SUMMARY OF THE PARAMETERS USED TO DETERMINE 

THE TEST SECTION AERO-THERMAL CONDITIONS IN THE IRT 

Parameter Program 

Name 

Description Units Source 

Average Bellmouth and D-Corner Conditions 

 PT,bm POBM Bellmouth total pressure psia Input 

ΔPbm DPBM Bellmouth delta-pressure psia Input 

 PS,bm PSBM Bellmouth static pressure psia Input 

 TT,davg TODC D-Corner total temperature °C Input 

 Prat,bm PRATBM Ratio of bellmouth static to total pressure none Calculated 

 Mbm MBM Bellmouth Mach number None Calculated 

 Pair SBAIRP Spraybar air pressure psig Input 

Test Section Conditions—Calibrated Parameters 

PS,ts PSTS Test section static pressure psia Calibration 

Mts MTS Test section Mach number none Calibration 

TT,ts TOTSC 

TOTSF 

TOTS 

Total temperature °C 

°F 

R 

Calibration 

(Units) 

(Units) 

Test Section Conditions—Calculated Parameters 

Prat PRAT Ratio of static to total pressure none Calculated 

PT,ts POTS Test section total pressure psia Calculated 

TS,ts TSTS Static temperature °R Calculated 

Uts VTS 

VTSK 

Airspeed ft/sec 

knots 

Calculated 

(units) 

ρts RHOTS Air density slugs/ft3

 Calculated 

μts MUTS Air viscosity slugs/(ft sec) Calculated 

Rets REFT Reynolds number 106/ft Calculated 

Recx RECX Reynolds number based on model length 106/ft Calculated 

γ GAMMA Ratio of specific heats = 1.4 none constant 

R R Gas constant = 1716 lbf-ft/(slug°R) Constant 

cx CX Model chord or other critical dimension feet PARAM File 

input 

NASA/CR—2010-216733 82
Appendix D.—2005 Interim Aero-Thermal Calibration of the IRT 

As mentioned in the Introduction section, there are three basic levels of calibration used in the IRT, 

full, interim and check calibrations. Full and interim calibrations occur in pairs and are designed to 

provide a level of redundancy to ensure consistent operation of the facility. The full and interim 

calibrations occur either following a facility modification or every 5 years if there are no facility 

modifications. The information in this appendix will provide the final test documentation of the 2005 

interim calibration of the IRT including the test matrix, test conditions, data and results. 

An interim calibration uses the same hardware as a full calibration, namely the 9-ft survey rake and 

the 2–D RTD array. For the 2005 interim calibration, testing using the 9-ft survey rake was completed in 

January (January 18, 2005), however, the total temperature calibration measurements using the RTD array 

was not conducted and completed until April (April 1 and April 4, 2005). The time between the 9-ft rake 

and RTD array testing was used for icing cloud calibration and diagnostics. 

The test matrices for the 2005 interim calibration are given in Table D.1 and Table D.2. These test 

matrices were designed to: 

• Complete data collection to check the total and static pressure calibration relationships and the 

centerline flow quality using the 9-ft survey rake at the test section centerline at the turntable 

station. 

• Complete data collection to check the total temperature calibration relationship and flow field 

surveys using a two-dimensional array of RTDs. The 2005 test matrix differs from the 2004 full 

calibration by the addition of a spraybar air pressure setting of Pair = 40 psig. The addition was 

made to begin a sensitivity study of the spraybar air pressure on the aero-thermal characteristics 

of the test section calibration and flow quality. 

In addition to the fundamental aero-thermal calibration goals, the 2005 interim calibration was used to 

conduct temperature stability studies between the D-corner and test section. The primary concern was the 

overall stability and settling time required to achieve uniform total temperature distribution in D-Corner 

and the test section. The 2–D RTD array was used to monitor the test section temperature distribution in 

conjunction with the D-Corner RTDs. The D-Corner RTDs are facility instrumentation mounted on the 

leading edge of the D-Corner turning vanes. The objective of the surveys was to document the 

settling/stability characteristics to the total temperature profiles in D-Corner and the test section and to 

determine guidelines for IRT operation to insure a uniform and stable total temperature in both D-Corner 

and the test section. The Escort system, the facility data acquisition and display system, was used to 

monitor the standard deviation of the D-Corner total temperature, σTT,DC, and the time variation in the 

D-Corner average temperature, ΔTDC . Details and results of the temperature stability study are given in 

(Ref. 3). 

TABLE D.1.—TEST MATRIX FOR THE IRT AERO-THERMAL CALIBRATION USING THE 9-ft SURVEY RAKE 

(2005 INTERIM CALIBRATION).

Rake 

height, Y, 

in. 

Total Temperature, 

TTdavg, 

 °C 

Test Section Airspeed, Uts, knots Spraybar air 

pressure, Pair, 

psig 

36 5 0, 43, 65, 87, 100, 130, 152, 173, 200, 217, 240, 260, 282, 304, 339 0 

36 5 0, 43, 65, 87, 100, 130, 152, 173, 200, 217, 240, 260, 282, 304, 339 40 

36 5 0, 43, 65, 87, 100, 130, 152, 173, 200, 217, 240, 260, 282, 304, 339 70 

NASA/CR—2010-216733 83
D.1 Total Temperature Calibration 

The total temperature calibration relationship is one of a set of three calibrations used to fully define 

the aero-thermal operating conditions in the IRT test section (the other two calibration relationships are 

static pressure and Mach number). The total temperature calibration simply relates the test section total 

temperature measured by the 2–D RTD (average of 49 measurements) array to the total temperature 

measured in D corner (average of 24 measurements). Prior to 2005, a single curve-fit was used to 

represent the total temperature calibration across the operating range of the IRT. This single curve-fit did 

not include data collected at low airspeed with the spray bar air turned on due to the amount of scatter in 

these data relative to the data collected at higher airspeeds. 

During the 2005 interim calibration, an analysis was done to determine if a family of calibration 

curves was needed for the total temperature to account for airspeed and spraybar air pressure effects. The 

final conclusion was that there is a break-point between 50 and 100 knots such that above 100 knots, 

spraybar air pressure no longer has a significant effect on the total temperature calibration. The result was 

to create a set of four calibration curves, one for airspeeds greater than 50 knots and three for airspeed less 

than 50 knots based on the spraybar air pressure setting (0, 40 or 70 psig). The resulting total temperature 

calibration curves are shown in Figure D.1. Table D.3 provides the details for the calculation of the 

calibrated test section total temperature from the measured D-corner total temperature. In order to provide 

some degree of vernier on the spraybar air pressure at the low speed conditions, the slope and intercept 

from each of the low speed calibration curve-fits were plotted against the spraybar air pressure so that 

calibration curve coefficients could be estimated for other spraybar air settings (Figure D.2). The resulting 

curve-fit equations for estimating CTT,0 and CTT,1 for Uts < 50 knots are listed below. 

For all Pair settings not specifically listed in Table D.3: 

2 CTT 0, 06550710.0 01720280.0 Pair 00004323.0 Pair = + +⋅ ⋅

For 0 < Pair < 40 psig: 

CTT Pair = 98917044.0 − 00001881.0 ⋅ 1,

TABLE D.2.—TEMPERATURE CALIBRATION TEST MATRIX 

USING THE 2–D RTD ARRAY (2005 INTERIM CALIBRATION).

Test section 

static 

temperature, 

TS,ts, 

 °C 

Test 

section 

airspeed, 

Uts, 

knots 

D-corner total 

temperature, 

TT,davg, 

 °C 

Spraybar 

air 

pressure, 

Pair, 

psig 

Test section 

static 

temperature, 

TS,ts, 

 °C 

Test 

section 

airspeed, 

Uts, 

knots 

D-corner total 

temperature, 

TT,davg, 

 °C 

Spraybar 

air 

pressure, 

Pair, 

psig 

–30 43 –29.8 0, 40, 70 –5.6 43 –5.3 0, 40, 70 

–30 100 –28.7 0, 40, 70 –5.6 100 –4.2 0, 40, 70 

–30 130 –27.8 0, 40, 70 –5.6 130 –3.3 0, 40, 70 

–30 200 –24.5 0, 40, 70 –5.6 200 –0.3 0, 40, 70 

–30 217 –23.8 0, 40, 70 –5.6 217 0.7 0, 40, 70 

–30 304 –17.8 0, 40, 70 0 43 0.2 0, 40, 70 

–17.8 43 –17.6 0, 40, 70 0 100 1.3 0, 40, 70 

–17.8 100 –16.4 0, 40, 70 0 130 2.2 0, 40, 70 

–17.8 130 –15.6 0, 40, 70 –0.6 200 4.7 0, 40, 70 

–17.8 200 –12.5 0, 40, 70 –1.7 217 4.6 0, 40, 70 

–17.8 217 –11.6 0, 40, 70 –7.2 304 4.9 0, 40, 70 

–17.8 304 –5.6 0, 40, 70 4.4 43 4.7 0, 40, 70 

3.3 100 4.7 0, 40, 70 

NASA/CR—2010-216733 84
For 40 < Pair < 70 psig: 

CTT Pair = 99428850.0 − 00014676.0 ⋅ 1,

These calculated values for CTT,0 and CTT,1 are then used in the calibration equation (Table D.3) to 

determine TT,ts at the low speed conditions. 

Based on the 2005 interim calibration, the subroutine used to determine the aero-thermal operating 

conditions in the test section (IRTAT) was updated (revision 6, dated April 11, 2005). The changes were 

the inclusion of the spraybar air pressure, Pair, as an input to IRTAT (the computer code name used was 

SBAIRP) and to generalize the equation for calibrated test section total temperature 

, TTtsT TCCT ,1,0, davgTTT ⋅+=

Where CTT,0 and CTT,1 are determined as defined above and in Table D.3. 

TABLE D.3.—GENERAL EQUATION AND CURVE FIT COEFFICIENTS USED IN THE 

CALCULATION OF THE CALIBRATED TEST SECTION TOTAL TEMPERATURE.

, TTtsT TCCT ,1,0, davgTTT = + ⋅

Test section airspeed, 

Uts, 

knots 

Spraybar air 

pressure, 

Pair, 

psig 

Curve fit intercept 

term, 

CTT,0, 

°C 

Curve fit slope term, 

CTT,1, 

°C/ °C 

> 50 ≥ 0 –0.19470324 0.99014859 

< 50 0 0.06550710 0.98917044 

< 50 40 0.82278309 0.98841811 

<50 70 1.48151795 0.98401532 

D.2 Flow Quality Comparison 

The interim calibration data are also used to check the consistency in the test section flow quality 

from the previous full aero-thermal calibration. For most of the flow quality parameters this becomes a 

comparison of centerline data since the 9-ft survey rake is only mounted on centerline for the interim 

calibration; however, for total temperature the entire test section cross section is compared as the RTD 

array provides complete coverage during both the interim and full calibrations. 

Figure D.3 shows the centerline Mach number comparison data for the 2004 full aero-thermal 

calibration and the 2005 interim aero-thermal calibration. The Mach number data collected from the 9-ft 

survey rake were normalized by the bellmouth Mach number. Data were also included to show the 

spraybar air injection effects. The y-axis scale on the Mach number ratio plots was expanded to show very 

minute detail in the flow field. A ΔM band was added to each plot to show the absolute variation in the 

data. 

In general, there is no significant difference in the centerline Mach number ratio distributions between 

the 2004 and 2005 calibration tests. For most every case there is nearly an exact match core portion of the 

test section between the data in the sets from each year for the Mach number and spraybar air setting. 

There are two points of slight difference between the 2004 and 2005 data. First is near the inside test 

section wall where the 2004 data is consistently lower than the 2005 data. However, the difference is only 

0.001 to 0.003 in Mach number. The other area is in the right-center of the test section (60 to 90 in. from 

the inside wall) where the 2004 data is slightly higher than the 2005 data, but again by only 0.001 to 

0.003 Mach number. Overall, the 2005 data shows a flatter Mach number profile for all test conditions. 

NASA/CR—2010-216733 85
Total temperature survey data measured using the 2–D RTD Array is shown in Figure D.4. In general 

the trends in the total temperature data are fairly consistent between the 2004 and 2005 tests. There is 

some improvement seen in the temperature distributions in terms of reduced spread in the data measured 

during the 2005 test, particularly at lower airspeed settings. Since there were not changes to the tunnel, 

this improvement is attributed primarily to allowing more time for the temperature conditions to stabilize. 

Figure D.5 compares the standard deviation in the total temperature measurements over the survey plane. 

This data also indicates little difference between the 2004 and 2005 data sets, except again for the low 

speed settings. 

Figure D.6 shows the resultant flow angle vectors collected using the 9-ft survey rake at 36-in. above 

the test section floor (centerline). This data provides a comparison between the 2004 and 2005 test results 

over the airspeed range of the IRT and at spraybar air pressure settings of 0 an 70 psig; in addition, Pair of 

40 psig data collected in 2005 is also included. 

A qualitative inspection of the flow angularity data indicates no significant differences in test section 

flow quality between 2004 and 2005, although the flow angle vector magnitude does appear larger in the 

2005 data. In general, both the magnitude and direction of the flow angularity vectors is consistent 

between the 2004 and 2005 data for like test conditions, particularly for airspeed settings of 100 knots and 

higher. There are some noticeable differences in the flow angle vectors at the lower airspeed settings 

(below 100 knots), but at these low airspeeds, the flow angle probes are not as accurate. 

To provide a more qualitative comparison, average values and standard deviations were computed for 

the pitch and yaw flow angle components (Figure D.7 and Figure D.8). The data were reviewed first by 

using the entire data set (11 points across the test section) then by looking only at the core data (the seven 

center probes along the survey line). The following points were drawn from the comparison of the 

average pitch and yaw flow angle data (Figure D.7): 

• The average yaw angle data is repeatable across the IRT airspeed range for both the spraybar air 

off and 70 psig pressure settings. Even the large yaw angles recorded at the low airspeeds with 

the spraybar air pressure at 70 psig show a close match between the 2004 and 2005 tests. 

• The core data indicates that the average yaw angle is about –0.02° for airspeed settings of 

100 knots and higher for both the 2004 and 2005 data. 

• There is a mismatch in the average pitch flow angle magnitude between the 2004 and 205 tests 

seen in both the full-span and core data sets. The core data from 2004 indicate that the average 

pitch flow angle is about 0.1° over the airspeed range; however the 2005 data shows that the 

average pitch angle increases form about 0.1° at 100 knots to about 0.5° at 300 knots. The cause 

for this change in average pitch angle was not determined. 

• Spraybar air effects do not have a significant effect on the average flow angle values (pitch or 

yaw angles) for airspeeds of 130 knots and higher. 

Evaluation of the standard deviation in the flow angle data indicates that there were no significant 

changes in the amount of variation in the pitch and yaw flow angle distribution between 2004 and 2005. 

The pitch and yaw standard deviations from 2004 and 2005 are very consistent over the airspeed range of 

the IRT (Figure D.8). These data also indicate for most the cases; the variation in flow angle is much 

lower in the core of the test section. This is particularly true for the yaw angle data with the spraybar air 

pressure set at 70 psig and the pitch angle data with the spraybar air off. The yaw angle deviation with 

spraybar air off is essentially the same for the full-span and core data. 

Comparison of turbulence measurements made in 2004 and 2005 are shown in Figure D.9. Similar 

turbulence levels and trends over the airspeed range are noted between the two data sets. The only 

discrepancy is at the lowest airspeed setting with spraybar air injection, where the 2005 measurements 

indicated a much higher turbulence level. 

NASA/CR—2010-216733 86
D.3 Calibration Curve Comparison 

Figure D.10 shows the aero-thermal calibration curves with flow field data from the centerline of the 

test section from both the 2004 and 2005 tests. The calibration curves were developed based on the 2004 

data. The flow field data is the average of the center seven probes on the 9-ft survey rake with the rake 

positioned at the test section centerline (vertical height of 36 in. above the test section floor). There is a 

very favorable comparison for all three calibration curves (Mach number, static pressure and total 

temperature). Figure D.10(c) also provides a comparison to the total temperature calibration curves based 

on the 2004 and 2005 calibrations. The 2004 calibration curve was based on entire data set (covered the 

entire airspeed range), while the 2005 calibration uses a family of curves to model the total temperature 

and take into account changes in the model over the airspeed range. For simplicity, only the 2005 total 

temperature calibration curve for airspeed greater than 50 knots is included in the comparison. 

° 

NASA/CR—2010-216733 87
Figure D.1.—Total temperature calibration curves for the IRT based on the 2005 aero-thermal interim 

calibration data. 

NASA/CR—2010-216733 88
Figure D.2.—Total temperature calibration curve coefficients for Uts <50 knots. This information is used to interpolate 

slope and intercept values for spray bar air pressure settings between 0 and 40 and between 40 and 70 psig. 

NASA/CR—2010-216733 89
Figure D.3.—Comparison of Mach number ratio distribution data across the tunnel at test section centerline. (a) Mnom= 0.066, Unom= 43 knots. 

(b) Mnom= 0.100, Unom= 65 knots. (c) Mnom= 0.133, Unom= 87 knots. (d) Mnom= 0.152, Unom= 100 knots. 

NASA/CR—2010-216733 90
Figure D.3.—Continued. (e) Mnom= 0.200, Unom= 130 knots. (f) Mnom= 0.234, Unom= 150 knots. (g) Mnom= 0.266, Unom= 175 knots. (h) Mnom= 0.306, 

Unom= 200 knots. 

NASA/CR—2010-216733 91
Figure D.3.—Continued. (i) Mnom= 0.336, Unom= 220 knots. (j) Mnom= 0.370, Unom= 240 knots. (k) Mnom= 0.402, Unom= 260 knots. (l) Mnom= 0.435, 

Unom= 280 knots. 

NASA/CR—2010-216733 92
Figure D.3.—Concluded. (m) Mnom= 0.470, Unom= 300 knots. 

NASA/CR—2010-216733 93
Figure D.4.—IRT test section total temperature distribution data collected using the 2–D RTD Array. (a) Uts=49.02 mph (42.6 knots), 

TT,davg=–29.8 °C, Pair=0 psig. (b) Uts=55.2 mph (48.0 knots), TT,davg=–29.8 °C, Pair=70 psig. (c) Uts=49.14 mph (42.7 knots), TT,davg=–17.6 °C, 

Pair=0 psig. (d) Uts=54.89 mph (47.7 knots), TT,davg=–17.5 °C, Pair=70 psig. 

NASA/CR—2010-216733 94
Figure D.4.—Continued. (e) Uts=49.71 mph (43.2 knots), TT,davg=0.2 °C, Pair=0 psig. (f) Uts=49.94 mph (43.4 knots), TT,davg=0.3 °C, Pair=40 psig. 

(g) Uts=49.59 mph (43.1 knots), TT,davg=0.2 °C, Pair=70 psig. 

NASA/CR—2010-216733 95
Figure D.4.—Continued. (h) Uts=116.46 mph (101.2 knots), TT,davg=–28.4 °C, Pair=0 psig. (i) Uts=118.76 mph (103.2 knots), TT,davg=–28.4 °C, Pair=70 

psig. (j) Uts=116.46 mph (101.2 knots), TT,davg=–16.3 °C, Pair=0 psig. (k) Uts=118.30 mph (102.8 knots), TT,davg=–16.3 °C, Pair=70 psig. 

NASA/CR—2010-216733 96
Figure D.4.—Continued. (l) Uts=116.57 mph (101.3 knots), TT,davg=–4.3 °C, Pair=0 psig. (m) Uts=116.80 mph (101.5 knots), TT,davg=–4.2 °C, 

Pair=70 psig. (n) Uts=116.34 mph (101.1 knots), TT,davg=1.3 °C, Pair=0 psig. (o) Uts=116.34 mph (101.1 knots), TT,davg=1.3 °C, 

Pair=70 psig. 

NASA/CR—2010-216733 97
Figure D.4.—Continued. (p) Uts=151.91 mph (132.01 knots), TT,davg=–15.4 °C, Pair=0 psig. (q) Uts=174.53 mph (151.6 knots), TT,davg=0.06 °C, Pair=0 

psig. 

NASA/CR—2010-216733 98
Figure D.4.—Continued. (r) Uts=233.38 mph (202.8 knots), TT,davg=–24.4 °C, Pair=0 psig. (s) Uts=233.61 mph (203.0 knots), TT,davg=–24.7 °C, Pair=70 psig. 

(t) Uts=235.10 mph (204.3 knots), TT,davg=–12.3 °C, Pair=0 psig. (u) Uts=234.87 mph (204.1 knots), TT,davg=–12.4 °C, Pair=70 psig. 

NASA/CR—2010-216733 99
Figure D.4.—Continued. (v) Uts=230.5 mph (200.3 knots), TT,davg=0.0 °C, Pair=0 psig. (w) Uts=231.19 mph (200.9 knots), TT,davg=0.0 °C, 

Pair=70 psig. 

NASA/CR—2010-216733 100
Figure D.4.—Continued. (x) Uts=253.11 mph (219.95 knots), TT,davg=0.67 °C, Pair=0 psig. (y) Uts=292.07 mph (253.8 knots), 

TT,davg=0.8 °C, Pair=0 psig. (z) Uts=291.61 mph (253.4 knots), TT,davg=0.7 °C, Pair=70 psig. 

NASA/CR—2010-216733 101
Figure D.4.—Concluded. (aa) Uts=338.33 mph (294.0 knots), TT,davg=–17.4 °C, Pair=0 psig. (bb) Uts=338.1 mph (293.8 knots), TT,davg=–17.7 °C, 

Pair=70 psig. (cc) Uts=338.9 mph (294.5 knots), TT,davg=–5.5 °C, Pair=0 psig. (dd) Uts=339.6 mph (295.1 knots), TT,davg=–5.6 °C, Pair=70 psig. 

NASA/CR—2010-216733 102
Figure D.5.—IRT test section total temperature standard deviation data collected using the 2–D RTD Array at the center line 

of the RTD Array. 

NASA/CR—2010-216733 103
Figure D.5.—Concluded. IRT test section total temperature standard deviation data collected using the 2–D RTD 

Array for the full RTD Array. 

NASA/CR—2010-216733 104
Figure D.6.—Comparison of test section centerline flow angularity vectors from the 2004 full calibration and 2005 interim calibration at 

various Pair settings. (a) Mnom = 0.066, Vnom= 43 knots. (b) Mnom= 0.100, Vnom= 65 knots. (c) Mnom= 0.133, Vnom= 87 knots. 

(d) Mnom = 0.152, Vnom = 100 knots. 

NASA/CR—2010-216733 105
Figure D.6.—Continued. (e) Mnom = 0.200, Vnom= 130 knots. (f) Mnom= 0.234, Vnom= 150 knots. (g) Mnom= 0.266, Vnom= 178 knots. 

(h) Mnom= 0.306, Vnom= 200 knots. 

NASA/CR—2010-216733 106
Figure D.6.—Continued. (i) Mnom = 0.336, Vnom= 220 knots. (j) Mnom= 0.370, Vnom= 240 knots. (k) Mnom= 0.402, Vnom= 260 knots. 

(l) Mnom= 0.435, Vnom= 280 knots. 

NASA/CR—2010-216733 107
Figure D.6.—Concluded. (m) Mnom = 0.470, Vnom= 300 knots. 

NASA/CR—2010-216733 108
Figure D.7.—Comparison of average data from flow angularity measurements made in the IRT test section (horizontal survey at centerline) during the 2004 and 

2005 test entries. (a) Full-span data (all eleven probes included). (b) Core data (center seven probes only). 

NASA/CR—2010-216733 109
Figure D.8.—Comparison of standard deviation data from flow angularity measurements made in the IRT test section (horizontal survey at centerline) during the 

2004 and 2005 test entries. (a) Full-span data (all eleven probes included). (b) Core data (center seven probes only).

NASA/CR—2010-216733 110
Figure D.9.—Comparison of 2004 and 2005 turbulence data (horizontal survey at centerline). 

NASA/CR—2010-216733 111
Figure D.10.—Comparison of the aero-thermal calibration curves created based on the 2004 and 2005 test results. 

(a): Mach number calibration curve comparison with 2004 and 2005 test data. 

NASA/CR—2010-216733 112
Figure D.10.—Continued. (b) Static Pressure calibration curves for the IRT based on the 2004 and 2005 calibration data. 

NASA/CR—2010-216733 113
Figure D.10.—Concluded. (c) Total temperature calibration curves for the IRT based on the 2004 and 2005 calibration data. 

NASA/CR—2010-216733 114
References 

1. “Calibration and Acceptance of Icing Wind Tunnels,” SAE Standards ARP5905, 2003. 

2. “Calibration of Subsonic and Transonic Wind Tunnels,” AIAA Recommended Practice R–093–2003, 

2003. 

3. Arrington, E.A. and Gonsalez, J.C., “Improvements to the Total Temperature Calibration of the 

NASA Glenn Icing Research Tunnel,” NASA/CR—2005-213875, AIAA–2005–4276. 

4. Gonsalez, J.C., Arrington, E.A., and Curry III, M.R., “Aero-Thermal Calibration of the NASA Glenn 

Icing Research Tunnel (2000 Tests),” NASA/CR—2001-210685, AIAA–2001–0233, 2001. 

5. Leary, William M., “We Freeze to Please: A History of NASA’s Icing Research Tunnel and the Quest 

for Flight Safety,” NASA/SP—2002-4226, 2002. 

6. Irvine, T.B., Kevdzija, S.L., Sheldon, D.W. and Spera, D.A., “Overview of the Icing and Flow 

Quality Improvements Program for the NASA Glenn Icing Research Tunnel,” AIAA–2001–0229, 

2001. 

7. Soeder, R.H., Sheldon, D.W., Ide, R.F., Spera, D.A. and Andracchio, C.R., “NASA Glenn Icing 

Research Tunnel User Manual,” NASA/TM—2003-212004, 2003. 

8. Gonsalez, J.C., Arrington, E.A., “Five-Hole Flow Angle Probe Calibration for the NASA Glenn Icing 

Research Tunnel,” NASA/CR—1999-202330, AIAA–96–2201, 1999. 

9. Ames Aeronautical Laboratory: “Equations, Tables and Charts for Compressible Flow,” NASA TR–

1135 (NASA TN–1428), 1953. 

10. Gonsalez, J.C., and Arrington, E.A., “Aerodynamic Calibration of the NASA Lewis Icing Research 

Tunnel (1997 Tests),” NASA/CR—2001-210809, AIAA–98–0633, 2001. 

11. Bruun, H.H.: Hot-Wire Anemometry, Principles and Signal Analysis, Oxford University Press, 1999. 

12. Assessment of Experimental Uncertainty with Application to Wind Tunnel Testing. Reston, VA, 

American Institute of Aeronautics and Astronautics, 1999. AIAA S–071A–1999. 

13. Assessing Experimental Uncertainty - Supplement to AIAA S–071A–1999. Reston, VA, American 

Institute of Aeronautics and Astronautics, 2003. AIAA G–045–2003. 

14. Coleman, H.W. and Steele, W.G. Experimentation and Uncertainty Analysis for Engineers. New 

York, Wiley, 1989. 

NASA/CR—2010-216733 115
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-07-2010 

2. REPORT TYPE

Final Contractor Report 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Aero-Thermal Calibration of the NASA Glenn Icing Research Tunnel (2004 and 2005 Tests) 

5a. CONTRACT NUMBER

NAS3-00145; NNC05CA95C 

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Arrington, E., Allen; Pastor, Christine, M.; Gonsalez, Jose, C.; Curry, Monroe, R., III 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 122711.03.07.03.02 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

Sierra Lobo, Inc. 

Cleveland, Ohio 44135 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-17310 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/CR-2010-216733 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 09 

Available electronically at http://gltrs.grc.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

A full aero-thermal calibration of the NASA Glenn Icing Research Tunnel was completed in 2004 following the replacement of the inlet 

guide vanes upstream of the tunnel drive system and improvement to the facility total temperature instrumentation. This calibration test 

provided data used to fully document the aero-thermal flow quality in the IRT test section and to construct calibration curves for the 

operation of the IRT. The 2004 test was also the first to use the 2-D RTD array, an improved total temperature calibration measurement 

platform. 

15. SUBJECT TERMS

Wind tunnel; Calibration; Flow quality 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

125 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18