DOT/FAA/AR-06/60 Air Traffic Organization NextGen & Operations Planning Office of Research and Technology Development Washington, DC 20591

Propeller Icing Tunnel Test on a Full-Scale Turboprop Engine Paul Pellicano Chris Dumont Tim Smith James Riley March 2010 Data Report This document is available to the U.S. public through the National Technical Information Services (NTIS), Springfield, Virginia 22161.

U.S. Department of Transportation Federal Aviation Administration

NOTICE

This document is disseminated under the sponsorship of the U.S. Department of Transportation in the interest of information exchange. The United States Government assumes no liability for the contents or use thereof. The United States Government does not endorse products or manufacturers. Trade or manufacturer's names appear herein solely because they are considered essential to the objective of this report. This document does not constitute FAA certification policy. Consult your local FAA aircraft certification office as to its use. This report is available at the Federal Aviation Administration William J. Hughes Technical Center’s Full-Text Technical Reports page: actlibrary.act.faa.gov in Adobe Acrobat portable document format (PDF).

Technical Report Documentation Page 1. Report No.

DOT/FAA/AR-06/60

2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle

PROPELLER ICING TUNNEL TEST ON A FULL-SCALE TURBOPROP ENGINE

5. Report Date

March 2010

6. Performing Organization Code

7. Author(s)

Paul Pellicano*, Chris Dumont, Tim Smith, and James Riley

8. Performing Organization Report No.

9. Performing Organization Name and Address

Flight Safety Group Paul Pellicano* Federal Aviation Administration Federal Aviation Administration William J. Hughes Technical Center Atlanta ACO

10. Work Unit No. (TRAIS)

Atlantic City Int’l Airport, NJ 08405 1895 Phoenix Blvd., Suite 45 Atlanta, CA 30349

11. Contract or Grant No.

12. Sponsoring Agency Name and Address

U.S. Department of Transportation Federal Aviation Administration Air Traffic Organization NextGen & Operations Planning Office of Research and Technology Development Washington, DC 20591

13. Type of Report and Period Covered

Data Report

14. Sponsoring Agency Code ACE-111

15. Supplementary Notes

16. Abstract This document presents data from a propeller icing test performed at the McKinley Climatic Chamber at Eglin Air Force Base, Florida. The test was proposed and sponsored by the Federal Aviation Administration Small Airplane Directorate. The test was designed to provide certification guidance to account for propeller performance in Title 14 Code of Federal Regulations Part 25 Appendix C icing conditions and in supercooled large drop conditions (SLD) in anticipation of SLD rulemaking. 17. Key Words

Propeller, Icing, Test, Thrust loss

18. Distribution Statement

This document is available to the U.S. public through the National Technical Information Service (NTIS), Springfield, Virginia 22161.

19. Security Classif. (of this report) Unclassified

20. Security Classif. (of this page) Unclassified

21. No. of Pages 80

22. Price

Form DOT F 1700.7 (8-72) Reproduction of completed page authorized

ACKNOWLEDGEMENTS

Several organizations and individuals contributed to the planning, implementation, and reporting of the full-scale propeller icing test. Without their dedicated efforts, the value of the test program would not have been fully realized. Hartzell Propeller, Inc. provided a custom propeller governor and propeller for the test at no cost to the Federal Aviation Administration (FAA). Mr. Brian E. Meyer provided technical guidance and was the designated Hartzell lead during the planning, teleconferences, and test. Ms. Suzanne Borden also participated in the planning and was present for part of the test. Mr. Eric Reinhart performed the test matrix calculations for the Hartzell propeller angle of attack, torque, and pitch angle, and contributed to appendix B. The MT Propeller Company provided several propellers for the test at no cost to the FAA. Mr. Gerd Muehlbauer, President of MT Propeller, was the MT Propeller lead during the planning meetings and the test. Mr. Martin Albrecht performed the angle of attack, torque, and pitch angle calculations for the MT propellers for the test matrix. Hamilton Sunstrand Division of United Technologies Corporation, represented by Mr. Chad Henze, and McCauley Propeller Systems of the Textron Corporation, represented by Mr. Scott Randle, participated in the planning of the test and provided valuable technical input on the test setup. Goodrich Corporation, represented by Mr. Alan Farhner, provided technical guidance for the deicing boot-related equipment during the planning meetings and teleconferences and provided technical input during the test. The unique capabilities of the McKinley Climatic Laboratory at Eglin Air Force Base were used for the test. Mr. Dwayne Bell and an outstanding team of engineers and technicians designed and built the engine test stand, put in place the complicated test configurations, calibrated the spray, and operated the chamber, tunnel, spray system, and other systems throughout the test. Mr. Bell wrote appendix A to this report. Aerospace Testing, Engineering, & Certification (AeroTEC), under contract to the FAA, supplied and installed equipment to measure and record the propeller blade angles during the test. They also recorded and displayed engine parameters. The blade angle measurements and the engine parameters were displayed real time during the test and were used to set the initial test conditions and to monitor the change. Mr. Jack Wolda and Mr. Kent Baines of AeroTEC did an outstanding job during the test, and AeroTEC president Mr. Lee Human, made valuable contributions during the planning of the test. Intercontinental Jet Service Corporation supplied the Garrett TPE-331 engine, along with the test stands, including controls and instrumentation, and installed cowling under contract to the FAA. Mr. Mark James was the lead during the planning phases, and Mr. Keith Chester Linnell was present during the tests and made valuable contributions.

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Mr. Vincent Reich, Mr. Chris Lynch, and Mr. Jim Firak of RS Information Systems, Inc. provided outstanding imaging support arranged under the FAA Interagency Agreement with National Aeronautics and Space Administration Glenn Research Center (NASA-GRC). In addition to the post-run photographs, the stop-action photography during the runs made a unique and extremely valuable contribution to documenting the propeller icing and shedding process. Dr. David Anderson of the Ohio Aerospace Institute, whose participation was also arranged using the FAA Interagency Agreement with NASA-GRC, performed the scaling calculations for the test runs planned with no ice protection. Dr. Andy Broeren of the University of Illinois at Urbana-Champaign performed the scaling calculations for the test runs with the thermal ice protection turned on. In addition, they played an important and valuable role in the planning of the test, and participated in planning meetings and teleconferences.

TABLE OF CONTENTS

Page EXECUTIVE SUMMARY xi 1. INTRODUCTION 1

2. BACKGROUND 1

3. TEST DESCRIPTION 4

3.1 Test Facility 4 3.2 Instrumentation and Data Acquisition 4

3.2.1 Imaging 4 3.2.2 Ice Accretion Measurements 5 3.2.3 Thrust Measurement 7 3.2.4 Engine and Propeller Parameters 7 3.2.5 Icing Conditions 8

3.3 Test Articles 8 3.4 Test Procedures 11 3.5 Test Conditions 13

3.5.1 14 CFR Part 25 Appendix C Icing Conditions 13 3.5.2 Supercooled Large Drop Icing Conditions 15

4. TEST RESULTS 15

4.1 Data Runs 15 4.2 Drag Due to Ice Accretion on Engine Stand 17 4.3 Spinner Lip Effect on Ice Accretion 17

5. SUMMARY 18

6. RECOMMENDATIONS 20

7. REFERENCES 20

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APPENDICES

A—McKinley Climatic Laboratory Support for the Federal Aviation Administration Large Propeller Icing Test B—Computation of Slipstream Drag Corrections for use With Icing Test Data Taken on the Hartzell 4-Blade 10282-5.3R Propeller C—Propeller Icing McKinley Laboratory Project Plan D—Run Summary

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LIST OF FIGURES

Figure Page

1 Video Capture From MU-2 Flight Test Showing Ice Accretion Along the Entire Span of the Propeller Blades 2

2 Incremental Drag From Clean Baseline Required to Match Flight Data Recorder-Measured Airspeed Loss for Three Turboprop Icing Events Compared to Simulated Ice Shape Flight Tests 4

3 Cutting and Tracing the Ice Contour 6

4 Tracing Locations of the Hartzell and MT Propellers 6

5 Blade Angle Measurement Equipment 8

6 Hartzell Service Condition Blade 1 9

7 Hartzell Service Condition Blade 3 10

8 Illustration of Test Procedure for Run 25 (A—engine start, B—wind on, C—torque set at 65%, D—torque set at 100%, E—torque set at 50%, F—spray on, G—deicing boots on, H—deicing boots off, I—wind/spray off) 12

9 The Effect of the Spinner Lip on Ice Formation 18

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LIST OF TABLES

Table Page

1 Planned Test Articles 8 2 Propeller Assemblies Tested 9 3 Deicing Boots Installed on Propeller Assemblies 10 4 Run Summary Matrix—Short Form 16 5 Summary of Thrust Reductions 20

LIST OF SYMBOLS AND ACRONYMS

α Angle of attack β Blade angle β0 Stagnation collection efficiency n0 Freezing fraction Ac Accumulation parameter AFB Air Force Base AMU Air make-up CFR Code of Federal Regulations EGT Exhaust gas temperature FAA Federal Aviation Administration gpm Gallons per minute IRT Icing Research Tunnel JW Johnson-Williams KCAS Knot calibrated airspeed KTAS Knot true airspeed LWC Liquid water content MCL McKinley Climatic Laboratory MHIA Mitsubishi Heavy Industries America MVD Median volume diameter NASA-GRC National Aeronautics and Space Administration Glenn Research Center NI National Instruments NTSB National Transportation Safety Board OAT Outside air temperature PMEL Precision Measurement Equipment Laboratory psig Pounds per square inch gauge rpm Revolutions per minute SLD Supercooled large drop TTL Transistor-transistor logic UIUC University of Illinois at Urbana-Champaign Vdc Volt direct current VFD Variable frequency drives

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EXECUTIVE SUMMARY This report presents data from a collaborative propeller icing test performed at the McKinley Climatic Chamber at Eglin Air Force Base, Florida, in November 2006. The test was proposed and sponsored by the Federal Aviation Administration (FAA) Small Airplane Directorate. The test was designed to provide certification guidance to account for propeller performance in Title 14 Code of Federal Regulations Part 25 Appendix C icing conditions and in supercooled large drop (SLD) conditions in anticipation of SLD rulemaking. The collaborators in the research included the FAA, Hartzell Propeller, Inc., MT Propeller Company, Eglin Air Force Base, Goodrich Corporation, the National Aeronautics and Space Administration, Ohio Aerospace Institute, and the University of Illinois at Urbana-Champaign (UIUC). Hamilton Sunstrand, a Division of United Technologies Corporation, and McCauley Propeller Systems of Textron Corporation made valuable contributions to the planning of the test. A Honeywell TPE-331 turboprop engine was tested with several propeller combinations on a thrust stand while exposed to a simulated in-flight icing environment. Liquid water content, median volume diameter, ambient temperature, and airflow velocity were recorded. Thrust and propeller blade angle were measured continuously. Engine revolutions per minute (rpm), torque, and other parameters were also recorded. Ice shapes were documented in three ways: (1) ice shapes tracings were made at the 50%, 75%, and boot center locations, (2) photographs were taken of the resultant ice shapes, and (3) stop-action video of the rotation propeller were taken in real time. There were two significant facility limitations of the test configuration. The maximum airflow velocity for the test configuration was limited to 100 knots. The pressure was not flight altitude pressure but simply sea level pressure. Since the airspeed for the test condition was lower than the desired flight reference condition, the power was reduced to achieve the reference blade angle of attack at the blade radius of interest while maintaining the reference rpm. The rpm was required to be the same for both the reference and test conditions so that the G-field would be the same. Due to these limitations, the conditions actually tested had to be scaled from the reference flight conditions. For the runs with deicing off, the David Anderson-Paul Tsao method was followed, with David Anderson performing the calculations. For the runs with deicing on, a method developed by Cessna, and modified by UIUC, was followed, and the calculations were performed at UIUC. This full-scale propeller icing test provided a unique and extensive data set for studying propeller icing. In addition to the numerical data, the exceptional imaging data obtained made a valuable contribution to the documentation of propeller icing.

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The test results supported the assumption of a nominal thrust penalty on the order of 10% for icing certification. The test results also showed that by simulating SLD conditions, it was possible to approximately match the ice accretion on a nearly full blade span.

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1. INTRODUCTION.

This report presents the results of full-scale propeller icing tests that were performed to achieve the following objectives: 1. Document propeller ice accretions in Title 14 Code of Federal Regulations (CFR) Part 25

Appendix C icing conditions.

• Leading edge accretions—size and shape, span location, and shedding frequency as a function of icing conditions, blade material, blade condition, and revolutions per minute (rpm)

• Runback ice accretions—size, shape, and location as a function of icing

conditions, blade material, blade condition, rpm, deicing schedule and power 2. Document propeller ice accretions in supercooled large drop (SLD) icing conditions and

determine if a propeller ice accretion from a Mitsubishi MU-2 flight test propeller in suspected SLD icing conditions can be approximately duplicated. (See section 2.)

3. Determine differences in thrust between non-iced and iced propellers. 4. Evaluate differences between reference- and scaled-icing conditions. 5. Determine if a nominal propeller thrust reduction can be proposed for certification. In November 2006, icing tests were conducted with various propellers on a Honeywell TPE-331 turboprop engine in the McKinley Climatic Chamber. The test matrix had to be substantially truncated and revised due to the time required for setup and calibration, which took several days longer than had been scheduled. Consequently, although the main test objectives were met, others were not. 2. BACKGROUND.

Airplane icing certification research has historically used simulated ice shapes on fixed parts of the airframe, such as the wing and tail leading edges, to evaluate airplane performance in icing conditions. The performance of the propeller in icing conditions had not been evaluated during certification unless a significant problem was identified during flight tests in natural icing conditions. However, separating performance effects due to propeller icing from those due to airframe icing is difficult for flight tests performed in natural icing conditions. Certification data for propeller icing included an analysis to justify the chordwise and spanwise extent of propeller deicing boots. The service ceiling in icing conditions on some propeller-powered airplanes, as measured by simulated ice shapes on the airframe, has been shown to be less than minimum enroute altitudes of some low-altitude airways in the western U.S. Therefore, it is important that the contribution of propeller ice accretions to airplane climb capability in icing conditions be understood.

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In 2000, Mitsubishi Heavy Industries America (MHIA), Hartzell Propeller, and Goodrich briefed the Federal Aviation Administration (FAA) Small Airplane Directorate on flight tests conducted by MHIA on the MU-2 model in natural icing conditions. The test objectives were to evaluate the airframe pneumatic deicing boot activation methods. During one icing encounter, the airspeed decayed dramatically, losing 40 knots in 1 minute and 25 seconds [1]. The flight crew assessed the airframe ice accretion to be small. A postflight analysis of the in-flight video showed ice accretion along the entire span of the propeller blades, as shown in figure 1. There was one area on the outboard portion of the deicing boots where it appeared as if the ice was ready to shed. An analysis by Hartzell estimated propeller efficiency loss to be on the order of 15%-20%, and the radial extent of ice had the biggest performance impact. Although the airplane was not equipped with drop size measuring equipment, the icing conditions were suspected to be SLD conditions based on visual cues.

Ice Accretion

Figure 1. Video Capture From MU-2 Flight Test Showing Ice Accretion Along the Entire Span of the Propeller Blades

A literature review conducted by the Small Airplane Directorate revealed more information and issues relating to propeller icing, as discussed below. 1. A presentation [2] to the SAE AC-9C Aircraft Icing Technology Subcommittee,

September 18-22, 1989, detailed experiences on the Saab model 340. In the winter of 1984-1985, Saab received comments from Saab model 340 operators such as “unable to climb above 12,000 feet in icing” and “aircraft decelerated in level flight and we had to descend to avoid stalling.” Saab conducted full-scale propeller ice shape tunnel tests and natural icing flight tests. The natural icing flight tests showed that at certain temperatures, runback ice behind the deicing boot occurred on both sides of the blades. The tunnel tests showed that 1-mm-high roughness at the leading edge resulted in a 3% efficiency loss, whereas the efficiency loss due to the runback ice behind the boots was on the order of 17%-20%. The Saab model 340 propeller deicing timing was modified to be a function of outside air temperature (OAT) as a result of these tests.

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2. Turboprop-powered airplane manufacturers have had similar experiences in certification. The propeller deicing system on one turboprop airplane had dual settings, based on temperature.

3. The Airplane Flight Manual for another regional turboprop states: “While in climb or

cruise flight, propeller icing may cause a significant decrease in indicated airspeed or climb rate, even though ice accretion on the leading edges of the wings, the empennage, or the engine inlets may be relatively light. At constant power an airspeed loss up to 25 knots is normal.”

4. Propeller runback ice was documented by photographs on at least one other regional

turboprop in service, although the effects on performance were not indicated. 5. Some turboprop-powered airplane accident investigations have indicated loss of control

that resulted from insufficient acceleration in airspeed during stall-warning recovery and drag that was higher than could be accounted for by airframe ice accretions. For example, Comair Flight 5054 lost 40 knots in about 3 minutes and stalled after an icing encounter in cruise at 17,000 feet altitude. Figure 2 shows how the drag count had to be increased (from clean airplane drag) as a function of time to simulate the data recorded on the flight data recorder for the event. The same drag increase for two other events on the same turboprop, Comair Flight 3272 and Westair, are shown in figure 2. The incremental drag due to simulated ice shapes representing critical, 14 CFR Part 25 Appendix C icing conditions, as determined from flight test data at 160 knot calibrated airspeed (KCAS), is shown in the lower right corner of figure 2. This simulated ice shape incremental drag ranged from 75 to 89 counts and cannot account for all the performance degradation that was experienced on Comair Flight 5054 and the other events. The National Transportation Safety Board (NTSB) stated that the probable cause of this accident was the flight crew’s failure to maintain airspeed during an encounter with severe icing. The NTSB accident report also states that the airplane may have encountered SLD conditions. Although airframe ice accretions caused by icing conditions outside appendix C, propeller ice accretions should not be ruled out.

Prior to this test, all SLD experimental research work was associated with ice accretions on fixed airfoils. Developing SLD engineering tools for propellers was essential to the practical implementation of anticipated new SLD regulations under consideration by the Ice Protection Harmonization Working Group of the Aviation Regulatory Advisory Committee.

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Del

ta D

rag

(Cou

nts)

Figure 2. Incremental Drag From Clean Baseline Required to Match Flight Data Recorder-Measured Airspeed Loss for Three Turboprop Icing Events Compared to Simulated Ice

Shape Flight Tests [3] 3. TEST DESCRIPTION.

This section describes the test facility, instrumentation and data acquisition system, test articles, test procedures, and icing conditions. 3.1 TEST FACILITY.

The test was conducted in the McKinley Climatic Laboratory at Eglin Air Force Base (AFB) in Florida. The laboratory, the complex propeller test setup, and the calibration procedures are described in appendix A. 3.2 INSTRUMENTATION AND DATA ACQUISITION.

3.2.1 Imaging.

Two time-lapse digital cameras were used to record the tests: a Vision Research Phantom V9.0 high-speed digital camera and a Pulnix TM-1320-15CL camera. The camera had a monochrome 1600- by 1200-pixel CMOS sensor with a minimum exposure time of 2 microseconds. While this is a high-speed camera by design, the camera was operated in “frame sync” mode. This allowed the camera to capture one frame for every transistor-transistor logic (TTL) pulse it received. The majority of the runs were captured at 720 by 720 pixels, with an interval of one picture per second. Depending on the cloud thickness, propeller angle, and propeller color, the exposure time of the camera was between 25 and 50 microseconds. A Nikon 85-mm lens provided a close-up view of the deicing boot, with a field of view of approximately 3 by 3 feet.

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The Pulnix TM-1320-15CL camera had a monochrome 1300- by 1030-pixel progressive scan charge-coupled device sensor with a minimum exposure time of 1/16,000 second (62.5 microseconds). The camera was also operated in a similar frame sync mode, and received a standard TTL pulse from National Instruments (NI) DASYlab software. The majority of the runs were captured at 1296 by 1018 pixels with an interval of one picture every 2 seconds. The exposure time remained constant at 62.5 microseconds. This camera was used to document the entire propeller length. The camera used an adaptor to mount a 35-mm Nikon lens onto the C-mount Pulnix camera body. This provided a field of view of approximately 6 by 6 feet. Both cameras were housed in enclosures to prevent ice accumulation and minimize vibration due to wind. The cameras were synchronized with the standard pulse signal output normally used for propeller synchronization. This provided a low-voltage electronic pulse once per revolution of the propeller. The cameras required a standard TTL ±5 volt signal for triggering, so the pulse was conditioned using a custom program written with NI DASYlab software. The propeller blades were illuminated using four Arri 1200-watt compact (Hydrargyrum medium-arc iodide) HMI gas discharge arc lamps. These lamps have a daylight balanced color temperature of 6000 K and use a Fresnel lens to concentrate large amounts of light on a small area. For photographic documentation following each test run, a Kodak SLR/n digital camera with a Nikon SB-800 flash was used. The camera had a 3000- by 4500-pixel sensor capable of producing 14-MB RAW files. All images were captured at full resolution and later saved as JPEG files. A 105-mm Nikon macro lens was used for posttest close-up imaging of the ice shapes. For general documentation, a combination of zoom lenses was used. 3.2.2 Ice Accretion Measurements.

Tracings and thickness measurements of the ice accretions on the Hartzell and MT propellers were taken at the conclusion of most runs. Tracings were accomplished by first cutting a chordwise slot into the ice with a thin (0.09 inch) sheet of heated copper. A template of the blade’s leading edge was then inserted into the slot. Using a pencil, the adjacent ice contour was traced onto the template, as shown in figure 3. Tracings were typically taken at midboot, 0.5 radius, and 0.75 radius on the Hartzell propeller, and only at the midboot location on the MT propeller (figure 4). There were two exceptions: on run 4C, a tracing was taken at 0.55 radius, and on run 26, an additional tracing was taken near 0.75 radius (noted as 0.75 “inner” on the tracing diagrams and photograph of these two runs).

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Blade Template

Heated Copper

Figure 3. Cutting and Tracing the Ice Contour

Hartzell propeller

Tracing at 0.75 radius

Tracing at 0.5 radiusTracing at

mid-boot

MT propeller

Tracing at mid-boot

Figure 4. Tracing Locations of the Hartzell and MT Propellers

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Ice thickness measurements were taken with a caliper. The thickest part of the ice was measured near the blade’s leading edge where the tracing slots were cut. The ice tracings are contained in the data. Ice roughness thickness measurements were also taken along the blade at various locations. 3.2.3 Thrust Measurement.

The engine assembly was mounted on a thrust table. See sections A.1.3 and A.2.1 in appendix A for a detailed discussion. 3.2.4 Engine and Propeller Parameters.

On the turboprop engine, the engine torque, rpm, and exhaust gas temperature (EGT) were monitored, and the torque and rpm were recorded. The engine rpm was measured from an electric pulse signal generated from the propeller synchronizer output, which generates one pulse for every revolution of the engine. The engine torque was measured using the engine’s standard torque output signal. The EGT was measured using a thermocouple. Engine vibration was monitored during the test for safety reasons. Vibration was measured using a three-axis accelerometer mounted on the engine stand. Vibration was never an issue for any of the runs. The propeller beta angle, which is defined as the angle between the chordline of a reference blade section and the plane of rotation, was measured in real time using string pots mounted in the propeller hub, as shown in figure 5. The reference blade station for the Hartzell propellers was the 30-inch (76.2-cm) radius, and for the MT propellers, it was 37.008 inches (94 cm). The string pots were calibrated to measure the beta angle at these locations. The data from the string pots were transmitted via a wireless transmitter mounted on the forward spinner backing plate. This system was calibrated for the full range of propeller blade travel and was found to be accurate to within 0.5° throughout its range. This calibration took into consideration the test temperatures and the catenary curve on the string pot wire due to the rotation speed of the engine. The blade angle measurement was used at the start of each run to adjust the torque to achieve the reference local angle of attack at the blade radius of interest. The blade angle was also monitored in real time as the ice accreted to determine how much pitch reduction was required for the propeller governor to maintain the rpm.

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Figure 5. Blade Angle Measurement Equipment (Courtesy of AeroTEC) 3.2.5 Icing Conditions.

Target liquid water content (LWC) and median volume diameter (MVD) were determined by setting the spray system water flow rate, atomizing air pressure, and water pressures, as defined by the calibration maps. (See section A.2.7 in appendix A.) The water flow rate and the water and air pressures were monitored and recorded. Chamber static temperature and humidity were also recorded. Wind speed was based on the fan calibrations, as described in appendix A. 3.3 TEST ARTICLES.

The planned test articles are listed in table 1.

Table 1. Planned Test Articles

Engine Maximum Propeller

(rpm) Propeller Blades

Deicing Schedule

2 aluminum 90/90 3 composite 90/90

TIO-540-J2BD 2500

4 composite 90/90; continuous 4 aluminum 34/34/68; 10/60;

20/60; 90/90 TPE331-10-511M 1591

5 composite Continuous

Due to delays in setup, the amount of time allotted for the test was reduced from 12 to 7 days. As a result, the TIO-540-J2BD reciprocating engine was not tested, and ice accretion as a

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function of rpm could not be evaluated. All TPE331-10-511M turboprop engine test configurations were accomplished, except those with 10/60 and 20/60 deicing schedules. Table 2 lists the Hartzell and MT propeller assemblies that were tested on the Honeywell TPE331-10-511M turboprop engine.

Table 2. Propeller Assemblies Tested

Hub Model Blade Model No.

Blades Blade Material Diameter (inches)

HC-B4TN-5( )L LT10282NSB-5.3R 4 Aluminum 98 MTV-27-2-E-C-F-R(G) CFRL250-55 5 Composite 98

The propellers are constant-speed propellers, and a propeller governor maintains the rpm set by the operator. Increasing engine throttle primarily results in an increase in torque but also results in an increased blade angle as the propeller governor maintains the rpm. The propellers are typically operated at 1591 rpm, or 100% rpm, for most phases of flight, although in some cruise conditions, the propellers are operated at 96%, or 1520 rpm. The original plans were to conduct all runs at 96% and repeat certain runs at 100%. These repeats could not be accomplished due to the shortened schedule. Since climb is usually accomplished with 100% rpm, a 100% rpm was targeted during the ice off and on thrust measurements at 65%, 80%, and 100% torque, while 96% was targeted during most of the test. This was accomplished for most runs. The MT propeller was a new propeller. Both new and used blades were installed in the Hartzell propeller assembly to evaluate the effects of surface roughness on ice adhesion, due to normal blade erosion in service. Blades in the same condition, new or “service condition,” were opposite one another. The blades were numbered for identification. Blades 2 and 4 were new blades, and blades 1 and 3 were service condition blades. As shown in figures 6 and 7, blade 1 had light paint erosion at the extreme leading edge but no other damage. Blade 3 had slightly more erosion than blade 1. Both service condition blades were considered to be in above-average condition for midtime blades (1468 hours into a 3000 flight-hour overhaul interval).

Figure 6. Hartzell Service Condition Blade 1

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Figure 7. Hartzell Service Condition Blade 3 Table 3 lists (in the order they were tested) the deicing boots that were installed on the propeller assemblies. These propeller deicing boots are electrothermal deicing boots, consisting of resistive elements sandwiched between neoprene layers that are bonded to the inboard leading edge of each blade. The deicing boots are designed to heat the surface above freezing to debond the ice from the surface and facilitate shedding. The Goodrich deicing boots cycle off and on, which allows ice to build and then shed once power is applied. For example, the Goodrich P/N 4E1188-7 deicing boot inboard zone is heated for 34 seconds, then the outboard zone for 34 seconds, and then the heat is turned off for 68 seconds. For the Goodrich 4E2837-10 deicing boot, two opposing blades are heated for 90 seconds, and then the remaining two opposing blades are heated for 90 seconds. The MT deicing boots operate continuously.

Table 3. Deicing Boots Installed on Propeller Assemblies

Propeller Assembly

Deicing Boot P/N

Deicing Timing Deicing Schedule

Hartzell Goodrich 4E1188-7

34/34/68 Inboard zone/outboard zone

MT Propeller Goodrich 4E4215-4

Continuous All blades

Hartzell Goodrich 4E2837-10

90/90 Opposing blades

The first test article represented the certificated configuration of the Mitsubishi MU-2B airplane. The second test article was not a certificated configuration, although the MT propeller planned for the reciprocating engine test was a certificated configuration. The third test article represented the certificated configuration of the Dornier 228 and Piper PA-42-1000. Deicing power was supplied by two 28 Volt direct current (Vdc) NI relay modules controlled by a LabVIEW® program. The program allowed the user to first select the proper deicing on/off times, then start the deicing when the cloud spray was turned on. The NI relay modules switched larger relays, located on the engine, that were connected directly to the deicing heating units. Boot1 (Relay 1) controlled the outboard boot and Boot2 (Relay 2) controlled the inboard boot.

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3.4 TEST PROCEDURES.

The turboprop engine test procedures were as follows: • Because the Eglin AFB fire guard was required to be present when the engine was

started, a fire truck was located directly outside the McKinley Climatic Chamber. A personnel count was required prior to starting the test. The operator of the Chamber air make-up system was notified when the test was about to start.

• The engine was started, the wind was turned on, and the engine was stabilized at a cruise

setting, typically at 36% torque. • After the wind reached maximum velocity, thrust measurements were taken at various

torque levels with 100% propeller rpm, to obtain clean propeller thrust values. (Note, in the initial runs this was not done, as the objective in those runs was to document ice accretions and record the thrust decrement.)

• Starting at run 15A, the fifth run, it was decided to record thrust measurements at or near

full torque, which makes it easier to determine a decrement due to thrust. Initially, the thrust was measured at 80% and 100% torque. Starting at run 4A, thrust measurements were recorded at 65% and 100% torque instead. The rationale was that these torque levels at tunnel airspeed equated to the 0.75 radius angle of attack. They were very close to flight values with 100% torque at 130 KCAS (MU-2 automatic autopilot disconnect airspeed) and 180 KCAS (MU-2 AFM minimum icing airspeed), respectively.

• The engine torque was reduced to achieve target blade angle at the target rpm, usually

96% rpm. The engine parameters during cloud spray represented cruise values for the MU-2. Thrust was noted when the engine parameters stabilized and was recorded continuously during the test.

• The spray and propeller deicing were turned on simultaneously. • The boot and blade images were monitored. It was easy to get a specific blade isolated in

the images for several minutes. The cloud spray was continued until three to four sheds had been observed. The one exception was run 18, which was an intermittent maximum icing condition.

• In the runs up to run 4A, the cloud spray was turned off and thrust measurements were

taken at various torque levels. Starting at run 21B, the measurements were taken with the spray still on. The rationale was that it represented an attempt to climb out of an icing cloud. After run 21A, thrust with propeller ice accretion was measured with the spray off. The spray was turned back on and the thrust was measured again.

• The engine torque was reduced to a cruise power setting. The cloud spray, wind,

propeller deicing, and the chamber air make-up systems were turned off.

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• After all run data were obtained, ice was removed from the propellers, nacelle, and engine stand with warm shop air. It was also necessary to sweep out the open jet nozzle, which accumulated ice crystals. The chamber temperature was either maintained or adjusted to the planned next point. Overnight, the chamber temperature was maintained at 40°F, allowing the ice accretions on the thrust table to melt.

The procedure is illustrated in figure 8 for run 25.

-10-505

10152025303540

Beta, Deg

0102030405060708090

100110120

331 Torque, Percent

800900

100011001200130014001500160017001800

Engine RPM , RPM

-0.5

0.0

0.5

1.0

1.5Wind, 1=ONSpray, 1=ON

-0.5

0.0

0.5

1.0

1.5

0:00 5:00 10:00 15:00 20:00 25:00

Boot1, 1 = Deice ON

A

D

F

C

B

C

D

E

I

-500

0

500

1000

1500

2000

11:09:36 11:16:48 11:24:00 11:31:12 11:38:24 11:45:36

0

0.2

0.4

0.6

0.8

1

1.2

H G

Beta Angle (deg)

Torque (%Max)

RPM

Wind/Spray

Deicing Boots

A

Thrust (lbs)

D D

C C

C

D

C

D

Figure 8. Illustration of Test Procedure for Run 25 (A—engine start, B—wind on, C—torque set at 65%, D—torque set at 100%, E—torque set at 50%, F—spray on, G—deicing boots on,

H—deicing boots off, I—wind/spray off)

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3.5 TEST CONDITIONS.

3.5.1 14 CFR Part 25 Appendix C Icing Conditions.

This section explains the relationship between the matrix and the summary. The planned test matrix is provided in appendix C. The run summaries are given in appendix D. There are two reference true airspeeds for the turboprop engine in the test matrix. The 186 knot true airspeed (KTAS) represented the minimum icing airspeed for the MU-2, and the 240 KTAS represented the true airspeed during a flight test encounter in suspected SLD. The reciprocating engine reference speed (135 KTAS) represents the minimum icing airspeed for the Piper Malibu. The Piper Malibu airspeed was used because the reciprocating engine and the three propellers were all approved at that speed. The tunnel velocity was 100 KCAS. As a result, a reference blade radius had to be chosen to scale the test point. Two reference radii were chosen—one at midspan of the deicing boot and one at 75% radius. The 75% radius was chosen because it is near the span location of maximum thrust. The engine and propeller test conditions were determined for the test matrix as follows: • The desired flight conditions (altitude, KCAS) were specified for the aircraft being

simulated.

• The torque (expressed as percent of maximum torque for the turboprop) and rpm were determined for the desired flight conditions from the Pilot Operating Handbook (POH) for the aircraft being simulated.

• The propeller manufacturer calculated the angle of attack (α) of the blade at a specified location for cruise at the given flight conditions. For the test runs with deicing, the specified location was the center of the deicing boot. For the test runs without deicing, the specified location was 75% radius.

• The propeller manufacturer also calculated the torque and the blade angle (β) (referred to as beta angle) required to achieve the given α for the test conditions at the given rpm. β was measured continuously throughout the test run. Once the test began, the torque calculated by the manufacturer was adjusted to achieve the desired β. Generally, only small adjustments were needed.

• The test run was conducted at sea level pressure and 100 KTAS in the McKinley climatic chamber. For the test run, rpm and α at the radius of interest matched the corresponding values at the flight conditions.

Two types of runs were planned: propeller deice-off and propeller deice-on. Propeller deice-off runs were designed to document ice accretions along the span of the blade. For these runs, the reference radius was 75%, and the scaling method used was the same method

13

developed for unprotected fixed wings [4]. Calculations were done at the stagnation point. Test LWC was set the same as the reference; test MVD was found by matching Stagnation Collection Efficiency (β0); test temperature was found by matching freezing fractions (n0); and test spray time was found by matching accumulation parameter (AC). The scaling does not take into account either aerodynamic or centrifugal shedding. Note that no scaling technique has been validated for propeller icing. The differences between scale and reference values were small for the 186 and 130 KTAS runs. Since many runs resulted in zero freezing fractions, predicting no ice accretions, calculations were also performed for the 50% radius. Propeller deice-on runs were designed to document ice accretions on and behind the deicing boot. For these runs, the reference radius was the midspan location of the deicing boot, and the scaling method used was the same method developed for a fixed-wing, bleed air, anti-icing system [5 and 6]. Altitude must be considered when testing thermal systems in a sea level tunnel, since heat and mass transfer coefficients decrease with altitude. The scaling method could not be validated because the reference altitude could not be run in the tunnel. However, the runback ice accretions observed in the tunnel compared well to flight test observations on a business jet wing. The applicability to propellers or thermal cyclical deicing systems is unknown. Airspeed, MVD, and LWC were adjusted to match the scaling parameters. The convective heat transfer coefficient, the modified droplet inertia parameter, water catch rate, and the relative heat factor served as the scaling parameters. Matching the convective heat transfer coefficient resulted in a range of test airspeeds, lower and higher than the 100-KTAS maximum tunnel speed. It was decided to perform all runs at the maximum tunnel speed, since the resulting difference in the test and reference convective heat transfer coefficient was small. Local velocities are also driven by propeller rpm, so changes in tunnel speed have less effect compared to fixed-wing tests. For the runs with the 186 and 240 KTAS reference speeds, a second scaling technique was proposed. The test MVD and LWC from the first scaling technique were used, and the test temperature was varied to match the test freezing fraction with the reference freezing fraction. Since the test was delayed due to setup, many planned runs were not accomplished. Therefore, it was decided early in the test to perform all runs with the propeller deicer on for the following reasons: (1) Thrust decrement due to ice accretion was a primary data output, and propeller deice-on represented an operational configuration. (2) The primary objective of the propeller deice-off runs was to document ice accretion along the span as a function of rpm, freezing fraction, and other parameters. The results from the initial runs showed that this could be done with the propeller deicer on, since ice accreted on and outboard of the deicing boots during these runs. Some of the conditions that had been planned for deicer-off were run with deicer-on. For example, run 3 from the test matrix had been planned for deicer-off, but it was performed with deicer-on. The convention in the run summary was to add a letter to the original run number, so it was called run 3A. Test runs that were originally planned for propeller deicer-on were performed at the reference icing (MVD, LWC, OAT, and spray time) conditions rather than the scaled icing conditions in the test matrix. This was done because the boot midspan was the reference radius for these runs. Since the 75% radius was critical for determining thrust loss, the reference icing conditions were used because scaling showed the calculated icing conditions to be very close to the reference

14

conditions at 75% radius. Although the reference airspeed and pressure were not accounted for (since the tests were limited to 100 KTAS at the altitude of the test facility), the effects on ice accretion at 75% radius were predicted to be negligible. However, for the heated ice protection system, airspeed and pressure may have more of an effect than for unprotected surfaces, and the effects of these parameters (airspeed and pressure) on the resulting ice shapes on the boot region is unknown. Starting with the repeats for run 15, the effects of ice accretion on propeller thrust were recorded by the test conductor by momentarily running the engine at a higher torque value than for the test condition. 3.5.2 Supercooled Large Drop Icing Conditions.

As discussed in section 2, an in-flight encounter with suspected SLD icing conditions caused ice to form along the entire radius of the propeller blades, which contributed to a large performance loss. One of the objectives of the test program was to spray SLD size drops on the propeller to determine if ice accretion and thrust losses were different than in 14 CFR Part 25 Appendix C icing conditions. The discrete point calibration method, as described in appendix A, was used to determine the SLD conditions that could be generated. A calibration point that achieved conditions of 96 μm MVD and 0.36 g/m3 LWC was chosen because it approximated the proposed Appendix X freezing drizzle MVD and LWC, with MVD>40, as defined in the Ice Protection Harmonization Working Group report [7]. However, the drop size distribution that was tested could not duplicate the bimodal distribution characteristic of the Appendix X MVD>40. Two static temperatures were arbitrarily chosen—12°F to simulate the MU-2 encounter and 24°F to simulate a warm SLD condition. 4. TEST RESULTS.

Due to the voluminous nature of the propeller icing test data, more detailed information is provided in an accompanying DVD titled, “Propeller Icing Tunnel Test on a Full-Scale Turboprop Engine Data Disk.” This DVD is available upon email request to the FAA William J. Hughes Technical Center Reference and Research Library (actlibrary@faa.gov). 4.1 DATA RUNS.

Table 4 shows the short form of the run summary matrix. The full run summary matrix is given in appendix D, which shows the parameters at the initiation and conclusion of spray and during the thrust measurements.

15

Table 4. Run Summary Matrix—Short Form

Spray Time Run Date

LWC (g/m3)

MVD (μm)

OAT (°F) Deice min sec

16 11/15/06 1.04 22 12 34/34/68* 11 45 15 11/15/06 1.04 22 12 34/34/68* 14 10 21 11/15/06 0.36 96 12 34/34/68* 11 30 18 11/16/06 2.44 15 12 34/34/68* 0 39

15A 11/16/06 1.04 22 12 34/34/68* 14 0 15B 11/16/06 1.04 22 12 34/34/68* 15 32 15C 11/16/06 21A 11/16/06 0.36 96 12 34/34/68* 12 25 3A 11/17/06 0.33 16.5 04.6 34/34/68* 10 35 3B 11/17/06 0.33 16.5 04.6 34/34/68* 4 15 4A 11/17/06 0.57 16.5 15.2 34/34/68* 11 53 21B 11/18/06 0.36 96 24 34/34/68* 11 59 19A 11/18/06 0.10 40 12 34/34/68* 15 1 21C 11/18/06 0.36 96 12 34/34/68* 13 34 4B 11/18/06 0.45 20 12 34/34/68* 14 34 4C 11/18/06 0.52 20 22 34/34/68* 13 54 26 11/20/06 Wind off static thrust on MT Prop

26A 11/20/06 0.40 20 12 On Continuous

14 4

26B 11/20/06 0.52 20 22 On Continuous

14 35

26C 11/20/06 0.36 96 24 On Continuous

13 36

25 11/21/06 0.36 96 24 90/90** 13 37 25A 11/21/06 Off 24 11/21/06 0.66 15 19 90/90** 15 1

24A 11/21/06 0.40 20 12 90/90** 17 20 24B 11/21/06 0.36 96 12 90/90** 11 14

*34/34/68 = 34 sec. inboard boots on then 34 sec. outboard boots on, then 68 sec. both boots off. **90/90 = 90 sec on/off alternate boots.

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4.2 DRAG DUE TO ICE ACCRETION ON ENGINE STAND.

On two runs, attempts were made to quantify the drag due to ice accretion on the engine stand. After run 25, the ice on the stand and nacelle was cleared. No tracings were done. Ice was also swept from the open jet tube. (This was done after every run, but it was also important to do it before restarting the wind to prevent the ice crystals from eroding the propeller or thrust stand ice accretions.) The engine was restarted and the wind turned on. Thrust levels were obtained at 65% and 100% torque. After run 24B, the ice was cleaned off the blades and the spinner, the engine restarted, the wind turned back on, and the thrust levels were recorded at 65% and 100% torque. During the engine shutdown and restart, the video was monitored to determine if any propeller ice was shed. None was observed. For run 25, after engine shutdown, the nacelle remained warm, and the ice shed from the nacelle. In the warm temperature condition, it was preferable to leave the ice on the propeller rather than on the stand and nacelle. Data were also obtained with the wind on and the propeller feathered to determine drag due to the stand and nacelle. Prior to run 15B, with the wind on and the propeller feathered, a clean stand drag of 300 lb was recorded. After run 15B, the propeller was feathered, and the wind was turned back on. An iced stand drag of 330 lb was recorded. This resulted in a drag increase due to stand and nacelle ice of 30 lb at 100 KCAS with a feathered propeller. Hartzell used this data to calculate the effect of ice on the engine stand with the propeller generating thrust, i.e., the “scrub drag” due to the stand being subject to higher local velocity in the propeller wash. The Hartzell analysis is in appendix B. 4.3 SPINNER LIP EFFECT ON ICE ACCRETION.

One of the interesting things observed during the test was the effect of the rolled spinner lip on ice accretion. The rolled lip caused the droplets to swirl behind the lip and impinge slightly inside the lip, as shown in the top of figure 9. On the composite spinner, which did not have rolled lips, the droplets impinged directly in the blade shank and did not go inside the spinner dome, as shown in the bottom of figure 9.

17

Figure 9. The Effect of the Spinner Lip on Ice Formation 5. SUMMARY.

The full-scale propeller icing test provided a unique and extensive data set for the study of propeller icing. In addition to the numerical data, the exceptional imaging data obtained is a valuable contribution to the documentation of propeller icing. The following summarizes the test objectives and how they were met.

1. Document propeller ice accretions in 14 CFR Part 25 Appendix C icing conditions.

a. Leading edge accretions—Size, shape, span location, and shedding frequency as a function of icing conditions, blade material, blade condition, rpm

This was accomplished for several icing conditions and the data is presented in this report; however, the effect of blade material and rpm were not evaluated since the reciprocating engine was not tested. A negligible effect of blade condition was observed. Ice shedding frequency seemed to be independent of the propeller and deicers tested, and averaged a shed every 3-4 minutes. An end-of-run inspection revealed that the blades did not shed all at once, and that a blade shed event was not apparent on the real-time thrust data.

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b. Runback ice accretions—Size, shape, and location as a function of icing conditions, blade material, blade condition, rpm, deicing schedule, and power

Runback ice accretions and the resulting 20% thrust losses documented in reference 1 were not observed. The runback ice accretion phenomena described in reference 1 occurred at a small, critical temperature range. It was not known if the test articles would exhibit runback ice accretions, and therefore, a range of temperatures was tested. A possible explanation may be that the temperature increments tested were too large and a critical temperature, where runback may occur, was not tested. On the propeller in which two deicing boots were evaluated, the deicing time did make a difference on inboard radius ice accretion at an intermediate ambient temperature of 12ºF. This ice accretion existed on the entire boot, even at the stagnation. Not enough tests were conducted at warmer temperatures to evaluate runback ice. The effect of blade material and rpm was not evaluated since the reciprocating engine was not tested. However, the continuous heating scheme had the largest icing efficiency losses. It is not known if this was due to runback ice.

2. Document propeller ice accretions in SLD icing conditions and determine if a propeller

ice accretion from a Mitsubishi MU-2 flight test propeller in suspected SLD icing conditions can be approximately duplicated.

One SLD icing condition that approximated the LWC and MVD for the proposed Appendix X, which will be in 14 CFR Part 25, freezing drizzle with MVD>40, was evaluated. Thrust penalties in the SLD condition were higher than the 14 CFR Part 25 Appendix C icing conditions tested. The ice accretion and estimated propeller efficiency loss of the MU-2 flight test event were approximately duplicated.

3. Determine differences in thrust between non-iced and iced propellers.

This was accomplished and documented in the report. Drag due to ice accretion on the engine stand and nacelle was not negligible. The drag due to engine stand ice accretion was tested for the SLD icing conditions and analytically determined for the 14 CFR Part 25 Appendix C icing conditions.

4. Evaluate differences between reference and scaled icing conditions (LWC, MVD, OAT, and spray time).

This was not accomplished. The planned test matrix was shortened because of delays.

5. Determine if nominal propeller thrust reductions can be proposed for certification. Thrust values were recorded continuously. The propeller thrust was comparable to the drag of the test stand while at test condition; as a result, the measured thrust was low. On several early runs, this resulted in the measured thrust reducing to zero midway through the run. To overcome this limitation, the engine was run up to 100% torque to measure

19

20

thrust values just before each test began and again immediately after each test. The Hartzell corresponding engine stand drag (see table B-5 in appendix B) was then added to the 100% torque-measured thrust values to obtain the total thrust values. The total iced thrust value was subtracted from the total clean thrust to find the thrust difference. The delta thrust was divided by the total clean thrust to find the percentage loss. The thrust reductions shown in table 5 were determined using this method. The thrust reductions for the SLD runs averaged 13.4% and had a maximum of 21.2%. The thrust reductions for 14 CFR Part 25 Appendix C icing conditions averaged 5.9% and had a maximum of 13.4%. A nominal thrust penalty on the order of 10% for 14 CFR Part 25 Appendix C icing certifications was proposed, unless another value can be substantiated. Earlier independent flight test results [8 and 9] suggest a value of 10% would cover most Appendix C icing encounters.

Table 5. Summary of Thrust Reductions

Deicer SLD 12°F SLD 24°F 14 CFR Part 25

Appendix C 34/34/68 21.2% 11.3% 13.4% Continuous 10.8% 10.5% 90/90 16.6% 9.0% 5.6%

6. RECOMMENDATIONS.

1. Conduct additional propeller icing tests to complete the original test objectives.

2. Conduct additional tests to measure thrust stand drag due to ice in 14 CFR Part 25 Appendix C icing conditions.

3. Validate the spanwise accretion prediction of existing analytical methods.

4. Empirically measure lift and drag of propeller sections with ice shapes to simulate the accretions observed in this test and calculate propeller efficiency losses with the measured lift and drag.

7. REFERENCES.

1. Timmons, L., “Icing Investigations and Product Development on MU-2B Airplanes,” SAE 2003-01-2088, FAA In-Flight Icing/Ground Deicing International Conference, Chicago, Illinois, June 16-20, 2003.

2. Rodling, S., “Experience From a Propeller Icing Certification,” presentation at the SAE Aircraft Icing Committee meeting in Zurich, Switzerland, September 18-20, 1989.

3. National Transportation Safety Board, Office of Research and Engineering, Group Chairman’s Aircraft Performance Study, Comair Flight 5054, Embraer 120-RT, West Palm Beach, Florida, March 19, 2001, NTSB Accident Number DCA01MA031, Docket Item 13 (Washington, DC: NTSB, 2002). (Contact NTSB at pubinq@ntsb.gov).

4. Anderson, D., “Manual of Scaling Methods,” NASA/CR—2004-212875, March 2004.

5. Whalen, E., Broeren, A., and Bragg, M.,“Runback Ice Characteristics for a Bleed Air, Anti-Ice System,” presented at SAE AC-9C Icing Technology Subcommittee meeting in Hartford, CT, April 20, 2004.

6. Whalen, E.A., Broeren, A.P., and Bragg, M.B., “Characteristics of Runback Ice Accretions and Their Aerodynamic Effects,” FAA report DOT/FAA/AR-07/16, April 2007.

7. Ice Protection Harmonization Working Group, “Task 2 Working Group Report on Supercooled Large Droplet Rulemaking,” submitted to the Transport Airplane Engine Issues Group, September 2005.

8. Preston, G.M. and Blackman, C.C., “Effects of Ice Formations on Airplane Performance in Level Cruising Flight,” NACA TN 1598, May 1948.

9. Neel, C.B. and Bright, L.G., “The Effect of Ice Formation on Propeller Performance,” NACA TN 2212, October 1950.

21/22

APPENDIX A—MCKINLEY CLIMATIC LABORATORY SUPPORT FOR THE FEDERAL AVIATION ADMINISTRATION LARGE PROPELLER ICING TEST

In early 2005, the Federal Aviation Administration (FAA) contacted the McKinley Climatic Laboratory (MCL) to discuss the feasibility of conducting in-flight icing on large-diameter propeller simulations. This appendix provides an overall discussion of the facility and simulation, with specific attention to the calibration procedures used to evaluate the accuracy of the simulated environments. A.1 FACILITY EQUIPMENT DESCRIPTION. A.1.1 MAIN CHAMBER. The MCL facility, shown in figure A-1, is the world’s largest climatic test chamber, measuring 250 feet wide, 260 feet long, and approximately 70 feet high at the center of the chamber. It is essentially a large aircraft hanger that is heavily insulated under the concrete floor, behind all walls and doorways, and above the ceiling. The facility uses three R-22 refrigeration units, each rated at 330 tons at -93°F (a total of 990 tons). The facility also has three large natural gas-fired boilers to provide steam heat.

Figure A-1. McKinley Climatic Laboratory The chamber’s air-handling system is a closed-loop system using eight return plenums spaced evenly along the south wall; six cold return plenums (two for each refrigeration unit), and two hot return plenums. Each return plenum houses a large centrifugal fan that discharges through cooling or heating exchange coils respectively. The conditioned air is then directed into a common plenum and returned to the chamber through large ducts and 24 diffusers located in the ceiling of the chamber. A.1.2 AIR MAKE-UP. Separately from the air-handling system, the chamber also has two independent Air Make-Up (AMU) systems to provide conditioned air during periods when large volumes of air are being discharged from the chamber, which occurred during the propeller tests. The AMU systems use approximately three-quarters of a million gallons of calcium chloride for dehumidification and

A-1

approximately the same amount of methylene chloride (R-30) as a heat sink to reduce the temperature of the air replenishing the chamber. Prior to supporting AMU operations in the chamber, one of the refrigeration units is used to direct cold R-22 to the heat exchangers, through which these chemicals are passed to chill them. It takes approximately 24 to 36 hours (depending on outside ambient conditions and the initial temperature of the chemicals) to fully regenerate the fluids. To support AMU operations in the chamber, the chilled fluids are pumped to a separate set of 40-foot-high bank of heat exchangers through which outside air is channeled by large fans. The conditioned AMU air is then directed through 13-foot-square ducts to openings in the ceiling of the main chamber. Each AMU system can deliver a maximum of 500 lbm/sec of air at -65°F for approximately 1 hour (a total of 1000 lbm/sec). When either the volume flow demands are less than the maximum or the required air temperature is warmer than -65°F, the duration of AMU operations is increased. The duration of the two AMU systems was sufficient to support the test schedule, about 3.8 runs per test day, with an average engine running time of 23.3 minutes. A.1.3 THRUST TABLE. For this test, one of the primary objectives was to obtain direct measurements of thrust degradation with the propellers iced. A thrust table suitable for this purpose (figures A-2 and A-3) was designed and built at the MCL. The table was designed as a rigid frame onto which a floating platform was suspended by sheet metal straps. Four vertical restraint straps (one at each corner) supported the weight of the platform, engine support stand, and engine. Two additional straps were used as horizontal restraint along the leading and trailing edges of the platform. The six restraint straps restricted the linear motion of the platform in all axes, except fore and aft. The platform was linked to the rigid frame with two electronic load cells configured for measurements in compression and tension. Ball joint rod ends and miniature turnbuckle devices were used to mount the load cells so that the loads could be equalized between the two load cells and any bending moments could be eliminated.

A-2

Figure A-2. Calibration of Load Cell Mounted on Thrust Table

Figure A-3. Engine, Stand, and Thrust Table

A-3

A.1.4 OPEN-CIRCUIT WIND TUNNEL. A large open-circuit wind tunnel was used to create the high-speed wind to present the icing cloud to the propeller (figure A-4). The wind tunnel has seven fans, which are 7 feet in diameter, configured in parallel at the tunnel inlet (figure A-5). Each fan is driven by a 480-volt, 300-hp electric motor, which in turn, is driven by a variable-frequency drive, allowing for precise speed control. This stack of fans is approximately 25 feet high and wide. A forward-curving edge is used around the perimeter of the fan stack. Additionally, each fan has its own inlet bell. The gaps between fan inlet bells are spanned with a sheet metal filler. In this manner, the seven fan inlet bells smoothly merge with each other and the edging to act as one large inlet bell to promote smooth airflow into the fan inlets. Each fan discharges into a diffuser, and the seven diffusers smoothly merge into a common plenum with fiberglass transitions between, again to promote smooth airflow. The common plenum has a hex-shaped cross-section, approximately 25 feet in diameter. The common plenum then discharges into a large converging nozzle that gradually transitions from the 25-foot hex to a 11.9-foot, eight-sided cross-section. The discharge opening is not a regular octagon, but instead, has slightly different length sides opposing each other to accommodate uniform fitment to the icing spray bar nozzle pattern. The centerline of the tunnel discharges at 14 feet above the floor.

Figure A-4. Open Jet Setup

A-4

Figure A-5. Stack of Seven Fans at Tunnel Inlet A.1.5 AIRFLOW SETUP. An icing frame composed of 12 evenly spaced, airfoil-shaped spray bars (approximately 2.25 inches thick by 14 inches in chord length) was mounted to the tunnel discharge nozzle (figure A-6). Immediately downstream of the icing nozzles, a 24-foot by 105-inch-diameter containment duct was used to prevent the cloud from expanding in physical dimensions until the last 8 feet of travel. The thrust table was positioned so the propeller was located approximately 32 feet downstream of the nozzles and approximately 8 feet from the discharge end of the containment duct. With the icing spray bars and containment duct installed downstream of the tunnel nozzle, the tunnel created wind speeds of approximately 100 knots at the propeller. A duct located aft of the engine/propeller assembly carried engine exhaust and most of the icing plume outside the chamber.

A-5

Figure A-6. Spray Bar Frame Mounted Upstream of Containment Duct A.1.6 SPRAY BARS. The spray bars have an airfoil-shaped cross-section, which has a maximum thickness of 2 1/4 inches and a chord length of 14 inches. The spray bar has four channels running spanwise, separated by internal bulkheads. Inside the nose, there are two separate channels that carry heated, deionized water. The foremost channel at the leading edge carries the heated water at a high flow rate specifically to prevent the water from freezing in the channel behind it. The water in the heating circuit channel does not mix with the water that will become the icing cloud, except in the storage vessel before being pumped to the chamber. The second water channel carries the pressurized, heated, deionized water that will be directed through the spray nozzles and ejected as an icing cloud. The third channel is an atmospheric pressure channel with a removable upper plate. Small solenoid valves that isolate individual nozzles (there is a solenoid for each nozzle) are inside the atmospheric pressure channel. The solenoid is plumbed with small tubing into the spray water channel immediately forward on the upstream side and is plumbed to a water tube assembly on the downstream side. The aft-most channel carries the atomizing air through the spray bar and provides pressurized air to the nozzles. The water tube assembly passes through the air channel to allow the water from the solenoid to be supplied to the back of the spray nozzles.

A-6

A.1.7 NOZZLES. Spraying Systems Company air-atomizing nozzles were used, which are composed of a three-part design incorporating a fluid cap, an air cap, and a retaining ring to secure the air cap onto the fluid cap (figure A-7). Water is supplied to a central fluid port and is directed through a single central orifice that projects the water flow into an internal mixing chamber. The atomizing air is supplied to a number of holes located in an annular ring. The number and size of individual air holes depends on the specific nozzle design. Air is channeled through these holes into the mixing chamber. The atomized mixture is then ejected through one or more holes into the atmosphere. The nozzle selection is determined during calibration to provide the necessary spray conditions. The nozzles most often used are custom designs that have been developed over several years.

Figure A-7. Spray Nozzle A.1.8 SPRAY WATER SYSTEM. The spray water system has the following subsystems: water deionization, water heating, pumping, flow measurement, pressure control, and spray bars. Each subsystem is briefly discussed in the following sections. A.1.8.1 Water Deionization. The water for the icing system is supplied from the facility’s water deluge main supply feed line at a pressure of approximately 50 psig. The water is passed through deionization canisters filled with resin beads, which have been chemically treated to absorb the minerals from the water (figure A-8). The canisters are arranged in parallel rows of two canisters. For this test, six canisters were used. When the canisters are replaced, the resistivity of the water increases to approximately 14 million ohms per centimeter (pure water molecules have a theoretical

A-7

resistivity of just over 18 million ohm/cm). As the water is passed through the canisters and the minerals are extracted, the resin beads’ ability to continue extracting additional minerals declines. The water deionization system has a meter that monitors the capacitance (inverse of resistivity) of the water. Typical icing tests will diminish the resistivity to levels far below the initial state, but still high enough that evidence of mineral deposits inside operating jet engines (where the water quickly evaporates) is minimal. Deionization also prevents the spray nozzle orifices from clogging over time.

Figure A-8. Water Deionization Canisters A.1.8.2 Water Heating. After deionization, the water goes to a direct-contact, 8-million Btu/hr hot water heater. The heater receives the water through a solenoid valve that is controlled by a float switch in a storage vessel at the base of the system. The solenoid valve opens and closes to maintain a constant water level in the vessel. When the solenoid opens, water is sprayed from a nozzle and cascades through a vertical stack of stainless steel packing rings. Simultaneously, a natural gas-fired boiler projects a flame through a horizontal leg that tees into the vertical stack. As the water cascades through the packing rings and the hot gas rises through the packing rings, heat is transferred to the water. The gas escapes through an exhaust stack at the top while the heated water is retained in the vessel at the bottom. If the level is sufficiently high, the solenoid valve will close to prevent additional water from entering the system. If the temperature is below a specified value, a small pump circulates the water from the vessel through a jacket around the horizontal burner section, and the burner is fired to continue heating the water. In this manner, the heater maintains a specified water level in the receiver at a specified temperature. The heater

A-8

maintains both the water level and temperature, with flow rates through the system as high as 120 gallons per minute (gpm). For the vast majority of icing tests, the water demand on the system is typically no more than 15 gpm. An additional steam-to-air heat exchanger is used inside the chamber to compensate for heat lost while the water travels through the insulated supply hoses in the chamber (as much as 150 feet at a very slow flow rate). The secondary water heater is located close to the spray bar control system, and the connection between this water heater and the control system is reduced in size to minimize heat loss between this station and the spray bars. A.1.8.3 Water Pumps. The icing spray system uses two pumps to deliver heated, deionized water from the heating system’s storage vessel to the spray bars in the chamber. One pump delivers the water that becomes the icing cloud (at a volume flow determined by the required liquid water content (LWC)), while the second pump delivers water (at a much higher volume flow) through a separate circuit, which is solely used to heat the spray bars. The spray pump was specifically selected with an extremely low net-positive suction head required value to prevent cavitation of the water at the pump inlet. The pump discharge piping has a recirculation leg that allows the majority of the water to be pumped immediately back into the storage vessel (usually around 300 gpm). The recirculation leg has a control valve that can adjust the water delivery pressure to the chamber. The delivery pressure is usually set to about 105 psig. The heat pump was selected to move higher water volume at moderate pressure (approximately 50 gpm at 50 psig). A.1.8.4 Water Flow Measurement. The water is directed to a flow measurement section that has three parallel flow paths. The first flow path bypasses the flow meters and is used only when high-volume flow rates are required (e.g., initial flooding of the spray system). The second flow path is fitted with a high-range flow meter (capable of measuring up to 60 gpm), and the third flow path is fitted with a low-range flow meter (capable of measuring up to 10 gpm). For this test, most conditions required the low-range flow meter. However, there were several conditions that required the high-range flow meter. Electrically activated isolation valves were used to select a specific flow path. A.1.8.5 Water Pressure Control. The water is directed to the pressure control system in the chamber (figure A-9). This system is a plumbing network housed in an insulated aluminum box, approximately 3 by 4 feet wide by 12 feet tall and is located at the upstream end of the spray bars. Inside this box the water is supplied to a header, which is common for all spray bars. Each spray bar has its own electronic pressure controller that receives a set point signal (desired pressure) from a computer running LabVIEW® software and compares that set point signal to a feedback signal from a pressure transducer mounted in the upstream end of the spray bar. The spray bar water supply circuit (between the header and the spray bar) uses two air diaphragm-actuated control valves in parallel (one with flow coefficient (Cv) = 0.05; the second with Cv = 2.0). The diaphragms on the control valves are set so that the low-range valve will open for a 3- to 9-psig air signal. A 9 to 15-psig air signal will open the high-range control valve as well. The pressure controller

A-9

determines the relative pressure needed for the air diaphragms to open or close to maintain the desired pressure in the spray bar. Before spray activation, the water travels through the spray bar and, on the downstream end, through a large solenoid valve and a control valve before discharging into an atmospheric pressure dump header that is fitted with a hose leading to a floor drain. The downstream control valve is positioned so that it provides a similar resistance to that which is experienced when water flow is directed through the nozzles upon spray activation. This configuration (prior to spray activation) allows for the presetting of proper water pressure and flow rate simultaneously inside the spray bar. Upon spray activation, the large downstream solenoid valve is closed and, simultaneously, a small solenoid behind each individual nozzle is opened, altering the water flow path from passing through the spray bar (but not the nozzles) to passing through the nozzles (but not the downstream end of the spray bar). Since the pressure and flow rate were preset to the appropriate values, the instantaneous transient spray bar pressure and flow rate are minimized. Stabilization typically is achieved in approximately 1 to 5 seconds, depending on the absolute values of pressure and flow and also on the accuracy of the preset values.

Figure A-9. Water Pressure Control System

A.1.9 ATOMIZING AIR SYSTEM. The atomizing air system has the following subsystems: air compression, pressure control, and air heating. Each of these subsystems is briefly discussed in the following sections. A.1.9.1 Air Compression.

A-10

The atomizing air must be pressurized to specific values and accurately controlled, since the atomization can affect both the droplet size (median volume diameter (MVD)) and the LWC in the cloud (this occurs because the nozzles typically used are internal mix nozzles and increases in atomizing air pressure induce a higher resistance for the water flow path). Rented portable air compressors are used to compress outside ambient air and to supply it to the atomizing air plumbing circuit (figure A-10). Outside ambient air typically has a much higher moisture content than the air inside the chamber, but the amount of moisture is typically not significant compared to the volume of water intentionally sprayed, and is therefore ignored. Atlas Copco PTS916 compressor units (by Prime Energy) are used. They are rated for 1600 cubic feet per minute (cfm) and 125 psig (but not both values at the same time—high-volume flows are delivered at lower pressures or the highest pressure may be delivered at a lower-volume flow). The units are diesel-powered and have a liquid crystal display touch screen for operator control. The unit has an internal pressure sensor that displays the measured output pressure. The operator selects the desired output pressure, but may or may not be able to achieve it, depending on the resistance of the output connection. The unit has an onboard air pressure-regulating valve, which is controlled by a process controller taking input from the operator control panel.

Figure A-10. Air Compressors A.1.9.2 Air Pressure Control. For most icing tests, multiple PTS916 units are required to achieve the necessary volume flow rates at the desired pressures. The units are plumbed in parallel into a common header. Downstream of the header, the airflow is directed through a series of parallel tee’d legs with manual blow-off control valves. These manual valves are required when the desired airflow and pressure combination is in a range beyond adequate control of the PTS916 onboard air pressure-regulating valve (usually low volume flow). When the PTS916 regulator attempts to operate at a condition that positions its valve near its seat (almost closed), the regulator tends to oscillate, repeatedly opening and closing as it attempts to control pressure. The manual blow-off valves

A-11

allow increased volume flow at the regulator, which stabilizes the oscillations and simultaneously dumps (i.e., blowing off) the excess volume flow away from the spray bar system. The spray bar air supply header has a pressure transducer, whose output is displayed to the compressor operator via a separate computer terminal. The operator positions both the PTS916 pressure set point, and, if necessary, the position of the manual blow-off valves to achieve the desired air pressure in the spray bar air header. A.1.9.3 Air Heating. After passing through the manual blow-off section, the air is directed through flexible, insulated hoses to a set of steam-to-air heat exchangers. The air header at the spray bars has a temperature probe, whose output is displayed on the compressor operator’s terminal. The steam provided to the heat exchangers is manually adjusted by the compressor operator to control the output air temperature. The output air temperature is a function of the heat input from the steam circuit and a function of the volume of air through the heat exchanger. After leaving the heat exchangers, the flow again is passed through flexible, insulated hoses, and bifurcates so that it can be supplied simultaneously to the headers at both ends of the spray bars. A.2 CALIBRATION. Since the MCL icing capability consists of temporary hardware (i.e., not permanently installed), the cloud conditions must be calibrated for each setup. The calibration effort includes verifying the instruments used to measure the clouds as well as correlating the pressures and flows needed to create the desired clouds. Additionally, a thrust table was used to obtain direct measurements of thrust changes with the propellers iced. Therefore, the calibration effort also included verifying the performance of the thrust table. A.2.1 THRUST TABLE CALIBRATION. The thrust table was calibrated prior to the test by attaching it to anchors in the chamber floor and using another fixture to load the table in the same way it would be during the test. A temporary frame was mounted to the top of the table so that loads could be applied in the forward thrust direction at an elevation of 6 feet above the platform (duplicating the thrust centerline with engine and support stand mounted). The temporary frame also used a crossbar so that a dead load could be applied to induce a known torque at the thrust point (torque also applied at 6 feet above the platform). Then, a rigid vertical structure was mounted approximately 10 feet in front of the table and attached to the temporary frame with a horizontal length of chain in conjunction with a turnbuckle and digital crane scale, which was previously calibrated by the Eglin Air Force Base (AFB) Precision Measurement Equipment Laboratory (PMEL). The turnbuckle was used to apply horizontal loads, measured with the crane scale and compared to the sum of the outputs from the two thrust table load cells. Comparison measurements exceeding 2000 pounds were taken, demonstrating a linear thrust response with approximately a -4% difference (figure A-11). Compensations were made to eliminate this difference before load values were displayed. A 1500-pound torque application indicated no significant change to the thrust measurement. Although the static thrust measurement accuracy for this thrust table and load cell configuration was estimated to be less than 1% of full load (in

A-12

practice), with a reciprocating engine fitted with a large propeller operating on the table, thrust measurement values appeared to rapidly and continually fluctuate by as much as 5% or more due to engine vibrations, wind speed fluctuations, electronic noise, and other factors. Consequently, all thrust measurements that were displayed and recorded for the test were running time averages of approximately 10 to 15 seconds.

2000

Thru

st T

able

Loa

d C

ell L

oad,

lb

1500

TorquePerfectLoad Cells

1000

500

0 0 500 1000 1500 2000

Applied Known Load (Crane Scale), lb

Figure A-11. Thrust Table Calibration A.2.2 VELOCITY PROFILE MEASUREMENT. A pitot-static probe was used to traverse the test area at the elevation of the tunnel centerline and was used for velocity measurements along the cloud centerline for all subsequent icing cloud calibrations. The velocity must be measured in real time to calculate LWC. The pitot-static probe itself is not a calibrated device, but rather, the pressure transducer, which measures the total and static pressures produced in the pitot-static probe, is calibrated by the Eglin AFB PMEL. Typically, for open jets, only the total pressure is measured from the pitot probe. The static pressure is provided by a separate barometric pressure transducer that senses chamber pressure. The velocity profile was not measured in the vertical dimension, but was assumed to be similar to the horizontal measurement. The velocity profile demonstrated that the velocity was uniform over approximately a 6-foot-diameter region, centrally located, and decreased by approximately 15% at the propeller tips (figure A-12). Removing the containment duct was proposed to achieve a uniform velocity profile over the full propeller span, but this potentially could have limited the maximum achievable LWC (i.e., the cloud would have grown much larger than the extraction duct and contamination of the chamber air from ice crystals becomes a concern). Since the primary region of interest on the propellers was at radial locations up to 75% span, and not at the tips (where the significantly higher g forces and lower freezing fractions typically prevent ice from forming in all but the heaviest of icing conditions), and the test local velocity at the tip is dominated by propeller rpm and not plume velocity, the FAA decided to accept the velocity profile as demonstrated and proceed with the test.

A-13

Figure A-12. Velocity Profile, 17 Degree Blade Pitch, 66°F MC Temp (View Looking Down)

Velo

city

, mph

Position, feet8 10 12 6420-2-4-6-8-10 -12

0102030405060708090

100110120130

A.2.3 FLOW METER VERIFICATION. The water spray system is controlled by pressure, not flow. However, for a given nozzle configuration and wind speed, the water flow rate may be used directly to determine the LWC in the cloud. Therefore, it is very important to have accurate water flow rate measurements, despite the fact that it is not directly controlled. Flow Technologies’ turbine wheel flow meters with electronic output are used to measure this. To verify the flow output, the spray system is configured to direct water flow through the spray bar flow path (actually flowing water through the spray bars). A water pressure and downstream control valve position are selected to achieve a desired flow rate and the discharge from the end of the flow path (immediately before it would typically dump into the floor drain) is timed as it goes into a bucket and then weighed on a large balance scale, which is calibrated by Eglin AFB PMEL. The temperature of the water is measured to account for density variations. During the time the water is flowing into the bucket, the LabVIEW system writes the flow meter data to an output file. Displayed and recorded values are time-running averages of approximately 30 seconds. At the end of the timed water flow, flow rate values in the output file are averaged over the entire period and compared to the flow rate calculated from the buckets water volume per unit time. For this test, flow rates were measured up to over 30 gpm, easily covering the full range of expected flow conditions for the test. All flow measurements below 10 gpm were taken using the low-range flow meter, and flow measurements above 10 gpm were taken using the high-range flow meter. As expected, the flow meter output compared well with the time bucket measurements (figure A-13). This process is used not only as verification of the flow meter, but also to ensure that there are no leaks in the system anywhere downstream of the flow meter.

A-14

High Flow MeterPerfectLow Flow Meter

Tim

ed B

ucke

t gpm

LabVIEW gpm30252010 1550

30

25

20

15

10

5

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Figure A-13. Flow Meter Check

A.2.4 JOHNSON-WILLIAMS LWC PROBE CALIBRATION. A Science Engineering Associates Johnson-Williams (JW) LWC probe was used for all LWC measurements. This instrument has no field calibration technique whereby a spray with known LWC is presented. Instead, the sense and compensating wires are checked to determine their relative difference over the range of wind speed (which determines the convective cooling rates for the instrument) that is expected for the test. Before taking LWC measurements, this instrument is presented with a discrete wind speed (no spray), the anti-icing elements activated, the instrument allowed to thermally stabilize, and the power consumption of the sense and compensating wires are noted. After repeating this procedure for several wind speeds, the relationship between the wires’ power consumption was determined, and the correlation used in the LWC calculations shown in figure A-14.

Linear (Series 1)Series 1

Sens

e W

ire P

ower

, wat

ts

20 0.5 1 1.5Compensating Wire Power, watts

7 6 5 4 3 2 1 0

y = 3.527x + 0.491

Figure A-14. The JW Calibration

A-15

A.2.5 MALVERN CALIBRATION. A Malvern Spraytec 5000 ensemble particle characterization laser interferometer is used for measuring particle size distribution from which the MVD is determined. For calibration, the instrument is fitted with an air-atomizing nozzle that was modified in the following manner: • A hole was drilled through the top of the nozzle body, which is normally used as the

atomizing air cavity.

• A small funnel was brazed onto the top of the body so glass microspheres can be poured into the hole.

• A compressed air line with a manually operated, intermittent hand valve was connected to the nozzle body port.

Upon activation of the hand valve, compressed air is blown through the nozzle, and the glass microspheres are sucked in and then ejected out the nozzle tip. With this apparatus rigidly mounted to the front of the Malvern instrument and aimed at the laser beam, various-sized glass microspheres may be presented to the instrument and compared to the instruments output. Glass microspheres ranging from 15 to 79 microns were presented to the instrument and the output compared to the actual sizes to demonstrate the instruments ability to accurately measure particle size. As expected, over this range of sizes, the instrument agreement with the known particle sizes presented to it was quite good (figure A-15).

(-10%)(+10%)PerfectSeries 1

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vern

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Duke Scientific Beads, μm806040200

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Figure A-15. Malvern Verification

A-16

The Malvern has a known tendency to estimate MVD values too small at MVD values larger than approximately 28 microns. This occurs because the instrument optics can only measure particles smaller than approximately 200 microns. Typical water sprays are composed of particle size distributions that include particles at least this large. For larger MVDs, the particle size distribution is typically shifted to the right and the right-hand tail of the size distribution gets truncated. Since the volume of the largest drops is proportional to the cube of their diameter, the largest particles may contain a significant percentage of the total volume of water; therefore, the calculated MVD is too low. The MCL Malvern instrument was tested in the NASA Icing Research Tunnel (IRT) to compare Malvern indicated MVD to IRT calibration for MVD over a wide range of MVD values between 10 and 190 microns. The data clearly show that, for MVD below 28 microns, there is no appreciable difference between the Malvern and IRT measurements, and for MVDs larger than 28 microns, the Malvern MVD underreads. For the requested conditions of this test, which were within Title 14 Code of Federal Regulations (CFR) Part 25 Appendix C icing conditions, the Malvern indications were used unmodified. Several conditions of large MVDs were requested as well as the Malvern-IRT correlation data used to correct the measurement (figure A-16).

Malvern - 23.313IRT = 1.9349*MVDMalvern > 25um, MVD

MalvernIRT = MVDMalvern < 25um, MVDFor MVD

For MVD

45 50 55 60 65 70 75 80 85 90

Note of Caution: This relationship MAY be a function of the particle size distribution from which the MVD is calculated and particle size distributions may vary for different spray nozzles. This effect has not been evaluated.

4035 30 25 20 15 10 5 0 5

0

10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95

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IRT

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, μm

(Tru

th)

To Be Corre edctPerfect Line Linear (To Be Corrected)

Use Uncorrected

Malvern MVD, μm

Figure A-16. Malvern—IRT Correction Chart (Ignoring first eight rings)

A.2.6 TOTAL TEMPERATURE CALIBRATION. MCL uses a Harco Laboratories total temperature sensor that uses an internal resistance temperature detector and by design, allows a small fraction of the stagnated air inside the measurement section to slowly bleed away to minimize thermal effects from bringing the flow to

A-17

rest. The sensor has a manufacturer’s suggested recovery factor of 0.97. The factory-calibrated total temperature sensor is part of a system designed and built by Particle Measurement Systems. The total temperature sensor output is observed during icing cloud sprays to obtain an approximation of the degree to which the cloud droplets are supercooled (i.e., on initial spray activation, provided the temperature output either remains constant or decreases, but does not increase, the droplets are assumed to be supercooled). A.2.7 CALIBRATION MAPS VS CALIBRATING FOR DISCRETE POINTS. Once the instruments are calibrated, the correlation between spray conditions and measured cloud conditions may be made. If the test requires a wide range of icing conditions (and especially if there is a possibility that the requested spray conditions may change in the middle of the test), sufficient sprays must be made (usually as many as 50 sprays per wind speed) to gather enough data to generate calibration maps, which are then used to select the appropriate spray conditions needed to generate the requested icing condition. If the test requires only a few icing conditions and there is little likelihood that the requested icing conditions will change, then only the specific spray conditions necessary to create the requested icing conditions are required (i.e., the maps are not created). This is known as calibrating for discrete points. In that case, an arbitrary spray is generated and the icing cloud measurement instruments are observed while the spray conditions are modified in real time until the requested icing condition is observed. At that point, the necessary spray conditions are noted and repeated for the test. The advantage of calibrating for discrete points is that the calibration can be accomplished more quickly. The disadvantage is that once the calibration work is completed and testing begins, the only icing conditions that can be generated are those that were specifically calibrated. If an additional test point needs to be accomplished or one slightly modified, the test has to stop and the calibration performed again. On the other hand, taking the additional time to create the calibration maps allows any icing condition (which is within the range of conditions calibrated) to be created with high confidence by interpolating from the maps. A.2.7.1 Chamber Temperature vs Cloud Temperature. Initially, the chamber temperature must be determined for a given spray condition. Since the open circuit tunnel is inside the chamber, the air that is drawn into the fan inlet is at static air temperature, which is measured, displayed, and recorded as chamber temperature. However, no fan is perfectly efficient, and there is some heat rejection from the fan motor. Additionally, the flow passes over the spray bars, which are intentionally heated to prevent freezing. The hot spray bar surfaces convect heat into the air stream. The nozzle ejects an atomized spray of heated water and air, which experiences a pressure and temperature drop when ejected from the nozzle. This causes a temperature differential between the spray and the air stream, with further heat exchange. Additionally, in an open jet, there is some degree of mixing between the jet plume and the surrounding air, especially at the edges of the plume. Consequently, the static temperature measured in the cloud is measurably different than the original static air temperature of the chamber. Several sprays are made at arbitrarily low and high flow rates with the chamber temperature near the warmest icing point. The chamber temperature is then lowered by about

A-18

15° or 20°F (or to the lowest icing test temperature, whichever is nearer to the original temperature), and the same low and high flow sprays are repeated. During each spray, the recovery temperature from the total air temperature probe is noted, and the corresponding static air temperature is calculated. The chamber temperature is plotted against either the static or the recovery temperature. This relationship is subsequently used for all icing to determine the appropriate chamber temperature (usually colder) necessary to achieve the desired static temperature in the cloud. For most low-speed icing tests, this adjustment may only be 1° to 2°F, but for many high-speed tests, the adjustment may be as much as 6° to 8°F. Note that the relationship depends on wind speed, so this procedure must be done for each wind speed.

A.2.7.2 The LWC Map.

When generating calibration maps, arbitrary sprays are made with different water pressure and atomizing air pressure combinations (usually water pressure is accomplished at 10-psig increments up to approximately 60 psig, and at least four air pressures for each water pressure are created). For each spray condition, both the measured LWC from the JW probe and the total water flow to the spray rig (gpm) is noted. The LWC created depends not only on the water and air pressure, but also on wind speed and the number of nozzles operating. In cases (usually for low wind speed) when it is difficult to achieve a low LWC value even at the lowest controllable water and air pressure, the number of nozzles that are spraying may be reduced. In rare cases, the number of nozzles may be reduced to so few that several iterations of nozzle selection (position within the array) are necessary to ensure adequate uniformity. Once proper nozzle selection has been achieved, the gpm and the measured LWC parameters are plotted against each other and the best curve fit drawn through all the points. The JW probe is generally accepted to be accurate to approximately 10%. The wind speed typically fluctuates as much as 3%, and the LWC calculation is directly and inversely proportional to the wind speed. For some spray conditions, (especially when some nozzles are intentionally turned off to reduce the LWC), the spray plumes from individual nozzles may not mix as well as expected, and the relative position of the JW probe within the cloud may become important. Additionally, the spray system is controlled by pressure, not flow. The pressure controllers allow the water flow rate to fluctuate up or down to maintain the requested pressure in the spray bars. All these factors tend to cause some amount of scatter on the gpm versus LWC chart. Typically, the vast majority of measurements will fall within ±10% of the curve fit line. Occasionally, sprays may be created that fall outside those bounds. Once calibration is completed, the best fit curve line is used as the basis for selecting the appropriate gpm to produce the requested LWC (figure A-17). Note, this LWC map is only applicable for a single wind speed. If additional wind speeds are required, additional LWC maps must be generated.

A-19

Figure A-17. An LWC Map, 55 GE-1 Nozzles, 100 Knots, 9′8″ Diameter

A.2.7.3 The MVD Map.

During the calibration sprays, the measured MVD is also recorded. For each water pressure, a line is created on a plot of gpm versus atomizing air pressure. At each measured point, the drop size is noted. Once these constant water pressure lines are placed, additional lines of constant MVD are superimposed over this plot (figure A-18). Note that the MVD map is only weakly dependent on wind speed and is generally applicable for a wide range of wind speeds.

Figure A-18. The MVD Rig Map, 55 GE-1 Nozzles, 100 Knots, 8′9″ Diameter

y = -6.302785931702E-04x + 1.549270656060E-023 x + 8.684716944712E-02x2

00.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 11.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 22.1 2.2 2.3 2.4 2.5 2.6 2.7 2.8 2.9 3

JW L

WC

, g/m

3

Data-10%10%Poly. (Data)

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 5.5 6 6.5 7 7.5 8 8.59 9.5 10 10.511 11.512 12.513 13.514 14.515 15.516

Total gpm to Rig

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14um 15um

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Pw=1Pw=2Pw=3Pw=5Pw=10Pw=15Pw=18Pw=20Pw=22Pw=25Pw=30Pw=40Pw=5014um15um16um17um18um19um20um25um30um40um50um60um

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A-20

A.2.8 PLUME UNIFORMITY.

A grid was placed at the propeller plane location to check uniformity of the plume (figure A-19).

Prior to testing, there was a question as to whether the propeller capture area at high power would be larger than the plume diameter, thereby reducing the density of water drops near the propeller tips. Figure A-20 shows the entire propeller immersed in the cloud plume.

Figure A-19. Grid to Check Uniformity

Figure A-20. Propeller Immersed in Cloud Plume

A-21

A-22

A.2.8.1 Water Temperature and Atomizing Air Temperature. Several test setups were analyzed using the AEDC1DMP code (Arnold Engineering Development Center) to consider the effects of initial droplet temperature on the final droplet temperature at the propeller test location. For nearly every case, and especially for 14 CFR Part 25 Appendix C size droplets, the code indicates that droplet temperatures rapidly approach the static air temperature before reaching the propeller test location. However, based on the results of previous tests, the following approach was adopted. For typical 14 CFR Part 25 Appendix C icing conditions, both water and atomizing air temperatures are maintained at approximately 130°F. However, in general, for high LWC conditions (>2 g/m3) or for large MVD sprays (>40 microns), both the water and atomizing air temperatures can be lowered by approximately 20°F to ensure supercooling. Also, if the requested static air temperature is below 0°F, the water and air temperatures can be increased by approximately 20°F to prevent the spray equipment from freezing. Another code based on NASA/CR-2004-212875, Manual of Scaling Methods, can be run that estimates the freezing fraction and ice thickness for the given spray conditions to determine whether the ice accretions observed appear to be correct (this estimate is based on a stationary item, like a wing, within the icing cloud and is not necessarily relevant to jet engine or propeller icing). A.2.8.2 Spray Procedure. To achieve a specific spray condition, the wind speed, static temperature, LWC, MVD, and duration must be specified. Initially, the requested static air temperature and wind speed are used to determine the appropriate chamber temperature. The chamber static air temperature is set to this value, which in most cases, will be several degrees colder than the requested in-cloud static air temperature. If the calibrating for discrete points process was adopted, the water and atomizing air temperatures and pressures will be already known. If the calibration maps process was adopted, the LWC map is used to determine the necessary total water flow (gpm) to the spray rig at the requested LWC. This gpm value is used with the MVD map to determine the necessary water and atomizing air pressures to produce the requested MVD. The water and atomizing air temperatures are adjusted, and then the water and atomizing air pressures are applied to the spray bars. The downstream control valves are positioned to adjust the total water flow rate to the spray rig so that it agrees with the gpm value determined from the LWC map. At this point, the spray bars are at the correct spray pressures and flows and only need the nozzle solenoids to be activated. The fan variable frequency drives (VFDs) are preset to the correct value to produce the requested wind speed. Once the customer is ready for wind, the fan VFDs are started and the fans brought to the correct speed. When the customer is ready for cloud initiation, spray activation is commanded, the spray bar downstream solenoids close and simultaneously the nozzle solenoids open, initiating the cloud. The spray control system is observed. The water and air pressures will generally stabilize within 5 seconds. However, it will generally appear to take approximately 30 seconds (due to time averaging) for the water flow rate to stabilize. This stabilized flow rate value is compared to the LWC map flow determined earlier. If the values disagree by more than approximately 10%, the water pressure is adjusted as necessary to achieve the required flow rate. In most cases, the adjustment will only be a few tenths of 1 psig. This process generally ensures that the LWC and MVD produced in the cloud will be within the accepted accuracies of the instruments that are used to calibrate those values.

APPENDIX B—COMPUTATION OF SLIPSTREAM DRAG CORRECTIONS FOR USE WITH FEDERAL AVIATION ADMINISTRATION ICING TEST DATA TAKEN ON THE

HARTZELL 4-BLADE 10282-5.3R PROPELLER B.1 INTRODUCTION. During a Federal Aviation Administration (FAA) icing test at Eglin Air Force Base in November 2006, thrust measurements were recorded for the Hartzell 4-blade 10282-5.3R propeller. These measurements included drag on the test stand due to the propeller slipstream during propeller operation and, therefore, are not directly indicative of propeller performance without a correction for test stand drag. To understand the performance of the isolated propeller as it collects ice, corrections for the drag of the test stand were computed. The isolated propeller thrust values will also be more useful for future analytical correlations. First, the effective flat plate area, f, of the test stand was computed using a method described in this report. Using the value of f, corrections could be estimated for each test data point to more closely determine the actual performance of the propeller. The analysis does not account for partial immersion of the stand in the propeller slipstream and assumes that it is 100% immersed. The correction, although somewhat simple, can serve as a reasonable, consistent, and documented method to aid in the determination of the isolated propeller thrust. Note that the correction does not include the influence of the stand blockage on the inflow velocity distribution for the propeller. B.2 COMPUTING TEST STAND EFFECTIVE FLAT PLATE AREA. Drag data were recorded with the propeller in the feathered configuration with the wind tunnel at full operating speed, 100 knot true airspeed (KTAS). The measured drag is viewed as the sum of two components:

(B-1) total prop standD D D= + Equation B-1 describes the total recorded drag as the sum of the drag due to the feathered propeller assembly (including the spinner area), and the remainder of the test stand. Since the stand drag will be the parameter of interest (to be corrected for) in a run with the propeller in operation, the feathered propeller drag was estimated to eliminate this unknown from equation B-1. To compute the drag of the feathered propeller, ten blade cross-sections were analyzed in XFOIL, given the known conditions of the drag tare run (airspeed, blade angle, air density, etc.). XFOIL is a panel method viscous flow solver that can provide reasonably accurate lift and drag coefficients of a given airfoil.

B-1

The results of the XFOIL runs are shown in table B-1.

Table B-1. Feathered Blade Drag Computation

r/R Radius

(in) Station

(in)

Twist Angle (deg)

AOA (deg)

Chord (in)

Chord (m)

Reynolds No.

Cd1 (XFOIL)

Cd (normalized)

Drag (lb)

0.081 7.900 7.150 59.19 30.63 3.602 0.091 389572 0.900 0.2499 0.200 9.770 9.020 56.54 27.98 4.442 0.113 480422 0.800 0.1801 7.50 0.300 14.655 13.905 49.45 20.89 5.860 0.149 633785 0.610 0.1041 12.95 0.400 19.540 18.790 42.47 13.91 6.830 0.173 738695 0.170 0.0249 5.88 0.500 24.425 23.675 35.88 7.32 7.370 0.187 797098 0.070 0.0095 1.57 0.610 30 29.250 30.56 2 7.650 0.194 827381 0.060 0.0078 0.90 0.700 34.195 33.445 27.68 -0.88 7.760 0.197 839278 0.055 0.0071 0.58 0.800 39.080 38.330 24.98 -3.58 7.830 0.199 846849 0.060 0.0077 0.67 0.900 43.965 43.215 22.81 -5.75 7.758 0.197 839062 0.050 0.0064 0.64

1 48.850 48.100 21.10 -7.46 7.718 0.196 834736 0.050 0.0065 0.59

The first station listed in table B-1, station 7.15, was included since it is located just outboard of the spinner dome. Beyond station 7.15, 10% radius ratio increments were chosen arbitrarily for analysis, including station 29.25, which corresponds to the location where the feather angle is referenced and set (and is therefore known). The drag coefficient values computed in XFOIL were averaged between each consecutive station and applied to the length of that corresponding blade span in the drag calculations, which are listed in the last column (Drag) of table B-1. Summing the tabulated section drag values yields an estimated drag of about 31.27 lb per blade. The feathered propeller drag in the 100 KTAS will be 4 times the per-blade value, plus 10% to account for the hub/spinner drag and three-dimensional (3D) effects (small amount of induced drag due to blade twist).

( )4 31.27 1.1 137.6 lb⋅ ⋅ =

Given this drag value for the feathered propeller in the 100 KTAS wind (for the stand with no ice on it), equation B-1 is solved for drag due to the test stand:

stand total propD D D= −

(B-2)

Run 15B on the FAA test matrix included a drag measurement of the stand without ice on it, with the propeller feathered, which gave a value of 300.6 lb. Substituting values into equation B-2 gives:

= 300.6 137.6stand

D −

B-2

163 lbstandD = From aerodynamic theory, drag can be expressed as follows: D q f= ⋅ (B-3) where is a reference area multiplied by a characteristic drag coefficient, which is often referred to as an effective flat plate area.

f

df C S= (B-4) The dynamic pressure, q, is given by

21

ρ2

q V= ⋅ (B-5)

where V is the flow velocity and ρ is the air mass density (which is a function of known air temperature and pressure). Given equation B-5, q was calculated based on the temperature and pressure of the test chamber during run 15B and the 100 KTAS tunnel speed.

( )210.00262 168.8

2q = ⋅

237.28

lbq

ft=

Solving equation B-3 for f

D

fq

= (B-6)

Known values are substituted into equation B-6 yielding

24.37cleanf ft=

Again, this value is for the test stand without ice on it.

B-3

Immediately after data was recorded for run 15B, per the FAA test matrix, the propeller was feathered and a drag measurement was taken again with the iced stand. The total drag value was recorded as 330 lb. The effective flat plate area of the stand after a 20-minute exposure to supercooled large drop icing conditions is computed in the same manner as . cleanf

330 137.6stand iceD = − 192.4 lb

icestandD =

237.28

lbq

ft=

D

fq

=

25.16icedf ft=

Note that the same estimated drag value due to the feathered propeller and spinner is used for both the clean and iced cases to compute the effective flat plate area of the stand. This approximation was made for the following reasons: • Because of the blade twist, much of the inboard portion of all blades are stalled, even

without the presence of ice. Adding scattered ice accretions to stalled regions of the blades should not significantly affect drag contributions.

• Insufficient test data is available to determine the effects of spinner ice on the drag due to the spinner.

B.3 COMPUTING AREA-WEIGHTED SLIPSTREAM DYNAMIC PRESSURES. Using the Hartzell PROP code, values for the ratio of the far wake velocity to the freestream velocity may be computed at discrete points along the blade radius for a given condition. Values for dynamic pressure downstream of the propeller were computed analytically at radius ratios of r/R = 0.2, 0.3, 0.45, 0.6, 0.7, 0.8, 0.9, and 0.95. These values were then used to compute a single area-weighted dynamic pressure, as follows: 1. q’s are computed for each local wake velocity given. A local disc area can be computed

for the vicinity of each q, viewing the propeller disc area as a set of concentric disc areas (each having an unique associated q), as shown in figure B-1.

B-4

2. The values of q are multiplied by their respective disc areas.

3. The values of q* area are summed and divided by the total disc area of the propeller.

Figure B-1. Example of Propeller Disc Delineated Into Concentric Areas This procedure yields an area-weighted value of dynamic pressure, which may be used in conjunction with the f value computed in the previous section to approximate test stand drag. Table 2 is a sample of the spreadsheet used to perform the area-weighted q calculation.

Table B-2. Sample Area-Weighted q Calculation

r/R W_bar

Axial Velocity

(ft/s)

Disc Area (ft2)

Section Area (ft2)

Section q

(lb/ft2)

Weighted Avg. q (lb/ft2)

0 0.00 0.00 0.00 0.1 0.52 0.2 12.43 45.34 2.08 2.73 2.64 7.22 0.25 3.25 0.3 25.03 87.89 4.69 4.07 9.93 40.37 0.375 7.32 0.45 37.24 129.09 10.54 7.03 21.41 150.50 0.525 14.35 0.6 38.84 134.51 18.74 7.65 23.25 177.79 0.65 22.00 0.7 38.64 133.82 25.51 7.29 23.01 167.73 0.75 29.28

B-5

B-6

Table B-2. Sample Area-Weighted q Calculation (Continued)

r/R W_bar

Axial Velocity

(ft/s)

Disc Area (ft2)

Section Area (ft2)

Section q

(lb/ft2)

Weighted Avg. q (lb/ft2)

0.8000 37.80 130.99 33.32 8.33 22.05 183.66 0.85 37.61 0.9 39.16 135.59 42.17 6.93 23.62 163.73 0.925 44.55 0.95 42.36 146.39 46.99 7.52 27.54 206.99 1 52.06

Sum of weighted q’s divided by total area 21.09

B.4 METHOD APPLIED TO RECORDED THRUST DATA. The preceding results, for both the iced and non-iced test stand, were used to compute slipstream drag correction values at three power settings (35%, 80%, and 100%), as these settings will be very close to the actual power settings used during the test runs.

Three different temperatures (4°, 10°, and 15°F) were used to compute more accurate values at each of the power settings and each for an rpm setting of 1591 (100% torque), and 1520 (96% torque). For each test run, the corresponding combination of power setting, ambient temperature, and rpm should be located on the tables below, or the values corresponding to a condition most closely matching the actual. The tabulated data should be used with some judgment in regard to the amount of ice that may have been present on the test stand at the time that the data being corrected was recorded. For example, after 20 minutes of exposure to very low liquid water content icing conditions and small droplet sizes, the correction factor closer to the non-iced stand rather than the iced stand should be chosen. Once an appropriate drag correction factor is located in tables B-3, B-4, and B-5, it should be added to the actual recorded thrust data to reach a value that more closely represents the actual performance of the propeller (isolated propeller thrust).

Table B-3. Slipstream + Stand Drag Corrections for 35% Torque

Iced Stand T = 4F T = 10F T = 15F

1591 rpm 308 306 304 1520 rpm 314 312 309

Non-Iced Stand T = 4F T = 10F T = 15F

1591 rpm 261 259 257 1520 rpm 266 264 262

Table B-4. Slipstream + Stand Drag Corrections for 80% Torque

Iced Stand

T = 4F T = 10F T = 15F 1591 rpm 463 461 458 1520 rpm 475 473 471

Non-Iced Stand T = 4F T = 10F T = 15F

1591 rpm 392 390 388 1520 rpm 402 401 399

Table B-5. Slipstream + Stand Drag Corrections for 100% Torque

Iced Stand

T = 4F T = 10F T = 15F 1591 rpm 529 526 524 1520 rpm 543 540 538

Non-Iced Stand T = 4F T = 10F T = 15F

1591 rpm 448 446 444 1520 rpm 460 458 455

B-7/B-8

APPENDIX C—PLANNED TEST MATRIX

C-1

Plan

ned

Test

Con

ditio

ns

Run

N

umbe

r R

PM

MP/

%

Torq

ue

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de

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(lb

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ce

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de

Rad

ius

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(min

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36

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70

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100

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57

16

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2 96

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36

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70

8 26

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FF

0.75

S.

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100

9.9

0.57

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3

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36

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5 26

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708

260

OFF

1.

75

S. L

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0 4.

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57

16

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85

4 96

%

36

0.5

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70

8 26

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FF

0.75

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L.

100

15.2

0.

57

16

22.4

3 0.

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5 96

%

36

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70

7 26

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FF

0.75

S.

L.

100

20.5

0.

64

16

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6 96

%

36

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0.75

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18.0

0.

64

15

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7 96

%

36

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100

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0.

71

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0%

TBD

TB

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96%

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9 23

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402

150

34/3

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ra

d cn

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. 10

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6.5

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19

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68

rad

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boot

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0.

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360

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ra

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68

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354

133

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68

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68

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0.

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(Con

tinue

d)

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r rp

m

MP/

%

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ue

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a @

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=30"

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(lb

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co

ntin

uous

ra

d cn

tr bo

ot

S. L

. 10

0 TB

D

TBD

TB

D

20

TBD

46

22

00

co

ntin

uous

ra

d cn

tr bo

ot

S. L

. 10

0 TB

D

TBD

TB

D

20

TBD

47

25

00

90

/90

rad

cntr

boot

S.

L.

100

TBD

TB

D

TBD

20

TB

D

48

2200

90/9

0 ra

d cn

tr bo

ot

S. L

. 10

0 TB

D

TBD

TB

D

20

TBD

C-3

A

OA

= A

ngle

of a

ttack

Fr

z Fr

ac =

Fre

ezin

g fr

actio

n H

P =

Hor

se p

ower

K

TAS

= K

not t

rue

airs

peed

LW

C =

Liq

uid

wat

er c

onte

nt

MP

= M

anifo

ld p

ress

ure

MV

D =

Med

ian

volu

me

diam

eter

O

AT

= O

utsi

de a

ir te

mpe

ratu

re

rad

cntr

boot

= R

adia

l cen

ter o

f boo

t rp

m =

Rev

olut

ions

per

min

ute

S.L.

= S

ea le

vel

TBD

= to

be

dete

rmin

ed

Ref

eren

ce C

ondi

tions

Run

N

umbe

r M

P/%

To

rque

A

ltitu

de

KTA

S O

AT

LWC

M

VD

Spra

y Ti

me

(min

) En

gine

1

58%

5,

500

186

12

0.57

15

20

TP

E331

-10-

511M

A

2 58

%

5,50

0 18

6 6

0.57

15

20

TP

E331

-10-

511M

A

3 58

%

5,50

0 18

6 0

0.57

15

20

4 58

%

5,50

0 18

6 12

0.

57

15

20

TPE3

31-1

0-51

1M A

5

58%

5,

500

186

18

0.64

15

20

TP

E331

-10-

511M

A

6 58

%

5,50

0 18

6 18

0.

64

15

20

TPE3

31-1

0-51

1M A

7

58%

5,

000

186

24

0.71

15

20

TP

E331

-10-

511M

A

8 M

CP

M

ost c

ritic

al o

f ru

ns 2

, 2a,

2b,

4,

6

20

TPE3

31-1

0-51

1M A

9

58%

8,

500

186

-22

0.2

15

20

TPE3

31-1

0-51

1M A

10

58

%

8,50

0 18

6 -2

2 0.

2 15

20

TP

E331

-10-

511M

A

11

58%

5,

500

186

12

0.57

15

20

TP

E331

-10-

511M

A

12

58%

5,

500

186

18

0.64

15

20

TP

E331

-10-

511M

A

13

58%

5,

500

186

18

0.64

15

20

TP

E331

-10-

511M

A

14

58%

5,

500

186

18

0.64

15

20

TP

E331

-10-

511M

A

15

58%

16

,100

24

0 12

0.

57

15

20

TPE3

31-1

0-51

1M A

16

58

%

16,1

00

240

12

0.57

15

20

TP

E331

-10-

511M

A

17

58%

16

,100

24

0 12

2.

44

15

0.65

TP

E331

-10-

511M

A

18

58%

16

,100

24

0 12

2.

44

15

0.65

TP

E331

-10-

511M

A

19

58%

16

,100

24

0 12

0.

1 40

20

TP

E331

-10-

511M

A

20

58%

16

,100

24

0 12

0.

1 40

20

TP

E331

-10-

511M

A

21

58%

16

,100

24

0 12

TB

D

SLD

20

TP

E331

-10-

511M

A

22

58%

5,

000

186

24

0.71

15

20

TP

E331

-10-

511M

A

23

58%

5,

500

186

12

0.57

15

20

TP

E331

-10-

511M

A

24

58%

5,

500

186

12

0.57

15

20

TP

E331

-10-

511M

A

25

58%

16

,100

24

0 12

TB

D

SLD

20

TP

E331

-10-

511M

A

C-4

Ref

eren

ce C

ondi

tions

(Con

tinue

d)

Run

N

umbe

r M

P / %

To

rque

A

ltitu

de

KTA

S O

AT

LWC

M

VD

Spra

y Ti

me

(min

) En

gine

26

58

%

M

ost c

ritic

al o

f 8-1

8

15

20

TP

E331

-10-

511M

B

27

58%

Mos

t crit

ical

of 2

2-23

15

20

TP

E331

-10-

511M

B

28

20 in

Hg

8,50

0 13

5 -2

2 0.

2 15

20

TI

O-5

40-J

2BD

A

29

29 in

Hg

8,50

0 13

5 -2

2 0.

2 15

20

TI

O-5

40-J

2BD

A

30

26 in

Hg

6,00

0 13

5 12

0.

57

15

20

TIO

-540

-J2B

D A

31

26

in H

g 6,

000

135

12

0.57

15

20

TI

O-5

40-J

2BD

A

32

26 in

Hg

6,00

0 13

5 18

0.

64

15

20

TIO

-540

-J2B

D A

33

29

in H

g 6,

000

135

18

0.64

15

20

TI

O-5

40-J

2BD

A

34

20 in

Hg

5,00

0 13

5 24

0.

71

15

20

TIO

-540

-J2B

D A

35

26

in H

g 5,

000

135

24

0.71

15

20

TI

O-5

40-J

2BD

A

36

26 in

Hg

8,50

0 13

5 -2

2 0.

2 15

20

TI

O-5

40-J

2BD

A

37

26 in

Hg

6,00

0 13

5 12

0.

57

15

20

TIO

-540

-J2B

D A

38

26

in H

g 6,

000

135

18

0.64

15

20

TI

O-5

40-J

2BD

A

39

26 in

Hg

6,00

0 13

5 18

TB

D

SLD

20

TI

O-5

40-J

2BD

A

40

26 in

Hg

M

ost c

ritic

al o

f 26-

33

15

20

TIO

-540

-J2B

D B

41

26 in

Hg

8,50

0 13

5 -2

2 0.

2 15

20

TI

O-5

40-J

2BD

B

42

26 in

Hg

6,00

0 13

5 12

0.

57

15

20

TIO

-540

-J2B

D B

43

26

in H

g 6,

000

135

18

0.64

15

20

TI

O-5

40-J

2BD

B

44

29 in

Hg

M

ost c

ritic

al o

f 39-

41

15

20

TIO

-540

-J2B

D C

45

26

in H

g

Mos

t crit

ical

of 3

9-41

15

20

TI

O-5

40-J

2BD

C

46

26 in

Hg

6,00

0 13

5 18

TB

D

SLD

20

TI

O-5

40-J

2BD

C

47

29 in

Hg

Sa

me

as 4

2

15

20

TI

O-5

40-J

2BD

C

48

26 in

Hg

Sa

me

as 4

3

15

20

TI

O-5

40-J

2BD

C

C-5

KTA

S =

Kno

t tru

e ai

rspe

ed

M

P =

Man

ifold

pre

ssur

e

OA

T =

Out

side

air

tem

pera

ture

LW

C -

Liqu

id w

ater

con

tent

MV

D =

Med

ian

volu

me

diam

eter

Pr

opel

ler S

pecs

R

un

Num

ber

Hub

B

lade

Ty

pe

Bla

de

Con

ditio

n Pr

op D

ia

Inch

es

Prop

Dia

fe

et

1 H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 2

HC

-B4T

N-5

( )L

LT10

282N

SB-5

.3R

4-

alum

inum

N

ew/O

ld

98

8.2

3 H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 4

HC

-B4T

N-5

( )L

LT10

282N

SB-5

.3R

4-

alum

inum

N

ew/O

ld

98

8.2

5 H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 6

HC

-B4T

N-5

( )L

LT10

282N

SB-5

.3R

4-

alum

inum

N

ew/O

ld

98

8.2

7 H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 8

HC

-B4T

N-5

( )L

LT10

282N

SB-5

.3R

4-

alum

inum

N

ew/O

ld

98

8.2

9 H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 10

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 11

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 12

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 13

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 14

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 15

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 16

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 17

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 18

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 19

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 20

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 21

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 22

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 23

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 24

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 25

H

C-B

4TN

-5( )

L LT

1028

2NSB

-5.3

R

4-al

umin

um

New

/Old

98

8.

2 26

M

TV-2

7-2-

E-C

-F-R

(G)

CFR

L250

-55

5-co

mpo

site

N

ew

98

8.2

C-6

Pr

opel

ler S

pecs

(Con

tinue

d)

Run

N

umbe

r H

ub

Bla

de

Type

B

lade

C

ondi

tion

Prop

Dia

In

ches

Pr

op D

ia

feet

27

M

TV-2

7-2-

E-C

-F-R

(G)

CFR

L250

-55

5-co

mpo

site

N

ew

98

8.2

28

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

29

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

30

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

31

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

32

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

33

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

34

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

35

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

36

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

37

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

38

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

39

HC

-I2Y

R-1

BF

F807

4 2-

alum

inum

N

ew

80

6.7

40

HC

-I3Y

R-1

E 78

90K

3-

com

posi

te

New

80

6.

7 41

H

C-I

3YR

-1E

7890

K

3-co

mpo

site

N

ew

80

6.7

42

HC

-I3Y

R-1

E 78

90K

3-

com

posi

te

New

80

6.

7 43

H

C-I

3YR

-1E

7890

K

3-co

mpo

site

N

ew

80

6.7

44

MTV

-14-

B

195-

30a

4-co

mpo

site

N

ew

77

6.4

45

MTV

-14-

B

195-

30a

4-co

mpo

site

N

ew

77

6.4

46

MTV

-14-

B

195-

30a

4-co

mpo

site

N

ew

77

6.4

47

MTV

-14-

B

195-

30a

4-co

mpo

site

N

ew

77

6.4

48

MTV

-14-

B

195-

30a

4-co

mpo

site

N

ew

77

6.4

C-7

C-8

Prop

Ice

Spec

s R

un

Num

ber

Dei

ce P

/N

Goo

dric

h K

it N

o.

Vol

tage

/Cur

rent

Ti

m

er S

eque

nce

1 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

2 4E

1188

-8

65-1

66

28 V

DC

1

8-22

A

Dei

cer o

ff

3 4E

1188

-9

65-1

67

28 V

DC

1

8-22

A

Dei

cer o

ff

4 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

5 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

6 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

7 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

8 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Dei

cer o

ff

9 4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

10

4E11

88-7

65

-165

28

VD

C

18-

22A

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

11

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

12

4E11

88-7

65

-165

28

VD

C

18-

22A

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

13

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

14

4E11

88-7

65

-165

28

VD

C

18-

22A

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

15

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

16

4E11

88-7

65

-165

28

VD

C

18-

22A

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

17

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

18

4E11

88-7

65

-165

28

VD

C

18-

22A

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

19

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

20

4E11

88-7

65

-165

28

VD

C

18-

22A

D

eice

r off

21

4E

1188

-7

65-1

65

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

22

4E11

88-7

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

23

4E11

88-7

28 V

DC

1

8-22

A

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

24

4E28

37-1

0

28 V

DC

op

posi

ng b

lade

s 25

4E

2837

-10

28

VD

C

oppo

sing

bla

des

26

4E42

15-4

N

/A

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

27

4E42

15-4

N

/A

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

Pr

op Ic

e Sp

ecs (

Con

tinue

d)

Run

N

umbe

r D

eice

P/N

G

oodr

ich

Kit

No.

V

olta

ge/C

urre

nt

Tim

er S

eque

nce

28

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

29

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

30

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

31

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

32

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

33

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

34

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

35

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

36

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

37

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

38

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

39

4E22

00-3

67

-585

28

VD

C

8-1

2A

All

Bla

des S

imul

tane

ous

40

4E39

97-1

5E

2627

28

VD

C

16-

20A

A

ll B

lade

s Sim

ulta

neou

s 41

4E

3997

-1

5E26

27

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

42

4E39

97-1

5E

2627

28

VD

C

16-

20A

A

ll B

lade

s Sim

ulta

neou

s 43

4E

3997

-1

5E26

27

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

44

2x4E

3278

-4 &

2x

4E42

15-4

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

45

2x4E

3278

-4 &

2x

4E42

15-4

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

46

2x4E

3278

-4 &

2x

4E42

15-4

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

47

2x4E

3278

-4 &

2x

4E42

15-4

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

48

2x4E

3278

-4 &

2x

4E42

15-4

28 V

DC

1

6-20

A

All

Bla

des S

imul

tane

ous

C-9/C-10

P/N

= P

art n

umbe

r Pr

op D

ia =

Pro

pelle

r dia

met

er

APPENDIX D—RUN SUMMARY

D-1

Run

N

umbe

r D

ate

(m/d

/y)

Engi

ne

Star

t (h

r/min

/sec

)

Engi

ne

Off

(h

r/min

/sec

)

Win

d O

n (h

r/min

/sec

)

Win

d O

ff

(hr/m

in/s

ec)

Boo

ts

Spra

y O

n (h

r/min

/sec

)

Boo

ts

Spra

y O

ff

(hr/m

in/s

ec)

LWC

(g

/m3 )

MV

D

(µm

) O

AT

(F)

Dei

ce

16

11/1

5/06

95

455

1016

15

1000

00

10

0318

10

1503

1.

04

22

12

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

15

11/1

5/06

13

3335

14

0255

13

4258

1347

53

1402

03

1.04

22

12

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

21

11

/15/

06

1654

00

1721

30

1701

47

1709

41

1709

33

1721

03

0.36

96

12

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

18

11

/16/

06

1151

57

1201

06

1155

14

11

5937

11

0016

2.

44

15

12

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

15A

11

/16/

06

1330

55

1354

33

1336

28

13

3936

13

5336

1.

04

22

12

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

15B

11

/16/

06

1607

53

1640

26

1615

59

1622

34

1622

28

1638

00

1.04

22

12

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

21

A

11/1

6/06

19

2036

19

5634

19

2231

19

4115

19

2834

19

4059

0.

36

96

12

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

se

cond

spra

y 19

5130

19

5427

3A

11

/17/

06

1026

10

1052

44

1029

59

1050

28

1037

23

1047

58

0.33

16

.5

4.6

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

3B

11/1

7/06

13

0817

13

2218

13

1148

13

2100

13

1645

13

2100

0.

33

16.5

4.

6 A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

4A

11

/17/

06

1505

31

1533

41

1507

40

1531

38

1516

14

1528

07

0.57

16

.5

15.2

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

21

B

11/1

8/06

90

538

9264

5 90

723

9251

7 91

313

9251

2 0.

36

96

24

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

19A

11

/18/

06

1111

14

1135

59

1112

49

1134

35

1119

29

1134

30

0.1

40

12

Alte

rnat

e 34

Inbd

/ 34

Out

bd /

68 O

ff

21C

11

/18/

06

1349

01

1409

57

1350

37

1408

55

1356

15

1408

49

0.36

96

12

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

4B

11

/18/

06

1603

07

1627

43

1605

10

1612

24

1611

50

1626

24

0.45

20

12

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

4C

11

/18/

06

1815

24

1841

27

1821

44

1839

30

1826

12

1840

06

0.52

20

22

A

ltern

ate

34 In

bd /

34 O

utbd

/ 68

Off

26

11

/20/

06

Win

d of

f sta

tic th

rust

on

MT

Prop

26A

11

/20/

06

1429

26

1501

27

1435

40

1459

30

1445

26

1459

30

0.4

20

12

ON

Con

tinuo

us

26B

11

/20/

06

1612

05

1635

51

1613

49

1634

12

1619

37

1634

12

0.52

20

22

O

N C

ontin

uous

26

C

11/2

0/06

17

3351

17

5829

17

3700

17

5619

17

4243

17

5619

0.

36

96

24

ON

Con

tinuo

us

25

11/2

1/06

11

1505

11

4122

11

1749

11

3928

11

2551

11

3928

0.

36

96

24

90O

n/90

Off

Alte

rnat

e O

ppos

ite B

lade

s 25

A

11/2

1/06

12

2435

12

3340

Tare

dra

g - i

ce o

n pr

op, n

one

on st

and

OFF

24

11

/21/

06

1413

00

1637

55

1414

50

1435

52

1420

51

1435

52

0.66

15

19

90

On/

90O

ff A

ltern

ate

Opp

osite

Bla

des

24A

11

/21/

06

1610

18

1637

48

1612

00

1635

55

1618

35

1635

55

0.4

20

12

90O

n/90

Off

Alte

rnat

e O

ppos

ite B

lade

s 24

B

11/2

1/06

17

4619

18

0808

17

4747

18

0616

17

5502

18

0616

0.

36

96

12

90O

n/90

Off

Alte

rnat

e O

ppos

ite B

lade

s 24

C

11/2

1/06

18

3725

18

4519

Tare

dra

g - i

ce o

n st

and,

non

e on

pro

p O

FF

D-2

LWC

= L

iqui

d w

ater

con

tent

M

VD

= M

edia

n vo

lum

e di

amet

er

OA

T =

Out

side

air

tem

pera

ture

Spra

y Ti

me

Run

N

umbe

r m

in

sec

Spra

y To

rque

(%

) rp

m

Bet

a Th

rust

(lb

)

STA

RT

23

1480

24

.5

76

NO

ICE

16

11

45

STO

P 20

14

90

22.5

0

ICE

STA

RT

30

1480

25

15

0 N

O IC

E 15

14

10

ST

OP

20

1500

22

0

ICE

STA

RT

23

1450

25

.1

47

NO

ICE

21

11

30

STO

P 21

14

50

24

0 IC

E ST

AR

T 21

15

20

24.6

7

NO

ICE

18

0 39

ST

OP

20

1520

24

.6

12

ICE

STA

RT

33

1520

26

.6

200

NO

ICE

15A

14

0

STO

P N

o da

ta

No

data

23

0

ICE

STA

RT

38

1600

26

30

5 N

O IC

E 15

B

15

32

STO

P N

o da

ta

No

data

25

25

IC

E ST

AR

T 45

16

00

26.2

45

0 N

O IC

E ST

OP

36

1600

25

50

IC

E 21

A

12

25

IC

E ST

AR

T 39

14

50

26.2

35

0 N

O IC

E 3A

10

35

ST

OP

39

1450

26

20

0 IC

E ST

AR

T 34

14

80

26.2

27

0 N

O IC

E 3B

4

15

STO

P 34

14

80

26

116

ICE

STA

RT

42

1520

26

.2

415

NO

ICE

4A

11

53

STO

P 42

15

20

26

220

ICE

STA

RT

36

1540

26

.2

230

NO

ICE

21B

11

59

ST

OP

32

1545

25

30

IC

E ST

AR

T 35

15

35

26.2

26

0 N

O IC

E 19

A

15

1 ST

OP

37

1535

26

.2

250

ICE

D-3

rpm

= R

evol

utio

ns p

er m

inut

e ST

AR

T =

Para

met

ers s

et a

t sta

rt of

spra

y ST

OP

= Pa

ram

eter

s at e

nd o

f spr

ay

Sp

ray

Tim

e R

un

Num

ber

min

se

c Sp

ray

Torq

ue

(%)

rpm

B

eta

Thru

st

(lb)

STA

RT

39

1540

26

.2

280

NO

ICE

21C

13

34

ST

OP

29

1545

23

.9

-200

IC

E ST

AR

T 36

15

60

26.2

30

0 N

O IC

E 4B

14

34

ST

OP

35

1560

25

10

0 IC

E ST

AR

T 33

15

20

26.7

22

0 N

O IC

E 4C

13

54

ST

OP

32

1520

26

.7

50

ICE

26

STA

RT

40

1485

15

.2

330

NO

ICE

26A

14

4

STO

P 30

14

85

13.8

30

IC

E ST

AR

T 42

15

20

15.2

40

0 N

O IC

E 26

B

14

35

STO

P 37

15

20

15

190

ICE

STA

RT

42

1520

15

.2

400

NO

ICE

26C

13

36

ST

OP

35

1520

14

.5

75

ICE

STA

RT

50

1600

26

.2

510

NO

ICE

25

13

37

STO

P 50

16

00

26.2

41

0 IC

E 25

A

IC

E O

N P

RO

P O

NLY

ST

AR

T 42

15

96

26.2

43

6 N

O IC

E 24

15

1

STO

P 39

15

95

26

267

ICE

STA

RT

45

1595

26

.2

490

NO

ICE

24A

17

20

ST

OP

35

1520

26

11

0 IC

E ST

AR

T 36

15

30

26.1

29

0 N

O IC

E 24

B

11

14

STO

P 32

15

40

24.8

-1

0 IC

E 24

C

N

O IC

E

D-4

rp

m =

Rev

olut

ions

per

min

ute

STA

RT

= Pa

ram

eter

s set

at s

tart

of sp

ray

STO

P =

Para

met

ers a

t end

of s

pray

Mea

sure

d Th

rust

at

Thru

st lo

ss b

ased

on

Har

tzel

l met

hod

Perf

orm

ed o

n 10

0% th

rust

mea

sure

men

ts

65

%

80%

10

0%

R

un

Num

ber

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Har

tzel

l C

orre

ctio

n (lb

)

Prop

elle

r Th

rust

(lb

)

Del

ta

Thru

st

(lb)

Thru

st

Loss

%

NO

ICE

ICE

16

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

NO

ICE

ICE

15

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

NO

ICE

ICE

21

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

NO

ICE

ICE

18

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

No

thru

st m

easu

rem

ents

take

n

NO

ICE

ICE

15A

N

o th

rust

mea

sure

men

ts ta

ken

N

o th

rust

mea

sure

men

ts ta

ken

N

o th

rust

mea

sure

men

ts ta

ken

N

O IC

E

10

30

31.7

80

16

00

1200

34

10

0 16

00

446

1646

22

0 13

.37%

IC

E 15

B

730

32.5

N

o da

ta

1590

90

0 34

10

0 15

90

526

1426

N

O IC

E

10

22

31.3

80

16

00

1200

33

.5

100

1600

44

6 16

46

190

11.5

4%

ICE

740

31.4

80

15

95

930

33.6

10

0 15

95

526

1456

IC

E

21A

700

31.4

80

15

95

920

33.6

10

0 15

95

NO

ICE

1020

31

.3

82

1510

12

55

33.3

10

0 15

15

458

1713

-8

-0

.47%

IC

E 3A

94

2 31

.1

81

1520

11

81

33.4

10

0 15

05

540

1721

N

O IC

E

12

70

33.6

10

0 15

20

458

1728

38

2.

20%

IC

E 3B

11

50

33.1

10

0 15

20

540

1690

N

O IC

E 78

5 20

.7

64

1520

99

8 30

.5

80

1520

12

30

32.5

10

0 15

20

455

1685

37

2.

20%

IC

E 4A

72

4 29

.9

65

1520

92

0 31

.8

80

1520

11

10

33.7

10

0 15

20

538

1648

N

O IC

E 76

0 30

.3

65

1540

12

10

33.6

10

0 15

95

444

1654

18

7 11

.31%

IC

E 21

B

585

29.5

65

16

00

943

33.9

10

0 15

90

524

1467

N

O IC

E 79

0 29

65

15

90

1250

33

.1

100

1590

44

6 16

96

40

2.36

%

ICE

19A

68

0 30

65

16

00

1130

34

10

0 15

85

526

1656

N

O IC

E 78

0 29

.5

64

1600

12

25

33.2

10

0 16

00

446

1671

35

5 21

.24%

IC

E 21

C

425

29.5

65

15

90

790

33.2

99

15

90

526

1316

N

O IC

E 80

0 29

.7

65

1600

12

25

33.5

10

0 15

90

446

1671

12

5 7.

48%

IC

E 4B

61

0 28

.5

65

1600

10

20

33.2

10

0 15

90

526

1546

D-5

rp

m =

Rev

olut

ions

per

min

ute

Mea

sure

d Th

rust

at

Thru

st lo

ss b

ased

on

Har

tzel

l met

hod

Perf

orm

ed o

n 10

0% th

rust

mea

sure

men

ts

65

%

80%

10

0%

R

un

Num

ber

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Thru

st

(lb)

Bet

a To

rque

(%

) rp

m

Har

tzel

l C

orre

ctio

n (lb

)

Prop

elle

r Th

rust

(lb

)

Del

ta

Thru

st

(lb)

Thru

st

Loss

%

NO

ICE

777

29.2

11

50

31.4

44

6 15

96

36

2.26

%

ICE

4C

690

30.1

66

16

00

1034

33

.8

1034

33

.8

526

1560

26

NO

ICE

770

17.7

64

15

20

1275

21

.6

100

1515

45

8 17

33

173

9.98

%

ICE

26A

58

0 18

.2

64

1510

10

20

22.1

10

0 15

05

540

1560

N

O IC

E 77

0 18

.1

65

1510

12

85

21.9

10

0 15

10

455

1740

18

2 10

.46%

IC

E 26

B

590

18.7

65

15

00

1020

22

.4

100

1500

53

8 15

58

NO

ICE

800

18

64

1520

12

70

22

100

1520

45

5 17

25

187

10.8

4%

ICE

26C

62

2 18

.6

64

1520

10

00

23.3

10

0 15

20

538

1538

N

O IC

E 77

5 28

.4

65

1595

12

30

32.8

10

0 15

95

444

1674

15

0 8.

96%

IC

E 25

57

5 28

.4

65

1595

10

00

33.7

10

0 15

95

524

1524

N

O IC

E 73

0 29

.2

65

1595

11

50

33.3

10

0 15

90

11

50

ICE

25A

N

O IC

E 82

0 29

.5

65

1600

13

00

33.4

10

0 16

00

444

1744

96

5.

50%

IC

E 24

70

8 29

.6

65

1595

11

24

33.7

10

0 15

95

524

1648

N

O IC

E 79

5 29

.5

65

1600

12

65

33.9

10

0 15

95

446

1711

95

5.

55%

IC

E 24

A

650

29.7

65

15

95

1090

33

.7

100

1590

52

6 16

16

NO

ICE

790

29.5

64

15

95

1275

33

.7

100

1585

44

6 17

21

285

16.5

6%

ICE

24B

44

0 29

64

15

95

910

34

101

1590

52

6 14

36

NO

ICE

24C

67

0 29

.5

65

1590

10

80

33.5

10

1 15

90

D-6

rpm

= R

evol

utio

ns p

er m

inut

e

R

un

Num

ber

LWC

M

VD

O

AT

Bla

de

Num

ber

Ice

Thic

knes

s at

Bla

de R

adiu

s

0.

28

0.5

0.75

O

ther

s

16

1.04

22

12

3

0.

06

0.07

2

0.65

0.

085

0.22

15

1.04

22

12

1

21

0.36

96

12

2

0.12

5

3 0.

125

0.14

5 0.

18

4

0.12

0.

165

18

2.44

15

12

2

0.06

0.

11

0.17

5

3

0.

15

15

A

1.04

22

12

2

0.

15

0.12

7

3

0.35

0.

03

15

B

1.04

22

12

2

0.12

0.

175

0

3 0.

4 0.

195

0.3

21

A

0.36

96

12

1

0.7

0.

07

3

0.

01

0.06

2

0.29

Rou

ghne

ss @

.5R

= .2

6 3A

0.

33

16.5

4.

6 4

0.

06

0.07

3 0.

69

0.08

5

1

0.

29

3B

0.33

16

.5

4.6

4 0.

29

0.46

5 0.

255

Feat

her h

eigh

ts =

0.1

15

3 0.

265

0.17

5 0.

235

Feat

her h

eigh

ts =

0.1

2 4A

0.

57

16.5

15

.2

3 0.

18

4

0.13

5

Ed

ge T

hick

ness

@ .6

5R =

0.0

3 21

B

0.36

96

24

3

0.

11

0.04

R

ough

ness

: In

ner B

oot =

0.4

, Out

er B

oot =

0.2

9, .4

R =

0.4

8, .6

5R =

0.2

2

4

0.

08

0.02

5 R

ough

ness

: M

id B

oot =

0.3

4, .6

5R =

0.2

5

2

0.

06

0.02

D-7

D-8

C

M

R

un

Num

ber

LWV

D

OA

T B

lade

N

umbe

r Ic

e Th

ickn

ess

at B

lade

Rad

ius

19

A

0.1

40

12

2 0.

145

0.21

0.

065

Rou

ghne

ss: M

id B

oot =

0.2

6, .5

R =

0.1

45

1 0.

095

0.14

5 0.

09

Rou

ghne

ss: M

id B

oot =

0.1

85, .

5R =

0.1

45

21C

0.

36

96

12

4 0.

04

0.16

0.

25

Thic

kest

ice

on tr

ace

at .5

R =

0.2

8

3 0.

035

0.16

5 0.

1 Th

icke

st ic

e on

trac

e: .

5R =

0.2

0, .7

5R =

0.2

85 -

Rou

ghne

ss

at M

id B

oot =

0.4

5

4B

0.45

20

12

2

47

13

0.24

5 0.

0.R

ough

ness

: .5R

= 0

.095

, .75

R =

0.01

-> 0

.02

- At w

hite

st

ripe

= 0.

18

4 0.

39

0.24

5

Inbo

ard

Whi

te S

tripe

= 0

.245

1

0.24

0.

05

0.02

4C

0.52

20

22

1

0.

105

0.02

Th

icke

st ic

e on

trac

e at

mid

boo

t = 0

.225

2

0.13

0.

19

0.08

Th

ickn

ess i

ce o

n tra

ce @

.55R

= 0

.49

26A

0.

4 20

12

1

Ed

ge th

ickn

ess:

Mid

Boo

t = 0

.34,

24.

75"

= 0.

205,

13

" = 0

.12,

4" =

0.3

6

2

Edge

thic

knes

s: 3

8" =

0.2

95, 2

8" =

0.5

9, 1

5.5"

= 0

.235

, 3.

25" =

0.1

0

4

Edge

thic

knes

s: M

id B

oot =

0.1

7, 2

8.5"

= 0

.70,

22"

= 0

.325

, 12

.25"

= 0

.10,

3"

= 0.

075

26B

0.

52

20

22

3

Edge

thic

knes

s: M

id B

oot =

0.1

8, 2

2.5"

= 0

.275

, 17

.25"

= 0

.345

, 13"

= 0

.105

1

Edge

thic

knes

s: M

id B

oot =

0.1

7, B

oot E

dge

@

26.8

75"

= 0.

372,

17.

75" =

0.1

05

5

Edge

thic

knes

s: M

id B

oot @

36.

75"

= 0.

197,

21

.5" =

0.1

32, 1

3" =

0.1

25

26C

0.

36

96

24

2

Rid

ge h

eigh

ts:

Inbo

ard

Boo

t = 0

.485

, Out

boar

d B

oot =

0.2

32

3

Rid

ge h

eigh

t: .7

5R =

0.1

16

5

Rid

ge h

eigh

t: In

boar

d B

oot =

0.6

40

25

0.36

96

24

N

O tr

acin

g ca

rd

R

un

Num

ber

LWC

M

VD

O

AT

Bla

de

Num

ber

Ice

Thic

knes

s at

Bla

de R

adiu

s

24

0.66

15

19

4

0.

335

Ed

ge th

ickn

ess:

.65

R =

0.2

15

1

0.

14

Edge

thic

knes

s: 7

.75"

= 0

.087

2

0.

255

Ed

ge th

ickn

ess:

8"

= 0.

005

3

0.31

4 0.

09

Edge

thic

knes

s: .

25 B

oot =

1.3

05, 8

.875

" =

0.11

5 24

A

0.4

20

12

2

0.35

0.

19

Edge

thic

knes

s: 4

" =

0.07

8

3

0.

37

0.11

4

0.42

6 0.

85

Edge

thic

knes

s: 6

" =

0.10

5, 3

5.75

" = 1

.20

1

Edge

thic

knes

s: 1

5.75

= 0

.223

, 2.5

" =

0.06

9 24

B

0.36

96

12

N

O tr

acin

g ca

rd

LWC

= L

iqui

d w

ater

con

tent

M

VD

= M

edia

n vo

lum

e di

amet

er

OA

T =

Out

side

air

tem

pera

ture

D-9/D-10