NASA CR-120,992
`LYCOMING 105.22.21
`
`DESIGN STUDY OF AN AIR PUMP AND
`INTEGRAL LIFT ENGINE ALF-504
`USING THE LYCOMING 502 CORE
`
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`Avco lycoming Division
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`SERVICE
`: NATIONAL AERONAUTICS AND SPACE ADMINISTFORMATION
`NASA Lewis Research Center
`
`NATIONAL TECHNICAL INFORMATION SERVICE
`
`GE v. UTC
`IPR2016-00952
`GE-1027.001
`
`

`

`1. Report No.
`NASA CR-120, 992
`4. Title and Subtitle
`DESIGN STUDY OF AN AIR PUMP AND INTEGRAL 'LIFT ENGINE
`ALF-504 USING THE LYCOMING 502 CORE
`
`2.-Government Accession No.
`
`3. Recipient's Catalog No.
`
`5. Report Date
`July 1972
`6. Performing Organization Code
`
`7. Author(s)
`
`Dale Rauch
`_
`_
`
`_Lycoming
`
`9. Performing Organization Name and Address
`~~~~~~~~~~~~Avco
`~11.
`Lyoming
`Stratford, Connecticut 06497
`
`Division'~
`
`12. Sponsoring Agency Name and Address
`National Aeronautics and Space Administration
`Washington, D.C. 20546
`
`15. Supplementary Notes
`
`8. Performing Organization Report No.
`
`Report
`10. Work Unit No.
`
`No. 105. 22. 21
`
`Contract or Grant No.
`NAS 3-15696
`13. Type of Report and Period Covered
`Contractor Report
`
`14. Sponsoring Agency Code
`
`Prepared in cooperation with Project Manager, Laurence W. Gertsma,
`NASA Lewis Research Center, Cleveland, Ohio
`
`16. Abstract
`
`Design studies were conducted for an integral lift fan engine (ALF-504) utilizing the Lycoming
`502 fan core with the finalMQT power turbine. The fan is designed for a 12.5 bypass ratio
`and 1,25:1 pressure ratio, and provides supercharging for the core. Maximum sea level
`static thrust is 8370 pounds with a specific fuel consumption of 0. 302 lb/hr-lb. The dry
`engine weight without starter is 1419 pounds including full-length duct and sound-attenuating
`rings. The engine envelope including duct treatment but not localized accessory protrusion
`is 53. 25 inches in diameter and 59. 2 inches long from exhaust nozzle exit to fan inlet flange.
`
`Detailed analysis includes fan aerodynamics, fan and reduction gear mechanical design, fan
`dynamic analysis, engine noise analysis, engine performance, and weight analysis.
`
`17. Key Words
`
`(Suggested by Author(s))
`
`High bypass ratio fan engine
`
`18. Distribution Statement
`
`Unclassified-unlimited
`
`19. Security Classif. (of this report)
`
`20. Security Classif. (of this page)
`
`Unclassified
`
`Unclas sified
`
`21. No. of Pages
`
`22. Price*
`
`171
`
`$3. 00
`
`l
`
`* For sale by the National Technical Information Service, Springfield, Virginia 22151
`
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`PRiCEDIANG PAGE BLANK NOT FiLb~
`
`FOREWORD
`
`The work reported herein was conducted at Avco Lycoming Division,
`Stratford, Connecticut under NASA Contract No. NAS3-15696. The study
`was conducted under the management of the NASA Lewis Research Cen-
`ter with Mr. Laurence Gertsma as project manager.
`
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`~DIpAGE BLANOK NOT
`
`ABSTRACT
`
`Design studies were conducted for an integral lift fan engine
`(ALF-504) utilizing the Lycoming 502 fan core with the final MQT power
`turbine. The fan is designed for a 12. 5 bypass ratio and 1.25:1 pres-
`sure ratio, and provides supercharging for the core. Maximum sea
`level static thrust is 8370 pounds with a specific fuel consumption of
`0. 302 lb/hr-lb. The dry engine weight without starter is 1419 pounds
`including full-length duct and sound-attenuating rings. The engine
`envelope including duct treatment but not localized accessory protrusion
`is 53. 25 inches in diameter and 59. 2 inches long from exhaust nozzle
`exit to fan inlet flange.
`
`Detailed analysis includes fan aerodynamics, fan and reduction
`gear mechanical design, fan dynamic analysis, engine noise analysis,
`engine performance, and weight analysis.
`
`Preceding page blank
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`fb~PAG BbLNe -NOT FIMD
`
`TABLE OF CONTENTS
`
`FOREWORD
`
`ABSTRACT
`
`LIST OF ILLUSTRATIONS
`
`LIST OF TABLES
`
`SUMMARY
`
`INTRODUCTION
`
`PRELIMINARY STUDIES OF SUPERCHARGED AND NON-
`SUPERCHARGED INTEGRAL LIFT ENGINE AND AIR PUMP
`
`FAN AERODYNAMIC DESIGN
`
`MECHANICAL DESIGN
`
`DYNAMIC ANALYSIS
`
`NOISE ANALYSIS
`
`ENGINE PERFORMANCE
`
`POWER TURBINE ANALYSIS
`
`ENGINE WEIGHT
`
`SCHEDULE AND COST ESTIMATE OF INTEGRAL LIFT
`ENGINE
`
`APPENDIXES
`
`I. Aerodynamic Influence of the Part-Span Shroud
`II. Axisymmetric Flow Solution for the Fan
`III. Dynamic Analysis Methods
`IV. List of Symbols
`
`REFERENCES
`
`BIBLIOGRAPHY
`
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`7
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`LIST OF ILLUSTRATIONS
`
`Figure
`
`Page
`
`1
`
`2
`
`3
`
`4
`
`5
`
`6
`
`7
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`8
`
`9
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`10
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`11
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`12
`
`Rotor Polytropic Efficiency
`
`Fan Meridional Flow Path
`
`Fan Rotor Velocity Triangles
`
`Fan Stator Flow Conditions
`
`Basic Single Stage Performance Map for Fan Duct
`Section (Experimental Transonic Stage)
`
`Estimated Performance Map for Duct Flow Fan A
`(21% Surge Margin at 100% N//8)
`
`Estimated Performance Map for Duct Flow Fan B
`(14% Surge Margin at 100% N//e)
`
`Measured Basic Single Stage Performance Map for
`Supercharger Fan Section (301 Fan Supercharger)
`
`Estimated Performance Map for Fans A and B
`(Supercharger Section)
`
`ALF-504 High-Bypass Fan Engine
`
`ALF-504 Fan Engine Variable Nozzle
`
`ALF-504 Fan Engine Installation Drawing
`
`13
`
`Fan Blade and Disc Stresses
`
`14
`
`Fan Rotor Blade Airfoil Composite, Scale Approxi-
`mately ZX
`
`15
`
`Midspan Shroud
`
`16
`
`Root and Groove Stresses
`
`10
`
`12
`
`13
`
`19
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`22Z
`
`23
`
`24
`
`25
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`27
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`28
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`29
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`30
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`33
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`34
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`35
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`37
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`Figure
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`43
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`44
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`45
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`46
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`47
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`48
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`49
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`50
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`51
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`52
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Standard Day
`
`A 1 8 = 1190
`at 5000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Standard Day
`
`A 1 8 = 1190
`at 10, 000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Standard Day
`
`A 1 8 = 1190
`at 15, 000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Standard Day
`
`A 1 8 = 1190
`at 20, 000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`on Tropical Day (900 F)
`
`A 1 8 = 1190
`at Sea Level
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`on Tropical Day (700 F)
`
`A 1 8 = 1190
`at 5000 Feet
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (510 F)
`
`A 1 8 = 1190
`at 10, 000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (320 F)
`
`=1190
`A
`at 15, 000
`
`Fan Configuration B (SM = 14 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (12 F)
`
`= 1190
`A 1
`at 20, 000
`
`Estimated Fan Performance Map for Configuration A
`Showing Operating Lines With Two-Position Fan Ex-
`haust Nozzle
`
`Page
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`82
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`83
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`Figure
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`53
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`54
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`55
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`56
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`58
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`59
`
`Fan Configuration A (SM = 21 Percent at 100 Percent
`N/e8) Estimated Performance at Sea Level on
`Standard Day With Two-Position Fan Exhaust Nozzle
`
`Modified MQT Power Turbine First Stage Velocity
`Triangles for 8. 5 Percent Reduced Inlet Flow
`Function
`
`Modified MQT Power Turbine Second Stage Velocity
`Triangles for 8.5 Percent Reduced Inlet Flow
`Function
`
`ALF-504 Engine Program
`
`Measured Flow Conditions Downstream of Rotor
`(Reference 1 - "Some Studies of Front Fans With and
`Without Snubbers")
`
`Radial Surveys Downstream of ALF- 502 Fan Rotor,
`Test -02
`
`Analysis of Part-Span Shroud on Axial Velocity Pro-
`file Downstream of Fan Rotor
`
`Pagj
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`92
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`94
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`95
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`98
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`Figure.
`
`17
`
`18
`
`Reduction Gear
`
`Ring Gear Stresses
`
`19
`
`Z0
`
`21
`
`22
`
`23
`
`24
`
`25
`
`26
`
`Z7
`
`28
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`29
`
`30
`
`31
`
`Reduction Gear Stress Analysis
`
`Rotor Blade Frequency Analysis, N = 5245 RPM
`
`Rotor Blade Vibration Interference Diagram
`
`Rotor Blade Flutter Criteria
`
`Fan-Power Turbine Torsion Analysis
`
`Fan Rotor Deflection
`
`Integral Lift Engine Noise at 200 Feet Radius and
`110 Degrees
`
`Engine Stations Used in Performance Evaluation
`
`Fan Configuration A (SM = 21 Percent, A 1 8 = 1190
`Square Inches) Estimated Performance at Sea Level
`on Standard Day
`
`Fan Configuration A (SM = 21 Percent, A 1 8 = 1190
`Square Inches) Estimated Performance at 5000 Feet
`on Standard Day
`
`Fan Configuration A (SM = 21 Percent, A 1 8 = 1190
`Square Inches) Estimated Performance at 10, 000
`Feet on Standard Day
`
`Fan Configuration A (SM = 21 Percent, A 1 8 = 1190
`Square Inches) Estimated Performance at 15, 000
`Feet on Standard Day
`
`Fan Configuration A (SM = 21 Percent, A 1 8 = 1190
`Square Inches) Estimated Performance at 20, 000
`Feet on Standard Day
`
`ix
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`Page
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`Figure
`
`Page
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`34
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`38
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`39
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`40
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`41
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`42
`
`Fan Configuration A (SM = 21 Percent,
`Square Inches) Estimated Performance
`on Tropical Day (900 F)
`
`A 1 8 = 1190
`at Sea Level
`
`Fan Configuration A (SM - 21 Percent,
`Square Inches) Estimated Performance
`on Tropical Day (700 F)
`
`A 1 8 = 1190
`at 5000 Feet
`
`Fan Configuration A (SM = 21 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (510 F)
`
`A 1 8 = 1190
`at 10,000
`
`Fan Configuration A (SM = 21 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (320 F)
`
`A 1 8 = 1190
`at 15,000
`
`Fan Configuration A (SM = 21 Percent,
`Square Inches) Estimated Performance
`Feet on Tropical Day (120 F)
`
`A1 8 = 1190
`at 20,000
`
`Estimated Fan Performance Map for Configuration
`A With Fixed Fan Exhaust Nozzle (A 1 8 = 1190 Square
`Inches) Showing Operating Lines
`
`Estimated Supercharger Performance Map With
`Fixed Fan Exhaust Nozzle (A 1 8 = 1190 Square Inches)
`Showing Operating Lines
`
`Gas Ge'nerator Performance Map Showing Operating
`Line
`
`Estimated Fan Performance Map for Configuration
`B With Fixed Fan Exhaust Nozzle (A 1 8 = 1190 Square
`Inches) Showing Operating Lines
`
`Fan Configurations A and B (A 1 8 = 1190 Square Inches)
`Estimated Performance at Sea Level on Standard Day
`
`Fan Configuration B (SM = 14 Percent, A1 8 = 1190
`Square Inches) Estimated Performance at Sea Level
`on Standard Day
`
`70
`
`71
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`72
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`73
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`75
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`LIST OF TABLES
`
`Basic Laminar-Type Profile, 10 Percent
`Thickness
`
`Fan Rotor Blading Design Data, Conical Sections
`(Z = 32 Blades)
`
`Fan Stator Blading Design Data, Conical Sections
`
`Integral Lift Engine Sideline Perceived Noise Level
`at 500 Feet
`
`Integral Lift Engine Polar Perceived Noise Level
`at 500 Feet
`
`Polar Noise Field, Untreated Bypass Duct
`
`Sideline Noise Field, Untreated Bypass Duct
`
`Table
`
`I
`
`II
`
`III
`
`IV
`
`V
`
`VI
`
`VII
`
`VIII
`
`Attenuation of the Bypass Duct Treatment
`
`IX
`
`X
`
`XI
`
`XII
`
`Polar Noise Field, Treated Bypass Duct
`
`Sideline Noise Field, Treated Bypass Duct
`
`Turbofan Engine Design Cycle Data
`
`Turbofan Engine Design Efficiency and Loss
`Assumptions
`
`XIII
`
`Turbofan Engine Station Cycle Data
`
`Page
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`15
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`18
`
`18
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`52
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`52
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`54
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`55
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`57
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`61
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`62
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`SUMMARY
`
`Preliminary studies and final design studies of an integral lift fan en-
`gine ALF- 504 were conducted using in all cases the Lycoming 502 fan engine
`core and MQT power turbine. The fan full-speed design pressure ratio
`is i. 25:1 at sea level standard conditions as specified by NASA. Pre-
`liminary studies were conducted to determine the relative engine perfor-
`rnance when supercharging with the fan alone and with the addition of
`two core inlet duct supercharging stages. The results of the prelim-
`inary studies show an increase in maximum sea level static thrust of
`10 percent for the additional supercharged engine and an increase of
`approximately 10 percent in velocity leaving the power turbine. Consid-
`erations of the small thrust increase and added noise contribution of the
`hot exhaust jet* resulted in a decision by NASA to base the final design
`studies on the engine without additional supercharging. Another impor-
`tant decision as a result of the preliminary studies resulted in an
`increase of fan tip speed from 985 ft/sec to 1100 ft/sec. This tip speed
`increase reduced the rotor blade hub overturn (turning past axial in the
`relative system) from 25 degrees to 10 degrees in the final design, and
`gives reasonable assurance of flow stability into the core compressor.
`The final design studies includes detail analysis of fan aerodynamics,
`fan and reduction gear mechanical design, fan dynamic analysis, engine
`noise analysis, engine performance, and weight analysis.
`
`Aerodynamic design of the fan includes consideration of the splitter
`shape and location as this affects local streamline curvature and static
`pressure gradient. A solution is realized when the splitter stagnation
`streamline static pressure shows the same value for the duct and core
`flows. Semi-empirical analytical technique making use of test data
`from the Lycoming 502 fan has been included in the axisymmetric flow
`solution to account for the effects of part-span shroud pumping, wake
`profile and higher losses behind the rotor shroud.
`
`Important design parameters of the final fan aerodynamic sea level
`standard design point condition for the ALF-504 fan engine are:
`
`1. Total fan flow 421. 1 lb/sec
`
`2. Bypass ratio 12. 5
`
`3. Fan pressure ratio 1. 25:1
`
`A diffuser after the power turbine was not used in either case.
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`4. Fan design efficiency 88 percent polytropic
`
`5. Fan tip speed 1100 ft/sec
`
`6. Fan tip diameter 48. 0 inches
`
`7. Fan inlet hub-tip ratio 0. 392
`
`8. Rotor hub overturn 10 degrees
`
`9. Rotor inlet relative tip Mach number 1. 15
`
`10. Absolute Mach number at inlet of core stator 0. 74
`
`Stress and dynamic considerations of the fan rotor blade and disc
`based on use of titanium material and a part-span shroud located at an
`18-inch radius show the steady stresses to be low compared with the
`material yield stress. The tangential stress at the disc bore is 43 ksi
`compared with 132 ksi yield, and the hub blade centrifugal stress is
`26. 5 ksi. First bending frequency is free from first-, second-, and
`third-order excitation above 80 percent speed, and therefore should
`provide adequate inlet distortion margin from a mechanical standpoint.
`Blade flutter analysis shows adequate design margin in torsion and
`bending when compared with NASA criteria.
`
`Fan noise analysis without inlet treatment based on a modified Smith
`and House method shows a maximum 500 feet sideline noise of 98. 6
`PNdB at 70 degrees from the inlet for this component.
`
`The total 500 feet sideline maximum engine noise without any acous-
`tical treatment is 104. 5 PNdB at 100 degrees from the inlet. Use of two
`sound-attenuating rings in the bypass duct in addition to wall treatment of
`this duct reduced the 500 feet sideline engine noise at 100 degrees from
`the inlet to 97.0 PNdB. This 7. 5 PNdB engine noise reduction requires
`an increase of 3. 8 inches in bypass duct diameter, produces a minimum
`of 2. 5 percent loss in sea level takeoff thrust, and adds 58 pounds in
`weight to the engine because of the attenuating rings alone.
`
`Sea level maximum static thrust is 8370* pounds with a 0. 302 lb/hr per
`lb specific fuel consumption. Use of a two-position variable duct exhaust
`nozzle gives 16 percent higher net thrust at 0. 4 flight Mach number sea
`level and 15 percent improved specific fuel consumption
`
`Does not include added pressure loss of sound-attenuating rings in by-
`pass duct.
`
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`The nozzle goes to the closed position (12 percent closed) when a
`flight Mach number of 0.4 is reached, to rematch the fan at the sea
`level takeoff point. Engine acceleration time from a steady-state con-
`dition of 80 percent maximum thrust to 93. 2 percent (66 percent of 80
`to 100 percent) is estimated to be 2.5 seconds in the sea level flight
`Mach number range of 0 to 0. 4.
`
`The dry engine weight without starter but including bypass duct sound-
`attenuating rings and wall treatment is 1419 pounds.
`
`INTRODUCTION
`
`The relative simplicity and the advantages and disadvantages of
`VTOL and STOL aircraft have been studied and debated in this country
`and in Europe for more than a decade. Indeed, many system concepts
`have evolved and a large variety of powerplant types and aircraft sys-
`tems have been built and tested. Since these aircraft have been devel-
`oped largely for military application, the noise by-product was of little
`consequence. However with application of VTOL and STOL to commer-
`cial aircraft, noise generated during takeoff and landing has become of
`paramount importance, with maximum effort and considerable funding
`being expended both to quiet existing engines and to better understand the
`fundamentals of the interplay between basic aerodynamic design of fans,
`compressors and turbines, and noise generation.
`
`As an effort to demonstrate the relative quiet of STOL aircraft em-
`ploying internal or external wing augmentation and aircraft of the same
`seating capacity with VTOL capability, NASA (Lewis) has sponsored a
`design and study program for lift cruise fan engines with Lycoming based
`on the 502 fan core. These engines would provide power for a demonstra-
`tor aircraft. It is intended that the outcome of this demonstrator program
`would lead to the design, development, qualification, and procurement of
`larger commercial VTOL aircraft and more powerful engines which
`would be phased into commercial use in such size, range, and cruise
`speed as to be economically viable.
`
`The Lycoming 502 fan engine core was chosen by NASA for the dem-
`onstrator program because of its relatively short length and low weight,
`the many development and flying hours of the basic core, and the current-
`ly intense fan program sponsored by the Air Force for the AX close-
`support aircraft.
`
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`The design studies reported herein have all been based on the 502
`fan core, and demonstrate the relative ease of converting the 502 fan
`core to a high bypass ratio (BR = 12. 5) low pressure ratio (Pr = 1. 25:1)
`fan for lift cruise application. The final design studies include detail
`analysis of fan aerodynamics, fan stress and dynamics, reduction gear
`design and stress analysis, fan and engine noise analysis, engine per-
`formance evaluation, and weight estimation.
`
`PRELIMINARY STUDIES OF SUPERCHARGED AND NONSUPERCHARGED
`INTEGRAL LIFT ENGINE AND AIR PUMP
`
`General
`
`Prior to the final design phase of the integral lift engine, parallel
`studies were made to determine the relative size, weight, performance,
`and fan aerodynamics for the nonsupercharged engine and a design
`having two supercharging stages following the fan. Reduction gearing
`design studies for each engine were also conducted to show any signif-
`icant differences involved. The results of these studies were presented
`to NASA in February, and the decision was made to continue the final
`design effort on the nonsupercharged engine. Its greater simplicity,
`lower development cost, and reduced hot jet velocity after the power
`turbine* (which may be a major noise contributor) more than offset
`the 10 percent lower takeoff thrust compared with the supercharged
`engine.
`
`Engine Thermodynamic Performance
`
`In this phase of study, only design point sea level static maximum
`power performance was considered. The use of two additional super-
`charger stages produced a higher cycle pressure ratio and increased
`mass flow in the core to increase the engine power. The following
`data summarize the performance results of the two engines studied:
`
`In each case the MQT T55-L-11 power turbine is used without exhaust
`diffuser or jet nozzle.
`
`*
`
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`Fan Alone
`
`Fan Plus Two
`Supercharging Stages
`
`Wa tot - lb/sec
`
`W core - lb/sec
`
`Bypass Ratio
`
`Prf
`
`Total Supercharging Pr
`
`Overall Pressure Ratio
`
`Ftot
`
`- lb
`
`SFC - lb/hr-lb
`
`406*
`
`31. 2
`
`12.02
`
`1.25
`
`1. 25
`
`9.73
`
`8000
`
`0. 307
`
`449
`
`34.4
`
`12. 02
`
`1.25
`
`1. 5
`
`10.7
`
`8800
`
`0. 300
`
`Fan Aerodynamic Design
`
`Two fans were studied, and they reflect the requirements of the
`supercharged and nonsupercharged engines. These fans were based on
`tip speeds of 985 ft/sec, which was later increased to 1100 ft/sec with
`concurrence of NASA to reduce the hub "overturn" (turning past axial
`in the relative system) with the belief that the added noise contribution
`would be small and would be as much affected by the blade aerodynamic
`loading as the increased tip speed. A summary of the more important
`aerodynamic parameters follows:
`
`Prf
`
`U t - ft/sec
`
`N- RPM
`
`Dt - in.
`
`Fan Alone
`
`Fan Plus Two
`Supercharging Stages
`
`1.25
`
`985
`
`4770
`
`1.25
`
`985
`
`4550
`
`47. 244
`
`49. 606
`
`Hub/Tip Ratio, Inlet
`
`Wa tot - lb/sec
`
`Rotor Tip Rel Mach No.
`
`0. 385
`
`406.0
`
`0.80
`
`Rotor Hub Overturn - deg
`
`25
`
`Final design based on 421. 1 lb/sec
`
`0. 385
`
`449.0
`
`0. 80
`
`25
`
`5
`
`GE v. UTC
`IPR2016-00952
`GE-1027.016
`
`

`

`Reduction Gearing
`
`Reduction gearing design studies were made for the supercharged and
`nonsupercharged fan designs with appropriate reduction ratios. The
`study included designs utilizing a rotating planet and carrier with fixed
`ring gear as a possibility to reduce the gear envelope and weight and to
`allow positioning of the gear further downstream in the inlet housing to
`reduce the overall engine length. The results indicated the bearing loads
`of the rotating carrier to be excessively high for the package required
`for this application, and this approach was, therefore, abandoned.
`Gearing having fixed planets and a rotating ring gear (the type used in the
`Lycoming 502 engine) was used as the basis of the final analysis. Re-
`sulting design parameters for the appropriate gearing of the nonsuper-
`charged and supercharged fan engines are as follows:
`
`Fan Alone
`
`Fan Plus Two
`Supercharging Stages
`
`Reduction Ratio
`
`Output Torque-ft/lb
`
`Output Speed - RPM
`
`Number of Planets
`
`Face Width of Sun and
`Planets - in.
`
`3. 56
`
`9080
`
`4770
`
`5
`
`2. 85
`
`Air Pump
`
`3.73
`
`9520
`
`4550
`
`4
`
`3.12
`
`In this study major emphasis has been directed to the integral lift
`engine powerplant; therefore, only a cursory study has been conducted
`for the air pump design concept as agreed upon by NASA.
`
`The fan design was based on a pressure ratio of 3. 5:1 with the bypass
`ratio dependent upon the static pressure desired after the power turbine
`and associated exhaust nozzle configuration. The fan design could be
`effectively accomplished from an aerodynamic standpoint with either
`three or four stages.
`
`The optimum number of stages would depend upon design trade-off
`studies showing the effect of relative tip speed (higher for the 3 stage
`design) and number of stages on fan noise, gear reduction ratio size and
`weight, and fan weight.
`
`6
`
`GE v. UTC
`IPR2016-00952
`GE-1027.017
`
`

`

`The fan supercharges the 502 core and, therefore, gives a significant
`temperature, pressure, and flow increase into the core of the fan engine.
`The study revealed that because of the temperature increase at the core
`inlet, the referred speed of the core compressor is 83. 5 percent of
`design value. At this speed the core compressor requires operation with
`the bleed port open to prevent surge.
`
`Two solutions to this problem are available:
`
`1.
`
`Increase of the core rotor speed above the present 19, 260 rpm
`value, which requires reset of the turbine nozzles and a review of
`the modifications required by the increased stresses.
`
`2. Removal of one or more of the front stages of the core compressor
`to lower the overall pressure ratio and allow acceptable operation
`without requiring the bleed port to be open.
`
`Both solutions are practical but require more extensive modifications
`than compatible with the minimum modification approach of this study.
`Accordingly, further analysis of the air pump design was not conducted.
`
`FAN AERODYNAMIC DESIGN
`
`Design Point Conditions and Data
`
`A fan with 1.25:1 SLS (sea level static) total pressure ratio has been
`selected for the present study. This design pressure ratio corresponds
`to maximum SLS power setting of the 502 core at maximum rating tur-
`bine inlet temperature. The corresponding fan mass flow rate and by-
`pass ratio follow from the power delivered by the supercharged engine
`core, the fan efficiency, and the condition of ambient static pressure
`level at exit of the power turbine (no turbine exit diffuser). An 88 per-
`cent polytropic fan efficiency is assumed as a realistic target value for
`a low hub tip ratio transonic fan stage with moderate pressure ratio and
`low supersonic relative tip Mach number. Initially, the fan design tip
`speed was set at 985 ft/sec. This speed, however, resulted in a specific
`work input coefficient Ah/U2 considerably larger than 1 at the hub sec-
`tion and in a positive slope of the r - cp operating characteristics for
`the entire supercharging section of the fan rotor blade. The tip speed
`was subsequently increased to 1100 ft/sec in order to minimize the risk
`of core flow instability.
`
`7
`
`GE v. UTC
`IPR2016-00952
`GE-1027.018
`
`

`

`The final fan SLS design conditions are summarized below:
`
`Total Stage Pressure Ratio P/P = 1.25:1
`
`Total Mass Flow Rate Watot = 421. 1 lb/sec
`
`Cor Engine Mass Flow Rate Wa = 31.2 lb/sec
`
`Bypass Ratio BR
`
`= 12. 5:1
`
`Tip Speed U t = 1100 ft/sec
`Target Total Polytropic Efficiency lPtot = 0. 88
`
`Aerothermodynamic Design Concept
`
`The main aerothermodynamic problem consists of designing the core
`supercharging fan section and matching its flow path with the existing
`50Z fan-core transition duct without increasing the engine length. This
`latter condition limits the rotor exit hub diameter and the hub work
`input capacity. With the tip speed increased to 1100 ft/sec, the hub
`1; - cp characteristics still has a slight positive slope with a specific
`work input coefficient (Ah/U 2 )hub = 1.22.
`
`Overturning gradually disappears over the channel height and the flow
`conditions at the supercharger upper section are conventional with
`(h/UZ)tip = 0. 8. The aerodynamic conditions in the supercharger flow
`region are influenced by the meridional curvature of the core flow path
`and by the location, orientation, and thickness of the core-duct flow
`splitter. The meridional curvature of the core flow path raises the
`meridional velocity level near the inner wall. In order to minimize the
`core stator inlet Mach number, both the tangential and the meridional
`velocity components at rotor exit should be minimized in the inner wall
`region. Minimizing the tangential component requires an increase of
`the hub radius, which, however, results in an increase of the core
`channel curvature and the hub meridional velocity component. Similar-
`ily, the location, orientation, and thickness of the flow splitter deter-
`mine the annulus area at inlet of the core stator and the average stator
`inlet velocity. The overall curvature of the core flow channel and the
`stator hub inlet velocity are also affected by these variables. The
`channel and splitter configurations thus markedly influence the flow
`conditions in the critical supercharger fan section, and their interaction
`must be studied in arder to optimize the aerodynamic fan design. The
`
`8
`
`GE v. UTC
`IPR2016-00952
`GE-1027.019
`
`

`

`core flow path matches the 502 core at the upstream flange of the fan
`support casing. Therefore, the present design allows use of the 502
`core cast inlet housing and bearing support structure.
`
`The flow conditions are calculated with Lycoming's IBM Program
`R1 36, which solves the complete radial equilibrium equation for the
`axisymmetric flow field of a turbomachine with a bypass flow splitter.
`The splitter streamline separating the core and the fan duct flows is
`subject to two conditions at the splitter stagnation point, perpendicular-
`ity to the splitter nose and vanishing of the streamline curvature.
`Both flows domain upstream and downstream of the splitter nose are
`treated simultaneously in the iterative computation procedure, which
`usually converges within 50 iterations.
`
`In addition to the fan design data specified above, the following as-
`sumptions are made for the calculation of the flow conditions:
`
`The compression process through the fan rotor blading is charac-
`terized by a polytropic efficiency that varies along the blade span
`according to Figure 1. The stator losses are defined by a total
`pressure loss coefficient w = 0. 05, which is constant over the radius
`except in the vicinity of the inner and outer walls, where the stator
`losses are increased to account for additional wall boundary layer
`and secondary flow effects. The rotor work input is determined in
`conjunction with the assumed rotor efficiency and the stator losses
`in such a way as to produce a constant overall fan total pressure
`ratio P/P = 1.25:1, except at the wall regions, where the additional
`stator losses are superimposed and result in a slight total pressure
`deficit. It will be seen in Figure 1 that the fan rotor polytropic
`efficiency has been decreased on the part-span-shroud streamline
`in order to take into account the effect of the shroud wake. The
`semi-empirical procedure used to simulate that effect is based on
`published test data (I) and in-house tests by Lycoming and is described
`in Appendix I. Finally global flow blockage effects of 1 percent at
`fan rotor exit and 2 percent at both fan duct and core stator exit
`stations have been assumed.
`
`The fan hub flow conditions are critical because of the low hub ro-
`tational speed resulting from the low hub tip ratio and tip speed limit-
`ation. However, for a given tip speed and a given rotor exit hub radius,
`the rotor hub speed can be increased by decreasing the fan annulus area,
`i. e., by increasing the fan specific flow capacity. The fan inlet Mach
`
`9
`
`GE v. UTC
`IPR2016-00952
`GE-1027.020
`
`

`

`1.0
`
`0.9
`
`0.8
`
`0.7
`
`0.6
`
`Uz U L
`
`L
`LU
`
`Ua
`
`-
`
`0 -
`
`j0a
`
`-
`
`0I
`
`0,
`
`--
`
`0 5 10
`
`20
`
`40
`
`50
`
`70
`
`80
`
`90
`
`HUB
`
`MASS FLOW PERCENT
`
`100
`
`TIP
`
`Figure 1. Rotor Polytropic Efficiency.
`
`10
`
`GE v. UTC
`IPR2016-00952
`GE-1027.021
`
`

`

`number consequently has been set at the highest level compatible with
`a favorable overspeed flow margin potential, namely 0. 55 as an average
`value. For aerothermodynamic and noise reasons, the fan is designed
`without inlet guide vanes. Moreover, the core-duct flow splitter is
`located downstream of the rotor blading in order to allow for increased
`rotor-stator spacing in the duct section without increasing the length of
`the fan-core transition duct. In addition the fan duct stator tip is leaned
`in the downstream direction to further minimize wake noise. Figure 2
`shows the fan meridional flow path with the stations used in the IBM
`program R136 calculation and the streamline pattern.
`
`The results of the aerodynamic design optimization are illustrated
`by the velocity triangles shown in Figure 3. The first three triangles
`describe the flow conditions in the supercharger section. With the tip
`speed increased from 985 to 1, 100 ft/sec, both the rotor flow overturn-
`ing and the Mach number at entrance of the stator have been reduced to
`favorable levels ( Zhub = 10.4 degrees and MVZhub = 0.715, as compared
`with 25 degrees and 0.75 to 0.8 with Utip = 985 ft/sec). As a result of
`the slight relative flow overturning, the highest rotor flow deceleration
`rate does not occur at the hub section but in the splitter region. The
`stator hub section, however, is subjected to the highest deceleration
`rate, which is somewhat above usual practice for stator design. The
`corresponding velocity ratio (V5/V4)hub = 0.712 is accordingly slightly
`lower than desirable for a shrouded design.
`(See "Aerodynamic Blading
`Designs ")
`
`The last three triangles describe the flow conditions through the upper
`fan section, which are typical of conventional transonic design practice.
`
`IBM Program R136 output section given in Appendix II contains a
`complete set of flow data that substantiate the basic aerodynamic design
`concept.
`
`Aerodynamic Blading Design
`
`Rotor Blading. - The main blading design problem consists of select-
`in a favorable compromise between the conflicting aerodynamic, weight,
`and acoustic design requirements. For the tip section, a profile with
`thin leading edge and maximum thickness located at 50 to 60 percent
`chord station is required. A double-circular-arc profile would be ad-
`equate.
`
`11
`
`GE v. UTC
`IPR2016-00952
`GE-1027.022
`
`

`

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`
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`
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`
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`
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`
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`
`Figure2.FanMeridionalFlowPath.
`
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`