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NI\S/\ NASA CR - 165185 L YC NO. 80-27
NASA-CR-165185 19810023604
FINAL REPORT ~_------
DESIGN AND EVALUATION OF AN INTEGRATED
QUIET CLEAN GENERAL AVIATION TURBOFAN (QCGAT)
ENGINE AND AIRCRAFT PROPULSION SYSTEM
by
Jon German, Philip Fogel and Craig Wilson
AVCO LYCOMING DIVIS~"~N~ f.r~V :~npy 550 South Main Street
Stratford, Connecticut 0649'fEB 24 1981
Prepared for
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
NASA Lewis Research Center Cleveland, Ohio 44135
Contract NAS 3-20584
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NI\S/\ FINAL REPORT
NASA CR - 165185 L YC NO. 80-27
DESIGN AND EVALUATION OF AN INTEGRATED
QUIET CLEAN GENERAL AVIATION TURBOFAN (QCGAT)
ENGINE AND AIRCRAFT PROPULSION SYSTEM
by
Jon German, Philip Fogel and Craig Wilson
AVCO LYCOMING DIVISION 550 South Main Street
Stratford, Connecticut 06497
Prepared for
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
NASA Lewis Research Center Cleveland, Ohio 44135
Contract NAS 3-20584
1 Report No I 2 Government Accession No 3 Recipients Catalog No
4 Title and Subtitle 5 Report Date
DESIGN AND EVALUATION OF AN INTEGRATED Aprll 19BO
QUIET, CLEAN GENERAL AVIATION TURBOFAN (QCGAT) 6 Performing Organization Code
ENGINE AND AIRCRAFT PROPULSION SYSTEM
7 Author(s} 8 Performing Organization Report No J. German, P. Fogel, C. Wllson LYC 80-27
10 Work Unit No 9 Performing Organization Name and Address
Avco Lycoml.ng DlvlSlon 550 South Mam Street 11 Contract or Grant No
Stratford, COnnectlcut 06497 NAS 3-20584
13 Type 01 Report and Per,od Covered
12 Sponsorong Agency Name and Address NASA - LeW1S Research Center
Contractor pec. 76 - Apr.BO
2100 Brookpark Road 14 Sponsorong Agency Code
Cleveland, Ohio 44135
15 Supplementary Notes
Program Manager, O.K. S,evers NASA - LRC, Cleveland, OhlO
16 Abstract
Avco Lycomlng partlclpated ln the NASA QUlet, Clean, General AViatlon Turbofan (QCGAT) englne program by deSignlng a small turbofan englIle ln the 7000 N(1600 lbf ) thrust class. The engl.ne and nacelle system deslgn was to demonstrate the apphcablhty of large turbofan englIle technology to small turbofans sUltable for the general aVlatlon market.
The Lycomlng deslgn was based on the LTS-IOI engl.ne famlly for the core englIle. A high bypass fan deSlgn (BPR,.9. 4) was lncorporated to provlde reduced fuel consumptlon for the deslgn mlSSlon.
All acoustlC and pollutant emlSSlons goals were achieved.
A d19cusSlon of the prehmlnary deSign of a buslIless Jet sUltable for the developed propulslon system 19 also ,ncluded ln this report.
Large engl.ne technology can be sucessfully apphed to small turbofans, and nOlse or pollutant levels need not be constralnts for the des>gn of future small general aVlatlon turbofan englnes.
17 Key Words (Sug~ted by Author(s}} 18 Distribution Statement - General AVlatlon - EmlSSlons
~ DlStnbutlo ded - Turbofan - Hlgh Bypas. Per-
Source :~ - QCGAT formance - NOlSe Reductlon ? .trla! Appu-..-, Center
19 Security Oasslf (of thiS report) 120 Securoty Class,l (01 thos page} fl No 01 Page< 122
Pnce-
UNCLASSIFIED UNCLASSIFIED 219
FOREWORD
The work described hereln was conducted by the LycoITling Gas Turbme DlVislOn of Avco CorporatlOn with support from Beech Alrcraft Corporation In the prehITllnary aircraft deslgn, Lockheed Aircraft Corporatlon In the acoustlcal duct treatment evaluatlOn, and the Aerostructures DlvislOn of Avco In fabrlcahon of the englne nacelle.
The support and technlcal experhse of the se subcontractor s
m conJunctlon wlth the varlOUS Lycomlng englneerlng department were lnstrumentalln attalnlng the contract goals addressed In thls docuITlent.
The author s wish to acknowledge Mr. Kelth Slever 8,NASA QCGAT PrograITl Manager, and hlS staff for their gUldance and
support.
iii
1.0
2.0
3.0
4.0
TABLE OF CONTENTS
FOREWORD ....................................... LLST OF ILLUSTRATION ........................... LLST OF TABLES ................................. SUMMARY
INTRODUC TION ................................... AIRCRAFT AND ENGlli"E DESIGN OPTIMIZATION •••••
3. 1 3.2
3.3
3.4
Aircraft Design Approach ••••••••••••••••••• Eng:i.rle Design •••..•••••••••••••••.••..•••• 3.2.1 Design Considerations ••••••••••••••• 3.2.2 Engine Hardware Description ••••••••• Nacelle Design Approach •••••••••••••••••••• 3.3. 1 3.3.2 3.3.3 Mixer
Aerodynamic De sign ••••••••••••••••• Acoustic Considerations ••••••••••••• Structural Design •••••••••••••••••••
Nozzle Design •••••••••••••••••••••••
SUBSYSTEM TEST RESULTS ••••••••••••••••••••••••
4.1
4.2
4.3
4.4
4.5
4.6
Fan Blade Analysis •••••••••••••••••••••••• 4. 1. 1 Test Program •••••••••••••••••••••• 4. 1.2 Fan Blade Test Results •••••••••••••• Low-Pressure Turbine Blade Analysis •••••••• 4.2.1 Power Turbine Blade Results ••••••••• Ring Gear Frequency Analysis • • •••••••••••• 4.3.1 Ring Gear Test Results •••••••••••••• Flow Divider and Fuel Manifold System Analysis ................................. . Fan Component Test Rig •••••••••••••••••••• 4.5. 1 Fan Component Test Results •••••••••• 4.5.2 Distorted Inlet Test Results ••••••••••• Combustor Module ••••••••••••••••••••••••• 4.6.1 4.6.2 4.6.3
Test Rig and Facilities •••••••••••••• Test Sequence ......••.•........... Test Results •...•..................
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2
3 15 15 17 48 51 63 68 79
84
84 84 84 84 90 90 90
97 97
102 108 114 116 116 119
5.0
6.0
TABLE OF CONTENTS
ENGrnE/NACELLE SYSTEMS PERFORMANCE ......... 5. 1
5.2
5.3
Overall Performance •••••••••••••••••••••••• 5. 1.1 Component Performance •••••••••••••• 5. 1.2 Full Engine Tests •••••••••••••••••••• Emission Test Results ••••••••••••••••••••••• 5.2.1 5.2.2
Design and Emissions Projections •••••• Emis sions Sampling ••••••••••••••••••
5. 2. 3 Emis sions Summary •••••••••••••••••• Acoustical Performance ••••••••••••••••••••• 5.3.1 5.3.2 5.3.3 5.3.4 5.3.5 5.3.6 5.3. 7
Background .......•..••••.........•• Engine Design and Noise Prediction •••• Sound Treatment Design •••••••••••••• Win.g Shieldin.g •..••.................. Acoustic Test Phase •••••••••••••••••• Data Arlalysis •......•••.••........... Conclusion ..........................
SUMMARY OF RESULTS ............................ 6. 1 6.2
6.3
Objectives Test Results
................................. ...............................
6.2.1
6.2.2 6,2.3
Overall Engine Performance Demons trated .......• e •••••••••••••••
Em.is sians •.•.....•.•....•.........•. ACoustlCS, ...•••••••••••••••••••••••••
ConcludIng ReITlar ks .•..•••.......•.•...•.....
Page 121
121 121 126 146 146 152 152 155 155 159 165 173 173 180 204
207
207 207
208 208 209 209
REFERENCES. . . . . . . . . • • . . . . . . . • . . . . . . . . . . . . . . • . . . .. 210
DISTRIBUTIONLIST •••••••••.••••••••••••••••••••••• 212
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Figure
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18
LIST OF ILLUSTRATIONS
New General Aviation Aircraft for the 1980 Decade . • • • • • • • • . . . . . . . • • . . • . . • . • • • . . • • • • . . . . . . . . 5
QCGA T Aircraft Design ......................... Design Areas Selected for Composite Structures .... Wing Selection .................................. Best Range Versus Payload ....................... Aircraft Stall Speed and Field Capacity •••••••••••••
Cycle Optimization for Specific Fuel Consumption
Selected Fan Pressure Ratio ...................... Engine Configuration ............................. EngiIle Module s ••••••••••••••••••••••••....•••.•.
Gas Generator Module ••••••••••••••••••••••••••••
Circumferentially Stirred Combustor Flow Fa ttern .....................••••..............•
Airblast Fuel Injector System ..................... Turbine Arrangement ............................ Cooling of First Stage Rotor Blades and Turbine Nozzle ••.•..•..............••......••........•.
Power Turbine Module ........................... Turbine Section Meridional Flow Path ............. Interturbine Duct Inner and Outer Wall Velocity Distribution at Sea Level Takeoff •••••••••••••••••
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29
31
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34
36
37
LIST OF ILLUSTRATIONS (Cont. )
Figure
19 Accessory Drive Module Installation ••••••••••••••• 39
20 QCGAT Accessory Gearbox ••••••••••••••••••••••• 40
21 Power Control and Fuel SysteITl ScheITlatic •••••••••• 42
22 Fuel Control Assembly ••••••••••••••••••••••••••• 44
23 Dual COITlpressor Flow Fence ScheITlatic ........... 45
24 Flow Fence Actuator .........•..••..•............ 46
25 Fan Overspeed Trip System ••••••••••••••••••••••• 47
26 Cross Section of Engine with Cutaway •••••••••••••• 50
27 QCGAT Nacelle External Geometry •••••••••••••••• 52
28 QCGAT Nacelle Front Cowl Geometry ............. 53
29 QCGAT Nacelle MaxiITlum Drag Divergence Flight Spe ed ••••.................••.•.•..•........... 54
30 QCGA T Upper Cowl Lip Spillage Margin •••••••••••• 55
31 QCGA T Inlet Flow Incidenc e Limits •••••••••••••••• 57
32 Inlet Pressure Recovery ......................... 58
33 QCGA T Nacelle Duct Areas and Mean Mach ....................................... 59
34 QCGAT Flight Nacelle with Sound Treated Panels IDS tailed ••......................•.•............ 60
35 QCGAT Lockheed Fan Exhaust Noise Treatment Design ........................................ 61
36 Sample from Lockheed's Materials Handbook for Metal ..•..•.•.•...... 0 •••••••••••••••••••••••• 67
Vlll
LIST OF ILLUSTRATIONS (Cont. )
Figure
37 Nacelle Sectional View ........................... 69
38 QCGA T Nacelle Components ...................... 70
39 Nacelle Materials and Weights ••••••••••••••••••••• 72
40 Nacelle Fire Prevention/Compartment
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Ventila tion •••••••.••••••.••••••••••••••••••.••.• 73
Test Nacelle Assembly •••••••••••••••••••••••••••
QCGA T Flight Nacelle Simulation .................. Replaceable Inlet Lips ••••••••••••••••••••••••••••
Bellmouth and Flight Nacelle Inlets ................ QCGAT Exhaust Nozzle System ••••••••••••••••••••
Core Engine Mixer Nozzle GeoITletric Definition ..... Lycoming QCGA T Mixer Nozzle Design
Mixer Nozzle Installation •••••••••••••••••••••••••
General Test Setup Holographic Analysis
Fan Blade Root Fixture •••••••••••••••••••••••••••
Fan Blade Natural Frequencies and Associated Mode Shapes - First and Second Bending Mode, First Torsional Mode •.....•..•...•...•.•..•..........
Fan Blade Natural Frequencies and As sociated Mode Shapes - Second Torsional Mode, Third and Fourth Bending Mode ..•.......••..••••••••••...•••.•..
Fan Blade Excitation Diagram ••••• 0 ••••••••••••••
Power Turbine Segment Mounted to Shaker ••••••••••
LX
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LIST OF ILLUSTRATIONS (Cont.)
Figure Page
55 Power Turbine Natural Frequencies and Associated Mode Shapes . . . • . • . . . . . . . . . . . . • . • . . . • . • . . . . . . . . . . 92
56 Power Turbine Blade Excitation Diagram at Ambient
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59
60
61
62
Tem.peratur e .•.•.............•.•.•••.•........•. 93
Power Turbine Blade Excitation Diagram at
1250°F ....................................... .
Ring Gear Clamping Arrangement and Natural Frequencies and Associated Mode Shapes •••••••••••
Ring Gear Natural Frequencies and Associated Mode Shapes .•••••••••.•........•••••••••••......••..
QCGA T Fan Module Test Rig ••••••••••••••••••••••
Fan Module Instrumentation Planes - Nomenclature ••
Aerodynamic Performance Instrumentation •••••••••
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100
63 Fan Module Inlet Instrumentation Planes and Distortion Screen Position •••••••••••••••••••••••• 101
64 QCGAT Fan Module - Testway Installation •••••••••• 103
65 QCGAT Fan Module Baseline Test - Bypass Overall Perform..anc e Map ..............•................ 104
66 QCGAT Fan Module Baseline Test - Supercharger Overall Performance Map. • • • • • •• • •• • • • • • • • • • • • • • • 105
67 Bypas s Exit Plane Profile Data •••••••••••••••••••• 106
68 Supercharger Exit Plane P:.:ofi1e Data •••••••••••••• 107
69 Area Blockage (15 Percent) Inlet Distortion Plate with Resultant Total Pressure Contours •••••••••••• 110
70 Distortion Index as a Function of Total Referred Airflow ........................•............... 112
x
LIST OF ILLUSTRATIONS (Cont.)
Figure
71 Bypass Overall Performance with Inlet Distortion 113
72 Comparison of the LTS 101 Combustor with the
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QCGA T Combus tor ••••••••••••••••••••••••••••••• 115
Circurnferentially Stirred Combustor Flow Pattern •••
QCGAT Annular Combustor Test Rig and Test Cell IDS talla tio n .•. • . • . . . . . . . . . . . • • • • • . . • . • . . . . . . . . . .
NO Versus Combustor Inlet Temperature •••••••••• x
Pr edic ted Influenc e of the Mixer •••••••••••••••••••
QCGA T Referee Configuration - Inlet ••••••••••••••
QCGA T Referee Configuration - Exit •••••••••••••••
QCGAT Test Nacelle .•...•.••.••••.•..••..•......
Replaceable Inlet Lips .......................... .
Details of the Test Nacelle
Low Spool Speed Versus Net Thrust ............... High Spool Speed Versus Net Thrust
Thrust Specific Fuel Consumption Versus Net Thrus t .•...•................••.••...••.•......
Low Spool Speed Versus High Spool Speed •••••••••••
Fan Pressure Ratio Versus Fan Bypass Airflow •••••
Total Engine Airflow Versus Low Spool Speed
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88 Low Spool Speed Versus Net Thrust - Nacelle fustalled ..•................. $ •• • • • •• • •• • • • • • • • • • 140
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Figure
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101
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103
104
105
106
LIST OF ILLUSTRA TrONS (Cont.)
High Spool Speed Versus Net Thrust - Nacelle IDS tailed ............................•...........
Thrust Specific Fuel Consumption Versus Net Thrust -Nac elle IDS tailed •...............................
Low Spool Speed Versus High Spool Speed - Nacelle IDS taIled ......•....................•..........•.
Fan Pressure Ratio Versus Bypass Airflow - Nacelle fus taIled •••.•.••.•...........••••..•........•..
Total Engine Airflow Versus Low Spool Speed -Nac el1e Ir1stalled ................•..•...........
QCGA T Combustor .............................. NO Versus Combustor Inlet Temperature
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Engine Emis sions Sampling Test Probes ••••••••••••
Cross Section of the Engine with Cutaway
QCGAT Noise Goals ............................. Noise Prediction Procedures ..................... Fan Noise Components ........................... Core Noise Model ............................... Sound Treatment Locations ••••••••••••••••••••••••
Sound Treatment Design .......................... Inlet Attenuation ................................ Exhaust Attenuation .............................. Sound Treatment Panels ..........................
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172
LIST OF ILLUSTRATIONS (Cont.)
Figure
107 Wing Shielding 174
108 Microphone Arrangement ......................... 176
109 Test Site ....................................... 177
110 Ins trumentation Setup for the Acquisition of the Acoustic Data ......................•............ 178
111 Barrier ........................................ 179
112 Probe Locations ................................. 181
113 Reference System · ............................. . 182
114 Measurements with Predicted Fan ................. 184
115 Measurements with Jet Added ..................... 185
116 Core Added to Measured •••••••••••••••••••••••••• 186
117 High Power Setting •••• 187
118 Core Comparison ................................ 188
119 Revised Core Comparison ........................ 190
120 Core Noise Dominance ........................... 191
121 Fan Directivity Plots ............................ 192
122 High Fan Directivity ............................. 193
123 Final Comparison · ............................. . 194
124 Sound Treatment · ............................. . 196
125 Insertion Loss - Inlet ............................ 197
126 Insertion Loss - Discharge ....................... 198
xiii
LIST OF ILLUSTRATIONS (Cont.)
Figure
127 Jet Spectra .•..•.............•.•••.............• 199
128 Flyover Procedures •••••••••••••••••••••••••••••• 200
129 Component Contribution - Approach Noise •••••••••• 202
130 Component Contribution - Takeoff Noise •••••••••••• 203
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132
Noise Levels - QCGA T Goals
Noise Levels - FAA Standards
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206
Table
1
2
LIST OF TABLES
Benefits from Advanced Aircraft Design ............ Engine Operating Parameters ••••••••••••••••••••••
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3 Results of Design Study •••••••••••••••••••••••••••• 122
4 QCGA T Performance Goals •••••••••••••••••••••••• 123
5 Engine Configurations Tested....................... 132
6 QCGA T Performance •••••••••••••••••••••••••••••• 147
7 Initial Estimated QCGAT Emissions ................. 150
8 QCGAT Emissions Results ••••••••••••••••••••• 0... 154
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0 SUMMARY
Avco Lycoming participated in the NASA Quiet, Clean, General Aviation Turbofan (QCGAT) engine program to design a small turbofan in the 7000N (1600 1bf) thrust class. Lycoming's engine IS a highbypass ratio, twin-spool design with a geared- fan. The core engine is a growth derivative of the Lycoming LTS 101 engine series being used in many turboshaft and turboprop applications.
The Lycoming demonstrator engine program accomplished the followmg goals which were the primary objectives of the NASA/QCGAT program:
o Large engine technology can be successfully applied to general aviation-size engines to reduce noise and pollutant emissIons verlfylng that these ltems are not a constralnt to the general aVlatlon market growth.
o Pollutant-emission design goals met or exceeded stringent requirements which have since been abandoned by the EPA as part of the Clean Au Act of 1970.
o Noise emissions were substantially below (-25 dB) the Federal Air Regulations, Part 36 limits for aircraft takeoff, approach, and side line conditions.
o Fuel efficiency was superior when compared with presently available moderate-bypass ratio engines in the thrust class up to 1l,000N (2500 lbf).
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2.0 INTRODUCTION
The general aV1atlOn fleet has shown slgn1flcant growth character1stics In the past decade, 1t is expected that th1S trend w1ll contlllue throughout the 1980's. In partlcular, the Slze of the multi-engllle general aV1atlOn fleet 1S expected to increase by 20 percent. These general aVlation aircraft typ1cally use suburban alrports unprotected by commercIal buffer zones. Therefore, there IS a potentlal for general aviatlon traffIc to create more wIdespread, adver se, commun1ty reactIon to n01se and pollut1on than that experIenced by the commer cia! air carrIer s.
A program lllitlated by NASA to addres s th1S segment of general aV1ation a1rcraft 1ndustry cu1m1nated in Avco-Lycomlng 's rece1vlllg a contract to des1gn, develop, and test a QU1et Clean General AVlatlOn Turbofan (QCGAT) engine. The obJectlVe of thIS program was to demonstrate the applicabIlity of ava1lable, large turbofan engIne technology to small general aviatlon eng1ne s so as to obtain slgn1flcant reductIons In noise and pollutant em1SSlons, while reduc1ng or maintalll1ng fuel consumptlOn levels.
Th1S report presents the LyComlllg- QCGAT engllle development study, coverIng des1gn through engine test and 1ncludlllg acoust1cal, pollutant emisslOns, and performance re suits, Also included 1S the prel1minary desIgn of an approprIate general aV1ation a1rcraft that could use the propuls1on system developed.
3,0 AIRCRAFT AND ENGINE I'ESIGN OPTWIZATION
The Avco Lycoming QCGAT engllle was des1gned to demonstrate the latest noise-control technology. The hIgh-bypass fan was opt1mized with an eX1stlng proven core engine and advances 1n em1ssion reduchon. The resultant design of the ancraft/propulslOn system used
the technical expertlse of the follow1ng subcontractors'
o Beech A1rcraft Corporatlon ••.•• Aucraft Des1gn
o Avco Aerostructures •••• Nacelle Mechanical Des1gn
o Lockheed Aircraft Corporatlon •.•• Nacelle Acoustlc Treatment
These studle s cuhninated in the prehtninary de sIgn of a twin-engme, six-place, executIve aircraft having an advanced propulsion systetn to tneet the needs of the general aviation industry for a quiet-clean, efficient aircraft in the current and future decades.
3. 1 AIRCRAFT DESIGN APPROACH
Beech Aircraft Corporation was enlisted under separate con tract wIth Lycotnmg to design an aIrcraft based on the proposed QCGAT engine. To guide the aIrcraft systetn design, five pritnary objectives were established.
1. Practical, direct application of technology without significant scaling.
2. The aircraft tnust offer attractIve range, fuel econotny, and flight speed. A target of 2593 krn (1400 NM) was established. ThIS exceeds the range of tnost current stnall business aircraft thereby provIding a nonstop capability between opposite extremes of high-density traffic areas.
3. A cruise Mach nutnber of 0.6, chosen as an optitnal cotnprotnise between titne and fuel econotny, prOVIdes a 40-percent greater crUIse speed than a turboprop with a 30-percent itnprovetnent In fuel econotny over operahon at 0.8 Mach nutnber.
4. A balanced field length of 762 tneters (2500 ft) was desired because it would pertnit safe operation frotn over 70 percent of all public airports, includmg those with unItnproved runways.
5. Close attention to ecologIcal characterIstics of an aircraft systetn desIgn was desIred SInce they tnIght well becotne pritnary competitive paratneter s for the General Aviation industry In future decades.
The initial step in aIrcraft prelitninary design was the selection of appropriate SIze and de SIgn. The vast tnaJority of general aVIation aircraft operating frotn airfields located m suburban COtntnUllltites are m the size class below 5450 kg (12,000 Ib) gross
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4
weight. In the lower extremity of the gross weight spectrum, small private aircraft in the range below 1800 kg (4, 000 lb) are generally powered by single reciprocating engines. It is expected that market constraints for very low-cost aircraft in this class will dlctate continued usage of reclprocating engines for the foreseeable future. It, therefore, follows that the greatest public ecological benefits can be realized by introduction of a quiet clean aircraft system in the 1814 -5433 kg (4, 000 - 12, 000 lb) gross weight class. Figure 1 shows the projected market volume for varlOUS size s of general aviation aircraft.
As with the passenger car trend toward smaller, more sophisticated car s to perform the same function, it is expected that the 1980's wlll see a slmilar general aviation trend towards reduced aircraft weight and smaller engines size for the same mission. Because noise, emissions, and fuel consumption reduce with engine size, subsequent improvement in ecological characterlstics can be anticipated. Using technologie s such as turbofan propulsion, high aspect-ratio super critical wing, and lightweight composite structures, it is expected that a new class of small general aviation aircraft will emerge in the eighties. A target of 30 percent weight reductlon was considered achievable.
For selectIon of aircraft size, our target was the largest segment of general aviation alrcraft where cost of turbofan propulslOn does not preclude ltS introduction.
F1gure 1 presents a composite plot of a1rcraft gross weight versus both "The Number of New Aircraft to be Built" and "The Current Estimated Nomlnal Aircraft Cost". The number of aircraft 1S based on General Aviation Manufacturer s Association data. The expected trend toward bghter we1ght and higher cost for the same mis sion has not been reflected to ensure conservatlve engine sizing. The range of 3175 - 4536 kg (7, 000 to 10, 000 lb) gros s we1ght appeared attractive, with 3629 kg (8, 000 lb) selected as our goal.
W1th the defined aircraft goals and Lycoming's estimates for engine performance, Beech Air craft Cor poration conducted parametric studies to optimize the aircraft preliminary design.
3.1.1 Fmal Aircraft Design Optimization
The evolved aircraft (Figure 2) is a sleek, advanced des1gn, six-place aircraft with 3538 kg (7800 lb) maximum gross weight.
U1
NUMBER OF
AIRCRAFT
10,000
5,000
4,000 6,000 8,000 10,000
NOMINAL AIRCRAFT
COST, $ x 106
2
COS" 1
12,000 (Ib) I I I I
2,000 3,000 4,000 5,000 Kg
AIRCRAFT GROSS WEIGHT
Figure 1. New General Aviation Aircraft For The 1980 Decade.
0'
III I I
/ j
TWIN FUSELAGE-MOUNTED QCGAT ENGINES
I •
~ /-
~-- ,
~,
WIDE USE OF COMPOSITE STRUCTURES
Flgure 2. QCGAT Aircraft Design.
~~---.') , ~
FULL SPAN/ FOWLER FLAPS
0Z ~ "" WINGLETS
It offers a 2778 k1lnmeter (1500 nautlcal mile) range with cruise speed of 0.6 Mach Inumber, and will takeoff and land on a majority of general aVIation alrflelds. Advanced feature s include broad application of composite materials and a super critical wmg design with wmglets. Full-span fowler flaps have been introduced to Improve landing capability. Engines are fuselage-mounted wIth inlets over the wmg to pr ovide shie lding of fan noise by the wing sur face s.
The high bypass-ratio QCGAT engine plays an Important role m shaping the alrcraft design. It offers a dramatic reductlOn m specific fuel consumption compared with current pure jets and low-to-moderate bypass ratlo turbofans. Table 1 provides this comparison, reflecting a 22 per cent improvement in fuel economy.
This lower fuel consumption may be used In either of two ways or in combination:
1. It can substantially reduce aIrcraft gross weight for the same range. The reduced weight provides compound interest on the fuel economy. It also requires lower thrust which favors reduction of nOIse and emissions.
2. 1£ preferred, the lower fuel consumption can be translated mto longer range for the orIginal gross weight.
We chose to reduce gross weight and favor ecologIcal char acterlstics.
CompOSIte structures have been used extensively m the aIrcraft prehmmary deSIgn to further reduce gross weIght. Areas selected by Beech for the apphcatlOn of composite materIals are shown In FIgure 3. Kevlar graphIte composites were used for aircraft weIght estimates. Further potentlal for weIght savings eXIsts In the engIne nacelles. A conventional deSIgn was used to reflect the low-cost test nacelle. Critical load carrying members, such as the wing spar, are of conventlOnal alumInum constructlon.
Approximately 40 percent of the structure is of flber epoxy or honeycomb-bonded structure. The use of composite structure In aircraft deSIgn provides a decreaslllg rate of benefIt as the application of compOSItes becomes more WIdespread in the deSIgn. Initlal selectlOn of apphcatlons IS in noncrItical areas. As the stress in selected areas Increase s, the de SIgn safety factor also Increase s to compensate for uncertalllties resultlng from advance technological use of the composite applicatlOn.
7
00
TABLE 1. BENEFITS FROM ADVANCED AIRCRAFT DESIGN
STRUCTURE AND AIRCRAFT FUEL PROPULSION GROSS
CONFIGURATION WEIGHT, kg WEIGHT, kg WEIGHT, kg (Ibm) (Ibm) pbm)
CURRENT 1370 2520 4522 AIRCRAFT (3,016) (5,551) (9,960)
INTRODUCE 992 2392 4017 QCGA T ENGINE (2, 186) (5,268) (8,848)
USE COMPOSITES 932 1979 3541 AND (2, 053) (4,360) (7,800)
SUPER CRITICAL WING
i
SAVINGS 32% 23% 22%
I
~
aD f~@:?:!~~
~\"
, ,
ftfFl GRAPHITE REINFORCED I2J KEVLAR SKIN
D ALUMINUM SKIN HONEYCOMB CORE
..c:::1L7AVCO LYCOMING DIVISION STRATFORD, CONN.
Figure 3. Design Areas Selected for Composlte Structures.
10
Structural des1gn is in accordance with Federal Aviation Regulation, Part 23, airworthiness standards for nortnal category an planes.
A 17 percent thickness-to-chord ratio supercr1tical wing shape was selected because It offer s a number of advantage s over the conventIonal 12 percent NACA shape. These advantages are summarlzed In Figure 4.
o From the cros s -sechonal comparislon shown here, it can be concluded that the supercrihcal wing prov1des larger volume for fuel storage for the same chord width. The thickness mcrease has the supplementary benefit of hlgher -sectlOn modulus, permitting hghter construction for equlvalent bending loads.
o The two shapes have comparable drag characteristics 1n the crU1se mode. Increasmg the NACA a1rfo11 thickness 1n an attempt to ach1eve slmilar volume 1S 1mpractical, because it re sults 1n a signif1cant reduction in useful flight speed combined with an overall drag increase at lower speeds.
o Iterative design studies show a 25 percent increase 1n fuel capac1ty comb1ned w1th a 3 percent decrease 1n aircraft gross weight. These savings are for an equivalent aspect ratlO of 10 and a des1gn wing loadmg of 2250 N 1m2 (47 lblsq ft) of wing area.
o Pr10r test data have shown an apprec1able 1ncrease 1n lift capability as dep1cted m this comparison. This prom1ses a more forg1ving aircraft for var1ations 1n angle of attack and enhance s safety. For equivalent sophistIcatIon of flap systems, reduced landmg speeds are ach1evable, thus re sulting 1n shorter landing -field length capability.
The airfoil selected by Beech 1S sim1lar to the NASA GA (W) -1 airfo1l but 1S tailored specifically for the high-speed, high fuel volume, and the high lift requirements of the QCGAT configuration.
Major lift parameters are summarized below. Establishmg ophmum flap setting s was beyond the scope of th1S study. However, experience 1ndicate s that a full flap deflechon of 40 degree s for
......
......
1-1- +20 illI +10 () (!) 0 a:w Ws -10 a.. -20
-=-- ............ ------- ..... iiiiii;;;'
NACA 12% \
NASA SUPERCRITICAL 17%
+25%
GROSS WEIGHT 1 -3%
0.12 0.17 THICKNESS/CHORD
Flgure 4. Wmg SelectlOn •
CD 17%
NACA
+ / --17% S.C. & 12% NACA
MN 0.5 0.6 0.7 0.8
C L ...L.! ___ ------,
17% S.C. ~CL MAX 35%
f 12% NACA
ex
landing and a take-off flap setting of 40 percent of full deflectlOn are appropriate for Fowler flap design. These values of C L re-present available state-of-the art with advanced airfoils. max
Flap Pos1tion C
L ~ CL @ =0 max
Up 0.132 0.088 1.6
40% (Take-off) 0.98 0.088 2.35
Down 2.13 0.088 3.45
Full-span fowler flaps and spoilers have been introducted to ach1eve the desired 762 meter (2500 feet) take-off f1eld length and landing distances w1th reduced wing area. Winglets have also been added to reduce actual span and w1ng structural we1ght, while ma1nta1n1ng a h1gh effectIve aspect ratIo.
Many drag -lnfluenclng de slgn deta1ls of the QCGA T a1r plane are not estabhshed at th1S tIme, because the a1rplane 1S as yet a prelim1nary des1gn study. For drag analys1s, ambIt10us estImates were made for the var10US Hems. Ach1evement of total air plane drag coeff1c1ents w1ll requ1re exactIng efforts 1n the practIcal development of the airplane. The result1ng QCGAT a1rcraft drag compares w1th that of LearJet Model 24, wh1ch 1S an extremely clean a1rplane. Allowances have been made for d1fferences 1n w1ng th1ckness, component size'; etc. Drag coeff1cients used are summar1zed below.
Total CD • P
Incremental
Incremental
Incremental
Incremental
12
Flaps and gear up
CD for land1ng gear P
for full flap
for T. O. fi<l:p
for one eng 1ne out
0.02661, O. 02534 crUlse
0.0164
0.04066
0.0163
0.01209
Four major a1rcraft var1ables cons1dered 1n the parametr1c study to optim1ze the wing conf1gurahon are:
1. Wmg area
2. Wmg aspect ratlO
3. Fuel Yle1ght
4. Takeoff we1ght
For each performance goal1n th1S study, the hm1tlng aspect raho versus w1ng area is plotted for several takeoff we1ghts, includ1ng the effects of w1ng geometry on wing we1ght. These hm1ts for each of the performance goals are then summar1zed on a graph so that the be st compromise could be selected. A de sign pOlnt of 15.33 square meters (165 square feet) wlng area and an effective aspect ratio of 10 were selected.
The expected weights are summarlzed for duel, structure-pluspropulslOn, and complete aircraft w1th payload for both convent1onal and QCGAT a1rcraft de signs 1n Table 1.
The first line of Table 1 represents a hypothetical aircraft of current vintage des1gn w1th low bypass turbofan propulsion. Introduction of a QCGAT high-bypas s turbofan reduce s fuel consumphon by 22 percent. When th1S savings is iterated through the aircraft des1gn, structure and gross weight reduce, providing an additlOnal 5.5 percent in fuel economy. Similar iterations with weight saVlng s from composite materials and supercntical wing result in an addltional4.4 percent sav1ngs 1n fuel The comb1ned engine and aircraft changes provide 32 per cent better fuel economy. The 22 per cent reduction in gros s welght permits the use of a smaller engine with 22 percent lower thrust and, therefore, lower absolute emisslons and n01se.
The a1rcraft study projected the maX1mum ranges shown in F1gure 5 for various payloads. Wh1le 1134 kg (2500 lb) is depicted as maximum payload for the alrcraft, only 753 kg (1660 lb) is required to accomodate six people with their baggage. At this payload, the ach1evable range 1S in excess of 2963 kilometers (1600 nautical m1les). Flight condltlons are 10058 meters (33, 000 feet) and an average fhght speed of approx1mately 0.5 Mach number.
13
>-"
*'"
Kg
1200
o 800 « 0 -1
~ 400 1
o
(Ib)
3000
2000
1000~
00
MACH NO 05 (MEAN) AT 107 Km (35,000 FT) MAX FUEL 975 Kg (2150 Ib)
FULL SEATS (6 PEOPLE) PLUS BAGGAGE
200 I I I I I , , I I
(NM)
o 400 800 1200 1600 2000 2400 2800 3200 Km RANGE
F1gure 5. Best Range Versus Payload.
In our QCGAT aIrcraft study, landlllg dlstance, rather than take-off capabllity, set the mmimum usable airfield length. IntroductlOn of
full-span fowler flaps with moderate wIng loading results in a very low "landing configurationil stall speed. The 32 meters / sec (62 knot) stall speed compares with 41 - 46 meters/sec (80--90 knots) for current typical jet and turbofan aircraft. Since landing distance is proportlOnal to stall speed squared I this low landing speed provlde s an attractive sea level FAR landmg field length of 811 meters (2660) feet.
Figure 6 shows a representative sample of general aviation airfields plotted on coordinates of field elevation and field length o The Beech QCGAT aircraft with full useful payload has a landing capability consistent with the majority of these fields 0
The expected stall speeds promise a very forgiving aircraft in both the take-off and landing modes where most accidents occur.
3 0 2 ENGINE DESIGN
The NASA QCGAT engine was designed using existing advance technology in the fields of noise and pollutant emissions control. Stringent noise goals that were set required both attenuation techniques and design considerations for the major noise sources 0 Similar goals were set for emissions levels based on the 1979 EPA standards. These standards were later abandoned by the EPA as being excessively strict; however I they were retained as the QCGAT program goals.
Work has been performed in these technologies for large commercial applications; however I little benefit has filtered through to the general aviation industry.
While not specifIcally stated as a program goal, low acquisition and operating costs are of unquestionable importance to the general aviation marketplace and are the key ingredients to widespread acceptance.
The Lycoming QCGAT engine was developed to specifically satisfy the program goals of low noise and emissions 0 This was accomplished without incurring the excessive complexity' and cost which normally are associated with advanced technology systems.
3.201 Design Considerations
In order to achieve all of the preceding goals, a multifaceted design approach was required to optimize each component for its task without placing an excessive penalty on other components 0
15
....... 0"
M/SEC (KCAS) 50 1001
SPECIFIC FIELDS SERVED BY LOCAL SERVICE AIRLINES IN U.S.A.
o w
40
w 30 Q. 00 ...J ...J ~ 20 I-00
10
o
M (FT)
801
3r10~1 --~--~--~--~--~--~--~
40f-1 --+--
2011---l---4--
o o o T"""
>< Z2 o I~ > W ...J W 1 o ...J W U.
01 1 1 5 6 7 9 (Ib) 0
I I I
2 3 4 Kg
WEIGHT x 1000
QCGAT 8 1 1 '< =-"'"
61 ,~ - I • ,+I---i
, ,
41 1 /- --=----1 ,--' 1 ' --10 ---I , ' , ,
" ,
21 fr' ': ':: ...:..' -i:...:..:-+----t----t--=-. ... . ... . .. ~ .. :.~.~... . -.-01 I I, ,': :'::'::';~:;':::::':',~:: ',I 1 o 2 4 6 8 10 12 14 (ft)
I 1 1 I 1 o 1 2 34M
FI ELD LENGTH x 1000
Flgure 6, Aucraft Stall Speed and Field Capacity,
3.2.1.1 Operating Cycle
Preliminary engine cycle definitions were based on the power requirements of the proposed airframe and the fuel consumption. Figure 7 illustrates the relation of fuel consumption-to-bypass ratio as a function of the compressor pressure ratio. Pressure ratios [PrHc)above approx1-mately 10.4 would have necessitated a multi-stage power turbine that would lmpact the eng ine complex1ty and cost. In addlhon, 1ncreaslng bypass ratio above 9.6 tends to result 1n reduced performance at alhtude caused by off-des1gn fan load1ng. By selectmg a bypass ratio of 9.4 (at 25,000 feet alhtude) and a compressor pressure raho of 10 0 2, eng1ne
slmpl1city can be malnta1ned with m1n1mum fuel consumption. The effect of bypass raho on fan pressure raho is glven in F1gure 8. Overall eng1ne slzlng 1S based on FAA regulahons which requ1re the capab1lity of an a1rcraft to successfully cl1mb after takeoff w1th one eng1ne inoperahve on a hot day.
This thrust level also ensures that the aircraft will be capable of serving a majority of small, general-aviation airfields, including many remote, unpaved runways.
3.2.1.2 Low Cost and Simplicity
The Avco Lycoming LTS 101 series engines selected as a buildmg block for the QCGAT engine is a strong contender in the generalaviation market in both turboshaft and turboprop configurations. The use of this engine as a base allows the maximum use of existing technology and proven design concepts. One feature, which is especially attractive to the general aviatIon operation, is the modular design concept. Although ideally no major maintenance would be specified for an engine during its service life, maintenance and repair are an inevitable necessity and can cause significantly rapid escalating downtime operating costs.
Modular maintenance allows the expeditious removal of any major component for inspection or repair without removal of the entire engine. Interchangeability of modules promotes a fast return to service and decreases the spares requirements and life-cycle costs 0
3.2.2 Engine Hardware Description
The final QCGAT engine configuration shown in Figure 9 consists of four basic modules 0 (See Figure 10.) These modules are the fan, the gas generator or core engine, the power turbine, and the accessory gearbox.
17
,..... (UNINSTALLED) 00
(Ibm/hr/lbf)
kg/hr/N 0.67 .068
~ rtro_ , AL TITUDE = 7620m(25,OOO FT)
0.66 1 ~..., --- , I MACH = 0.6 I
.067 U u..
O.65r en .066 I-
ADDITIONAL Q TURBINE REDUCED
0.641 STAGES CRUISE .065 L REQUIRED PERFORMANCE
0.63 1 1 1 1 1 1
8.4 8.8 9.2 9.6 10.0
BYPASS RATIO
FIgure 7. Cycle Optim.izatlOn for Specific Fuel Consumption.
(Ibm/hr/lbf) (UNINSTALLED) kg/hr/N
.69 r BYPASS RATIO = 8.2 .070
I .68
I .069
I .67l 9.0
(,) -
~ .068 r \ \ I OPTIMAL FAN PRESSURE
.66~ 9.6 \ "'----/ RATIO
.067 I \ \ / II
ALTITUDE = 7620m(25,000 FT}
.65, Y MACH = 0.6 .066 L CORE COMPRESSOR
PRESSURE RATIO = 10.2 I
, .64
1.0 1.1 1.2 1.3 1.4 1.5 FAN PRESSURE RATIO
...... Figure 8. Selected Fan Pressure Ratio • -.0
20
FAN MODULE
-CORE MODULE
: ~ ~t I I
ACCESSORY MODULE
0-" r'\ til nr--'"""'--l r' L_. V---, oj L, LV" , , ( , --- J "\/LJi)tJ''-' . /Ifl J u ~_
Figure 10. Engine Modules. N .....
POWER TURBINE MODULE
J}~J1 ," --1~ L.. - -~,-
3.2.2.1 Fan Module
Design of the QCGAT engine fan module is similar to that of the Lycoming ALF 502. This module compnses a main support frame that features provisions for engine mounting and supports the fan, reduction gearing, and the power producer. The fan 1S driven on 1tS splined shaft by a bell gear which is, in turn, driven by the power turbine rotor through a connecting shaft to a sun gear and five fixed planet gears. The power turbine input speed is reduced by a 3.102 ratio by the reduction gear system.
The relatively wide-chord, low aspect ratio (1.6) design of the fan blades avoided the need for midspan dampers. This elimination of dampers resulted in an efficient design, both from a noise reduction and performance viewpoint. The fan rotor consists of 24 multiple circulararc airfoil blades and has a tip diameter of 57 Q 7 cm (22.72 m.). The inlet hub -to-tip ratio 1s 0.45.
The single-stage fan is followed by a splitter in order to separate the airflow mto the bypass and core engine flow channels Q In the bypass channel, a single-row stator vane assembly having 59 double circular-arc airfoils is used to remove all whirl from the flow before it enters the main frame duct. The number of vanes was selected on the basis of noise considerations because large-engine technology indicated a quantity in excess of twice the rotor blade number 0 The bypass stator vanes have the unique capability, as does the fan rotor, to indIvidually replace either a single vane or all vanes if they are inadvertently damaged.
In the supercharger channel that leads to the core engine, a double-row stator assembly was used. This configuration reduced the inlet velocity to the supercharger duct; this reduced velocity, which could not have been achieved with a single-row assembly, allows higher diffusion and lower losses. There are 53 double circular-arc vanes in each row with an average inlet Mach number and d1ffusion factor of 0.76 and 0.35, respectively.
3.2.2.2 Gas Generator Module
A cross section of the gas generator module is shown in Figure 110 The main components are the high-pressure compressor, the combustor, and the high-pressure turbine assemblies. This gas generator, also referred to as the core engine, is a growth version of the proven
22
23
and certified LTS 101 basic design which has both turboshaft and turboprop configurations currently in production in the 500 to 750 horsepower class. Takmg advantage of the latest technology in compressor, lowloss cOITlbustor, and high-teITlperature turbine design, a viable core engine suitable for the turbofan was developed.
The inside rear surface of the fan fraITle provides the Interface for the gas generator module (Flgure 9). The compressor section, which is a stacked as seITlbly, contains two axial stages and one centri
fugal stage claITlped together on the gas producer shaft by a nut located at the rear of the centrIfugal stage. The gas producer turbine assembly, consIsting of a spacer seal assembly and a disc-blade asseITlbly I is located at the rear of the gas producer shaft and cOITlpletes the shaft assembly for the ITlodule.
The cOITlpressor assembly and the turbine asseITlbly are each individually balanced; this perITlits reITloval of the turbine and reinstallation and eliITlinates rebalancing of the total assembly.
Buildup of the rotor systeITl IS accoITlplished by sliding the rotational components of the compressor section onto the gas producer shaft and seating theITl in place, on their piloting surfaces, with the claITlping force applied by the nut located at the rear of the centrIfugal stage. This claITlped asseITlbly is then balanced as a unit.
When the ITlodular buildup is cOITlpleted I the cOITlpressor nut clamping force that was initially applied across the axial and centrifugal stages is relieved because the claITlping force applied by the nut at the rear of the turbine dISC causes enough stretch of the gas producer shaft to relieve the initial load of this interITlediate nut. The shaft then acts only as a bolt applying a claITlping force across the rotor interfaces.
3.2.2.2.1 High-Pressure COITlpressor
The high-pressure cOITlpressor for the QCGAT engine consists of a zero-staged LTP 101-700 unit. As such,it contains two axial stages and one centrifugal stage. "Zero staging" of the -700 cOITlpressor reqUIred SITlall ITlodifications to the original axial stage to cOITlpensate for inlet air angle variations caused by the addition of a preceding stage. Blade chord width was increased 16 percent so as to alter natural frequency and provide additional ITlargin froITl vibratory excitation. No changes were required for the original axial exit stator, centrifugal cOITlpressor rotor I or air diffuser.
24
The new first-stage (zero stage) was designed at modest stage pressure ratio to favor part-speed operation for the QCGAT design. The strip stock for the new stator IS identical to that used on the second stage.
The first-stage axial compressor IS an integral casting contain
ing multlple clrcular-arc blades deslgned to maxlmize the tlp sectlon efficiency at the high rotor tip relative Mach numbers at which the blades operate. This compres sor stage 1S cast in an alloy that is a proprietary
rnartensitlc, age-hardenable stalnless steel.
ThiS rotor IS followed by the first-stage stator which is a split and brazed assembly. These vanes are rnade of strlp stock for the beneflt of low cost and have double c1rcular-arc proflles. The stator 1ncldence angles were set based on an off-deslgn analysis to provide a w1de operatlng range.
The second aXlal stage 1S a slm1lar des1gn to the flrst stage, but at sl1ghtly dlfferent sol1dities.
The centrifugal compress.or stage following the two axial stages IS an integral casting and is ldentical to Unlts operating In the LTS 101 engines.
3.2.2.2.2 Combustor
Selection of the LTS 101 combustor style for a small, low-pollution turbofan engine was based in part on the unique features of the circumferentially stirred or "horseshoe" vortex annular combustor configuration.
Figure 12 Illustrates the aerodynamic concept embodied in the combustor design. Primary air is admitted through slots in the liner header to produce flow circulation about a circumferential mean line. Air Jets ("folding jets"), entering through the inner wall, force primary zone rec1rculation. Since the secondary holes exist only on the inner wall, the vortex fills the full annular height of the liner and produces adequate flame stabilization within a smaller cross -sectional space.
With the folding jet in line With the fuel injector, initial flow circulation is in a circumferential direction o The vortex is permitted to turn to the axial direction on either side of the folding jet. As a resuIt, the mean path of the combustion zone flow vortex takes the shape of a horseshoe centered on the injector and folding jet axial centerline.
25
N 0'
SECONDAHY (FOLDING) JETS
PRIMARY
FUEL INJECTOR
COMBUSTOR IIOUSING
I
PRIMARY
VORTEX CIRCULATION
PATTERN
HORSESHOE-SHAPED MEAN CIRCULATION PATH
SECONDARY (FOLDING JET)
~,./ -,~ d9 '8-\ ,·-1 GJ;
~~Q~ ! J J,-,t%t t PRIMARY
SLOT AIR ~:x FUEL
INJECTOR
Figure 12. Clrcumierentially Stirred Combustor Flow Pattern.
The number of fuel inJectors is reduced by one half, compared with normal practice, because of thIS unique combustor primary zone aerodynamic concept.
Primary air, entering in two stages from the outer liner, is directed along the dome of the liner to provide cooling and is then dlffused into the vortex created by the folding jets. The baffle which deflects the first-stage primary air is cooled by Jet lmpingement. The second-stage of primary au enters the burner on the injector centerline through a step formed by the double liner end.
Two spark igniters located in line with the lower two fuel injectors penetrate through the primary air baffle. Since this location has been found to be optimum in the LTS 101 engine, the QCGAT combustor housmg has been designed to maintain the location.
Eight airblast fuel injectors equally spaced in the liner end straddle the vertical engine centerline. Each injector has a self-aligmng air seal at the liner and penetrates past the second-stage primary air step. The fuel manifold consists of two halves to allow easy removal. A flow divider mounted under the engine has separate primary and secondary lines to feed each manifold half.
The liner walls are designed for maximum structural integrity and wall coohng efficlency. Splash-cooling rings are formed by the overlap of each cone, which provides a cooling effectiveness advantage over the standard LTS 101-type cooling ring inserts. Coohng air is metered through holes drilled in the outer cone so that air iets impinge on the splash ring, thus providing cooling of the splash ring lip; this air then spreads to become a lower velocity film at the discharge of the splash ring. Air quantity and conical step length are designed so that adequate cooling effectiveness persists until the next Joint is reached. To minimize buckling caused by uneven heating and thermal expansion, the joints are spot welded and then back-brazed to provide good conduction of heat into the colder, outer cone. In addition, the lips on the outer liner are slotted to prevent compres sion buckling.
Inner and outer liner-to-curl seals are the "fishmouth"-type used on most Lycoming engines to insure a seal regardless of local liner or curl warpage. These curls are cooled by air films, slmilar to those at the liner walls. In addition, the outer curl is of double-wall construc .. tion with turbine inner shrouds cooling air flowing through it, and the hot side wall is coated with thermal barrier coating. This additional cooling provides margin for the higher temperatures of the QCGAT, as compared
27
with the single-wall LTS 101. Liner mounting is accomplished by means of four radial mounting pms which make the liner free floating, while providing complete centering and axial retention.
The design configuration selected for the QCGAT liner produces a unique CO/NOx relationship which permits trade-offs to meet EPA emission requirements (low NOx at high power settings). Furthermore, the use of ALF 502 airblast injectors improves combustion efficiency at idle over the LTS 101-type dual-orifice injectors with a resulting reduction of UHC and CO. The fuel injection system used for this engine is similar to the airblast system used on the ALF 502 engine. Eight standard-production ALF 502 injectors are used. Flow divider orificing has been changed from the ALF 502 to accommodate the fewer number of injectors in the QCGA T engine. The fuel injector manifold has been split in two halves to provide easy removal and mstallation on the QCGAT engine.
The airblast injector flow system is depicted schematically in Figure 13. This system provides an optimum trade-off for fuel contamination resistance, cold starting capability, simplicity, and reliability. Fuel flow entering the flow divider, under starting conditions passes through a wash-flow filter and into the eight pilot injectors, which are low-flow/high-pressure drop orifices used to provide good starting fuel spray. As mlet fuel flow increases, the airblast portion of the injector starts to "cut-ln"; as this flow increases, the pilot flow decreases. This arrangement minimizes the lncrease in fuel pump back pressure, yet still keeps the pilot injector cool enough to minimize internal coking. The mixing of pilot and airblast fuel flows at idle was chosen to produce good combustion efficiency and minimize CO and UHC. A reference pressure line from the spring-side of the flow divider vents to engine fuel pump inlet pressure, thus preventing fluid lock of the piston.
During starting, there are, in effect, eight starting fuel injectors (pilot injectors) and, because they are placed in the center of the airblast injector (Figure 13), atomization and droplet distribution is superior to merely pressure-atomized spray. This system demonstrated an outstanding improvement in cold starting, altitude relight capability, and higher combustion efficiency at idle operation in other Avco engines.
This type of combustor is applied to the standard Avco Lycoming reverse-flow annular combustor design to effect a short, compac t, lightweight engine having no critical shaft speed in the operating range.
28
LEFT MANIFOLD
REFERENCE DRAIN (A/C BOOST)
ONE WAY
VALVE
PILOT WASH
FLOW FILTER
CONTROL -6 FITTINGS
...J W :::J U-
RIGHT MANIFOLD
INJECTOR FILTERS PILOT
SLOT 005- 007 WIDE
FILTER 002-004'
DIA
m AIR BLAST
it SLOT 026 DIA
~ PASSAGE
- fill! Rlt\SI SI'RM
:5 020- 025 ANN al g; FILTER 013-015' DIA «
000075 RADIAL CLEARANCE 000125
INLET FILTER 020/ 024 \ IN DIA HOLES
Flgure 13. Airblast Fuel InJector System.
29
3.2.2.2.3 High-Pressure Turbine
The gas generator turbine consists of a single -stage axial turbine that powers the compressor. This high-pressure turbine is similar to the LTS 101-600/700 gas generator turbine except for modifications that optirn.i.ze it for the higher cycle pressure ratlo and temperature of the engine. The turbine arrangement is depicted in Figure 14. The gas generator turbine stator and rotor are cooled by compressor discharge air. Stator cooling air is discharged through slots on the pressure-side near the trailing edge, while rotor cooling air is discharged by means of trailing-edge ejection.
The vane s ln the flr st-stage turblne nozzle have a generous leading edge radius and relatively thin trailing edge in order to obtain a low leading edge gas -side heat-transfer coefficient and low trailing edge wake losses. The vanes are cooled by a two-pass cooling configuratlon by means of internal convection cooling at the main body of the vane and external film cooling at the vane's trailing edge. The twopass cooling design provides a long in-line passage length for good use of the cooling-air thermal capacity. Passage flow areas are varied to maintain a high mass velocity flow rate to match the local heat input rate at the vane's outside surface to achieve a near-uniform vane temperature.
Cooling air to the internal air passages of the vane is supplied through an opening at the vane's leading-edge hub section. Cooling air in the vane's leading-edge passage flows radially outboard. At the vane tip, the air makes a 180-degree turn and enters the vane's tail passage. In the tail pas sage, the air flows inboard radially, and, at the same time, it discharges gradually to the gas stream through three slots in the vane's trailing edge pressure-side wall. A schematic representation is presented in Figure 15.
The gas producer rotor blades are cooled by a two-pass coohng configuration similar to that of the gas producer nozzle vane. However, two separate cooling air passages are used in the rotor blade cooling design, as can be seen in Figure 15.
One branch of coohng alr enters the blades' leadlng-edge passage through an opening at the front face of the blade root. ThlS alr flows radlally outboard ln the leadlng edge passage and turns to the tail passage at the blade tip. In the tall passage, the au flows radlally Inboard, and, at the same time, discharges gradually to the gas stream through a traihng edge "through slot". This flow path termlnates
30
31
Vl N
----.......--r--~::., c. (, I "" ,:; 1-'" """ ', .... ' .... c. I CI I , C'I~ I .. ,I ,,,, I CI't I '(I CI ,. ~ I ::., I \ CI (. I c. ",I \ I '. I '_)
.. I , ... C,
BLADE LEADING EDGE COOLING I"
PATH
BLADE ROOT COOLING AIRFLOW PATH
NOZZLE COOLING FLOW ---
Flgure 15. Coolmg of First Stage Rotor Blades and Turbine Nozzle.
Just short of the blade I s hub section. The other branch of cool1ng aIr enter s the blade at the blade root where it flows fir st aXIally rearward between the blade root and the dISC rim and then radIally outboard through the rear portion of the blade shank to the blade hub sectlon. At the blade hub section, the air discharges to the gas stream through the trall1ng edge "through slot". By flowing through the blade root and rear shank, the second branch serves to cool both the blade root and disc rIm. The blade coohng air passages are designed wIth variable flow areas to obtaIn optimal local mass velOCIty flow rate sand coohng effectiveness.
3.202.3 Power Turbine Module
The QCGAT low-pressure power turbIne (Figure 16) is a sIngle, axial stage wlth moderate stage loadl.ng and non-free vortex desIgn. An outward flowing, diffusIng duct connects the low an(1 high pre ssure turbmes and provides bearing support. The low-pressure rotor is integrally cast, wIth unshrouded, medium aspect ratlo, constant tlp diameter blade s.
The hIgh-speed, s ingle-stage configuration was chosen as the best combination of performance, mechanical simplicity, weight, and costa The reduction gear between fan and .low-pressure turbine allows rotational speeds nearly optimum for both components, while maintaining low fan tip speed and turbine blade passing frequency beyond the audible range for reduced engine noise o The low-pressure turbine design point is between sea level takeoff and M=0.6, 7620 m (25,000 ft) maximum cruise operating points.
The rear flange of the combustor casing on the gas producer module .Figure 14. provides the forward interface for the power turbine module,Figure 16. The two major components of this module are the rear bearing support housing and the power turbine rotor, which it supports. The rear bearmg support housing, which is the module's main structural member, is a segmented investment casting. This housing, which supports and transfers the power turbine loads, also functions as the main artery to ensure proper operation of the power turbine module.
An area-compensated aerodynamic diffuser flow path in this housing interconnects the gas producer turbine with the single-stage power turbine. This duct is divided by four hollow airfoil-shaped struts which serve as the access corridors for lubrication and ventilation, as well as the structural supports for the rear bearing housing.
33
F 19ure 16. Power Turbine Module.
34
The support housing's inner cylinders, which are supported by the access struts, contain the bearings and seals necessary for the rear of the compressor shaft and the power turbine stub shaft. Spring preloading of the aft bearings on both of these shafts assures a constant axial bearing load application. This bearing compartment is sealed from hot combustion gases by fore and aft heat shields, which extend from the diffuser duct inner wall to the bearing housings, and by fore and aft controlled gaps eals •
The periphery of this housing, which bolts to the rear of the gas producer module transmits the power turbine reaction loads, seals the burner cavity, and contains the necessary service bosses. The cast
power turblIle nozzle and brazed rear beat shield assembly. and at a larger radlus, the fuel manifold assembly, bolt to the rear of the castlng. Themocouples that measure the gas temperatures enterlng the power turbine lnsert radlally through bosses next to the module connect~on flange. Opening thIs maIn flange connection and 'the power turbine shaft lock, permits withdrawal of this module from the core englne for hot-end inspectlon.
Tl e power turbme rotor consists of a cylindrical stub shaft and
an integrally cast, inertia-welded rotor assembly. This assembly pllots on and IS clamped to the two spht ball bearing s in the rear bearIng support housing.
A schematic of the mel'idlOnal flow path of the turbine sections is shown in Figure 17.
3.2.2.3.1 Interturbine Duct
The interturbine duct moves from the high-pressure turbine exit gradually outward following the lower contour of the combustor to reach the larger diameter of the low-pressure turbine. A moderate diffusion is tolerated throughout the duct in order to enhance the flow acceleration rates across the low-pressure turbine blade rows.
Figure 18 shows in solid line the velocity distribution along mner and outer duct wall, respectively, as calculated by axisymmetric analysis. The dashed line indicates the strut surface velocity distribution obtained by superlmposltion of the two-dimensional cascade flow on the meridional flow velocity.
Four struts provide the load carrying structure for the turbine bearings and rigidly connect the duct walls. These struts have an axial
35
W 0'
C/) W I o ~ Z 0
10, 25
9
Sr 20
7 HIGH PRESSURE
~ 6 15 ~ TURBINE o « 5 a:
41 10
:~ 5
1
-5 o 5 10 15
LOW PRESSURE TURBINE
41 36 VANES BLADES
20 25 30 35 CM
I I I I I I I I I I I I I I I I I INCHES -2 -1 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14
AXIAL DISTANCE
Figure 17. Turbine Section Meridional Flow Path.
u UJ CJ)
~ 1000
900 -
>- 800 !:: u o -I
~ 700
600
500
u UJ CJ) ...... tl:
1000
900
~ 800 o o -I UJ > 700
600
500
300
250
200
150
CJ) ...... :E
300
250
200
150
UJ g UJ CJ Z ::;
~ I-
--- WAll
-- STRUT
UJ UJ J l-I ,
L-__ ~L_ __ ~I~I __ ~I~ __ ~I~~~I--~~IIL--.~I--~I. eM " 6 8 10 12 14 16 18 20
a: o 5 a: a: o .... « a: UJ z UJ CJ CIl « CJ
~~----~3----~~----~5~--~6~--~7~1----~8 INCHES
AXIAL DISTANCE
-- WAll
-- STRUT
UJ J UJ
L-� ____ L� ____ L�JI __ -L� ____ IL_ __ ~IL_ __ ~I~I __ -JI ____ J_I CM 4 6 8 10 12 14 16 18 20
2~1----~3----~J~--~5~1 -----6~1----~7----~8 INCHES
AXIAL DISTANCE
Flgure 18. Interturbine Duct Inner and Outer Wall Velocity DistnbutlOn at Sea Level Takeoff
37
length of 7.6 CITl (3.0 in.) and a ITlaxiITluITl thickness of 1.2 CITl (0.47 in.) to allow internal passage of the rear bearing cavity service lines.
To ITlininrlze strut blockage losses, channel contour in between the struts has been adjusted to cOITlpensate for strut blockage by locally tailoring the flow path in the strut region.
3.2.2.4 Accessory Drive Module
As there were no contractural requireITlents for the gearbox design, it was sized for available bearing size's and therefore represents a basic boiler plate design concept. Each particular airfraITle application would specify its accessory load requireITlents and the appropriate flight weight housing and gearing systeITl would then be designed.
The accessory drive ITlodule is ITlounted at the bottoITl of the far.. fraITle and connected to the gas generator core through a steel turret shaft as shown in Figure 19.
The ITlodule is easily reITloved by reITloving the plug, extracting the turret shaft and unbolting the housing (4 bolts). A cross section of accessory gearbox is depicted in Figure 20.
The gearbox housing is a two-part design and consists of the housing and a front cover. Scavenge oil, which drains frOITl the fan ITlodule through a hollow strut into the gearbox ITlodu1e, is reITloved froITl the gearbox by an externally ITlounted scavenge pump (oil pUITlP pad). PUITlP, scavenge, and drain ducts have been sized to perITlit safe oil removal under any operatlng condltlon o A rotatlng oil-au separator is used to dUITlp engine seal pressurization air overboard via a vent located in the gearbox forward face. All gears and bearings are oil ITlist lubricated.
The gearbox also contains the gas generator speed pickup as well as a chip detector and oil drain plug.
3.2.2.5 Engine Control SysteITl
The majority of fuel and power control cOITlponents are grouped around the ITlain fan fraITle and are acces sible for ground ITlaintenance.
The gas producer fuel control is installed directly on the fuel PUITlP, which in turn, is ITlounted on the accessory gearbox. This
38
TURRET SHAFT
Figure 19. Accessory Drive Module Installation.
39
>I>o
GAS PRODUCER SPEED PICKUP
FUEL CONTROL
A
--~-
SECTION A-A
FIgure 20. QCGA T Accessory Gearbox.
t A
external arrangerrlent elirrlinates vulnerable external pressure and return lines to the fuel rrletering section of the control and saves an additional pad and gearing on the accessory gearbox. Both the pUrrlp and control filter screen assemblies are designed so that they rrlay be easily rerrloved for cleaning or rep1acerrlent with the control rrlounted on the engine. Local screening is used within the control in order to provide protection for orifices and valves.
The ambient terrlperature signal at the engine inlet is directed to and frorrl the gas producer control by two flexible hoses rrlounted at the entrance to the engine fan stage.
The fuel and engine control systerrl for the QCGAT application were selected to achieve irrlportant design criteria such as reliability, durability, and sirrlplicity without cOrrlprOrrlising functional requirerrlents of the engine. The engine control systerrl perforrrls the bas1c functions of rrletering the required arrlount of fuel flow to the engine during starting, acceleration, deceleration, and steady-state operation. In addition, it also controls the operation of the cOrrlpressor inlet flow fence asserrlbly during steady-state and transient rrlaneuvers. The QCGAT engine control systerrl consists of the following rrlajor cOrrlponents: Fuel control, fuel PUrrlP, flow fence actuator and terrlperature cOrrlpensator.
The high-pressure fuel pUrrlp incorporates a pUrrlping e1errlent and gear set used on Lycorrling T53 engines which has been repackaged in a new housing for utilization on LTS/LTP 101 power plants.
The gas producer control and ambient terrlperature cOrrlpensator are substantially identical to systerrls that are in production for the Lycorrling LTS/LTP 101 engine.
The inlet flow fence actuator was specifically designed for the LTS/LTP 101 engine and with the exceptlOn of a few rrlinor changes, is directly applicable for the QCGAT application.
3.2.2.5.1 Fuel Systerrl Operation
A scherrlatic flow diagrarrl of the engine power control and fuel systerrl is shown in Figure 21. The fuel frorrl the airfrarrle supply systerrl is supplied to the fuel inlet port on the engine fuel pUrrlp. A 10-rrlicron barrier filter 1S required in the airfrarrle fuel systerrl to protect the engine fuel systerrl against contarrlination. The inlet pressure 1S increased by the engine-driven fuel pUrrlp to the level required for fuel nozzle injection. The pUrrlp output flow enters the gas producer control
41
*'" N
r-;~l-' I I I SUPPLY I L ___ J
r ;;l~;' I 1O-<{ L ___ .J
JET
T .IINOUCER STAGE
MFP 260 FUEL PUMP
LEGEND
FUEL
c:==!::> AIR
r===!> OIL
NH ------
_·_·-·_·_·_·_·-1
FUEL ! STRNNER !
ACCEl & DECEl
AMBIENT TEMPERATURE I<:'!=:::=J
SENSOR
NH
FUEL SCHEDUlING
lINKAGE 14"< I
i_._._._._.
PRESSURE
RELIEF
VALVE
fUEL
BYPASS VALVE
METERING
VAlVE
Ll.P-C
DP 52 FUEL CONTROL
fOOT
VAlVE
ENGINE COP
VALVE I
CONTROL POWER LEVER
AIRCRAFT POWER LEVER
TO OVEASPEEO DETECTOR
FUEL MANIFOLD & NOZZLES
OVERSPEED SOLENOID
VAlVE
a: ~
I OIL ~ I COOlER I rt r----1~
-~-~- ~LI ~~
DUMP
nANN
VALVE
Flgure 2l. Power Control and Fuel System Schematlc.
frorrl where the engine derrland flow is rrletered to the low pressure turbine fan overspeed solenoid valve, engine oil cooler, and finally into the fuel flow divider. Here the flow is split into a primary path and a secondary path injected into the combustor by eight dual orifice injectors. The pUrrlp flow in excess of the engine demand is internally returned from the control to the pump inlet.
The fuel control performs the following functions:
o Maintains the engine speed condition as demanded by inputs from the operator via the power lever.
o Schedules the proper amount of fuel flow for accelerations and decelerations.
o Schedules the fuel flow required for engine starting.
The ambient temperature compensator, which is physically a separate unit mounted in the engine inlet, is functionally a part of the control and serves primarily to bias the acceleration schedule with ambient temperature. The control is pictorially shown in Figure 22.
3.2.2.5.2 Inlet Flow Fence Control and Actuator
A pair of retractable rings are located in the engine inlet housing in front of the compressor. The rings are mechanically operated by a pneumatic actuator whose output stroke is scheduled by a closed-loop, integral pressure ratio (Pc/Pin or P3 /P2 .1) controller. Refer to Figure 23 for a schematic representation of the complete assembly. The actuator and controller are mounted on the compres sor diffuser and are shown in Figure 24.
At speeds up to approxirrlately 80 percent NH, the rings are extended into the inlet air stream to prevent low-speed rotating compressor stall. Above 80 percent NH, which corresponds to a particular engine pressure ratio (Pc/Pin), the actuator begins to move and gradually retracts the flow fence rings. Above 90 percent NH, which corresponds to another Pc/Pin, the fences are completely retracted out of the air stream. This action permits surge-free gas producer accelerations of approximately 5 seconds from flight idle and to achieve maximum rated thrust for takeoff conditions.
43
FUEL PUMP
Figure 22. Fuel Control Assembly.
>I:>U"1
ZERO STAGE FLOW FENCE
FULLY RETRACTED
MID POSITION
FULLY EXTENDED
FIRST STAGE FLOW FENCE
COMPRESSOR DISCHARGE PRESSURL P3
COMPRESSOR INLET ~ ~ ~ PRESSURE I P2.1 ~~~~CJ ~~%~
~\..~S~
F 19ure 23. Dual Compressor Flow Fence Schematic.
Figure 24. Flow Fence Actuator.
The unit is self-contained with mechanical closed-loop positioning of the actuator piston. The only input required is compressor discharge pressure (Pc).
3.2.2.5.3 Engine Overspeed Protection
The overspeed trip utilized here 1S the same unit, with modified trip point frequencies as that used on the Lycoming LTS 101 commercial engine. The design was taken in total from the engine protection/ sequence control developed for Lycomingls AGT 1500 gas turbine army tank engine.
The system particulars are as follows:
Trip Speed:
Reset Speed:
Response Time:
Input Power Supply:
Switched Output:
Input Signal Volts:
Temperature:
Altitude:
Vibration:
Shock:
EMI:
108% of Power Turbine, Takeoff Speed, RPM
95% of Power Turbine, Takeoff Speed, RPM
Less than 6 milliseconds for the trip, Ie s s than 45 millis e cond s for the total system
16 to 30 VDC per MIL-STD-704
1.5 amps
+4.5, -2.0 min at trip speed +2.5, -2.0 min at 1/2 test
_650 to +250o F
-1,000 to 50,000 feet
20 gls 5 to 500 Hz
30 gls 11 MS
Tested to MIL-STD-461, Notice 4
47
Engine power turbine overspeed protection is provided by an electronic overspeed trip unit via the engine's main fuel flow. The overspeed protection system. is schem.atically shown in Figure 25. Protection is accom.plished by m.onitoring the engine power turbine speed with a variable reluctance m.agnetic speed pickup which senses the shaft speed directly. The resultant output signal is an output pulse repetition rate which is proportional to the shaft speed and having a m.inim.um. voltage am.plitude. This signal is supplied to the overspeed trip unit which conditions it into a fixed geom.etry pulse train whose frequencyl shaft speed inform.ation has been carefully preserved.
3.3 NACELLE DESIGN APPROACH
A prelim.inary design of the flight nacelle was defIned to establish a realistic baseline from which a ground test nacelle could duphcate the im.portant feature s at reduced program. cost. However, only a ground test nacelle was fabricated.
The flight nacelle conception shown m Figure 26 com.prises the following sections:
1. An inlet duct to provide uniform. flow into the engine
2. A fan outer duct and core cowl to guide the bypass air around the engine
3. A mixer assem.bly to for ce the mIXIng of hot, higher velocity core engine exhaust with the cooler, lower velocity fan stream
4. A mIXIng cham.ber preceding the final nacelle eXIt nozzle
5. An aerodynamically shaped outer skm de signed to mmImIze drag at the higher flight speeds.
A mIxed-flow exhaust system was selected because it reduce s the peak exit velocity which improve s propUlsive effICIency and reduces jet noise.
Noise attenuatIon treatment in the form of perforated acoustic panels was introduced in the air intake section and the fan duct outer wall.
48
~ -.!)
AIRPAX SPEED PICKUP
FAN SPEED
AUTRONICS OVERSPEED DETECTOR
AIRCRAFT DC POWER SUPPLY I
I I
OVERSPEED
FUEL SOLENOID
Wt TO ENGINE
t
LOGIC SIGNAL SPEED & NF ACTUAL .1 SIGNAL SWITCHING (NF>4 5 VOLTS),
' C IRCU IT C IRCU IT I ........ ~_ MONOPOLE PICKUP
OVERSPEED SOLENOID
VALVE
HALF SPEED CHECK (NF~ 25 VOLTS)
t Wt FROM CONTROL
(BEFORE FAN REDUCTION GEARS) 6 MILLISECONDS RESPONSE 15 MILLISECONDS RESPONSE
TOTAL SYSTEM RESPONSE INCLUDING COMBUSTOR LAG IS LESS THAN 45 MILLISECONDS
Flgure 25. Fan Overspeed Tnp System
U"l o MIXING CHAMBER/FINAL NOZZLE
MIXER NOZZLE
FAN DUCT
Figure 26. Cross Section of Engine W lth Cutaway.
The fhght nacelle 1S designed for optimum crU1se performance at 0.6 to 0.65 Mach no. at 7620 meters (25,000 feet) altitude and 1S adaptable to pylon mountlng on the side of the fuselage or top mountlllg for underwing lllstallations.
ConventlOnal metal construction (mainly aluminum) that is consistent wIth simphcIty and the low cost requIred to general aVIatlon
application was used throughout the de slgn. The mam de slgn objectIve s have been low nOlse (low internal and external tones), ease of accessibility, low weight, and low cost. The air inlet is designed to provIde high-pressure recovery and inflow incidence tolerance.
3.3. 1 Aerodynam1c Design
The overall aerodynamic contour s of the flight nacelle are shown in F1gure 27 • A separate evaluation of drooped inlets having reduced curvature over the front lower contour indicated no Improvement m the external aerodynamic drag when compared with a stra1ght inlet. As a re sult, the lower cost aXlsymmetric mlet was chosen.
Basic de sign considerations and front cowl geometry are shown in Figure 28 The inlet throat denoted 1Zt 1n the figure has been designed for a low Mach number to efficIently accommodate up to 20 percent mass-flow growth. Front cowl external geometry ratlos
LEXT/RMAX' Roo IRHL and R MAX IRHL have been checked to en
:Jure zero compressiblhty drag divergence and zero spillage drag over the full range of cruise power setting s.
The drag divergence Mach number of the front cowl shown in Figure 29 1S substantially above 0.65 Mach number for the upper and lower contours of the nacelle. The criterion used was based on tests of various geometry NASA Series 1 cowl shapes that relate dimensions and cowl curvature to drag dIvergence Mach number.
Spillage drag margin shown in Figure 30 relates cowl curvature determining dImensions and mass flow ratio (Roo IR
HL) 2 to a limit
line that defines onset of sp1llage.
The crIteria shown for cruise on the critical upper hp at 0.65 Mach number at 7620 m (25, 000 ft) wIth the engme operatlllg at max1mum
crUIse power would permit an aircraft incidence of 2 degrees before the onset of spillage drag. Th1S incidence margin is higher than that WhICh would be experienced in steady-state cruise. Some spillage
51
U1 N
2289.81 1--1 --------- (90.15) -------J,...-.,I
NACELLE MAX VERTICAL PLANE
533.4 -l (21.0) 1 882.90
1(34.76) I
377.44 I 6146.8 l I (14.86)ll1R~ci~s ..-----====~r-:=- I I -J.
604.52 I
DIA (20.8) AN = 1438.71 SQ. eM - J-(23.8) 528.32~ _ i! I 1 I
1 DIA 30 49'-: (223 SQ. IN.) 40
I L L. -.J I I' - -;J==-----.--
2/1 ELLIPSE
I '
NACA-1 SERIES
CONTOUR
4673.6 (184) 180 24' RADIUS
NACELLE WETTED AREA 5.11 SQ. M (55 SQ. FT)
NOTE: Dimensions Are in MM (IN.) Excepted as Noted
F 19ure 27. QCGAT Nacelle External Geometry.
\Jl W
RMAX
Rlt
---- - -____ H ............. r
I --..j~---
DIMENSIONS CM (IN.) UPPER LOWER UPPER LOWER
RATIOS LIP LIP CONSIDERATION SYMBOL LIP LIP
LExT/RMAx 1 413 1 558 AVOIDANCE OF
RMAxlRHL 1 252 1 676 SPILLAGE OF COMPRESSIBILITY
(1)R IRI 0761 DRAG DIVERGENCE
INCIDENCE OF
RMAX 3774 5055 !
(14 86) (1990) I RI 2644 (1041)
RHL 3015 (11 87) I
RHURI 1 14 CROSSWIND Re 2908 (11 45) TOLERANCE LExT 5334 7874
Re/RI 1 1 PROVIDE INTERNAL
LINT/RI 2056 DIFFUSION WITH MINIMUM DISTORTIOI'L
(21) (31 ) LINT 5436 (21 4)
(1)R 2294 (903)
(1) Cruise Max Continuous Rating at 7,600 M (25,000 FT), 065 Mach No, Standard Day
Figure 28. QCGA T Nacelle Front Cowl Geometry.
~_L_E_X_T ____ ~~~---. ___ ____
~L ~ao I
RMAX
---4--~-~
9
0 Z :l: U « ::E w 8 u z w <!J a: w =:! 0 <!J « a: c
7
o 2
( : :~ ) / 1 - (RHL/R MAX) 2
Figure 29. QCGAT Nacelle Maximum Drag Divergence Flight Speed
S4
permlsslble for transient maneuvers will not impact missIOn fuel economy. The lower lip with its hIgh external lip curvature gives substantially more margin.
The incidence tolerance of the inlet for takeoff and landing approach conditions IS shown in FIgure 31. With a contrachon ratio (R /R) equal to 1.14, the Inlet maintains acceptable flow
HL t distrlbuhon to the fan for an inflow incidence over 30 degrees for approach and over 40 degrees for takeoff. This provides at least a 20.6 m/sec (40-knot) crosswlnd tolerance.
The ratio of the Inlet length (Flgure 28) to the throat radlUs LINT/Rt IS shghtly above 2. OJ this ensures mlnlmUm Inlet distortlOn and provides suffIcient space for effechve Inlet noise attenuahon panels. The Inlet dlffuslOn half angle IS below 4 degrees as shown in Figure 2~'
The inlet cowl shape which has a leadmg edge radIus' equal to 1.5 percent of the hl~)oIIrhght radlUs (R ) blends into a NASA Senes 1 outer cowl contour and a 2: 1 aspe~l-"ratlO elhpse that forms the mner lip shape.
Estimated inlet pressure recovery (PT1
/PT
co) at the fan lnlet face IS shown in Figure 32 with both nOlse attenuahon panels and hard wall panels. At flight Mach numbers greater than 0.15, the pressure recovery IS approximately 0.997 for a hard-panel duct and 0.995 wIth noise -attenuation panels.
The external geometry of the nacelle is shown in FIgure 27. The
boat tall angle at the lower fan contour has been limited to 18 degrees to ensure separatlOn-free operatlOn for steady-state flight throughout the aIr craft fhght envelope. The SIde and upper contour s have a boat tail angle of about 14 degrees and a curvature (Rc/R
MAX) of 16. The
total wetted area of the nacelle IS 5.1 square meters (55 square feet). The nacelle drag area IS .015 square meter (0.158 square feet) assumIng a drag coefflclent (C ) of 0.0024 based on total wetted area and a
Dwet" body flneness ratlO form ... actor of 1. 2.
Flow areas and mass-flow averaged Mach numbers along the air inlet and fan flowpaths are shown m Flgure 33 for sea level takeoff conditions. Duct mach number s at all stations are below 0.4. Fan and primary
exhaust flows are mIxed by means of a multI-lobed mlxer nozzle as indIcated m th~ F\gure 32 •
56
\.J1 -l
KG/SEC M2
180 DHL= 11 QCGAT INCIDENCE LIMIT Ot . I 1.2 35
160
30 INLET I 140
SPECIFIC REFERRED AIRFLOW, 25 ~120
( LB/SEC_)
FT2
20 100
80
/11.14/ QCGAT / ",1,:1 /
----I~~ / TAKEOFF / ,/111'/
/ // 1,'/\11 DHL
/1i',1 APPROACH ~ / i /
~
Ot
20 30 40 50 60
INLET ANGLE OF ATTACK, oC , DEGREES
Figure 31. QCGA T Inlet Flow Incldence LimltS.
U"l 00
1.0
.99~
i .93 ~ I
.97
o .1
WITH HARDWALL PANEL
'WITII llOISE
~PTI TREA TMENT PANEL
~ c:::= CRITICAL OEI PTco
CLIMB-OUT --~
.2 .3 .4 .5 .6 .7 .8
FLIGHT MACH NUMBER
Figure 32. Inlet Pressure Recovery.
U1 -.0
FLOW AREA
(IN 2) --- -"",...
500 CM2
3000
400~,~ 2000 r 300 I I
1000 200
100
u c: i= w ~ « 6 III U. I-~ U. (/) ::J0~
COMBINED FLOW
W...J 4 Z~W I'~ I « > u I- ~ 2, « @J
~ « W (/)
0
----------r-------Ii I
ENGINE DUCTING
FAN FLOW--
MIXING PLANE
MIXER NOZZLE
FINAL NOZZLE
--,.-=-- ----------
CORE FLOW COMBINED
FLOW
' ............. CORE FLOW
Flgure 33. QCGAT Nacelle Duct Areas and Mean Mach Numbers.
0' o
WWt
~,
"
t-~ ---'l~----
Figure 34. QCGAT Fllght Nacelle With Sound Treated Panels Installed.
18
16
In 14
"0
ui 12 U)
0 ...J
10 z Q I- 8 ffi U)
~ 6
• Designed at Approach Power Setting • Effective Off-DeSign Performance
FAN INLET DUCT ATIENUATION SPECTRA
- 125 FACE SHEET FlOW RESISTANCE - 23 mks rayls
-=FACE SHEET FLOW RESISTANCE - a7Pc""'; I I
PROGRAM GOALS
~TAKEOFF ~APPROACH
:t±:±11~111 20
18
16
In 14 "0
CJ) CJ) 12 0 ..J
Z 10 Q I-a:
8 w CJ)
~ 6
4
2
0
Figure 35.
31 5 63 125 250 500 11< 2K 4K 8K 16K
FREQUENCY. Hz
FAN EXHAUST DUCT ATTENUATION SPECTRA
I I I I I I I I I UR -IS
t- OPEN AREA - 5"- -BACKNG DEPTH - 16 mm (063 101
t- FACE SHEET FlOW RESISTANCE - 12'pc mks rayls ~ \7 !"""J.c, f\ ~AKI~OFF
~\ 1/ 1\ I ~
/1 /%
3t5 6 3 1~ ~5 21 50 5( 10 11, 2K 4K 8K
FREQUENCY. Hz
PROGRAM GOALS
~TAKEOFF ~APPROACH
QCGA T Lockheed Fan Exhaust Noise Treatment Deslgn.
61
TABLE 2. ENGINE OPERATING PARAMETERS
Parameter
Power Settlllg
Altitude
Forward An Speed
Number of
Fan Blades Inlet GUide Vanes EXit Gu ide Vane s Support Struts
RatlO of Distance Separating Blade s fr om Vane s to AXial Blade Leng th
Fan Rotor Speed
Blade Tip Re1ahve Mach Number
Inlet Temperature
EXit Temperature
Inlet Pressure
Exit Pressure
Average Mach No.
Inlet Duct
Discharge Duct
62
De sign POlllt
Approach
112.8m (370 it)
91 knots
24 o
59 8
2.35
5400 rpm
0.5117
99.9 kPa (14. 5 psia)
1 07. 9 kP a (1 5. 7 p s ia)
0.2
0.22
Off-De.ngn Point
Takeoff
1105.2 m (3626 ft)
103 knots
24 a
59 8
2.35
9928 rpm
0.995
88. ';' kPa (12.87 pSia)
116. 5 kPa (16.89 pSia)
0.36
0.38
3.3.2 Acoustic ConsideratlOns
It has been recogn1zed for somehme that the fan mlet and d1scharge ducts of the engine nacelle offer ideal locahons for installatlOn of sound treatment mater1al to absorb much of the noise generated by the fan. Absorptive materials are particularly eff1cient 1n absorbing sound energy in the h1gh-frequency region where much of the acoustical power radiated by the fan is concentrated. In add1hon, sound treatment can be accomplished in the use of fl1ght-worthy materials that add httle weight to the aircraft. Finally, the theory and experlence of deslgnlng sound-treatment panels are sufficiently sophlsticated to accurately predlct the results that will be achieved from a particular
design. Consequently, sound-treatment panels for the QCGAT engine nacelle were mvestigated to determme the beneflt that would be derived from their incorporahon in the aircraft deslgn. The treated areas are deplcted in Flgure 34.
The acoushcal problem statement comprlsed a set of attenuation requirements, a descr1ption of relevant engine and nacelle geometry and operating parameters, and est1mates of the acoushcal conditlOns 1n the ducts. The attenuation requirements are shown superimposed on the inserhon-loss predlction graphs and are broken down into 1/3
octave spectra for approach and take-off conditions. (See Figure 35 ). These spectra are dominated by the requirements at blade
passage fundamental. The geometry of the alr passages was presented as scalable line drawings, and the operatmg condihons definmg approach and takeoff are summarlzed in Table 2.
The llliet duct was modeled as a slmple cylinder, 533 millimeter s (21 inches) III dlameter. Lengths of 1 radlus and 1. 25 radlUs were conSldered. The 1nner wall 1S cons1dered to have an In-place acoustlc lmpedance Z, where Z is a funchon of alr space depth, faclllg sheet throughflow reslstance, lnertance, mean grazlng flow, the sound pressure level, and sound pressure spectrum.
The d1scharge duct was modeled as a stra1ght annular duct havmg larger dlameter, 610 mm (24 In.), and lnner dlameter, 3 S~ mm (14 In.). The inner duct wall is nonabsor phve. The outer wall has a fmite acoustic impedance Z. Inlet flow 1S de sCrIbed by 1ts mean Mach number (-M) and dlscharge flow by its mean Mach number (+M). The convention of slgned Mach number s is pecuhar to acoustical analysls.
63
The theory of the transm1s sion of sound m ducts containmg flow leads to a general governing equatlon known as the convected wave equation. Its solutions are infmite in number, but only a small group represent the propagation of energy along the duct, the so-called propagatlng modes. The remainder (nonpropagatmg modes) represents pressure d1sturbances that cannot propagate energy even along a nonabsorptive duct. These pressure disturbances decay along the duct, usually very rapidly. In the case of c1rcular or annular ducts, each mode, except the zero-order mode, is represented by a rotating pressure pattern having sinusoidal lobes (m) and radial nodes (q) called spin11.ing modes. Whether or not the m, qth mode can propagate at any given frequency is precisely calculable, and the details of the noise source mechanism determine the modal content of
the sound in the duct. Aerodynamic n01se lS hkely to contaln all possible propagatmg modes. The actual energy dlstrlbutlOn among the modes 1S seldom, 1f e~er, known, but a reasonable assumption is equipartltion among the allowed lobe counts 0 to m and equlpartition of the energy in the mth lobe count among 1tS q rad1al mode s.
The pure tones generated by rotor blade-stator vane interactlon, are a very restr1cted set of "allowed" modes. As first shown by Tyler and Sofrin (R eference 2), a mode is "allowed" only if m = nB -kV, where n lS the harmonlc number of blade passage frequency, k is any
pos1~ive or negative integer or zero, B lS the rotor blade count, and V lS the stator vane count. Negative m slmply means rotatlon Opposlte to shaft rotatlon. It also is shown that for blade-tip circumferential" veloc1ties up to sonic, propagation can occur only if nB.>m. Thus, so long as V:> 2B, the fundamental blade passage tone due to blade-vane lnteractlon. 1S suppressed. Any such tone appreclably present must then be due to an aerodynamlc process such as blade choppmg of mlet distortlon. Thls acts analogously to a slngle inlet gUlde vane V = 1, makmg all propagatlng modes posslble. This was the modal d1Str1butlon prevlOusly ascrlbed to aerodynam1c noise.
Aerodynamlc noise was asslgned the classlc equlpartltlon of energy among the ill propagatmg lobe counts. Th1S, ln turn, was subdivlded equally among the q radial mode s for each value of rr.. where
the bars signlfy maXlmum poss1ble values of the lndlces m or q.
Pure tone fundamentals of the blade passage were formally inserted lnto the wave equatlon, even though the presence of no propagatlng allowed modes was assured by the criteria prevlously mentloned. It was assumed that the fundamental tones were ascribed to lnlet distortion. They were, therefore, represented by the same modal energy dlS-
64
irlbutlon used for the broadband noise. The second harmomc, a propagating pure tone, was consldered. Shll higher harmonics of blade passage frequency were disregarded for two reasons. Fust, the harmonlcs are above the audible range and, hence, no attenuahon reqUl.rements were 1dentif1ed. Secondly, except for certa1n rare OCCl1rences that are not perhnent to thlS des1gn, the harmonlcs are ln the high frequency range that attenuate rap1dly.
In the absence of deta1led boundary layer defimtlOn, plug flow (zero boundary layer th1ckness) was initially assumed for all four cond1tlons. It is well known that shear flow which has a negligible effect on d1Scharge duct attenuatlOn, and what effect 1S present is constructive as sheared flow tends to refract sound outward mto the liner (Reference 3).
In additlon, marked dlfferences 1n optimum 1mpedance and moderate d1fference s 1n attenuation pred1ctlons can occur for ind1v1dual mode s at fa1rly high Mach number s (Reference 4). However, the var1atlons 1n optimum lmpedance and attenuatlOn pred1ctions for the ensemble of mode s that const1tute an aerodynamic n01se are much less pronounced. Futhermore, the effects of shear are Oppos1te to the effects of red1str1buting the modal energy due to roll-off at the h1ghest order modes. It is quite hkely that such a roll-off 1S actually present. It 1S, therefore, concluded that for the aerodynamic type of modal d1str1butlOn, wh1ch m the case of the subject engine 1ncludes the blade-passage fundamental tone, the effects of shear are less than the poss1ble effects of the uncertalnty of modal energy distribution (Reference 5). For these reasons, the plug-flow assumphon 1S considered justlf1ed.
The duct-analysls computer program used by Lockheed (Reference 1) solves the chosen form of the convected wave equation by an lteratlve method which calculates inserhon loss (ratio of sound power in a treated duct to that in an untreated duct in dB) for any glven round-duct geometry, flow, and in-place wall 1mpedance for each propagatmg mode. These indlvidua1 mode solutions are assembled into an attenuation predichon for any chosen model energy d1stribution.
Rarely, however, will material that is available to the de slgner exhib1t the properhes necessary to ach1eve the optlmum design. Rather, it becomes the case of selectmg the material that comes closest to the requirement. In the QCGAT des1gn, two materials exhib1t properties closest to the desired parameters. The f1rst 1S a f1bermetal, a mesh
65
of staInless steel fIlaments rolled and compressed Into a porous sheet a.ld lald over a perforated sheet that offers the advantage of approach1ng the Opt1ITlUITl de slgn where the change 1n attenuatlOn due to change 1n flow res1stance 1S less cr1tIcal. F1gure 36 dep1cts the flow res1stance of th1S ITlater1al. The second ITlater1al, a pla1n perforated sheet hav1ng a sITlall percentage of open area, offers the advantage of proven des1gn and SlITlpl1C1ty of fabr1catIon. The flgure also shows the flow res1stance of
th1S ITlater lal.
For reasons of fhghtworthlnes s, the plaln perforated-face - sheet was selected for the lnlet and d1scharge duct sound-treatITlent panels. F1gure 36 shows that for the relatively low flows and sound pressure levels ex pected 1n the lnlet duct, a low-percentage open-area perforate w1ll be required. A 5 percent open-area perforate 1S consldered the
lower practlcal hITllt for a perforate -open area. Below this value, the rapldly lncreaslng inertance renders the tun1ng both narrow and capriclous, ITlanufactur1ng tolerance then becoITle cr1tIcal. W1th sITlall per centage open-area perforates, hole Slze becoITle s a hITl1t1ng factor. L1ITlltatlOns ln the perforator's art preclude holes whose dlaITleter IS less than ITletal th1ckness. The perforate used 1S ITlanufactured froITl a 0.636 ITlITl (0.025 In. ) thlCk 2024 aluITllnuITl panel wIth 0.838 ITlITl (0.033 In.) dlaITletnc holes on 2.85 ITlITl (0.112 In.) centers. ThIS desIgn YIelds a 6.8 per cent preas seITlbled open area wIth 12 hole s per square centIITler (79 holes per square lnch).
The core wlll be fabrlcated of Hexcel 5056 F-40, wlth a fOll thlckness of 0.066 ITlITl (0.0026 In.). Blockage of the holes by the fall IS expected to reduce the open area to 5 percent. The Hexcel whIch has 0.95 cells per square centlITleter (6. 11 cells per square inch) yields approxlITlately 12 hole s per cell.
For the OCGAT deSIgn, 1t was possIble to accoITlphsh the dual tumng at 2160 and 3971 Hz with a hner de pth of les s than 25.4 ITlITl (1 m.). The optlITlUITl values of reactance are 0.6pc for approach and
O. 7pcfor takeoff. The 1nertance of the selected Inlet and d1scharge faCIng sheet 1S known to be less than 4.5 x 10- 5 seconds. Thus, values of -0.8 pc for approach and O. 1 pC for takeoff can be achieved wIth an airspace of 16 ITlITl (0.63 In.). These values are suff1clently near opbITlurn •
Usmg these desIgn parameters, the attenuatlOn graphs are reentered to pred1ct the Insert10n loss of the panels for var10US treatITlent lengths. A treatITlent length of 1. 25 tlITleS the average mlet duct rad1us and 1. 5 t1ITleS the average d1scharge duct radIus were selected to Yleld an InsertlOn loss wlth1n 2dB of the attenuation goals. The predIcted fan 1n-let and dlscharge duct msert10n los se s are shown 1n FIgure 35.
66
FLOW RESISTANCE R = Ro + RI (u)
1000~~~~~~~ ~ ~: - __ 1 ___ 1_:=-- 1-_ ==~ = r-t1l1 £!' 200 --- -- - - - - ---11- - - - - - t Line 01 Constant
I • ~ /R __ 100 _
'1 70 _ - :: :: - 1::- ~ Ii - -.I: - -u :,0 -!~ r- ,,- - ~- fl. (:u ~ - -. ----- - - - - - --- . r - - ~. -r .. ~-"' <~- -".>- - - -~ )0 - - - - - - - - ... - • ~- r:-.,.,1 -:! - -I- --~ 20 - ~~, - ~ _____ li~ :::: :y - Ul.t'.:~ i--- - - - - -
I J~ ---- 0'" rc 0 II,...-?-
10 ==v ~ =:; ;.3 o~ ;) ~ -=~ =~~ ~ l~' -:!J n -. = = = = = -.: -." 7 =====- _ -\'2: f~ .::: _ ;; -:. : -~i.~ ~.::: -:;;;:- ~ ~ ~ : - _ - - == == ~ -u ~ _ _ \' _ - ~l.l'_ ~ ~7 - - - - --- -- -- - - - -
~ ) 'l'" ;, ~ ;.:: ~ -::L ~ = = -=- - :=;==: ==- = = = -f-Ii 2 -~~~,.\--A----------' -~~--
li ~ - :§~~= ===c== t 5 --- - -- - = =- - - =====- === == = -= -: --- -=== = = = = .. -------------------- - --------!: .J - - - - - - ---- -- - - - - -o 2 - -- - - - ----- - - - - --- -- - - '- -n:
• '1----.,.0-k2;:--"O~),--J"""'o fs '" 7 0\ ( 2 () {5 ( fIIl~I"r~) 10 """~ £0' !Irurn~/crn)
FlBERMETALS
TUE RELATION DETWj:EN Ro AND R, fOR flOERMETALS Rei Lockheed Report IILR 28254
FLOW RESISTANCE R = Ro + Rqu)
'DOC 70(, 50C
1~ 1 lOG -
1 70 50 •
~ 30
--
--~-~~~~~~~~~~~~=~~-~-2<1----J--+-I -J--I-H-H----'-- - - - - - - -1----'-- - --
! Ie I
0 !: f:~~~~§f~F:--~'-~~~~~~~~~~~~~tf~~~~~~~~~~t~~ - - - -'- LineR of Constant
-~ J 5 2
" n 7 ~
L 5 .. C .3
11----1-- -- -- - - - - - - - r:::-J----\--,- - - - - - - --- - - - - - _ - ---- O----I---I-l~-_I_I
_ ... -- i----- -1-'-
1~~~~~~~~~-rull~P' · ~ I -""" ----: - -.-== ::::::-;;;:;l _~ ~ - - ~ -c:::. = ~ - - -!!l2! f2 !.t> - -- _ ::z ,-= ~ =- - - 1-1- 1-= I-- ....... "" -,tf\-- ---
,---+--,-+~.<~.J.=.-I \'!I'le'~ -- - - - - - - J---J---J--I--J-.f-II-I 0 2 '"
.002 00] 005 007 01 02 0) 05 07 2 1 "' 5 6., 91 RII~lope) on "",IS 0/ !IrOIO~/cm]
PERFORATES
F 19ure 36. Sample From Lockheed's Matenals Handbook for Metal.
67
68
3.3.3 Structural Design
MechanIcal design of the flight nacelle Incorporates current stateof-the -art lightweight skln-On-fraITle structure and acoushc suppre SSlOn panels that are Integrated mto the air mtake duct and fan exit duct.
The te st nacelle retains flow channel contour s Idenhcal to those of the flight nacelle. The structural desIgn IS based on the reqUlreITlent to readIly change duct flow panels froITl acoustic suppre SSlOn panels to hardwall and air intake hps froITl conventlOnal fhght lips to bellITlouth and approach siITlulator. ThIS requireITlent resulted In departures In
structural desIgn froITl those used in the fhght nacelle desIgn.
Except for the reITlovable inlet lips which are of fiberglass construction, the nacelle's basic structure conSIsts of ITletal skin and fraITle.
The fan's ITlaln fraITle, which contams four ITlain ITlountmg pads, serves as the foundation for the nacelle structure of both the flight and test configurations.
The ITlixer nozzle, bolted to the turbine exit casing, is ldenhcal for both flight and te st confIgurations; however, the sheet metal of the test nozzle is slIghtly thicker to allow for local ITlaterial thinnmg that is anticIpated for one of a kInd stretch-forITl ITlanufacturlng technIque s.
DeSIgn details of the flight and test nacelles are described below.
3.3.3. 1 Flight Nacelle
The de SIgn profIle of the fhght nacelle shown in Figure 34 illustrates the baSIC Internal and external structure. Figure 37 IS a sectional VIew of the ITlain engine ITlounhng stahon.
BaSIC cOITlponents of the nacelle shown in FIgure 38 follows:
are as
o The nose cowl contains the air mtake and nacelle forward cowl and IS bolted to the engine inlet flange on the fan front fr aITle.
o The cowl tall section contains the fan duct outer wall and the rear contour of the nacelle and is bolted to the mner aft flange of the fan ITlain fraITle.
· ~ Q) .-i
:>
69
-J o
\ . .
BOTTOM PANEL
I I~I
~ SIDE
PANEL
EASY ACCESS PANEL
CORE COWL COWL TAIL
SECTION
-=::=;~;; ttf1=~=-
\ NOSE COWL
SECTION
BOTTOM PANEL
F 19ure 38. QCGA T Nacelle Components.
o
o
o
The core cowl forms the inner fan flowpath and provides a firewall over the core engine combustion chamber and fuel nozzles. The cowl WhlCh is cantilevered from the aft flange of the fan maln frame, also contains the walls of a service strut across the fan flow channel.
The mixer nozzle forms the core eXlt nozzle and is cantilevered from the rear flange of the turbine casing.
The removable access panels are fastened to the flxed structure of the nacelle as shown in Figure 38.
The englne is installed ln the airframe to a lTIain engine lTIounting yoke which plcks up two of the engllle' s main mountlllg pads on the fan's lTIain fralTIe (Flgure 37 ). A front lTIounting strut attached to the containment ring assembly offers only lateral support. The structure of the nacelle front cowl and ln1et consists baslcally of a1umlnum skln frame wlth an lntegra1 honeycomb noise-suppression panel that forms the inlet duct wall. The inlet lip WhlCh lS de -iced by hot air from the englne compressor has sufflCient thickness to resist hail damage and erosion. The rear bulkhead forward of the accessorles serves as a firewall and support frame. The llltermedlate bulkhead lS used only for external skin support.
The forward portion of the tall cowl IS an integrated honeycomb noise-suppression panel that forlTIs the outer wall of the fan flow channel. The cowl bolted to the fan malll frame flange supports the tall sectlOn that comprlses the outer wall of the fan exit duct and the external boat tail cowl.
The cowl is split into two halve s that are fastened together by bolts at the top and clamped at the acce ss strut fairing at the bottom. All external acce s s panels are of formed and welded alumlllum sheet and equipped wlth quick-release fastener s for ease of removal.
The welght of the nacelle, excluding engine mounts, is 106. 7 pounds. A breakdown of the weight, materla1 se1ectlOn, and nacelle center -of-graVIty are shown In Figure 3-;).
3.3.3. 1. 1 Fire Prevention and Compartment Venh1ation
The section of the core engine between the rear flange of the engine main frame and the mixer nozzle attachment flange, as shown in Flgure 40 lS the prlme fire-protected area because lt contains a hot
71
-J N
HEXCEL HONEYCOMB ACOUSTIC PANEL
AL 2024
COWL NOSE AL 6061
WT75(165)
NOTE
ACCESS DOORS (2) AL 6061
WT / (154)
L- C
" ]
I
I , r ,~=J
ALL WEIGHTS ARE Kg (Ib) ALL DIMENSIONS ARE cm (Inches)
Flgure 39.
CORE COWL & ACCESS STRUT
300 SERIES STAINLESS WT 42 (93)
HEXCEL HONEYCOMB ACOUSTIC PANEL
AL 2024
COWL TAIL SECTION AL 6061
WT 140 (309)
MIXER NOZZLE ASSEMBLY INCONEL 718
WT 453(100)
TOTAL NOMINAL WEIGHT 483 (106 7)
Nacelle Mater lals and Weights.
r~"J
-..J VJ
Ventilation Rate - 100 Changes/Minute Surface Velocities> 1ft/sec
(12) 187 IN DIA HOLES AT 45° (RAM TYPE)
\ : L.1---
! -~\ --:5:::-"" ~- '--..-~-
t-'1 +~
--------
205
6 RAM scoops 25x117 IN LINE WITH MIXER LOBES
770 l 780 745 725 470 780 - 25 IN THICK METAL CLAD INSULATION
(J M CERAFELT OR EQUNALENT) SURFACE TEMPERATURES (OF)
FOR ENGINE AT TAKEOFF POWER, SEA LEVEL STATIC
Figure 40. Nacelle Fire PreventlOn/Compartment Ventllabon.
engine casmg and fuel hnes. For the critical hot-day stahc operation at takeoff power, surface temperatures over the combustion chamber casing are below 7800 F, and surface temperature s over the turbme caslng and rear portion of the mixer nozzle are kept below 78 OOF by thick metal-clad insulatlOn. The steel-core cowl forms a fire -re sistant wall around this porhon of the core engine caslng. Venhlative alrflow for the fan lS provided lnto the core zone by means of 6 scoop inlets located at the forward part of the core cowl and through 3/16-mch dlameter holes through the main frame struts. The airflow reenter s the fan air stream through the annulus formed between the rear edge of the core cowl a_nd the outer surface of the mixer nozzle. The ventilative anflow circulates circumferentially through the core cowl compartment at a rate of 100 air changes per minute; the lowe st surface velocity in the zone is greater than 1 foot per second. The se criteria are known to be safe in the prevention of fire. ThlS venhlahve alrflow reduces the ambient temperature m the core zone and represses 19nihon of leaklng flammables upon hot surfaces.
Consldermg the unlikely occurrence of fire in the core zone, a flye detectmg and an exhngulshlng system with airframe-mounted fire-suppresslOn flUld would be provlded.
3.3.3.2 Te st Nacelle
The test nacelle configurahon is shown m Flgures 41 and 42. All lnternal ducting contour s from the au In}et to the final exit plane of the nozzle are idenhcal to those of the flight nacelle.
74
Except for two short front and rear boat-tall cowl portions, external skins are excluded so as to reduce costs and improve accessibihty without affechng the attamment of QCGAT goals.
The deslgn permits adaptation h the test nacelle of three distinctly dlfferent air lnlet conflgurations -- one bellmouth, one re-gular inlet designed for minlmum loss at crUlse conditlOn, and one inlet designed to statically simulate fan inflow condltions corresponding to an inflight approach conditlOn (Flgures 43 and 4~. The acoustic suppresslon panels ln the an lntake and fan ducts are of the same constructlon as those used ln the fhght nacelle but are de slgned to to facll1tate replacement wlth hardwall panels.
-.J 111
,'" , ~--=-=== ----f- ~
I ~ " i
I
CORE COWL MIXER NOZZLE
---------
t-"
, I 'I ' ,.-
. I II i I I It- iEft - ! iii IT -j ~ - - -~ *= I I
. I ,'I j " ' '
I
I ~-- ~'-----~~:::
INLET COWL
i j' ~ - ____ __ _----:~: I -:::~----L ' ;- ~ ----------- ---
r---' --- - - ------- ----, .,------
i'
FIREWALL
Flgure 41. Test Nacelle Assembly.
TAIL COWL AND AFT STRUCTURE
\ SERVICE STRUT
Figure 42. QCGAT Flight Nacelle Simulation.
--:J --:J
FLIGHT NACELLE
d ~rrQ
~ ~
SYMETRICAL BELLMOUTH
Flgure 43. Replaceable Inlet Lips.
APPROACH SIMULATOR
-J co
ELL I I
Figure 44. Bellmouth and Flight Nacelle Inlets.
c I
Since ease of replacement of fan duct panels precludes mtegrating the panels Into the rear cowl support structure, the rear duct section IS supported by a frame comprised of barrel rmgframes and longerons. ThlS structure IS canhlevered from the rear flange of the fan main
frame.
The core cowl and mlxer nozzle and thelr attachment are similar
to those for the fhght nacelle deslgn.
The front cowl lips and air mlet duchng are supported by a barrel structure lhat IS canhlvered from the engine mlet flange. The acoustic suppression panel in the Inlet can be removed and replaced with a hardwall panel, and the mlet hps can be replaced without disconnectmg the barrel support structure. Other detalls are dlscussed m Paragraph 5.3.3.
3.4. MIXER NOZZLE DESIGN
The Lycoming QCGA T exhaust nozzle system shown in Flgure 4S comprlses a fan duct, a multilobe mixer nozzle, and a mixlmg chamber Iflnal nozzle. This type of exhaust mlxing system waR chosen because of ItS propulslve efflclency and reduced noise - beneflts. Mulhlobe mlxer nozzles such as that shown In Flgure 45 have been reported been reported (Reference 8 ) to yield a conslderable amount of noise suppresslOn when compared wlth the more conventlonal split-flow nozzles. The suppresslon IS believed to result from reduced jet turbulence levels and a reduction In the mean-relative jet-veloclty gradlents (Reference 2).
A parametrlc study was evaluated to optlmize the mlxer nozzle wlth conslderatlons for reduced nOlse emlS Slons and Improved crUlse fuel economy. These detalls are documented In Reference 8'. The resultlng deslgn 1S shown 1n F1gures 46 and 47. The mlxer nozzle 1nstallatlOn on the englne IS shown in F1gure 48.
Shaker test performed on the nozzle and turbofan engme straln gage testing Indicated satisfactory dynamic characterlstlcs.
79
00 o
Figure 45.
MIXING CHAMBER/FINAL NOZZLE
MIXER NOZZLE
FAN DUCT
QCGAT Exhaust Nozzle System..
00 ~
V)
=> .... CI
m (i n)
.251= 10
8 ~ SUPPORT RODS AS FAR AFT AS POSSIBLE
61- J:t=ll I E- -_~ 2 ~ afF:
9
~ .10
.05
CENTERBODY END OPTIONAL FULL RADIUS OR CUT AT (27.0)
30 (i n.)
I I I I I I I I I I .30 .35 .40 .45 .50 .55 .60 .65 .70 .75 m
ENGINE STATION
Flgure 46. Core Engine MlXer Nozzle Geom.etric Definltlon.
,.
TURBINE EXIT CASE
FLANGE (REF)
6.35 mm (.25 in.) CLEARANCE FROM
TURBINE CASE
ATTACHMENT BOLTS
SUPPORT ROD CENTER BODY
DAISY PETAL NOZZLE
Figure 47. Lycoming QCGAT Mixer Nozzle Design.
82
6 LOBES
Figure 48. Mixer Nozzle Installation.
4.0 SUBSYSTEM TEST RESULTS
As part of the NASA/QCGA T contract, component test rigs were used in the design development of the fan and combustor modules. Additionally, laboratory tests using laser-holography techniques were used to evaluate the frequency response and mode shapes encountered on the fan blade and of the reduction gear assembly.
4.1 FAN BLADE ANALYSIS
Aerodynamic design of the QCGAT fan blade is the same as that used in the successful ALF-S02 turbofan program. To determine and evaluate blade natural frequencies and mode shapes, a holographic interferometry analysis was conducted.
4.1.1 Test Program
Figure 49 shows the general test setup used. Englne level hardware was used as test items.
A spec.Lal fudure was made to allow clampmg of a fm.Lshed machined fan blade at the dovetail root and mounting the fl.Xture to the plezoelectr.Lc v.Lbrat.LOn exciter. The des.Lgn of this fl.Xture and the blade attached to the fl.Xture are deplcted .Ln F .Lgure 59. A real t.Lme hologram was made of the blade at rest and at predommant modes up to 5000 Hz. All mode shapes were ldentifled, photographed, and documented.
4.1.2 Fan Blade Test Results
The natural frequencies of the fan blade as determined experimentally, along with their as sociated mode shapes, are identified in Figures 51 and 52. A diagram is presented ill Figure 53. Measured frequencIes are identified at the ordinate. Frequency characteristics for other than "Zero" rpm condition are based on analytical predictions. The experimental analysis verified analytical prediction of second order of the first bending mode as a potential blade excitation.
4.2 LOW -PRESSURE TURBINE BLADE ANALYSIS
A similar procedure was used for the QCGAT power turbine blade. Since the blade and disc are integral, the test required cutting a blade with the attached disc section from a final tip-ground turbine wheel
84
00 U1
llOLOGRAM
FRONT SURFACE MIRROR
LENS PINIIOLE SPATIAL FILTER
, ODJECT DRAM
~----~ ----
PLATE 1I0I.DER------"
FRON'r SURFACE
VARIABLE BEAM SPLIT'rER
NE LASER MW 63281'.
PIEZQELECTRIC VIBRATION EXCITER AND MOUNTING FIXTURE
Figure 49. General Test Setup Holographic Analysis.
DOVETAIL ROOT AND TAB
FAN BLADE
~ DOVETAIL ROOT HOLDING FIXTURE
BASE SHROUD
SHAKER HEAD CLAMPING BOLTS
CLAMPING --~-- I SH1l\1S
I I
I ALIGNMENT DOWELS
Figure 50. Fan Blade Root Fixture.
86
1 s t Bending Mode (214 Hz)
2nd Bending Mode (538 Hz)
1st Torsional Mode (110 Hz)
Figure 51. Fan Blade Natural Frequencies and Associated Mode Shapes-First and Second Bending Mode, First Torsional Mode.
00 00
1804 Hz 3rd Bending
3140 Hz 4th Bending
4527 Hz 2nd Torsional
Figure 52. Fan Blade Natural Frequencies and Associated Mode Shapes-Second Torsional Mode, Third and Fourth Bending Mode.
N I ->a z w ::l o w II: u. w o « ....J co z « u.
QCGAT FAN BLADE
IDENTIFIABLE MODES ARE INDICATED
5000
4000
4th BENDING
3000
2nd TORSION
.. . ..."".- ---- --1st BENDING'- ___ ---:_----
o~--~~------~~--~~------~----~ __ --~~--o 2000 12,000
FAN SPEED (RPM)
Figure 53. Fan Blade Excitation Dlagram.
II: W o II: o Z o ~ « I-a x w
89
assembly. Thls segm.ent was then bolted dIrectly to the vibration exclter as shown m Figure 54. Real-tim.e hologram.s were obtained up to 24,000 Hz covering the turbine blade frequency spectrum.
4.2.1 Power Turbine Blade Results
The typlcal natural frequencIes and associated mode shapes of the power turbine blade are identified in Figure 55. Seve ... al m.odes could be identified in the frequency range under consideration (0 to 24,000 Hz). A blade-frequency dIagram of the power turbine blade is shown in FIgure !i6 for amblent test temperature condItions and Figure 57 for l250 0 F (engine operatmg conditIon).
4.3 RING GEAR FREQUENCY ANALYSIS
A siml1ar setup using the holographic mterferometry techniques
was used on the rlng gear wlth the exceptlon that dlfferent fIxtures were requlred. The rlng gear was bolted to the plezoelectrlc vlbratIon exclter by means of a mounhng adapter slzed to lnterface wlth the external sphne and lnternal bearlng shoulder at the shank end of the gear. Thls clamplng arrangement (Flgure 58) conslder s a slmply supported constralnt at the bearing locatIon and a free-free condltIon at the rlng gear's open end.
A real-time hologram was m.ade with the ring gear at rest, i.e., with no input excitationo The gear was excited axially (sine-wave excitation) at varying frequencies up to 24,000 Hz (hmit of present exciter setup), and then the predominant natural frequencies were deterrrllned and recorded. A tim.e-average hologram was taken at each of these recorded frequencies to identify the respective mode shapes. Each hologram. was then docum.ented by photograph.
4.3.1 Ring Gear Test Results
Many diametrical and circumferential modes can be encountered over the frequency range expenenced by the rlng gear. Example of several m.odes are shown in Figure 59 showlng front and ~ft Vlews.
90
Figure 54. Power Turbine Segment Mounted to Shaker.
91
92
MODE 3 8086 HZ MODE 5 11194 HZ
MODE 6 13175 HZ MODE 10 21734 HZ
Figure 55. Power Turbine Natural Frequencies and Associated Mode Shape s.
N I ~
> () Z w ::> a w a: LL w o <{ ....J (D
28~
24 ~
8~
FREQUENCIES DETERMINED
BY HOLOGRAPHY @ R T
41
139 I, II
II II
II I II ~I I
II
o 0 100 200 300 I
400
POWER TURBINE SPEED - RPM x 102
I 500
Figure 56. Power Turbine Blade Excltation DiagraIn at AInbient TeInperature.
93
94
N I ~
>() Z w ::J o w a: u. w o « -l a:l
FREQUENCY DETERMINED BY
HOLOGRAPHY ADJUSTED TO 1250 0 F
41
a: w Cl a: o z o I« I-C3 X W
3 I ~_ Oa~S -~~\'f...€ __ -I -~?-\,.: v
~-.",... \'<' ---- S 4 - .. - "---S,R\Ji
V ~-.",...-~~-~ ...... ----------.. ~ ...... . . .""", ......... _ ...... -'-----::..--- I ......... ~ .. -. , "j" ~ ~ I I
o 0 100 200 300 400 500
POWER TURBINE SPEED, RPM x 102
Flgure 57. Power Turbme Blade Excitation Dlagram at l250 oF.
6438 HZ. (2n;"'3C) LOOKING FWD LOOKING AFT
Figure 58. Ring Gear Clamping Arrangement and Natural Frequencies and Associated Mode Shapes.
95
96
18850 HZ (3D-7C) LOOKING FWD LOOKING AFT
LOOKING AFT
Figure 59. Ring Gear Natural Frequencies and Associated Mode Shapes.
4.4 FLOW DNIDER AND FUEL MANIFOLD SYSTEM ANALYSIS
The purpose of this component laboratory test was to determine the effect of usage on the QCGAT flow divider /fuel manifold assembly. This assembly consists of two manifold segments, each containing four ALF 502 airblast fuel injectors. The two manifold segments are made up of LTS 101 manifold bosses and tubing, except that the flow divider boss has been changed to accept the airblast flow divider.
The test was defined at 500 cycles with each cycle providing one step excursion from take-off fuel flow to below idle and one transient excursion to take -off and shut off. This series provides maximum wear and fatigue testing for the time involved. Results show the effect of wear and spring relaxation on plunger leakage, stickage, hysteresis, and flow schedule with either small-incremental or full-range transient flow changes. Injector spray quality was observed and monitored throughout the test cycles.
Flow divider performance, in terms of flow schedule, reference port leakage, and hysteresis was not adversely affected by this test. The fuel injectors also performed satisfactorily with excellent spray quality.
4.5 FAN COMPONENT TEST RIG
The purpose of the fan module subsystem test was to provide measurements of aerodynamic performance necessary for successful matching to the core engine and power turbine modules and to demonstrate mechanical integrity of the fan component, including satisfactory vib ra tion s tre s s levels.
A eros s -sectional view of the test rig is shown in Figure 60. All englne level hardware are used wlth the exceptlOn that the fan rotor was a dlrect drlve, ln heu of the reductlon gearlng, wlth the test facll1ty. Measurement plane s for detalled stage -performance evaluatlon were deflned as deplcted ln Flgure 61, and a breakdown of the lnstrumentatlOn lnstalled lS glven ln Flgure 62.
The inlet section to the test rig consists of an airflow measurement bellmouth, optlonal distortlon screen holder, and a Slx-strut distortion instrumentation housing, which also carries a freon-cooled slip ring for strain-gage testing. The rig bellmouth serves as a secondary air flow measurement and is calibrated with an ASME nozzle located at the entrance of the plenum chamber. Figure 63 is a schematic of the
97
-.0 00
II RfLOI
L[---=-= -=,_L I \I~l ~'
o B
Flgure 60.
o B 12 01 MlIM fRIME
QCGAT Fan Module Test Rig.
BIPISS VIlVE
- - - I - I
-.D -.D
RIG BE llNOUTH AND INLET DUCllNG
Note
1--_0
•
5DIA--l -FAN MODULE
"
(SURV[Y)
@ @ @ I?: ,I ~ _ I.~, r 1.1' (
I EXIIAUST DiffUSER
EXIIAUST 01 HUSER
Ctrcun1ferentwl Ortpnt.ltl()ll of extL plane instrumentatIon (Ref. 1) 1S trep-stream w1th respect to bypaf>s struts (12) and core channel struts (4).
Figure 61. Fan Module Instrwnentation Planes - Nomenclature.
1
I I
,
100
STATION DES I GNA TI ON LOCATION
-4 PLENUM NOZZLE
0 PLENU~l CHM1BER
0.6 RIG BELLMOUTH/DUCT
0.8 INLET DISTORTION HOUSING
12.00 I ROTOR I NL ET PLAN E 12.01 ROTOR EXIT PLANE 12.015
I ROTOR EXIT SURVEY PLANE
12.02 BYPASS STATOR INLET PLANE
12.04 BYPASS STATOR EXIT PLANE
2.05 EGV I A I INLET (SUPERCHARGER CHANNEL)
2.06 EGV I B I INLET 2.07 EGV I B I EXIT
2.08 ENGINE CORRELA7ION PLANE -SUFERCHARGER CHANNEL
2.1 SUPERCHARGER EX IT PLAN E
13.00 BYPASS EXIT PLANE
13.00 BYPASS EX IT -ENGINE CORRELATION
NOTE' (a) = OUTER WALL (c) = INNER WALL
INSTRUMENT
4 PT 8 Ps . 8 PT
12 TT 4 PT 8 Ps
4 TT (ENGINE ONLY)
I 72 PT (DISTN ONLY) 6 Ps (a)
4 Ps (a)
4 Ps (a)
I 3 RAD ACTUATORS
4 Ps (a) 4 Os (c)
I 4 Ps (a) 4 Ps (c)
4 Ps ( a ) 4 Ps ( c)
4 Ps ( a ) 4 Ps ( c)
4 Ps ( a ) 4 Ps ( a ) 4 Ps ( c )
2 PT 2 TT 2 Ps ( a )
40 PT 40 TT
4 Ps ( a) 4 Ps ( c )
80 PT 80 T~
I ( a ) 8 Ps
8 Ps (c)
4 PT (INiEG) 4 TT 4 Ps (a)
F 19ure 62. Aerodynamic Performance Instrumentation.
I I
I
I
I
I I
,.... o ,....
I (PLENUM)
AIRFLOW.
RIG DELLMOUTH
I... I.36DIA. __ I
I
DISTORTION SCREEN WITH 15"!o "LOCKAGE PLATE
(OPT. ) --- 0.86 DIA. ------~
DISTORTION INSTRUMENTATION IDUSING
Figure 63.
PLENUM CII1\HBr;R
Fan Module Inlet Instrumentation Planes and Distortion Screen Position.
inlet duct and plenum system. Downstream of the main fan support frame are low-loss diffusing ducts to exhaust the bypass and supercharger flows to ambient pressure. Both exhaust channels are independently throttleable to permit selection of any desired component loading thru cylindrical, actuator-driven exhaust valves 0 Loadings over the flow range from choke to stall can be evaluated. The QCGAT fan rig cell installation is shown in Figure 64.
4.5.1 Fan Component Test Results
The development test program that was conducted consisted of three phases: steady-state base performance, dIagnostic, and distorted inlet. The baseline performance for the bypass and supercharger is given in Figures 65 and 66 0 Both the bypass and supercharger were mapped from 50 to 105 percent of design referred speed. The overall test results are compared with design goals as tabulated below:
Design Test
Design Test
Bypass Performance (Sea Level Static)
Referred Flow Pressure Ratio Kg/sec(lb/sec)
1.380 1.380
33.70 (74.0) 33 0 75 (74 0 1)
Polytropic Efficiency
0.870 0.870
Referred Speed (rpm)
11,200 11,200
Supercharger Performance (Sea Level Static)
Referred Flow Polytropic Referred Pressure Ratio Kg/sec (lb/sec) Efifciency Speed (rpm)
1.350 3 0 63 (8.00) 0.850 11 ,200 1.308 3.63 (8.00) o. 715~:< 11,200
At Peak Efflclency
1.324 3.41 (7.50) 0.732 11,200
~:< The super charger was rede slgned and re sulted In a 8-polnt g aln In efflclency based on englne performance data.
The bypass component met or exceeded all design goals; whereas, the supercharger performance was lower than the desired goal. Bypass and supercharger exit profiles are shown in Figures 67 and 68.. The bypass channel exit distributions of total pressure, total temperature, and polytropic efficiency are compared with the design
102
BLADE STRAIN GAGE LEADS
ROTOR CASING
Figure 64. QCGAT Fan Module - Testway Installation.
INLET PLENUM
TEST - 01
100% N/f7i= 11200 RPM
1 7 .ro ~------~-------r------~--------------~------~ >-o ~
1 6 !d >- - .80 -+, ---f--~-I----~-:---t'--~--"'------I ~v,......;---r------t O~
I
ott Co LoU
~§ I 1 5 .70-+------+------+-------+------r~------r------1
1 4
1 3
1 2
1 1
1 0 20
I 10
104
SYMBOLS
50% -<) 60% -G 70% -~ 80% -0 9~o -0
100% -0 105% -X
30
105%
-----+-------+-----+--r-------4-tt------+---:-DES1GN POINT
~ 1
60% I 50%
I 40 50 60 90
I I 15 20 25 30 35 Kg/sec
BYPASS REFERRED AIRFLOW - (WI1ija)o (Ibm/sec)
Note Dashed lines indicate engine operating lines for sea-level static (MN=O ) and altitude cruise (MN=O 6) conditions
Figure 65. QCGAT Fan Module Baseline Test - Bypass Overall Performance Map
o Ia.. ...... ~
N
Ia..
o ~ « a: w a: => en en w a: a.. -J « Io I-
TEST - 01
100% NV(j=11200 RPM
17~----~------;------;------'------;------~-----r
>-o .74 'Q.
,c I
1 6 ~ >- -.72r---.-,..:------~t_~r-~r_-~__1--.;l"'~4~--__"'1 O~ L / "" w 1 >= V .70 i;) ~_ I Ot.: Q.. W
1 5 1----.68 1------4l--if-"""'---i-=~-__+---__:---_1
DESIGN
1 4 1-----~---!---~r----J--+----...,I-:/----+71 POINj SYMBOLS.
50% -0 / 60% -~ PREDICTED / 70% -6 STALL LINE / BOOS -0 90%-0
1 3
95"S -C] 100% -0 105% -Q ,
12~----~------t---~/~t--e~~~~~j-------~----4 /
/
7rFto
10L--____ ~ ______ ~ ____ ~ ______ ~ ____ ~ ______ ~ ____ ~
2 3 4 5 6 7 8 9 (Ibm/sec)
II 1 2 3 4 Kg/sec
SUPERCHARGER REFERRED AIRFLOW - (w..rel a) 0
Figure 66. QCGAT Fan Module Baseline Test - Supercharger Overall Performance Map
105
;-;
9 ~
>-.)
C U .... ..... ,
::2 C
>-
"" £'
c I , " ,-,
,-~ ...
" <
.... ~
< l..
;:;
-< -C
'g - f -, ~ ... [
'-'
< c::
... ;;; v ..: ... CI.
< '-0 ~
106
I 00
.90
.80
70
bO 0
Test - 01
'rest Condltlons 1" / ... /! = 100'"
PO:"YTROPIC E:rnCIE'-:C,
20
I I
KE"l :;; -PI ~ 100,,02 - PR = 1 42<: O-Pt Ii 100601 - PR = I 357
40 60
" BYPASS STREAMLIl\E
I BIP MAP I I
I
80 IO:!
TOTA:' .. EMPERA Tt: R=-__ ....... ___ +-__ ...;.. ___ .:-__ ......i._..",..,,.....:::::::....::"----ii R.ATIO
I
1.11- • I
~I ~-·--r_~~---~~~~~r 1.10 r:'
o 20 4C 00 Joe ,), BYPASS STREAMLI'-;E
42
20 60 80 100
O.W. o I. W.
40
t;, BYP'>SS STREAMLI!,\E
Figure 67. Bypass Exit Plane Profile Data
c
=
...l < -C
s < c::
... < fa f-
li
C
SiC MAP
Test -01 POlnt' 100602 - PR = 1 311
Test Cond,t,ons
~ 'ESIGN ponn ~ P~ = 1 3;
90
.80 t-
.70
I 60 r 50
o
N/,}; .100"',
POLYTROPIC EFFICIENCY
-1--1-7- --I-- -~-DESIGN EFFICIENC '" DISTRIBUTION
~ ~ 1:1
'1:1. 1,.---1.:1" -20 40 00
0/, SUPERCHARGER STREAMLINE
TOTAL TEMPERATURE
1-- 1----'--.,.
~- I'r=
8u 100
1 12 RATIO r-------~---------~-----~--------~-----~------~----__ ------___
:.::~~-l ~+-E-I-~k~L~k~=-t=--h== I o 20 40 bO 80 100
0/, SUPERCHARGER STREAMLINE
- TOTAL PRESSURE RATIO
1.30 ~ - ---- -- -.- _.- -.--- --~&:>-- &I
.. ~
t- V ./
V ~
t-' -
1.34
32
1 30
1 28 - ...
1 26 o 20 40 60 80
1. w. 0/0 SUPERCHARGER STREAMLDlE
Figure 68. Supercharger Exit Plane Profile Data
-1\
7
100
O.w.
107
profiles. The blade pressure levels in the 60 to 80-percent streamline portion are greater than design thereby Y1elding the higher than design efficiency.
This pressure profile trend was later corroborated wlth rotorexit survey measurements. The same profiles for the supercharger show a substantial deficit in total pressure from the hub to m1d-stream. Follow-on tests showed that the observed efficiency deficit is largely associated with the total pressure loss through the supercharger vane assen+b1y and _duc:;t syqte-r:n.
The two major factors influencing the supercharger total presure ratio are:
1. Geometry of the fan blade hub section velocity triangles 1S such that relative flow is turned by the rotor to the axial directlOn. A decrease in flow causes a small change in work output. The resulting flat map characteristic yields little or no pressure change with loading.
2. The splitter was placed relatively far aft of the rotor; this coupled with a high bypass ratio design allows air flow to divert to the bypass channel as loading is increased
in the supercharger.
The second test phase l.solated the supercharger perfor:rnance loss to be attributed to the vane assembly and interconnecting duct. The effl.ciency deficit was a result of sensl.tivlty to pressure loss at 10w_ pressure levels. Subsequently, the supercharger vane assembly was redesigned to :rnlnlmlze the se los se s.
4.5.2 D1storted Inlet Test Results
The response of a turbofan to inlet distortion is of prime lmportance from the viewpol.nt of aerodynamic performance and mechanical integrity of the blades. Sl.gmficant distortions frequently occur in aircraft lnstallations as a result of intake flow separation induced either by crosswinds or high angles of attack.
The bypass map characteristic is generally flat once the peak pressure ratio has been reached; moderate amounts of circu:rnferential distortion :may cause the stall1ine to deteriorate and simultaneously ral.se the engine operating line.
108
In order to produce a crosswind type of distortion artificially I a crescent-shaped solid plate was placed about 1.4 diameters upstream of the rotor. Area blockage for this plate was 15 percent. Figure 69 shows the relationship of this plate to the channel and also the resulting distortion pattern which was obtained at 90 percent speed. The Lycoming distortion index definition is defined as follows:
Distortion Index
Dr =(PT mean - PT low mean) [K~J PT mean
where PT mean = avgo total pressure at the measurement plane based on an area weighted average.
PT low mean
Kp
Kp
M
M
PT low min
E
E
R
R
ALhub
= avg. total pressure area-averaged over all regions where PTC: PT mean
= Factor accounting for profile and extent of distortion
= tiMER
= Magnitude of the peak distortion relative the average in the depressed region.
= 6 0 0 PT MEAN - PT low mean
PT mean - PT low min
= Minimum total pres sure level
= Extent of distorted region
=
= Total annulus area
= Radial distortion sensitivity factor
= The area extent of low pressure regions which fall in the inner (hub) 50% annulus area.
109
110
970
D.1. - 0.057
CONTOUR INTERVAL - • 005 I SIc MAP
~ PTMAX
:: 13.53
= .975 Test Conoltlons
90% N/../T
Figure 69. Area Blockage (15 Percent) Inlet Distortion Plate with Resultant Total Pressure Contours
U sing the Lycoming D.1. as a descriptor to quantify the level of distortlon, the QCGAT fan was tested up to an index of 0.06. Since this level is quite severe and well above limlts generally set for satisfactory turbofan performance, the testing was not continued beyond 90 percent speed. A plot of distortion index versus total airflow is shown on Figure 70.
The effect upon overall performance of the combined radial/ circumferential distortion pattern is shown on Figure 71.
At 80 percent speed, the stall margin (S .M.) for the bypass is reduced from 13 to 9.0 percent.
Percentage S.M. = [1 - (W/PR) stall ] x 100 (W /PR) op. line
The level of distortion at the 80 and 90 percent speed stall point wa sO. 03 D. 1.
Peak efficiency at 80 and 90 percent speed decreased approximately 4 to 5 points.
Blade stresses remained at acceptable levels throughout the entlre test.
In summary, the QCGAT rotor demonstrated satisfactory aerodynamic performance and excellent mechanical performance under inlet distortion conditions that are representative or in excess of those found in typical turbofan installations.
III
.....;
t:l
x '-'l t:l z ..... z 0 ..... f-o ~ 0 f-o U) ..... t:l
112
070
060 Key I .-~O% N/V"'G' IT ~-60% .-70%
050
040
.-80"fo II
.-90% _I .lIt -;
/ •
030 ~I
V 020
.( /
V" 010
/ L----
V -o
70 20 30 60 80 (Ibm/sec) 10 40 50
o 10 20 30 Kg/sec
REFERRED TOTAL AIRFLOW- [W-;,e ]08
Figure 70. Distortion Index as a Function of Total Referred Airflow
Z
I-a.. ....... 0 0 M ,....
I-a..
0 i= « a: w a: ::J en en W a: a.. ....J « t-O t-
DARK SYMBOL OVERLAY: TEST - 01
100% N/17i = 11200 RPM (INLET DISTORTION)
1 7 90 >--J
0 Q,.
c-
1 6 ~>--.80 c..u Oz C=:w ~->-u ..... -ott Q,. w
1 5 .70
~X 14 SYMBOLS·
50% -0 60% -~ 70% -~ 80% -0
1 3 90% -0 100"1c. -0 105% -X DISTORTED INLET
STALL LINE
1 2
1 1 r-----~------_r--~~~------~----~------~------~
~~/~ 60"1c.
10_~ ____ ~ _____ 1~ ____ ~ ____ ~ ____ ~ ____ ~ ____ ~ 20 30 40 50 60 70 80 Ibm/sec
10 15 20 25 30
BYPASS REFERRED AIRFLOW - (wriila)
Figure 71. Bypass Overall Perforrn.ance with Inlet Distortion
35 Kg/sec
113
4.6 COMBUSTOR MODULE
The Lycoming LTS 101 combustor has demonstrated exceptlonally low emissions by compartson with conventional combustors of similar size. Several design changes were required to adapt the combustor to the QCGAT turbofan design. A new outer curl with improved cooling was designed, and the liner-to-curl seals were redesIgned because of the higher pressure and temperature requirements of a fan engine design. Airblast fuel injectors from the ALF 502 fan engine were installed, and a new flow divider with the same relative flow schedule as the ALF 502 was desIgned and tested 0 A comparison of the basic LTS 101 combustor and the QCGAT development confIguration are shown schematically in Figure 72.
During the lterations that optimize an engine performance cycle, continuous attention is required to avoid adverse impact on emissions characterIstlcs. A summary of the primary causes for emissions in conjunction with engine parameters that have a beneficial influence are as follows:
Emissions
Unburned Hydrocarbons, UHC
Carbon MonoxIde, CO
OXIdes of Nitrogen, NOx
Smoke
Cause
Combus tion Inefficiency
Inadequate Residence Time, Temperature, Efficiency
High Residence Time/Temperature
Local Rich Zones
Unburned hydrocarbons and carbon monoxide emissIons are primarily a reflection of poor combustor efflciency at idle. Low combustor inlet temperature at idle aggravates the carbon monoxide emissions. To reduce these two constituents, one would strive for very high combustor efficiency at idle, combined with elevated combustor inlet temperature. To achieve the higher inlet temperature, a compressor with poor efficiency at low speed is desired. Whereas idle conditions have the primary influence on UHC and CO, take-off conditions predominate in the creation of NOx • Generally, the higher the combustor inlet temperature at take -off the more difficult the problem is with NOx • An lmportant axiom is that NOx and CO can usually be traded through combustor design modification. Either emission can be improved at the
114
LTS 101 COMBUSTOR
AIR ENTRY
QCGAT COMBUSTOR ASSEMBLY
Figure 72. Corn.parison of the LTSIOI Corn.bustor with the QCGA T Corn.bus tor
115
expense of the other to achieve the desired combination.
The NOx emission goal was considered the most difficult to achieve. Because these emissions increase with pressure ratio, the 10:1 pressure ratio was a main consideration for the QCGAT compressor deslgn. The hlgh-bypass ratio also favors lower emisslOns for a given thrust rating.
The final combustor configuration culminated in a design that incorporated the optimum cycle characteristics, along with a unique vortex recirculation pattern that results in a lower rate of NOx increase with increasing combustor inlet temperature than for conventional combustors. This characteristic permits selection of high combustor efficiencies at idle with resulting low UHC and CO values without exceeding NOx values at high power sethngs. TypIcal combustion flow pattern and air distribution are shown in Figure 73.
4.6.1 Test Rig and Facilihes
Limited development effort was required to meet the emission goals and obtam a good turbine inlet temperature distribution, while maintaining adequate liner life.
The combustor test rig is shown schematically in Figure 74; a view of a cell installation IS included. Instrumentation locations that can also be seen include pressure and temperature at the inlet and exit planes. A rotating drum arrangement was used in the exit plane as a traversing mechanism to obtain temperature data.
Airflow was supplied at engine operating pressures and temperatures by the facility compressors. Airflow measurement was accomplished with a standard ASME orifice arrangement. Fuel flow was measured WIth turbine flow meters. Exhaust emissions analysis equipment that was used complied WIth EPA Standard 40 CFR, Part 87 and was used to calculate combustion efficiency, as well as measure unburned hydrocarbon, carbon diOXIde, oxides of nitrogen, and carbon monoxide. Carbon balance calculations were made and compared with fuel and air measurements to verify a representative sample.
4.6.2 Test Sequence
The initial test was conducted with an existing LTS 10lliner to determine the optimal injector immersion and the flow-divider spht. Also, the effects of combustor airbleed at idle and liner wall temperature
116
..... ..... -.J
PRIMARY
FUEL INJECTOR
COMBUSTOR IIOUSING
I
PRIMARY
VORTEX CIRCULATION
PATTERN
HORSESHOE-SHAPED MEAN CIRCULATION PATH
SECONDARY (FOLDING JET)
~_/ .-,~ ;;rr.;> q?-,\ .,~ G1;
~~Q~ j J tj~t t PRIMARY
SLOT AIR ~:wx FUEL
INJECTOR
Figure 73. Circwnierentially Stirred Com.bustor Flow Pattern
Inlet Plenum
Figure 74.
118
Inlet Pres. & Temp. Station /'
"-r... ....... ,-~ Exit Pres. & Gas
~ Sampiling Station 8 Sampling Probes
Exit 4 Pressure Probes
Thermocouple Station
Cooled Exhaust
Exhaust
______ Plug Valve'-+----t--Assy
QCGAT Annular Combustor Test Rig and Test C ell Ins tall at ion
were evaluated. Test points were those defined by the EPA; these included idle, approach power, 90 percent and take-off powers.
The next test phase was conducted with a liner moduled in accordance with the initial test results. The effects of changes in the cooling flow network by air partitioning on emissions at all engine operating conditions were evaluated. Liner temperature measurements and temperature paint tests were made at take-off power. Liner and curl durability were evaluated after each high-pressure test.
4.6.3 Test Results
Emission calculations using the taxi-idle and take-off power test points gave EPAP values equivalent to 45 percent of the UHC I 88 percent of the CO I and 101 percent of the NOx requirements. FIgure 75 presents the effect of air-partition modifications on NOx • Initial tests indicated the NOx results were within design goals. However, as the combustor pressure drop was increased to reduce smoke, levels increased. The cooling flow network was then modified by air partitioning to meet the NOx ernis sion goal.
The QCGAT liner finally selected to meet design goal has a slightly steeper slope, as depicted 1n Figure 75. The Lipert corre-lation , (Reference 9) for conventional combustor s 1S for comparison.
119
..... N o
x w o
20
z 10 Z 9 o 8 en 7 en ~ 6 ~ 5 o Z 4
3
INCREASED P PRESSURE DROP /
LINER y'" LTS 101 \ /" /('1 LINER Y"~ ~ I I /:% II
~ ~
/" I I ~ I TAKEOFF
~ 90%
30%
aCGAT LINER
fl~«-' ~,O v<? [t.,V' O<?-<f.
C;
I I I
TAXI IDLE
2, I I I 200 ._- ~-- ~nA~'
I I I I
100 200 300 400°C
Figure 75.
COMBUSTOR INLET TEMPERATURE
NO Versus Combustor Inlet Temperature x
5.0 ENGINE/NACELLE SYSTEMS PERFORMANCE
5. 1 OVERALL PERFORMANCE
The QCGAT engine was configured to reduce overall emisslOns and noise levels w1thout ser10usly impacting the advanced performance goals for the cycle. The de sign objechve for the QCGAT eng1ne program was to provide a min1mal fuel consumption 1n a cruise condition of 7620m (25,000 it) altitude at a Mach No. of O. 6 without sacrificing one-engine inoperative capabilities. Design and trade-off studies were performed to define the optimum cycle in terms of n01se, em1S sions, and performance. The selected des1gn cycle, resulting from the study, 1S presented in Table 3.
The QCGAT englne installed performance goals for the two pnme fhght cond1hons are shown in Table 4. This installed performance is with the nacelle system 1nclud1ng the fhght hp, m1xer nozzle and acoustic treatment. The sea level stahc takeoff thrust is 7166 N \1611 lbf) and spec1f1c fuel consumptlOn 1S 0.037 kg/hr/N (0.363lbm/hr/ lbf). For the 7620 m (25,000 ft) Mach 0.6 cruise, the thrust 1S 2157 N (485 lbf) and specif1c fuel consumphon is 0.064 kg /hr /N (0.628 lbm/ hr /lbf).
A mixer nozzle, Reference 8~ was chosen for the eng1ne conf1guration because of acoushc and performance reasons. F1gure 76 presents the eshmated var1ations of spec1fic fuel consumption, along an eng1ne operahng hne, with total net thrust at the selected cruise cond1hon, for the split and for ced mixer exhaust systems. As shown, a potential performance ga1n, at the cruise thrust, of approx1mate1y 3.0 percent could be realized with a m1xer.
5. 1. 1 Component Performance
The Avco Lycomjng LTS 101 turboshaft englne was selected as the basic core for QCGAT engine. Core component modiflcations, reqUlred to meet QCGAT design goals, were Lycoming funded. The major components developed, under the NASA contract, were the fan module, reduction gearing, and the nacelle system wh1ch includes the forced m1xer nozzle. The fan and nacelle were des1gned wlth low noise as a primary cr1teria. In add1hon, combustor system modif1cahons were made, as required, to meet the emissions goals.
The core compressor was tested to establish mechanical and aerodynamic performance w1th the turbofan inlet duct. The compressor per-
121
-N N
TABLE 3. RESULTS OF DESIGN STUDY
ALTITUDE = 7620m(25,000 FT), MACH = 0.6
Fan Pressu re Ratio
Cycle Pressure Ratio
Core Compressor Pressure Ratio
Thrust/Total Airflow, N/kg/sec(lbf/lbm/sec)
Bypass Ratio
SELECTED DESIGN
1.36
13.7
10.3
113.7(11.6)
9.4
.... N VJ
TABLE 4. QCGAT PERFORMANCE GOALS
(STANDARD DAY, INSTALLED)
SEA LEVEL STATIC
Rating Takeoff
Thrust, N(lbf) 7166(1611)
SFC, kg/hr/N(lbm/hr/lbf) 0.0370(0.363)
7620m(25,000 ft) MACH = 0.6
Cruise
2157(485)
0.0640(0.628)
I-' N ~
(Ibm/hr/lbf)
.72 kg/hr/N
0.072 .70
U .68 u.. en
0.068 I
.66
I 0.066
I .64
I 0.064
.62 200
ALTITUDE = 7620m(25,000 FT) MACH = 06 UN INSTALLED
300 400 500 I I I I
1000 1500 2000 2500 N
Net Thrust
Figure 76. Predicted Influence of the Mixer
600 (Ibf)
formance and surge char~ r.ter1shcs w1th pressure distortion as measured during the fan component testing were also established. The rig test results showed that the compressor efficiency was wlthlll 1. 0 percent of the design goal. The compressor showed high tolerance to pressure dlstortlOn produced by the fan. Also, the turbofan lnlet duct caused a reduction III airflow to the compressor of 1. 0 percent at the QCGAT operatlng cond1tions.
Rlg te sts on the imtlal gas producer turbine hardware confirmed that the des1gn eff1ciency of th1s stage was met with1n 1. 0 percent. However, the nozzle s were substanhally larger in flow area than deslgn. An attempt was made to correct for flow Slze, by reduclng the annulus area formed by the 1nner and outer wall contour. Th1s corrected the flow area problem but caused cascade losses wh:lch reduced stage performance by approx1mately 3 p01nts.
In addition, the interturb1ne duct pressure losses increased because of a resulting change in the turbine exit swirl angle. A redesIgn of the nozzle and rotor, to recover gas producer effiCIency, was completed, and further component and engine performance veriflcatlOn program IS conhnuing.
An expenmental evaluation of the QCGAT fan module has shown that the bypass performance has exceeded des1gn goals. (See discusslOn in Sechon 4.5 ). At the des1gn pressure ratio (1. 38) and speed (11,200 rpm), stage polytrop1c eff1ciency of 0.875 was demonstrated. This exceeded the des1gn goal efficiency of 0.870. Bypass airflow at thlS POlnt was 33.7 kg /sec (74.3. Ibm/sec) compared with a goal of 33.6 kg /sec (74.0 Ibm/sec).
The low pressure turbine, which was not rig tested, appeared to perform as anticipated based on measured engine data.
Eng1ne performance eshmates obtained from math model slmulations, based upon component test results, showed that further component develpment of the core, was requ1red to ach1eve performance goaL;. Lycoming 1S conhnuing the core development.
However, as a re sult of the analys1s, lt was concluded that the Lycom1ng QCGAT engine was a v1able veh1cle for demonstrat1ng noise, emissions and spec1flc fuel consumphon improvements whlch were the program IS pr1me obJective s.
125
5.1.2 Full Englne Tests
5. 1. 2.1 Engine Configurations and Test Plan
FollowIng the component rIg te sts, the full engine and nacelle system tests were conducted. Two engine configurations have been te sted. The referee configuration consists of a calibrated bellmouth followed by a straIght inlet duct to the fan shroud as shown In Figure 77.
In the exhaust system, the bypass and core flows are physically separated, (See Flgure 78). Separate exhaust nozzle s permit Individual change of fan pressure ratio and variation of the power spht between the fan and core.
The QCGAT test nacelle confIguration, shown in Figure 79,has the flight Inlet lip and dlffusmg duct mounted to the fan shroud. The flight lip can be readIly Inter changed wIth the bellmouth or the approach simulator inlets.(See Figure 80) Details of the test nacelle are shown in FIgure 81. The diffusing duct following the inlet contains interchangeable hardwall or acoustically treated softwallliners. The nacelle rear section consists of a core cowl coverIng the core engine whIle providing a smooth aerodynamic Inner wall contour for the fan flow surroundIng the core. The common mixed exhaust nozzle clamps to the rear face of the fan frame and contains the removable duct portion of eIther hardwall or softwall panels.
Various combinations of the two basic engine configurations, the referee and test nacelle, were tested during the performance calibration sequence. Table 5 shows an overvIew of the seven prime engine configurations whlch were tested in order to determine the performance characteristIcs of the engine and nacelle system components. Prlor to these tests, a baseline engIne configuration was tested with a cahbrated bellmouth coupled to a constant area duct and split exhaust.
The first three confIgurations, listed Table 5 WIth the split" or referee exhaust system, were tested with the diffusing flight inlet duct and the various interchangeable Inlet lips.
All tests with the spht exhaust were performed without the acoustic panels. The referee configuration with a bellmouth inlet was also used for the emissions sampling.
126
..... N -.]
Figure 77. QCGA T Referee Configuration - Inlet
I-' N 00
Figure 78. QCGAT Referee Configuration - Exit
Figure 79. QCGAT Test Nacelle
...... w o
FLIGHT NACELLE
~~lld n.s. (~-
SYMETRICAL BELLMOUTH
r=F/
Figure 80. Replaceable Inlet Lips
APPROACH SIMULATOR
..... VJ .....
REMOVABLE INLET LIP
~1 'hr I
I
- ------- ................... -... ...
t-'I , 't" \ I ,
I 111 ! !,
Ii :J-hiL!EI GttiP'~
REMOVABLE ACOUSTIC
PANELS
Figure 81. Details of the Test Nacelle
I-' TABLE 5. ENGINE CONFIGURATIONS TESTED UJ N
(PERFORMANCE TESTS)
ENGINE CONFIGURATION ACOUSTIC TREATMENT i
INLET EXHAUST INLET BYPASS
REFEREE CONFIGURATION
*8ellmouth Split Hardwall Hardwall
Flight Split Hard\vall Hardwall
Approach Split Hardwall Hardwall Simulator
TEST NACELLE
8ellmouth Mixer Hardwall Hardwall
8ellmouth Mixer Softwall Hardwall
8ellmouth Mixer Softwall Softwall
Flight Mixer Softwall Softwall
*Emission Test Configuration
The test nacelle conflguration with the m1xed exhaust was 1n1tially tested, for performance purposes, only w1th the bellmounth inlet. First, tests were conducted w1th hardwall panels in the inlet and fan bypass exhaust. Then, acoushc panels were placed in the inlet only. Finally, the eng1ne was te sted with acoustic panels in both the 1nlet and fan bypass exhaust. The installed performance demonstration was with the flight nacelle 1nlet, mixer nozzle and full acoustic treatment.
5. 1. 2. 2 Engine Tests Results
The purpose of the in1tlal tests with the referee conf1gurahon was to evaluate mechan1cal eng1ne operation and stress levels on fan and gear components.
Subsequent te sts using the referee system, were conducted to evaluate overall eng1ne and component performance prior to evaluatmg losses assoc1ated with acoushcally treated nacelle system. Var1ahons 1n performance attributed to the mixer system were also to be deter m1ned.
The purpose of these tests were twofold: First, to establish a base calibration for determin1ng component performance. Second, to evaluate 1nlet pressure losses assoc1ated w1th the diffusing duct coupled to the var10US 1nlet l1ps.
Engine test data with the varlOUS mlet l1ps are presented In F1gure 82 through 87. As previously stated, the engine te sts, with the various inlet !lps, were conducted 1n the early phases of the test program. Although the data, obtamed from the initial tests, does not reflect the final performance characteristics of the engine, the data 1S vahd for evaluahng the impact of the inlet hps on the overall engine performance.
Deta1led analys1s of the test data has ind1cated that the diffusing duct and various inlet hps had a neglig1ble 1mpact on the overall engine performance.
Following the referee system performance and emissions te sts, the mstalled nacelle test sequence was conducted. The purpose of these tests was two fold: f1rst, to establish engine performance with a mixer nozzle; second, to evaluate the 1mpact of the inlet and fan bypass exhaust acoustical panels on engine performance. The engine test results, as shown in Figures 88-93, ind1cated that the acoushcal panels, used for nOlse reductlOn had a neglig1ble influence on the overall engine performance. After the performance evaluation tests, the eng1ne was
133
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O. 200 400 600 800 1000 1200 (fbf)
o 1000 2000 3000 4000 5000 6000 N
NET THRUST REFERRED TO FAN INLET
Figure 82. Low Spool Speed Versus Net Thrust
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o 1000 2000 4000 5000 6000 N
Figure 83.
NET THRUST REFERRED TO FAN INLET
High Spool Speed Versus Net Thrust
135
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NET THRUST REFERRED TO FAN INLET Figure 84. Thrust Specific Fuel Consumption Versus
Net Thrust
(Ibf)
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HIGH SPOOL SPEED REFERRED TO FAN INLET. PERCENT RPM
Figure 85. Low Spool Speed Versus High Spool Speed
137
o ~ a: w a: :J en en w a: a.. z ~
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1 28
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LOW SPOOL SPEED REFERRED TO FAN INLET, PERCENT RPM
Figure 87. Total Engine Airflow Versus Low Spool Speed
139
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NET THRUST REFERRED TO FAN INLET
Figure 88. Low Spool Speed Versus Net Thrust -Nacelle Installed
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11 ACOUSTIC PANELS, INLET ONLY
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NET THRUST REFERRED TO FAN INLET
Figure 89. High Spool Speed Versus Net Thrust -Nacelle Installed
141
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o 200 400 600 800 1000 1200 (Ibf)
o 1000 2000 3000 4000 5000 N
NET THRUST REFERRED TO FAN INLET
Figure 90. Thrust Specific Fuel Consu:mption Versus Net Thrust - Nacelle Installed
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HIGH SPOOL SPEED REFERRED TO FAN INLET, PERCENT RPM
Figure 91. Low Spool Speed Versus High Spool Speed -Nacelle Installed
143
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FAN BYPASS AIRFLOW REFERRED TO FAN INLET
Figure 92. Fan Pressure Ratio Versus Bypass Airflow -Nacelle Installed
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20 30 40 50 60 70 80
LOW SPOOL SPEED REFERRED TO FAN INLET, PERCENT RPM
Figure 93. Total Engine Airflow Versus Low Spool Speed -Nacelle Installed
145
transferred to the acoustic test slte for noise evaluation.
Table 6 shows a comparision between the demonstrated installed engine thrust and specific fuel consumptlon w1th the design goals. The measured static thrust and specific fuel consumption are 6485 N (1458 lbf) and 0.0400 kg Ihr IN (0.392 lbm/hr Ilbf). The crUlse performance was estimated based upon eng1ne statlc test data and component r1g te st re sults.
5. 1.2.3 Engine Performance Test Summary
The estlmated crU1se performance of the Avco Lycoming QCGAT engine, in terms of specific fuel consumptlon, is approx1mately a 10.0 percent 1mprovement over currently ava1lable small turbofan eng1nes m the 13,344 N (3000 lbf) or less thrust class.
The performance goals were arnbItlous and certaInly achievable withm today's eX1sting technology. Although the program performance goals were not ach1eved, the loss 1n eng1ne performance has been 1dentifled as defiCIencies in the turbine section of the core engine. A redesign of the affected hardware has begun under a separate Lycommg funded program, and further development testlng will be conducted as necessary.
~. 2 EMISSIONS TEST RESULTS
In 1970, Congre ss passed the Clean Au Act. This Act, which was to be effectlve in 1979, directed the Envlronmental Protection Agency to estabhsh em1SS1ons standards apphcable to alrcraft. These standards, Reference 10, for small turbofan aircraft, which have now been abandoned by the EPA, were kept as NASA goals for the QCGAT engine program. To achieve these em1SS1ons hmits, the basic combustor des1gn used 1n the LTS 101 eng1ne, References 11 and 12, were selected.
5.2. 1 De slgn and Em1SS1ons Projectlons
Th1s design, wh1ch 1S a circumferentlallystirred combustor, 1S shown in F1gure 94. In pr1nc1ple, the primary alr is adm1tted through slots in the liner header producing flow circulation about a circumferentlal mean line. A1r Jets, called "folding jets" entering through the 1nner wall re1nforce the primary zone recirculatlon, and the vortex fills the full annular height of the liner.
146
......
"'" -.J
TABLE 6. QCGAT PERFORMANCE
(STANDARD DAY, INSTALLED)
SEA LEVEL, TAKEOFF
Thrust, N(lbf)
GOAL DEMONSTRATED
7166(1611) 6485(1458)
SFC, kg/hr/N(lbm/hr/lbf) 0.0370(0.363) 0.0400(0.392)
DESIGN CRUISE, 7620m(25,000 ft) MACH = 0.6
Thrust, N(lbf) 2157(485) 1850(416)*
SFC, kg/hr/N(lbm/hr/lbf) 0.064(0.628) 0.074(0.723) *
*Estimated from Static Data
148
MEAN VORTEX • ~A CIRCULATION PATH Y~~
SECONDARY -----: 6(,~.€ c:::Il FOLDING JETS ____ ~
---...~~. ~ VORTEX
CIRCULATION PATTERN
AIR ENTRY --.1,,_
~. ~~i!f VIEW A-A
Figure 94. QCGAT Combustor
\ FUEL
INJECTOR
The vortex spreads c1rcumferentlally 1n both directions and 1S forced to turn in the axial direction on either side of the folding Jets and the mean path of the com.bustion zone flow vortex takes the shape of a horseshoe. The number of fuel injectors is thereby reduced by one half, compared with normal practice, because of this un1que combustor primary zone aerodynamic concept.
Em1ss1on measurements, for this type of combustor, attained from the LTS 101 engine were available for use In predicting em1SS1ons for the QCGAT performance cycle. Table 7 shows the estimated emissions values, for the QCGAT cycle, with the production LTS 101 combustor. These EPA parameter s were generated for a takeoff and landlng cycle for clas s T1 aircraft, (Reference 10).
The se emissions projections lndicated that further development of the LTS 101 combustor was required to reduce smoke.
Alrblast lnjector s, which replaced the dual or1f1ce injector s, were selected to reduce smoke. The introduction of the airblast lnjector s also 1ncreased combustor eff1c1ency and oxide s of n1trogen (NOx) at 1dle.
Increaslng the combustor pre ssure drop for temperature distr1bution control, also 1ncreased NOx and combustor efficiency while apprec1ably decreasing carbon monoxide and unburned hydrocarbons. This is typical of the improved primary zone m1x1ng, wh1ch re suits from the h1gher pressure drop. Air partition, modificatlOns were then made, as requlred, to meet the design goals for NOx. Figure 95 presents the effect of air partitlon modiflcatlons on NOx. Unburned hydrocarbons and carbon monox1de were w1thin the goals in all tests. In1tially, the NOx slope for the LTS 101 combustor was as predlcted, and met the goal. However, as the combustor pressure drop was increased to reduce smoke, NOx increased.
Air partitlon modif1catlons, as previously stated, were then made to meet the NOx em1S Slons goal.
The final selected QCGAT liner, which m.et the goal, has a slightly steeper slope than the initlal configuratlOn. The L1pfert correlation, Reference 9, for conventlonal combustor s lS shown for comparison.
149
..... U1 o
TABLE 7. INITIAL ESTIMATED QCGAT EMISSIONS
L TS 101 COMBUSTOR
Estimated Values*
NASA Goals*
UHC
0.034 (1.2)
0.045 (1.6)
co
0.238 (8.4)
0.266 (9.4)
*g/kNs (lbm/1000 Ibf thrust hr-cycle)
NOx
0.096 (3.4)
0.105 (3.7)
SMOKE NUMBER
70.0
45.0
I-' U"l I-'
x UJ o
20
~ 10 Z 9 o 8 ~ 7 ~ 6 UJ x 5 o Z 4
3
INCREASED /'-PRESSURE DROP ,,/
LINER Y" LTS 101 \ / /(i LINER ~;~ ~ I I /:% II
~ ~
.,/ I I ~ I TAKEOFF
~ 90%
I 30%
QCGAT LINER
f<<<-~ ~ .. p V~ ~VO~~
C;
I
Tlxl IDLE
2. I I I I I 200 ._- --- --- --- --I I I ,
100 200 300 400 °C
COMBUSTOR INLET TEMPERATURE
Figure 95. NO Versus Combustor Inlet Temperature x
5.2.2 EmIssions Sampling
Development and inItial emisslOns testing of the combustor was conducted in the laboratory. (See discussion in Section 4.2.) After the laboratory tests, the QCGAT liner was transferred to the engine for demonstrated emissions sampling.
The emissions test probes were installed as shown in FIgure 96. The probes, which are crucIform-shaped, were set at two angular posItions. One probe measured along the horIzontal and vertical axes. The other probe was rotated 45 degrees.
Table 8 is a comparison of the emIssions test results with the NASA goals. Measurements from the engine test showed that the unburned hydrocarbons were 60 percent lower than required. The carbon monoxide was 30 percent lower, oxides of nitrogen 1.0 percent hIgher and the smoke number 50 percent lower than the goal.
5.2.3 Emissions Summary
The emissions requIrements of the QCGAT engIne have been met and, In most cases, surpassed. The QCGAT combustor provides substantla1 margIn for carbon monoxide and unburned hydrocarbons emissIons while meeting the goal for NOx wIthm the scope of the program.
The combustor system modifIcatIons required to meet the emIssions goals had a negligible effect on engine performance.
152
Figure 96. Engine Emissions Sampling Test Probes
...... U"l ~
TABLE 8. QCGAT EMISSIONS RESULTS
Goal*
Engine Test*
Engine Test/Goal
UHC
0.045 (1.6)
0.017 (0.6)
0.4
co
0.266 (9.4)
0.193 (6.8)
0.7
*g/kNs (lbm/1000 Ibf thrust hr-cycle)
SMOKE NOx NUMBER
0.105 (3.7)
0.106 (3.75)
1.01
45
24
0.5
5.3 ACOUSTICAL PERFORMANCE
Avco Lycoming particlpated in the NASA QCGAT program by developing a fan module based upon an existlng turboshaft engine. The fan was designed using the latest in large engine nOlse control technology and a mixer that was added to reduce the already low exhaust-gas velocity. A nacelle incorporating sound treatment was also provided for the test englne. A noise prediction model was used throughout the design effort to evaluate various design alternahve. Acoushc te sts were then made to verify the predictlon and identify the noise characteristics of the fan, core, jet, and sound treatment. Analysis of the recorded data yielded close agreement with the expected results. As anticipated, core nOlse was the predommant source of noise produced by the QCGAT engine. Flyover noise predictions made indicated that the Avco Lycoming QCGAT engine would meet the goals defined for the QCGAT program.
5.3. 1 Background
The Avco Lycoming Quiet Clean General AVlation Turbofan engine program was designed to demonstrate the latest control technology for gas turbine noise ln a general aVlatlon size englne. A conslderable amount of effort was requued to ldentify the de sign feature s that offset the generation of nOlse. This work is shU in progress, as can be wltnessed by the complexlty of the facllitles at LeW1S Research Center and elsewhere. Most of this work, however, has been directed towards the commercial transport class of engines. The QCGAT program was designed to broaden the scope of subject effort to lnclude the general aviation size engine.
The slgnificant feature s of the QCGA T engine shown on Flgure 97, are:
o Low exhaust veloclty achieved by a hlgh bypass fan de sign
o Use of an exhaust mixer
o No fan 1ll1et guide vanes
o Subsonic fan blade de slgn
o Large fan blade to vane spacmg
155
Ave e IN E INE
Figure 97. Cross Section of the Engine with Cutaway
o H1gh fan blade to vane ratlo
o Acoustical lining of the fan inlet and discharge ducts.
Nacelle and a1rcraft configuratlons play an important role in 1ncor poratmg the above feature s in the overall acoustic de slgn. For example, the mixer 1S enclosed in a shroud formed by the nacelle. Also, the fact that forward airspeed m1tigates the amount of Jet noise generated has been factored mto the de sign. Based upon the re sults of the pred1ctlOn of the acoustlcal performance of the engine alrcraft system and the impact of each component on the overall des1gn, the above features were optim1zed for the QCGAT aircraft.
QCGAT n01se goals were se lected by NASA to create a de slgn that mcluded the latest noise control technology. Avco Lycom1ng's QCGAT engme de sign enta1led the add1tlon of a new fan de slgn module that 1ncorporates the latest noise techn1ques to an eX1stmg tur boshaft eng 1ne.
Or1g1nal estlmates of the engine n01se emissions, based upon that de slgn, are shown m F1gure 98 along w1th the relevant measurement locatlons. Th1S analysis 1nd1cated that takeoff noise levels would be 4 EPNdB below the goal, sldelme :) EPNdB below goal and approach to be 9 EPNdB. The pred1ctlOn also md1cated that the core would be the dom1nant source of noise at each measurement positlon, w1th the fan contr1butmg to the approach n01se and the Jet contr1butmg to the takeoff n01se levels.
Notlce that the goals are glven 1n terms of ail-craft flyover noise. From the pOlnt of brake release and WIth the aircraft flYIng dIrectly overhead, the takeoff measurement pOInt bes 6500m (3.5 nautIcal miles) down range. The sidelIne measurement pOlnt also lIes down range on a takeoff but 1S d1splaced 460m (1/4 of a nautlcal m1le) to the slde and cons1sts of a senes of p01nts 1n order to determIne max-1mum nOIse level. The approach measurement point IS located under the land1ng flIght path at a point one nautlcal mile from the runway threshold. Since the approach glide slope IS defIned as 3 degrees, the altltude of the aircraft over the measurement pOInt 1S fixed at 112m (370 feet), (Reference l3). Thus, aircraft performance had to be considered 1n the engIne des1gn. Beech A1rcraft Company was contracted to defIne the characteristIcs of a tWIn-engIne QCGAT-powercd aircraft. With respect to nOlse, the rate of climb at takeoff, the power reqUIred at approach, and the geometry of the wing were determined. AIrframe nOIse, however, was not Included In the nOIse estlmates.
157
...... \J1 00
APPROACH FL YOVER MEASUREMENT POINT
s<;;, ~"'<;;,~
~<;;, «' ,<Q«-.'f.
'"' ( < -' __ ----==--- 3.5 NM 7' 30 ~ ~
/2 1 NM ~ ~2 ~UNWAY .;fN: TAKEOFF FLYOVER
0~~ «'
o",Q 4 MEASUREMENT S~ 0 0 0 0 POINT
~~«-.«) MEASUREMENT POINTS TO DETERMINE ~~ MAXIMUM TAKEOFF SIDELINE NOISE
QCGAT ACOUSTIC PERFORMANCE
EPNL GOAL PREDICTED ENGINE CONDITION EPNdB EPNL,EPNdB
Takeoff Flyover 69.4 64.8
Takeoff Sideline 78.4 71.7
Approach Flyover 83.4 73.8
Figure 98. QCGAT Noise Goals
The de sign and performance of subject aircraft plays an important part In the nOlse emisslOns of the QCGAT engines. That is, the approach speed and takeoff performance of the aircraft can be varIed to meet market requirements. For example, a lower approach power could be used that would result In lower approach noise levels but would require more runway length. Because the approach noise levels were predicted to be low, a small penalty was accepted to reduce fleld length reqUlrements. This wIll allow the aircraft to be certified for use at most air fields In the Umted States.
Gas turbine engine noise -source identification and control (FIgure 99) starts with the engine. Given the geometric and performance characteristics of the engine, predIction of an engine IS
nOlse emissions can be made. Engine noise is subdivided into five
dIstinct nOlse generating mechanisms: 1) fan, 2) compressor, 3) combustion process, 4) power turbines, and 5) the turbulent mlXlng of the exhaust Jet with the ambIent air. The maJorlty of the work accomphshed to advance the state-of-the-art for gas turbine and aircraft noise ldentification and predlction was and is presently belng carried out by NASA as part of their Aircraft Noise Prediction Procedures (ANOPP) (References 14 thru 17). This work served as the basls of noise prediction efforts used in thls program. Certain modifications were made in order to more accurately reflect the experiences at Avco Lycoming with engine noise predlctions. Then by using thls aircraft performance and applying flight effects, alrcraft flyover nOlse s were calculated.
5.3.2 Engine Design and Noise Prediction
5.3.2. 1 Fan Deslgn
The fn st task reqUlred to de sign a fan module for an existing turboshaft engIne involved several iterahons to assess the design alternatives. Reduction of nOlse was achIeved through the use of a low-pressure raho fan to reduce blade loading and noise genera tion, which was part of the fan de slgn from its conception. Other de sign feature s also shown to have re suited in quieter fan de signs for the large turbofan engine s are as follows:
o Low Blade Loading
o SubsonIC Blade Tlp Speed
159
..... 0' o
I ENGINE CYCLE DATA I !
STATIC NOISE PREDICTION FOR:
• FAN ATMOSPHERIC • COMPRESSOR
MODEL FOR NOISE PROPAGATION • COMBUSTION PROCESS ,
• TURBINE
• JET
! FREE FIELD ENGINE
NOISE LEVELS
! AIRCRAFT CONFIGURATION : FL YOVER NOISE PREDICTIONS 1
AND PERFORMANCE 1 EFFECTIVE PERCEIVED
NOISE LEVELS
Figure 99. Noise Prediction Procedures
STATIC ENGINE NOISE
MEASUREMENTS
COMPARISONS TO IMPROVE PREDICTIONS
FLIGHT EFFECTS
o Blade to Vane Spacing Greater 2 Blade WIdths
o Vane to Blade Ratio Greater than 2
o No Inlet Guide Vanes
Because fan blade tip speed was desIgned to be subsonic, multiple pure tone s or "Buzz Saw Noise" were ellmlnated altogether. The design relahve-tfp Mach number for the QCGAT engine is 1. 13; this Y1elds a subsonic value at all sea-level operatIng points. The distance separating the fan blades from the fan exit gUIde vanes 1S great when compared w1th the blade width in order reduce rotorstator mteraction noise that is expressed as fan broadband noise. A value of 2.3 was used for this ratio. The ratlO of fan vane s -toblades optim1zed at a value of 2.5 elimmates what is known as splnn1ng modes that propagate at the blade passing frequency fundamental. In addItion, mlet guide vanes were not used in the fan de sign. To further insure that inlet turbulence was reduced, a long inlet duct was mc1uded in the nacelle de slgn. The se feature s were accounted for in the prediction of the fan noise levels. The nOIse prediction ind1cated that the fan would be a contributor, along with the core, to only the approach power levels. By identIfying the effect of the various alternative s w1th the aId of our prediction procedure s, a balance was mamtained that achieved a low-no1se slgnature at approach.
The model for fan nOIse pred1ction was derived from Reference 16, In which fan noise 1S considered to be composed of five sour ce s.
1. Fan miet broadband
2. Fan d1scharge broadband
3. Fan mlet d1screte tone
4. Fan discharge discrete tone
5. Fan 1nlet combIned tone.
Combmed tones do not propogate below cutoff and do not corne mto play for the QCGAT engme.
The baSIC features of the prediction program are outlmed in Figure 100, where each factor is part of the equahon
161
162
:Q "C
... : :u > :u
...:l :u '" ... :::)
Z .:.: '" :u p.. "C :u !: -'" a .. 0 Z
.. o U III ~
I':
~
70
60
50
o
~ -10
'" .. o U
1.6
0-1.0
MTRO ~ l.l~~
.5 1.0
Rotor Tlp Relative Mach Number, Mtr
:Q "C .10 ..
B u III ~
I':
~ u 0 Q) .. .. o U
- RSS < 100%
RSS ~ 100%
_ZOL-~~~~~ ____________ -Io~I--------~------~---
~IZ345 10 100 1000
:Q Harmonlc, k Rotor Stator Spaclng. %
Figure 100. Fan Noise Com.ponents (Ref. 16)
Lc = 20 log (
Temp R1se) Ref Temp
T 10 log (Mass flow ) + Ref Mas s flow
f (Relative T1p Mach No. ) + f (Blade -to-vane spacing) + D1rectivity
+ Spectrum Correction
The re sult was then extrapolated to the corre sponding free -field location for cornparison with test-stand data.
5.3.2.2 Jet Noise
Jet n01se is the second major element in the QCGAT engine design. A high-bypass fan design is used to reduce the exhaust velocity, thereby reduclllg the noise generated by the turbulent mixing of a highvelocity Jet.
To further reduce Jet noise, a six-element mixer was de slgned to m1X the core engine and fan exhaust gas to yield a slllgie low-veloc1ty exhaust Jet. (See Reference 8 for details of the mixer de sign.) The mixer, however, is not entirely free of side effects. Pre -and postrnix1ng turbulence can be an add1tional source of n01se that must be dealt with. The se noise sour ce s can be reduced by the additlOn of a shroud. The shroud effect was 1ncor porated by the extended m1xlllg section of the f1nal nozzle.
5.3.2.3 Core Noise
The high-bypass fan and mixer were des1gned to reduce the Jet n01se component to a noise level below that of the core when forward fl1ght effects cause a reduction to occur in the Jet noise that leaves the core n01se component. Core noise is the noise generated by the combushon process. Engine compressor and turbine noises were predicted to be above the audible range. Thus, the se noise source s which do not contr1bute to the perce1ved noise of the QCGAT engine were not considered 1n the des1gn.
Core nOlse models for the most part have been emplrically denved. The ANOPP routine has been found to be adequate for the core turboshaft engine (Reference 15). Thls predictlOn rnodel uses combustor mass flow, temperature rise, and pressure drop as the basls for predlcting core noise. (See Ftgure 101.) Empirical data also suggest a 7 to 10 dB reductton for the turbofan version of thts model (Reference 6). Core n01se 1S now recogniz:ed as a rnajor
163
.... 0' f!::.. (GOOD AGREEMENT FOR TURBOSHAFT ENGINES)
======= , COMBUSTOR INLET == I~ ..
COMBUSTOR EXIT
Noise a Function of:
• Mass Flow
• Temperature Rise
• Pressure Drop
Figure 101. Core Noise Model
FUEL NOZZLE
source In turbofan eng1ne noise and is the focal point of much research; however, core-nOIse control was not part of this program. As predicted, the core was a sIgnu1cant contributor to the noise characteristics of the aircraft. Consequently, further fan and jet n01se reductIon would have to be unwarranted.
5.3.2.4 Flight Effects
An a1rcraft engine operates differently in flight than when it is tied down to a test stand. Engine noise character1stics also change. In flight, the air inflow is streamlined because of the flight cleanup effects of the forward alrspeed and the absence of ground turbulence that 1nfluence the generation of fan noise, parhcularly the tone at the blade -pas sing frequency. Forward flight, however, has its greatest 1mpact on the generation of jet noise. In flight, the relahve veloc1ty between the exhaust and the amb1ent air is reduced. This plays an 1mportant part m the overall design of the engine a1rcraft system. For example, the a1rspeed at takeoff is, in part, determmed by ava1lable runway length. A longer takeoff roll would perm1t a higher takeoff speed. Consequently, the same Jet n01se level and relat1ve Jet veloc1ty could have been ach1eved by using a h1gher exhaust veloc1ty and a higher takeoff a1r speed.
As the aircraft fhes past the observer, the sound vane s both III hme and spectral content. Dynamic ampliflcation acts to increase the n01se level as the a1rcraft approaches and reduces 1n noise levels as it recedes. There is then, the doppler effect that 1mparts a frequency sh1ft to the nOlse spectrum as the aircraft flies by. These phenomena must be accounted for to accurately predicc the perceived noise of the a1rcraft.
5.3.3 Sound Treatment Design
It has been recognized for sometime that the fan inlet and discharge ducts of an engine nacelle (Figure 102) offer 1dea11ocahons for the mstallation of sound treatment materials are particularly efficient generated by the fan. Absorphve material are particularly efflcient in absorbmg sound energy in the high-frequency reglOn where much of the acoushca1 power radiated by the fan 1S concentrated. In add1hon, the treatment of sound can be accomplished by the use of fhght-worthy materials that add minimal weight to the aircraft. Finally, the theory and exper1ence of de signing sound-treatment panels are sufficlently soph1stlcated to accurately pred1ct the re suits that w1ll be
165
I
I
\ \ ' , i --\..jr--- I
\4 i~~
,f---- 1 , .
166
C/l $:l o
'1"'4 -:..> rd u o H
'8 :::I o
U)
. N o .-t
achieved with a particular de sign. Sound-treatment panels were, therefore, installed in the QCGAT engme nacelle to determine the benefits that would be derived from thelr use ln an aircraft design and to evaluate the manufacturmg tolerance s of the size s requlred for the QCGAT engine.
Sound attenuation requirements were determined by comparing the predicted noise levels with the QCGAT program goals. Alrcraft approach repre sented the only condition where the fan nOlse was
... 1redicted to contribute slgnificantly to the alrcraft noise levels In addition, the frequency of the blade-passmg tone at approach oc·curs in the more heavily weighted part of the audible spectrum. Consequently, the approach power pOlnt was selected for the de slgn of sound treatment. At other condltions, the fan does not contrlbute significantly to the aircraft noise levels.
The Lockheed California Company was contracted to design the sound treatment arrangement for the nacelle. Given the dimenslOnal hmltatlOns, the nacelle needlng sound treatment, and englne operahng parameters at approach, Lockheed generated a set of deslgn curves from WhICh the sound treatment was de signed (Reference 1).
These curves were based upon an analyhcal and emplrlcally derived solution to what are known as the convected wave equatlOns. The se equations de scribe the sound generated by the fan as mode s of acoustic energy rotating with and agamst the fan. This acoushc energy can only propagate under certaIn boundary condltions. The physical characterIstics and operatmg parameter s form the se boundary condihons and determine which modes will propagate. Lockheed performed this analysis and recommended a de sign.
Recommendations based on the Lockheed deslgn, shown ln Flgure 103, were for a sIngle degree of freedom panel for both the mlet and discharge ducts. The Lockheed deslgn conslsts of a solid backing plate held 16 mm (5/8 in. ) off an mner plate that lS perforated to achleve a 5 percent open area. A honeycomb cell-structure material separates the mner and outer plates. The inlet panel. 330 mm (13 In.) 10113. fulfllls the aval1able space ln the mlet duct. The discharge sound treatment conslsts of a 45.7 mm (18 m.) long panel on the outer duct wall. The mner duct wall formed by the core cowl was not treated. The dlscharge panel was termlnated before the start of the mlxer to slmphfy the de slgn. OtherWIse, the radIant heat from the mlxer would have requlred the selechon of more expenslve materlals and fabrlcahon technlque s.
167
..... '" ~
Fan Inlet
Fan Discharge
0.63 IN.
OPEN THICKNESS LENGTH AREA
0.63 IN.
0.63 IN.
13 IN.
18 IN.
SOLID BACKING PANEL
/---5% OPEN AREA
FACE SHEET
5%
5%
Figure 103. Sound Treatment Design
The pred1cted msertion losses for the fan mlet sound treatment panel at the approach and takeoff pomts are shown in Flgure 104. The sound treatment, as discussed earlier, was des1gned for the approach condition. At this power setting, the peak attenuation is made to cOlncide with the blade -passmg frequency. The insertion los s 1S higher at the takeoff cond1tion because of the increase in airflow through the engine. The blade-passing frequency at takeoff is also higher. The result 1S an attenuation that is approX1mately the same as that for the approach condition.
The pred1cted attenuatlon for the fan discharge treatment 1S shOV'rl m F1gure 105. The duct w1dth between the inner and outer wall makes the treatment more effective, even through the inner wall is not treated.
The test nacelle and sound treatment panels were fabr1cated by Avco Aerostructures in Nashville, Tenn. The test nacelle was designed to take msert panels m the fan inlet and discharge ducts where ordinarily the sound treatment would have been made integral w1th the nacelle. Two sets of 1nserts were fabricated. Each set was de signed to be of one piece for easy removal and installation during testing and to be rigid enough to maintain the des1red wall contours. The panels were of a sandwich-type construction with a honeycomb structure separatlng the 1nner and outer plates. The th1ckness of the honeycomb was determinted by Lockheed's sound attenuation requ1rements. One set was fabricated w1th a sohd mner plate, and one set (F1gure 106) was fabricated wlth an mner plate perforated to ach1eve a 5-percent open area. In this way, the engine could be tested with and without sound treatment m the nacelle.
The small radius of the inlet and discharge duct was limited to the depth of honeycomb that could be used without warping the cell structure walls. The selected honeycomb material used a small cell pattern in order to be flex1ble enough to accommodate the curvature. This, however, meant that there would be fewer holes per cell and potentially have more hole s blocked by the cell walls as the honeycomb was la1d over the perforated plate. The mmimum hole Slzes available dictated a w1de diversion of holes. A speclal adhesive was used that m1grated up the cell walls during the curing process and did not plug holes. The periol'ated plate was punched to a 6-percent open area. Then when th~ honeycomb was bonded to theplate, the open area was reduced to the designed 5 percent.
169
12 TAKEOFF
OJ 10 "0 -en en 8
0 -1
Z 6 0 .- 4 a: w en z 2
0 250 500 1 000 2000 4000 8000 16000
FREQUENCY, Hz
Figure 104. Inlet Attenuation
170
14
m 12 "0 -en 10 en
g 8 z o 6 f-a: w 4 en z
2
APPROACH
r I
I
O~~~~~~~------~
500100020004000800016000
FREQUENCY, Hz
Figure 105. Exhaust Attenuation
171
I LET L ISC GE EL
Figure 106. Sound Treatm.ent Panels
5. 3.4 Wmg Shielding
The program goals are given in terms of aircraft flyover nOlse parameters. Experience has shown that when the englne lS placed above the wlng, the wing serves as a barrier. A barrier -attenuation routine was included ln the aircraft noise model to account for this effect. As shown in Flgure 107, the wlng creates a shadow zone that moves in conjunction with the aircraft. Since only a small fraction of the noise is refracted around the leading and trailing edges of the wmg, the forward radiated fan noise wlll not reach the ground as the shadow zone passes an observer.
5.3.5 Acoustic Test Phase
The goals of the test program were to verify the noise predlctions by comparison wlth measured data to determlne the noise reductlon of the mixer and to determlne the effe chvene s s of the sound-treatment panels.
The normal method of recordmg noise emltted by an engme lS to record the 1/3 octave band sound pressure levels at nineteen postlOns located on an arc 30 meter s (100 feet) from the engine. A full set of data over an arc of 180 degrees can be obtamed wlth mlcrophones located every 10 degrees. Four power settmgs corresponding to the
operahng envelope of the englne were used. In addihon to the farfield microphones, acoushc probes were placed on the englne to aid in ldentifying core and mixer components and the nOlse reduction of the sound treatment. A barrler was also used during a part of the te shng to aid in lsolatmg, the fan inlet and dlscharge component sound levels.
Three separate engine conflgurations used during the acoustic te sting of the QCGAT engine were: 1) a split-flow exhaust nozzle configuration called the referee system: 2) the hardwall nacelle configuration in which the te st nacelle, mlxer, and hardwall fan inlet and discharge panels were used, and 3) the softwall nacelle configurahon m which the hardwall panels were replaced wlth the sound treatment panels. Each configuration was tested to record the engine noise at four power setting s. The QCGAT engine was mounted in a test frame and after a test cell series, lt was moved to the freeheld test site. Thls site is locat~d remotely from the plant in an area free of most nOlse intruslons and where testing does not affect the local communlty.
173
..... -..]
~
_-+ __ ENGINE/ _ NACELLE
FAN INLET NOISE SOURCE
WING .-.... - -. -.-..-
. .. e. o • • :-. ~
~ b
:....\
~ ~')l ,~
DIRECT RAY (NO WING)
Figure 107. Wing Shielding
The engine (In the nacelle and test frame) was Installed on a rotating test stand that lS capable of rotating a full 360 degrees. The normal method of testing is to record engine noise on an arc 30 meter s (100 feet) from the engine by means of five microphone s placed 10 degrees apart, as shown on Figure 108. By rotatlng the englne and repeating the test points, a full 180 degrees of noise can be obtained with some overlap. The microphones located at the 170-and l80-degree points ln exhaust stream were not used for most runs.
One-half lnch condenser mlcrophones fitted with wind screens placed on the ground were used as recommended by NASA. Thls allowed 6 dB correction to be used when correcting the data recorded over a reflecting plane to free-fleld conditions for comparlson wlth the predicted noise levels. The microphone placement is shown in Figure l09. Slgnal condlhonlng lnstrumentation was located in an acoustic data acquisltion trailer where the data were recorded on magnetic tape for later analysis.
The test data were recorded on magnetic tape in 2-mInute segments. The instrumentatlon settlp for the aqulslhon of the acoushc data is shown 0'1 Figure 110. The tape -recorded data were then played back through a 1/3-octave band dIgital frequency analyzer to obtain the 1/3-octave band sound pressure levels for each run. An averaging technlque was used to average out the random fluctuations In the data so as to yield a steady-state value for the 2-mInute sample. The data were fir st corrected to standard acoustic day condltions and free fleld condltions. They were then organized and tabulated by frequency-versus-angle from the englne inlet for each operating condltion and conflguratlOn.
At each te st point, a complete set of engine performance data was recorded for use in predlctlng the engine static nOlse levels for comparison with the measured sound levels. The amblent pressure, temperature, and relative humldity were also recorded.
Fan nOlse radlates both from the lnlet and the exhaust. Near the ln1et aX1S, lnlet fan n(:>lse domlnates, and near the exhaust axis exhaust fan nOlse dominates. However, near 90 degrees, the two blend together. To lsolate the fan Inlet nOlse from the fan dlscharge nOlse, a barrler was :::hyslcally placed between them. This was accomplished at the freefleld test slte wlth the barrier as shown on Figure Ill. The barrlel' was constructed of a flxed partltlOn 4.3 ITleter s (14 feet)
175
176
POSITION 4
~~ G\NE. ROTATION
POSITION POSITION 3 ? \,{:,(j~{:, \ / a~~~\'~ 400f' '{' '{
40~ 50° _~ . \ ~ 10°
/
POSITION 5
- -+ 0-- 100
0--+ . 10° ~100 O----i-
'Y -~ ~ FEET 9--- 10°
MICROPHONE 50° '\ ~ LOCATION
POSITION 6
I
Figure 108. Microphone Arrangement
Figure 109. Test Site
178
preampl~f~er power supply
~=d'men~----;:r ~ 12.7 :n-a (1/2") m~crophone
Sank of Sound Level Meters
o 0
FM Tape Recorder
8 eRO
Figure 110. Instrum.entation Setup for the Acquisition of the Acoustic Data
(ISOLATE FORWAR D AFT RADIATED FAN TONES)
Figure Ill. Barrier
h1gh by 6 meters (20 feet) long and a movable part1tion through Wh1Ch the eng1ne llliet protruded. Th1S effectlvely removed the fan d1scharge n01se from the measurements. Data were then recorded over an arc of 80 degrees. The movable part1tlon was then removed, and the engine was rotated 180 degree s so that the exhaust protruded through the barrier when it was moved back into position. The fan discharge n01se was then recorded without fan inlet noise contribuhons. Both of these tests were run at the same four-power setting with the hardwall and the softwall nacelles installed on the engine.
Locations of engine -mounted probe s are shown III F1gure 112. Half-inch condenser microphone s were located both upstream and downstream of the Inlet sound treatment to measure the noise reduction across the llliet sound treatment panels. Semi-infinite wave guide probes, supplied by NASA, were used to sample the acoushc pressure levels in the primary engine exhaust and at the mixer exhaust plane. These probes consisted of 6.35 mm (1/4 lllch) condenser microphone s III a sealed tube (See Reference 7 • ) A low-volume flow of nitrogen at a pressure Just above that in the duct provIded a gas seal to prevent hot exhaust gas from entering the tube where it could damage the mIcrophone.
The probes were designed to record the acoustic pressure levels
at the IndIcated probe locahons. They were to be used In coherence analyses 1£ It became necessary to determIne what part of the n01se 1n the far-held org1nated from w1th1n the eng1ne.
The split-flow nozzle configuratlOn with the semi-mfimte waveguide probe s mstalled in the primary exhaus't nozzle 1S shown on Figure 113. Th1S configuration was used to obtalll baseline data for comparison with the mIxer nozzle noise levels.
5.3.6 Data AnalysIs
5.3.6.1 Static Test Conditions - Durmg the individual test runs, engIne performance was monitored, and relevant ambient and operahng parameters were recorded. Using these data and the appropriate cycle sheet data, predictions of the expected sound pressure levels were computed. These were then compared point by point, frequency by frequency, and angle by angle with the measured sound pressure
180
.... 00 ....
0.5 IN. MICROPHONES
:-: -- -=--r=r=q j
SEMI-INFINITE WAVE GUIDE PROBES
I \ I! I .,:~., ! I : liD, - -------I L' I;~ _1 - --::-:: r ~ - U /r-:.-=-..:..:::-:-~~--l-_ _ _ '--'-- -">-.- '.:'-.. 'L.r----.I
, -~
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\
1 : ~l -1-- r _ L { 0
_ ' I : \ i .. ----- - ----:!-~ ------tnm ~ I " \: \ ---- I ----------=====::HrmlJ-L~" \. -. -III!! ~-------'- ~~--
"
'rJ~'t::-'~t C-nr~",--~ K
l,-~;U!~t~-__ Figure 112. Probe Locations
I-' 00 N
Figure 113. Reference System.
levels. In this manner, estimates of the contribution of each component to the overall nOlse levels at each power setting were made. The predictions were then adjusted to reflect thlS comparison, and the correlation was run again. The insertion loss of the sound treatment was determined along with the mixer noise reduction.
The indlvidual component contributlon to the overall noise levels was determined on a spectrum basls, as shown on Flgure 114. This plot consists of the one-third octave band sound pressure levels over a frequency range from 25 to 20,000 Hz. Th. predlcted fan noise contrlbution was then overlaid. The calculations correctly located the blade pas sing tone, its harmonics, and the br oadband component. The magnltude of the blade pas slng -tone fundamental, however, was underpredlcted. Next, the predlcted jet noise component was added as shown on Flgure 115. As expected, the Jet component does not contnbute dlrectly to the noise levels at the low-power sethng. When the predicted core nOlse component is added to the nOlse spectrum as shown on Flgure llG, the predlcted spectra match the measured spectral shape. The agreement, however, lS only falr m the mid-frequency region through the blade passmg-tone fundamental. Thls same analysis was carrled out for the soHwall and spht-flow conflgurahon. The analysls was also carrled out at each power settmg. The hlgh-power setting is shown on Figure 117. Notice that the agreement is only falr across the mld-and hlgh-frequency reglOns of the spectrum. The low frequency part of the spectra appear to be ln close agreement. Here, the Jet noise component lS predicted to be the predommant source. Based upon this comparisi.Jn and slmilar ones at other power sethngs and configuratlOn, it appear s that the Jet-nOl,se prediction model used lS adequate for the QCGAT program. Consequently, the predlcted Jet-noise levels could be analytically removed from the measured data. The remaining nOlse levels would then be those composed of the core and fan components. Once the Jet component has been removed, the sound power levels attributed to the core were then compared with the predlctedcore sound power levels, as shown in Figure 118 • Also plotted are the sound power levels derlVed from the acoushc probes located in the prlmaryexhaust. The probe data are shown more as a confirmahon of the slope, rather than the sound power levels correctly calculated. The se data mdicate that the core noise model underpre:ilctetes--the core nOlse level by roughly 3 dB. Thls under prediction appear s to be Independent to the power setting of the englne. Therefore, a slmple
183
en r r « ~
C\I ,.... I
0 T"""
w a: CD u
-' w > w -' a: w ~ 0 a.. 0 Z ::J 0 en
I LOW POWER SETTING - HARDWALL CONFIGURATION I 125
120
115
110
105
100
95
90
85
80
MEASURED
PREDICTED
FAN
31.5 63 125 250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 114. Measurements with Predicted Fan
184
(f)
l-I-<t: ~
C\I .-I
0 T"""
w a: m "C
....J w > w ....J
a: w ~ 0 a... 0 z ::> 0 (f)
I LOW POWER SETTING - HARDWALL CONFIGURATION I 125
120
115
110
105
100
95
90
85
80 31.5 63
- MEASURED -- PREDICTED
125 250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure US. Measurements with Jet Added
185
(f) t-t-« ~ N .....
I
0 T"""
w 0: en "0
-.J w > w -.J
0: w ~ 0 a.. 0 Z ::J 0 (f)
I LOW POWER SETTING - HARDWALL CONFIGURATION I 125
120
115
110
105
100
95
90
85
80
186
- MEASURED -- PREDICTED
31.5 63 125 250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 116. Core Added to Measured
I HIGH POWER SETTING - HARDWALL CONFIGURATION I
(/) lI« 3:
130
N 120 .,... I o ~
w a: m "U 110 -I w > W -I
ffi 100 3: o D-
O Z :::> 90 o (/)
MEASURED
PREDICTED
31.5 63 125 250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 117. High Power Setting
187
115 / / 100%
iC -/ /CORRELATION « r- / « 110 / 0 eJ 0 /
0 / W / a: / :J / en 105 /
:) TAILPIPE MIC « / W / ~ /
~ /
100 / 0- FARFIELD DATA 0 /
a: / LL /
/
95 /
95 100 105 110 115 PREDICTED*
*Sound Power Level, dB RE 10-12 Watts
Figure U8. Core Comparison
188
3 dB correchon factor could be apphed to the core nOlse predlction procedure s. After makmg thlS refmement to the core n')lse model, the predlcted-to--measured correlation was then rerun. F1gure 119 shows this comparlsion. The spectral agreement between the measured and predlcted data lS good over the frequency range of intere st. It can be seen that the sound levels m the band containing the tone at blade passmg are also m good agreement; this indicates that the core noise contrlbutes across the spectrum. The domlnance of the core noise can be seen ln F1gure 120. The n01se levels 1n the discharge quadrant are domInated by the core component to the extent that the fan component lS almost enhrely masked. For this reaso'l, the reduchon ln the fan noise levels by the sound treatment was dlfflcult to dlscern. When the core nOlse component lS removed from the 1/3 octave band contammg the blade pas smg tone, and the resultIng blade passmg tone 1S plotted agamst the angle from the mlet, as shown on Flgure 121, a fan-tone directivity pld is formed. The predIcted sound-pressure levels at the peak angles are also shown for the inlet and discharge quadrants. The expected re sults wlth the barrier m place come from the prediction procedures. Only when the barrier is in place will the measured data approach the se hne s, w hlch it doe s as can be seen by the dotte d line s. Th1s plot shows how the fan nOlse contributes to the forward-and aft-radiated engine noise levels. If an observer was to move past the engine, the nOIse levels exper1enced would fir st rIse and then fall off as the observer moved past. Once past the engine, the noise levels would then rIse again as the dlscharge fan noise reached the observer. Th1s 1S roughly how the static data were converted to observed-fhght sound levels. At the high power setting (F1gure 122), the core noise obscure s the aft fan tone from the analysis. A small adJustment was made to the fan noise model from whlch these data were derIved. Th1S adjustment had to do w1th the effect of relahve blade tlp des1gn Mach number. With th1S adjustment, we concluded from the agreement shown here alld on the previous figure that the fan noise model accurately computes the fan noise levels. The sharp dip at the 60-degree point re sults from the fact that the data from o to 40 degrees were recorded at sl1ghtly different power settmgs than the data from 50 to 90 degrees. The predicted data show this same dip. This IS an artifact of the data acquisitIon process and is not a character1stic of the fan noise. The individual component contributions appear to be adequately predicted once the noted corrections have been made. Figure 123 shows a fmal compar1son of the measured
189
I HIGH POWER SETTING - HARDWALL CONFIGURATION I 130
(J) r r « ~ N 120 ....
I
o ,..-
w a: m 1J 110 -' w > w -'
ffi 100 ~ o a.. ()
z :::> 90 o (J)
190
MEASURED
PREDICTED
31.5 63 125 250 500 1000 2000 4000 800016000
FREQUENCY - Hz
Figure 119. Revised Core Comparison
90. C/) 0: 85. w ..... w 2C/) 80. 0...J M« ~O 75. «C/) ...J« wo.. >0 70. wO: ...JO W2 65. 0: 0 ~C\J C/)
60. c/)w wo: O:m 0.."'0 55. 0 z ~ 50. 0 C/)
45.
• Fan Tone Masked By Core Noise
I HIGH POWER SETTING - HARDWALL NACELLE I EXHAUST QUADRANT
SPECTRUM
.r-L,
1
I L
31.5 63 125 250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 120. Core Noise Dominance
191
t-' -.0 N
C/) 0: W
tu 90 ~C/) 0...J ('f)« .-0 «C/)
...J~ Wo 80 >0: ~O W~ 0:
0 ::J('\J C/).
ill ~ 70 a:cn o..'U o Z ::J o
r LOW POWER SETTING - HARDWALL NACELLEI
o 0
WITHOUT BARRIER EXPECTED RESULTS FROM BARRIER TEST BARRIER TEST DIRECTIVITIES
PREDICTED PEAK SOUND LEVELS
? -. \ -~.-.. ~ ... " / . .... '-<. :
,/ ," 0
/
• 0
, "0 \
C/) 60~' ----~--~--~----~--~----~--~--~--~ o 20 40 60 80 100 120 140 160 180
ANGLE FROM INLET, DEGREES Figure 121. Fan Directivity Plots
..... ~ lJJ
C/) a: W l- 90 W ~C/) 0---' Ct)«
l-O «C/)
«
m~ 80 >a: Wo ---'-W~ a: o =>N C/).
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0
\HIGH POWER SETTING =-HARDWALL NACEL.LE)
0---0
,. \ .-/. \
---/ \ .. - / \ \/ /\ \
/ \
\
WITHOUT BARRIER - - -- EXPECTED RESULTS FROM
BARRIER TEST •••••• BARRIER TEST DIRECTIVITIES
o 0 PREDICTED PEAK SOUND LEVELS
120
ANGLE FROM INLET, DEGREES
Figure 122. High Fan Directivity
..... -.0 ~ C/)
I- 140 I-« ~
C\.I T""
« 6 135 I-or«w ocr: 00) W"O cr: -=>.....J C/)W «> wW ~.....J
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130
125
120
rHARDWALL NACELLE I
/ /
100% // CORRELATION
/
o C/) 115LI ____ ~ ____ ~ ____ ~ ______ ~ __ ~
115 120 125 130 135 140 PREDICTED SOUND POWER LEVEL,
dB RE 10-12 WATTS
Figure 123. Final Corn.parison
and pred1cted overall sound power levels. This plot was generated to verify the accuracy of the prediction techniques for the static case before proceed1ng to the flyover analysis. The agreement shown ind1cates that the updated n01se prediction model accurately reflects the static noise emis sions of the QCGA T eng ine.
As noted earlier, it was difficult to discern the n01se reduction of the sound-treatment panels from the far-f1eld data. Figure 124 shows the one-third octave band sound pressure levels at the upstream and downstream m1crophone locations in the inlet. Here, the acoushc energy 1S propagating against the airflow in the inlet duct. The upstream microphone then recorded the inlet n01se after 1t passed through the treated part of the mlet duct. Figure 125 shows the expected msertion los s and the insertion los s derived from the test data; these are the values that will be used in the flyover noise estimates. Figure 126 shows the expected and estimated inserhon loss for the fan discharge-duct sound-treatment panels. The discharge panels had no provision for microphone s and high core noise levels precluded a determination of its noise reduction. It is as sumed that the treatment was funchon1ng properly. The estimated values for the discharge sound-treatment panels are shown on Figure 124.
The Jet noise levels were predicted to be low because of the use of a high bypass-ratio fan. Figure 127 showns the d1fference between the noise spectra of such an engine fitted with the split-flow nozzle conf1guration and with the mixer nacelle configuration. The shaded area represents the stahc n01se reduction of the mixer.
Above 250 hertz, the :::ore noise-source starts to mask the jet noise, and above 1000 hertz the fan is dominant. When flight effects are added, both the mixed and split-flow Jet components will drop leaving the m1xed-flow Jet n01se levels below the core noise levels. The split-flow noise levels would drop and be roughly equal in magnitude to static Jet-noise levels.
5.3.6.2 Flight Predictions - The procedures employed (Figure 128) m the QCGAT program to assess the noise emissions of a QCGAT powered aircraft are the Federal Aviation Adminstration's certification procedure s for turbojet-powered aircr aft (Reference 13). This is a very rigorous method. Basically, the FAA requirements call for measuring the aircraft's noise everyone-half second as the aircraft flies over the measurement point. For this analysis, predicted data were subshtuted for the actual measurement. The dem-
195
196
150.
145.
140.
135.
130
125.
120.
115.
110.
105
FAN BLADE PASSING
TONE
DOWNSTREAM MIC
UPSTREAM MIC
250 500 1000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 124. Sound Treatment
..... ~ -J
12
Ol 10 "0
C/} C/} 8 0 -I
Z 6 0 I-
4 a: w C/}
z 2
0
• Treatment Estimated to Meet Design Specifications
• Analysis Limited By Low Fan Sound Levels
-- GOAL --- MEASURED
APPROACH TAKEOFF
500 1000 2000 4000 8000 16000 500 1 000 2000 4000 800016000
FREQUENCY - Hz
Figure 125. Insertion Loss-Inlet
..... ~ 00
m "'0
(f) (f)
0 -l
z 0 f-a: w (f)
z
14
12
10
8
6
4
2
0
• Treatment Estimated to Meet Design Specification
• Analysis Limited By Low Fan Sound Levels
-- GOAL
--- ESTIMATED
APPROACH TAKEOFF
500 2000 8000 500 2000 8000 1000 4000 16000 1000 4000 16000
FREQUENCY - Hz
Figure 126. Insertion Loss - Discharge
t-' -.0 -.0
~ :J a: ..-() W(J) 0.....-(J)..-a:« w~ ~~ 06 o...r-OW wa: NaJ .....Ju « ~ a: 0 z
0
-S.
-10.
-1S.
-20.
-2S.
-30.
-3S.
-40.
• Greater Than 5 dB
• Coincides with Predicted Results
I ENGINE AT HIGH POWER SETTING I
rJ'---""l.. J "II
r I~ rf...J FAN COMPONENT
-1\ c"l-l_
rJ CORE COMPONENT I
WITH MIXER AND NACELLE
31. S 63 1 2S 2S0 SOO 1 000 2000 4000 8000 16000
FREQUENCY - Hz
Figure 127. Jet Spectra
N o o
DEMONSTRATION ENGINE DESIGN AIRCRAFT CYCLE PERFORMANCE PERFORMANCE
~ t ENGINE PERFORMANCE
~ NOISE PREDICITON
~ AIRCRAFT PERFORMANCE
FOR FL YOVER CONDITION PROCEDURES FOR GIVEN ENGINE SETTING
~ I WING SHIELDING COMPONENT NOISE LEVELS I FLIGHT
EFFECTS t
AIRCRAFT NOISE LEVELS I 1
10.5 SECOND TIME HISTORIES I t
TONE CORRECTED PERCEIVED NOISE LEVELS (PNL T) AT EACH 0.5 SECOND
~ MAXIMUM PNL T AND LENGTH OF
TIME BETWEEN 10 PNdB DOWN POINTS
t EFFECTIVE PERCEIVED NOISE LEVEL~(E~N~)_
Figure 128. Flyover Procedures
onstrated engIne performance and the Beech aircraft desIgn were used to compute the individual test pomt performances. These data were ,then _entered into the prediction procedure s, The appropriate flight and wing shIelding effects were then applied to the individual component noise predictions. The aircraft noise signature was then derived by combining these into a table of aircraft noise. Then by analyhcally mOVIng the aIrcraft nOIse table past the measurement pomt, the hme hIstory of the flyover could be constructed for each half-second interval. These sound levels were then used to compute the tone corrected perceived noise levels for the flyout event. The maximum tone-corrected perceIved nOIse levels were then found along WIth the hme the aircraft nOIse was withing 10 PNdB of the maXImum. The effechve perceIved nOIse levels were calculated from these data.
Figure 129 shows the tone-corrected perceived noise levels versus time for the approach flyover. The maximum tone-corrected perceived noise level, labeled PNLTM, would occur after the aircraft had passed directly overhead. The time that the PNLT was within 10 PNdB of the value would be 8.5 seconds. This plot also shows that the fan inlet and discharge noise would be heard at separate times. The valley between the peaks was caused by the lower sound levels generated at the sideline positions. Wing shielding, the shaded l)ortion, would act to cut the inlet peak off early and make the valley deeper. The core noise component would be heard after the aircraft passed, as most of the core nOIse IS In the aft quadrant of the engine. Because of the duration correction, the fan component noise levels are higher and contribute more to the effective perceived noise levels.
The takeoff flyover tone-corrected perceived noise level time history is shown similarily in Figure 130. The time interval that the noise was within 10 PNdB of the maximum would be much longer. At the approach condition, the altitude at flyover would be 113 meter s (370 feet). For the takeoff condition, it would be 792 meters (2600 feet). Consequently, the time the aircraft requires to fly past would be considerably longer. The maximum tone-corrected perceived noise level would also occur much later as the sound would take longer to reach the observer because the dominant noise sources are the core and jet. These components radiate most of their acoustic energy to rear quadrants and, as such, it would not be heard until the aircraft has flown past the observer. Also shown here are the higher noise levels of a split-flow nozzle configured aircraft. Here the jet component would contribute more to the aircraft noise levels In both
201
N o N
o Wen >"'0 WZ 00.. CI:~ W~ 0.....J
80
OZ We:. 70 ~...J Ow w> CI: w CI:...J Ow
°00 w-zO OZ ~ 60
PNLTM
PNL T -10 dB
, , , , , , , , , , , I ,
I
PNL T -10 dB
I.. 8.5 SEC
I tzl1 WING SHIELDING 1 EPNL, EPNdB =
77.3 - OVERALL 72.2 - CORE 73.9 - FAN
-10 -5 0 5 10 15
TIME FROM FLYOVER, SECONDS
Figure 129. COIllponent Contribution - Approach Noise
N o w
o W(!) >-0 WZ 00... a:-W-0...1--1 OZ wo... 1--0-1 wW a:> a: W 0-1 OW
CJ) WzO OZ I-
70
60
PNLTM
." ."
'"
." ., ."
." ."
,"_/ , ,,
." -~----------
~ ~38SEC
· · · ---"'--;., ~A> : "IV(
~j:)",.( " ;>
d..? r-.s /:) t.. /">- "/:) I\t. I f:"t.. CIS
0*,. : ...... ~ : 4t/~ ...... 6'B • 1~' 0<1.1:-. ; () ;-, ,''.OI\t. · <01,,',,0'&
"1/,. ~, ~
PNL TM -10
50 I I I I , I I , I I ,
-15 -10 -5 0 5 10 15 20 25 30 35
TIME FROM FL YOVER, SECONDS "-----
Takeoff Flypver Noise Time History at Observer Location 3.5 Nautical Miles Down Range from Brake Release
*No Sound Treatment
Figure 130. Com.ponent Contribution - Takeoff Noise
204
magnitude and duration. The duration would be increased because the jet noise peaks further aft than the core noise. This means that the peak noise would occur later in the flyover. Thtls, the addition of the mixer not only reduces the aircraft flyover noise levels, the alrcraft nOlse d.oes not hnger as long.
5.3.7 Conclusion
For an aircraft powered by two Avco Lycommg QCGAT engines installed in a nacelle that includes a mixer, fan inlet and discharge sound-treatment panels, and mounted over the wing s, the effective perceived noise levels for the takeoff, sideline, and approach conditions would be 68.4, 70.6, and 77.3 EPNdB, respectively. These noise levels, shown in Figure 131, are below the limits set by the QCGAT program goals. In the analysis, the effect of several alternative engine configurations on the aircraft noise was assessed. For example, removal of the sound-treatment panels would add 2 EPNdB to the approach noise levels and still be below the QCGAT goals. The other positions would not be affected appreciably.
When the iterations are completed for this engine design, the increased thrust of the engine means that the aircraft will achleve an altitude of 3600 feet over the takeoff point ver sus the pre sent 2600 feet. ThlS will result in a 3 EPNdB reduction ln the takeoff noise levels and a 1 EPNdB reduction in the sldeline noise levels. In this case, the spht-flow exhaust nozzle configuration would be within I EPNdB of the QCGAT goals. FIgure 132 shows that the Avco Lycoming QCGAT engine's effectlve percelved noise levels plotted against the Federal AviatlOn Administration's Stage III noise standards and the high technology levels that were used by NASA for the QCGAT program goals. This demonstrates that the technology which has worked for the large engine can be applied to the general-aviation size engine.
In summary, large turbofan nOIse-control technology was successfully applied to a general-aviation size engine. The stringent program goals set by NASA forced a concept that required integration of a quiet fan design with the nacelle and aircraft to provide an effiCIent propulsion system.
Static noise tests have demonstrared that the QCGAT program goals can be met with the latest noise-control techniques wlthout incurring a performance penalty.
N o U1
aCGAT ACOUSTIC PERFORMANCE
CONDITION
Takeoff Flyover
Takeoff Sideline
Approach Flyover
EPNl GOAL DELIVERED ENGINE EPNdB EPNl,EPNdB
69.4 68.4
78.4 70.6
83.4 77.3
Figure 131. Noise Levels - QCGAT Goals
N 0 0'
CO "0 z a.. w --1 Z a.. w
SIDELINE 110
100
90
80
70
60
APPROACH 110
100
90
80
70
60 ............... ~ 3 10 100 1000 (Ib) I I I I
1.3 4.5 45 450 Kg GROSS TAKEOFF WEIGHT
x 1000
TAKEOFF
f-
~ f-
IfL. If /. . .. ........ ... .. ... f-
.... .... ...... ~ .. ··S
I I" I I I I I11I I I I IIII
3 10 100 1000 L I I !
1.3 4.5 45 450
U{{(UUU FAR-36 LIMIT
............... QCGAT GOAL
.... AVCO LYCOMING ~ QCGAT
Figure 132. Noise Levels - FAA Standards
6. a SUMMARY OF RESULTS
6. 1 OBJECTIVES
The objective of the NASA Quiet, Clean General Aviation Turbofan (QCGAT) engine program was to demonstrate the applicability of large turbofan engine technology to small general aviation engines. Specifically, reduction in the levels of noise and pollutant emissions, along with fuel consumption, were addressed.
Avco Lycoming participated in the NASA /:2CGA T engine program by designing and developing a 7,000 N (1,600 lbf) thrust turbofan that used the Lycoming LTS 101 as a core engine. The prime areas investigated in order to meet the NASA requirements were:
1. Define the engine and determine its characteristics and requirements for a QCGAT engine applicable to general aviation aircraft. (This included a preliminary design of an aircraft by the Beech Aircraft Corporation).
2. Design and fabricate the new and modified parts required to be used with an existing gas generator core in the turbofan engine.
3. Perform evaluation test of critical components
4. Perform evaluation tests of the QCGAT engine
5. Design and fabrication of an acoustically treated nacelle
6. Measure engine noise,emlsslonS, and sea level static overall engine performance to establish validity of predictions prior to engine delivery to NASA_
7. Deliver a quiet, clean, turbofan engine, an acoustically treated nacelle, and engine test support hardware to NASA.
6.2 TEST RESULTS
QCGAT engine test results compared with the predicted design performance goals are tabulated below.
207
6.2. 1 Overall Engine Performanc e Demonstrated
Sea Level Takeoff Deviation from (Standard Day Installed) Goal Test Design (%)
Thrust N 7166 6485 -9.5
(lbf) (1611) (1458 )
Specific Fuel Consumption
Kg/N-hr 0.0370 0.0400 +9.5
lbm/lbf-hr (0. 363) (0.392)
Design Cruise 7620 M (25,000 it), Mach 0.6
Thrust N 2157 1850 -'- -14.2 (416) -.' (lbf) (485)
Specific Fuel Consumption
Kg/N-hr 0.064 0.074 ~:~ +15. 1
lbm/lbf-hr (0.628) (0. 723)
* Estimated from Static Data
The overall performance did not meet the desired goals. However, subsequent component testing has identified the deficiency to be in the matching of the high-pressure turbine and the power turbine. Since the hardware design and procurement were not within the NASA tirneframe, the engine performance demonstration test was conducted with existing hardware. The design goals are viable and achievable with the Lycoming QCGAT engine.
6.2.2 Emissions
Goal~:~
Unburned Hydrocarbons (UHC) • 045 (1.6)
Carbon Monoxide (CO)
208
.266 (9.4)
Test * .017 (.62)
.193 (6. 8)
Deviation
-62.2%
-27.4%
Oxides of Nitrogen (NO) x
Smoke Number
* - Units gm/KNsec
• 105 (3. 7)
45
(lbm/1000 lbf Thrust hr-Cycle)
Test*
.106 (3. 75)
24
Deviation
+1%
-46. 7%
The Lycoming QCGAT engine test demonstrated that all emission goals were easily achievable with the exception of NO which was slightly over the design estimated goal. x
6.2.3 Acoustics
The NASA program goals were 15 to 20 EPNdB below the FAR Part 36 nOise levels. The high technology levels of the LycomIng QCGAT program demonstrated performance levels better than the goals, as tabulated below:
Noise
Condition Goal Measure Iluprovement {EPNLdB) (EPNLdB! {dB!
Cd,t/ Takeoff Flyover 64.4 68.4 1.0
Take off Side hne 78.4 70.6 7.8
Approach Flyover 83.4 77.3 6. 1
6.3 CONCLUDThlG REMARKS
Efforts of the NASA/QCGAT program have contributed to a new engine that is designed to serve the needs of the general aviation industry in the 1980 1 s. The engine and nacelle designs have demonstrated the primary program objectives. Technology to reduce engine n01se has been succ.essfully applied to the general aviatlon size engine to obtain the acoustic goals With margin. Any foreseeable ;!.coustic or pollutant emiSSion requirements Will offer no major constraint.
Challenging objectives for a fuel-efficient, ecological aircraft were set for the preliminary design to respond to our assessment of general aviation market needs for the next decade. The aircraft design achieves these objectives to provide a truly quiet, clean, six-place, long-range capability that will be attractive to both the user and suburban community.
REFERENCES
1. Wharton, H. E., AVCO LYCOMlliG QUIET CLEAN GENERAL AVIATION TURBOFAN (QCGAT) ENGlliE NACELLE SOUND ABSORPTION PANEL DESIGN, Lockheed California Co. Report LR 28254, July 20, 1977.
2. Tyler, J. M. and Sofrin, T. G., AXIAL FLOW COMPRESSOR NOISE STUDIES, SAE345D, 1961.
3. Ko, S. H., SOUND ATTENUATION lli ACOUSTICALLY LINED CIRCULAR DUCTS lli THE PRESENCE OF UNIFORM FLOW AND SHEAR FLOW, Journal of Sound and Vibration, Vol. 22 (2), pp. 193-210, 1972.
4. Rice, E. , SPINNING MODE SOUND PROPAGATION IN DUC TS WITH ACOUSTIC TREATMENT AND SHEARED FLOW, AIAA Paper 75-519, March 1975.
5. Prydz, R. A., BOUNDARY LAYER CONSIDERATIONS FOR OPTIMIZATION OF ACOUSTIC LINERS FOR AIRCRAFT ENGlliE DUCTS, SAE Paper 760896, November 1976.
6. Dunn, D. G. and Peart, N. A. , AIRCRAFT NOISE SOURCE AND CONTOUR ESTIMATION, NASA CR 114649, July 1973.
7. Laurence, J. C. and Benninghoff, J. M., TURBULENCE MEASUREMENTS IN MULTIPLE lliTERFAClliG Am. JETS, NASA TN 4029.
8. Hurley, J. F., I'Anson, L. and Wilson, C. A., DESIGN OF AN EXHAUST MIXER NOZZLE FOR THE AVCO LYCOMlliG QUIET CLEAN GENERAL AVIATION TURBOFAN (QCGAT), NASACR-159426, August 1978.
9. Lipert, F. W., CORRELATION OF GAS TURBINE EMISSIONS DATA, ASME 72-GT-60.
10. EMISSIONS STANDARDS AND TEST PROCEDURES, Title 40, CFR Part 87, published in the Federal Register, July 17,1973.
11. U. S. Patent 3,671,171, ANNULAR COMBUSTORS, Brian W. Doyle.
210
12. U. S. Patent 3,645,095, ANNULAR COMBUSTOR, Jerry O. Melconian.
13. FEDERAL AVIATION ADMINLSTRATION NOISE STANDARDS Title 14, Code of Federal Regulation, Chapter L Part 36 Amendment 7.
14. Stone, J. R., INTERIM PREDICTION METHOD FOR JET NOISE, NASA T MX-71618 (1974).
15. Huff, R. G., Clark, B. J., and Dorsch, R. G., INTERIM PREDICTION METHOD FOR LUW FREQUENCY CORE ENGINE NOISE, NASA TMX-71627 (Nov. 1974).
16. Heldmann, M. F., INTERIM PREDICTION METHOD FOR FAN AND COMPRESSOR NOLSE SOURCE, NASA TX-71763 (June 1975).
17. Gillam, R.E., etal,ANOPP USERS MANUAL, (July 1978).
211
212
Distribution Llst - QCGAT Flnal Report
Addressee Number of Cople s
1. NASA SClentlflc and Technlcal Information Faclhty p. O. Box 8757 Bal't/Wash. InternatlOnal Airport MD 21240 Attn: Acces siomng Dept.
2. NASA LeWlS Research Center 21000 Brookpark Road Cleveland, OhlO 44135 Attn: G. K. Siever s MS 301-2
-NASA Project Manager L. Schopen MS 500-305
-NASA Lewis Contractlng Officer Lewis Library MS 60-3 Report Control Office MS 5-5 J. F. McCarthy, Jr. MS 3-2 J. M Kline be r g MS 3 - 3 W. L. Stewart MS 3-5 D. Poferl J. R. Esterly C. E. Feller J. F. Gr oeneweg E. J. Rlce W. L. Jones J. G. McArdle U. H. von Glahn J. R. Stone
M. Reshotko J.A. Yuska W. H. Fenmng A. S. Moore D. D. Chrulskl R. J. Weber K. L. Abdalla D. L. Bresnahan
MS 500-207 MS 500-207 MS 54-3 MS 54-3 MS 54-3 MS 500-208 MS 500-208 MS 54-3 MS 54-3
MS 54-3 MS 500-208 MS 500-208 MS 500-208 MS 500-208 MS 501-10 MS 500-127 MS 500-127
25
1
1 2
1 1
1 I 1 1 1 1 1 1 1
1 1
1 1 1
1 1
1 10
1
Addressee
NASA LeW1S Research Center (cont'd)
3.
4.
E. A. KreJ sa D. A. Petrash R. E. Jones E. A. Lezberg J. S. Fear
M. A. Behelm D. Draln D. L. Nored M.R. Vanco R. Rudey M. J. Hartmann W. J. Ande r son L. P. Ludwlg D.P. Townsend C. L. Ball L. Reld R. D. Moore H.E. Rohllk R. Y. Wong F. S. Stepka R.C. B1U J. J. Coy P. L. Meltner C.P. Blankenship L. M. Hlbben E.R. Hersman L. A. POVlnelli
NASA Headquarters 600 Independence Avenue SW Washlngton, D. C. 20546 Attn: Assoc. Admln, OAST
W. B. Olstad -Dep. Assoc.
Admln W. S. Alken, Jr. H. Johnson R. Rose G. Banerlan
NASA Langley Research Center Hampton, Vlrglnla 23365 Attn: O. W. N1Cks
MS 54-3 MS 60-6 MS 60-6 MS 60-4 MS 54-6
MS 86-1 MS 100-1 MS 301-2 MS 301-4 MS 60-4 MS 5-3 MS 23-2 MS 23-2 MS 6-1 MS 5-9 MS 5-9 MS 5-9 MS 77-2 MS 77-2 MS 77-2 MS 23-2 MS 6-1 MS 77-2 MS 1 J5-I MS 500-211 MS 21-4 MS 86-1
Code R-1 Code RD
Code RJ-2 Code RJG-4 Code RJG-4 Code R TP-6
MS I03A
Number of Cople s
1 1 1 1 1
1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1
1 1
1 3 1 1
1
213
Addressee
NASA Langley Research Center (cont'd)
5.
J.P. Raney J. W. Stlckle B. J. Holrne s
NASA Ames Research Center Moffett Fleld, Cal1forma 94035 Attn: L. Ro~)erts
T. L. Galloway M. D. Shovl1n
6. NASA Dryden Fhght Research Center P.O. Box 273
7.
Edwards, Cal1forma 93523 Attn: I. T. Glllam
E;1vlronmental Protectlon Agency Crystal Mall, Bldg. 2 1921 Jeffer son DaVls Hlghway Arl1ngton, Vlrglnla 20460 Attn: H. J. NOZ1Ck
J. Shettlno W. Sperry
8. De pt. of Tr ans por tatlOn ARD 500 Washington, D. C. 20591 Attn: H. True
9. DOT/FAA 8~O Independence Ave, SW Washlngton, D. C. 20591 Attn: J. O. Powers
C. Foster R.N. Tedrlck
10. DOT/FAA
11.
214
Nat'l Av. Fac. Exp. A"lantlc Clty, New Jersey 08405 Attn: J. F. Woodall
AF Systems Command Llalson Offlce Lewls Research Center, NASA 2100 Brookpark Rd. Cleveland, OhlO 44135 Attn: Maj. A.J. Wllloughby
MS 461 MS 246A MS 247
MS 200-3 MS 237-9 MS 237-10
Number of Cople s
1 1 1
1 1 1
1
1 1 1
1
1 1 1
1
1
Addressee
PropulslOn Laboratory US Army Res o & Tech. Laboratories LewIs Re se ar ch Center, NASA 21000 Brookpark Road Cleveland, Ohlo 44135 Attn: J. Acurlo
13. E.E. Balley
14.
Te chnlcal Llals on AFAPL/DO Wrlght Patterson AFB, OhlO 45433
Wrlght Patterson AF Base Dayton, OhlO 45433 Attn: W. C. Elrod
E. Klenke E. Lake R. L. Sp'.::ncer
15. US Army Re s. & Techn. Lab Appl1ed Technology Lab
16.
Ft. Eustis, Vlrglnla 23604 Attn: C. A. Ell10t
H. Morrow
Naval Alr Development Center Warmlnster, Pennsylvanla IS 974 Attn: S. R. Shaplro
17. AlResearch Manufacturlng Co. 402 South 36 Street Phoenlx, Ar lzona 85010 Attn: F. B. Wallace
R. Heldenbrand W. Norgren L. Klsner W. Blackmore F. Lewls
IS. Garrett Corporatlon P. O. Box 9224S Los Angeles, Cal1fornla 90009 Attn: 1. E. Spear
J. Mason
19. Garrett Cor poratlOn 19201 Susana Road Compton, Cal1fornla 90221 Attn: G. Paden
AFIT/ENY ASD/ENFP AFAPL/POP 1 AFAPL/TBD
Number of Cople s
1
2
1
1 1 1
1 1
1
1 1 1 1 1 1
1 1
1
215
216
Addressee
20. Garrett Cor poratlon Dayton Offlce 333 West Flrst Street Dayton, OhlO 45402 Attn: B. Tyson
21. Garrett Corporatlon SUlte 419 1625 I Street, NW Washlngton, D. C. 20006 Attn: M. S. Rachhn
22. Boelng Commerclal Alrplane Co.
23.
P.O. Box 3707 Renton, Washlngton 98124 Attn: C. G. Hodge
The Boelng Company 3801 S. Ohver Street Wlchlta, Kansas 67210 Attn: M. D. Nelson
24. Beech Aucraft Corp. 9709 East Central Wlchlta, Kansas Attn:
67201 R. L. Beneflel F. C. Umcheldt W.D. Wlse
25. Bolt, Beranek & Newman, Inc. 21120 Vanowen St. Canoga Park, Attn:
Cahforma 91303 W. J. Galloway J. F. Wllby
26. Cessna Alrcraft Company P.O. Box 1521 Whlchlta, Kansas 67201 Attn: M. S. Harned
27. Cessna Alrcraft Company Pawnee D1V1Slon Wlchlta, Kansas 67201 Attn: A. Asher
28.. Cessna Alrcraft Company Wallace Dlvlsion P. O. Box 7704 Wlchita, Kansas 67201 Attn: E. Kraus
Number of Coples
1
1
1
1
1 1
1
1 1
I
1
1
Addressee
29. Detrolt Dlesel Allison Dlvlsion of General Motor s SUlte 312 333 West Flrst Street Dayton, Ohlo 45402 Attn: F. Walters
30. Douglas Alrcraft Company 3855 Lakewood Blvd. Long Beach, Cal1forma 90846 Attn: R. Pendley
31. GAMA 1025 CT Ave. NW #1025 Washlngton, D. C. 20036 Attn:. S.J. Green
32. Gates Learjet Corporation P. O. Box 7707 WlchIta, Kansas 67277 Attn: E. Schlller
R.C. Scott
33. General ElectrIC Co. P.O. Box 81186 Cleveland, OhlO 44181 Attn: M. H. Rudasll
34. General Electrlc Co. 1000 Western Ave.
35.
Lynn, Massashusetts 01910 Attn: A. J. Koschler
Grumman Aerospace Corp. Bethpage, New York 11714 Attn: C. A. Hoelzer
36. Gulfstream Amerlcan Corp. P. O. Box 2206 Savannah, Georgla 31402 Attn: R. J. Stewart, MS D-04
Number of Coples
1
1
1
1 1
2
1
1
1
217
218
Addressee
37. Learavla P. O. Box 6000 Reno, Nevada 89506 Attn: I. Gllchrlst
38. Lockheed Georgla Co. 86 South Cobb Drlve
39.
40.
Marletta, Georgla 30063 Attn: H. J. Rubel Dept. 72-73
W. A. French Dept. 72-47
Piper Alrcraft Corp. Lock Haven, Pennsylvanla 17745 Attn: C. L. Kuntz
M.J. Dees, Jr.
Pratt & Whltney Aircraft 20800 Center Rldge Road Rocky Rlver, OhlO 44116 Attn: J. C. Chew
41. Rockwell InternatlOnal Corp. General AVlatlon D1V1Slon 5001 North Rockwell Ave. Bethany, Oklahoma 73008 Attn: L. L. McHughe S
42. Rohr Corporatlon P. O. Box 878 Chula Vlsta, Cahfornla 92012 Attn: F. Horn
43. Solar Turblne Internatlonal P. O. Box 80966
44.
San Dlego, Cahforma 92138 Attn: G. Hosany MZR 1
Teledyne CAE 1330 Laskey Road Toledo, OhlO 43601 Attn: B. Slgnh
R. Snook H. Paget E. Benstein
Number of Cople s
1
1 1
1 1
2
1
1
1
1 1 1 1
Addressee
45. UnlVerslty of Oklahoma 865 ASP Norman, Oklahoma 73019 Attn: Dr. K. H. Bergey
46. Vought CorporatlOn P. O. Box 5907
47.
48.
49.
50.
Dallas, Texas 75222 Attn: S. Laden
Wllhams Research Corp. 2280 West Maple Road P.O. Box 95 Walled Lake, Mlchlgan 48088 Attn: E. Lays
Llbrary
Unlver Slty of Kansas Department of Aerospace Englneerlng Lawrence, Kansas 66045 Attn: J. Roskam
General AVlatlon Assoc., Inc. Basklng Rldge, New Jer sey 07920 Attn: J. W. Olcott
Alrcraft Owners and Pllots AssoclatlOn 7315 WlsconSln Avenue, N. W. Washlllgton, DC 20014 Attn: J. L. Baker
51. NatlOnal Buslness Alrcraft Assoclatlon, Inc. One Farragut Square Washlngton, DC 20006 Attn: J. H. Wlnant
52. Avco Aerostructures Dlvlslon P. O. Box 210 Nashvllle, Tennessee 37202 Attn: J. Burkard
53. Avco Lycomlng Englne Group 652 Ol1ver Street
54.
Wllhamsport, Pennsylvanla 17701 Attn: S. Jednzlewskl
J. Duke
Lockhead Cahfornl.a Company Burbank Cal1fornlcl. 91503 Attn: H. E. Wharton
Number of Cople s
1
1
1 1
1
1
1
1
1
1
1
1 219
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