Fuel conservation merits of advanced turboprop transport aircraft

TABLE OF CONTENTS Section Page FOREWORD iii LIST OF FIGURES vii LIST OF TABLES xi SUMMARY xiii INTRODUCTION xix ABBREVIATIONS/SYMBOLS/CONVERSIONS xxv 1 AIRCRAFT DESIGN EVALUATION 1-1 1.1 TURBOFAN AIRCRAFT 1-1 1.1.1 Baseline 1-1 1.1.2 1985 Technology Assessment 1-5 1.1.2.1 Design Range Increase 1-8 1.1.2.2 Cruise Speed Reduction 1-12 1.1.3 1990 Technology Assessment 1.14 1.1.3.1 Advanced Technology Turbofan Engine 1-14 1.1.3.2 Aircraft Optimization 1-14 1.2 TURBOPROP AIRCRAFT 1-16 1.2.1 RECAT Baseline 1-22 1.2.1.1 Revised Baseline 1-26 1.2.2 1985 Technology Assessment 1-28 1.2.2.1 Design Range Increase 1-40 1.2.2.2 Cruise Speed Reduction 1-42 1.2.2.3 Alternate Turboshaft Engine 1-42 1.2.2.4 Off Design Cruise Speed Effects 1-45 1.2.3 1990 Technology Assessment 1-45 1.2.3.1 Advanced Technology Turboshaft Engine 1-46 1.2.3.2 Aircraft Optimization 1-46 2 PROPFAN TECHNOLOGY BASE 2-1 v

-. PRECEDNG PAGE- BLANK

NOT FILMS

TABLE OF CONTENTS (Continued) Section Page 3 ACOUSTIC TREATMENT METHOD 3-1 3.1 TYPICAL TURBOFAN CABIN NOISE ENVIRONMENT 3-1 .3. 2 PROPFAN PASSENGER COMFORT CRITERION 3-2 3.3 PROCEDURE FOR DEFINING ACOUSTIC TREATMENT WEIGHT 3-4 3.4 CABIN WALL TRANSMISSION LOSS PREDICTION 3-4 3.4.1 Transmission Loss Assumptions 3-4 j.4.2 Structural Design Constraints 3-7 3.5 EXTERIOR NOISE DATA ' 3-14 3.5.1 Previous and Current Prediction Results 3-14 3.5.2 Exterior Near Field Noise Prediction 3-18 3.6 ACOUSTIC TREATMENT WEIGHT PENALTY DATA 3-19 3.6.1 Presentation of Acoustic Treatment Weight Penalty Data 3-19 3.6.2 Disk Loading Effects on Acoustic Treatment 3-19 3.7 POSSIBILITIES FOR FUTURE TREATMENT WEIGHT REDUCTIONS 3-39 3.8 CONCLUDING REMARKS CONCERNING ACOUSTIC TREATMENT WEIGHT PENALTIES 3-41 42 COST/BENEFIT COMPARISONS 4-1 4.1 PERFORMANCE COMPARISON 4-4 4.2 ECONOMIC COMPARISON 4-7 4.3 MISSION FUEL COMPARISON 4-7 5 CONCLUSIONS AND RECOMMENDATIONS 5-i 5.1 PROPFAN DESIGN 5-1 5.2 AIRCRAFT ACOUSTIC TREATMENT 5-2 5.3 AIRCRAFT CONFIGURATIONS 5-3 5.4 ADVANCED TECHNOLOGY ENGINES 5-3 REFERENCES R-1 vi LIST OF FIGURES Figure Page 1 Domestic Mission Flight Profile xxii 2 General Arrangement - Turbofan Aircraft 1-3 3 Selection of Baseline Turbofan Airplane Design 1-6 4 Turbofan DOC Versus t/c - 2000 nm Range 1-10 5 ASSET Crossplot - Turbofan DOC (2000 nm Range) 1-11 6 ASSET Crossplot - Turbofan DOC (0.75 Mach) 1-13 7 ASSET Crossplot - Turbofan DOC (STF 477 Engine) 1-17 8 Propfan Efficiency and Propulsion System Weight Trends with Disk Loading 1-20 9 TOGW, DOC, and Block Fuel Versus Propeller Disk Loading 1-21 10 General Arrangement - Baseline RECAT Propfan Aircraft, 1-23 11 Effect of Disk Loading on Propfan Efficiency 1-27 12 DOC Sensitivity to Propfan Efficiency 1-30 13 Block Fuel Sensitivity to Propfan Efficiency 1-31 14 Effect of Acoustic Weight on DOC Savings 1-32 15 Effect of Acoustic Weight on Block Fuel Savings 1-33 16 Propfan DOC Versus t/c - 2000 nm Range 1-34 17 ASSET Crossplot - Propfan DOC (2000 nm Range) 1-35 18 Effect of Mach Number on Engine SFC 1-36 19 ASSET Crossplot - Propfan DOC (0.75 Mach) 1-37 20 ASSET Crossplot - Propfan DOC (PD370-22 Engine) 1-38 21 ASSET Crossplot - Propfan DOC (STS 487 Engine) 1-39 22 Double "Limp Wall" Concept for Acoustic Treatment of Cabin Walls 3-3 23 Required Treatment Length with Relative Tip Clearance 3-9 24 Required Treatment Length with Relative Tip Clearance 3-10 vii

ORIGINAL

PAGE

IS

QUALITY POOR OF LIST OF FIGURES (Continued) Figure Page 25 Example of 5 Step Acoustic Treatment for a Relative Tip Clearance of Ay/D = 0.8 3-11 26 Acoustic Treatment Weight, 8 Blades, 800 fps, 3860 lb Thrust 3-20 27 Acoustic Treatment Weight, 8 Blades, 800 fps, 3000, lb Thrust 3-21 28 Acoustic Treatment Weight, 8 Blades, 700 fps, 3860 lb Thrust 3-22 29 Acoustic Treatment Weight, 8 Blades, 700 fps, 3000 lb Thrust 3-23 30 Acoustic Treatment Weight Required Versus Tip Speed 3-24 31 Propfan Efficiency Versus Tip Speed 3-24 32 Acoustic Treatment Weight, 10 Blades, 800 fps, 3860 lb 3-26 Thrust 33 Acoustic Treatment Weight, 10 Blades, 800 fps, 3000 lb 3-27 Thrust 3-28 34 External SPL Correction Versus Relative Tip Clearance 35 Blade Passage Frequency Versus Disk Power Loading 33-29

36 Propeller Diameter Versus Disk Power Loading 3-29

.37 External SPL at Blade Passage Frequency Versus Disk 3-30 Power Loading 38 Total Acoustic Treatment Area Required Versus Disk 3-31 Loading 39 Total Acoustic Treatment Length Versus Disk Loading 3-34 4o Segment Length/Diameter Versus Disk Loading 3-35 41 Segment Total Wall Weight Versus Disk Loading 3-36 42 Segment Area Versus Disk Loading 3-36 43 Acoustic Treatment Weight Increment Versus Blade 3-38 Passage Frequency

44 Modal Density for Acoustically Fast Modes 3-4o

45 Fuel and Cost Comparison for Baseline Aircraft of 4-2 Previous RECAT Study vii.

LIST OF FIGURES (Continued) Figure Page 46 Fuel and Cost Comparison for Revised Baseline Propfan Aircraft 4-3 47 Effect of Design/Mission Characteristics on Aircraft Cruise SFC 4-5 48 Effect of Design/Mission Characteristics on Aircraft 4-6 Empty Weight 49 Effect of Design/Mission Characteristics on Aircraft 4-8 (1500 nm) DOC 50 Effect of Design/Mission Characteristics on Aircraft 4-9 DOC (475 nm) 51 Effect of Design/Mission Characteristics on Block 4-10 nm) Fuel (1500 52 Effect of Design/Mission Characteristics on Block 4-1

am)

Fuel (475 53 Fuel Savings of Propfan Versus 1985 IOC Turbofan 4-13 Aircraft 54 DOC Savings of Propfan Versus 1985 IOC Turbofan 4-14 Aircraft ix LIST OF TABLES Page Table xiv 1 Propfan Savings for Study Conditions 2 Fuel Efficiency for Study Conditions xv xvi 3 Effect or Study Conditions on Aircraft Performance 4 Cost Factors xviii xxvi 5 Study Ground Rules 1-1 6 Study Matrix for Aircraft Configurations Characteristics 1-4 7 Turbofan Baseline Aircraft 8 Design and Performance Characteristics of Turbofan Aircraft 1-7 1-8 9 Engine Features for JT1OD-2 Turbofan Turbofan (2000 rm Range) 1-9 10 Parametric Study Matrix - 11 STF 477 Engine Parameters 1-15 Engine Installation Losses 1-18 1-24 Baseline RECAT Propfan Aircraft Characteristics Characteristics 1-25 14 Baseline RECAT Engine Characteristics - Propfan 15 Design and Performance 1-29 Aircraft 1-41 Parametric Study Matrix 1-44 Engine Data Comparison 18 Double Wall Mass Law - "Limp Wall" Theory 3-6 19 Wall Weight Constraints for 5 Step Double Wall Treatment 3-13 20 Comparison of Previous RECAT Versus Current Cabin Noise Treatment Methodology and Data 3-15 3-16 SPL Results 21 External 22 Comparison of Harmonic Levels of External SPL Data 3-17

PREQIG PAGE BLANK NOT FILN

xi SUMMARY This study is an extension of the previous RECAT propfan studies, documented by NASA Report CR 137926, which show savings in fuel and operating costs for a 1985 IOC propfan aircraft, at a 1500 nautical mile, Mach 0.8 design mission and 00/gallon fuel cost, of 17.8 and 8.2 percent respectively when compared to an equal technology turbofan aircraft. The objective of this study was the examination of further potential savings in fuel and operating costs for the propfan aircraft by incorporating: * New propfan data " Revised design mission profiles " Additional engine performance characteristics The results of this study are summarized in Tables 1, 2, and 3, and show that: o New propfan data does not alter the previous RECAT findings - impact of higher exterior noise levels is offset by measured directivity characteristics and higher propulsive efficiency. Fuel and DOC savings of the revised baseline propfan over the baseline turbofan is 17.6 and 7.8 percent respectively.

acoustic treatment requirements and " Reduced cruise speed decreases fuel efficiency of the propfan aircraft. Fuel and DOC increases savings of the propfan over the turbofan (both at Mach 0.75 cruise), is 21.0 and 10.0 percent respectively.

" Incorporation of the Allison PD 370-22 turboshaft engine confirms the the propfan over the RECAT results. Fuel and DOC savings of previous baseline turbofan is 17.8 and 10.1 percent respectively.

engines offer fuel and DOC savings * Incorporation of 1990 technology engines.

of 11 and 7.5 percent respectively over the 1985 technology and DOC savings of the 1990 propfan over the 1990 turbofan is Fuel 17.1 and 7.8 percent respectively.

xiii , NOT FIL.

PAGE BLANK

PRECEDENG

TABLE 1. PROPFAN SAVINGS FOR STUDY CONDITIONS SAVINGS OF PROPFAN OVER TURBOFAN--PERCENT DESIGN MISSION (100% L.F.) 475 N.M. (58% L.P.)

BLOCK STUDY CONDITION BLOCK FUEL DOG (6o FUEL DOC (600) Previous RECAT Baseline 17.8 8.2 20.4 8.2 (1500 N.M., Mach

o.8)

Revised Baseline - New Propfan Data 17.6 7.8 19.6 8.1 (1500 N.M., Mach 0.8) Design. Range Increased 16.5 8.0 18.5 8.3 (2000 N.M., Mach 0.8) Cruise Speed Reduced 21.0 10.0 22.9 10.2 (1500 N.M., Mach 0.75) Higher Press.

Ratio Engine - PD 370-22 17.8 10.1 19.8 10.3 (1500 N.M., Mach 0.80) Engine Technology 17.1 7.8 19.1 8.1 (1500 N.M., Mach 0.80) TABLE 2, FUEL EFFICIENCY FOR STUDY CONDITIONS DESIGN RANGE - 100% L.F. 475 NM, - 58% L.F.

TURBOFAN PROPFAN TURBOFAN PROPPAN STUDY CONDITION (LB/ASM) (LB/ASM) (LB/ASM) (LB/ASM 'Previous RECAT Baseline .0956 (.0786 .1808 .1439 Revised Baseline .0956 .0788 .1808 .1454 Design Range Increased .0957 .0799 .2414 .1967 Cruise Speed Reduced .0932 .0736 .1764 .1359 PD 370-22 .0956 .0785 .1808 .1449 1990 Engine Technology .0847 .0702 .1603 .1296 TABLE" 3, EFFECT OF STUDY CONDITIONS ON AIRCRAFT PERFORMANCE FOR DESIGN MISSION (100% L.F.) AND 60/GALLON FUEL PROPFAN SAVINGS - % TURBOFAN SAVINGS - STUDY CONDITION - FUEL DOC FUEL DOC Mission Effects Increased Design Range N/A 1.1 N/A 0.9 Decreased Cruise Speed 6.5 2.3 2.5 Technology Effects High Press. Ratio Engine 0.3 3.3 N/A N/A 1990 Technology Engines

10.8 7.4

11.4 7.5 Savings Are Relative to Baseline Design Airplanes as Follows: 1.

Propfan - 1500 NM., 0.8 Mach, STS 476, Revised Propfan Data 2. Turbofan - 1500 NM., 0.8 Mach, JT1OD The propfan and turbofan aircraft previously studied were restrained to a 1985 I0C with a design range of 1500 nautical miles, Mach 0.8 cruise speed, and a payload of 200 passengers. The engines employed were a rematched ver sion of the Pratt and Whitney STS 476 turboshaft using a Hamilton Standard eight bladed propfan operating at 800 feet per second rotational tip speed and a scaled version of the Pratt and Whitney JT10D turbofan.

Advanced technology incorporated into the airframe design included: o Supercritical wing * Active controls * Advanced composite materials for cost effective secondary structure Incorporation of the above resulted in a 4.5 percent reduction in aircraft empty weight. Direct operating costs were calculated using 1973 dollars and the cost factors shown in Table 4.

Maintenance factors, identical to those utilized for the previous RECAT study, were as follows: " Airframe Maintenance - Maintenance cost per cycle was reduced by 25 percent for the propfan due to decreased maintenance requirements for wheels, brakes, and landing gear.

" Engine Maintenance - Propfan engine maintenance was adjusted using factors previously provided by Pratt and Whitney for the turboshaft engine and by Hamilton Standard for the gearbox and propeller (Reference NASA CR 137926, Appendix A and B).

Maintenance labor cost per flight hour was reduced by 0.017 man-hours per engine flight hour from the baseline turbofan engine and then gearbox and propeller maintenance was added. No change was made for engine labor cost per cycle.

Turboshaft maintenance material cost per flight hour was adjusted using thrust relationships with gearbox and propeller cost added. Turboshaft main tenance labor cost per cycle was adjusted using thrust relationships but with no addition for gearbox and propeller.

of fuel savings and operating cost advantages was accomplished Assessment during this study at the following conditions: xvii TABLE 4. COST FACTORS 1973 DOLLARS PROPFAN TURBOFAN COMMENTS Cost Breakdown Flyaway Cost (Millions $) 14.15 13.39 Avg. unit cost based on Airframe 10.34 10.09 350 units Propulsion 3,31 2.80 Dev. cost amortized over Avionics 0.50 0.50 250 units-15% profit D.O.C.

Factors

* Flight Crew Cost ($/Hr.) 223 223 )

* Maintenance Factors 1973 rates - Labor Rates ($/Hr.) 6.10 6.10 - Maintenance Factors To adjust ATA formulas Airframe Labor/Cycle 0.57 0.60 and/Hour 0.57 0.60 Propfan brakes and wheels Airframe Material/Cycle 0.47 0.60 Airframe Material/Hour 0.75 0.75 Engine Labor/Cycle 0.60 0.60 Engine Labor/Hour 0.78 0.75 Includes engine, gearbox Engine Material/Cycle 0.49 0.Eo and propeller for and/Hour 0.65 -0.75 propfan/turboprop Burden (Factor) 1.8 1.8

* Fuel ($/Lb.) 0.088 0.088 600/Gallon

* Oil ($/Lb.) 1.0 1.0 * Insurance (%) 1.0 1.0 * Depreciation 00 Years 16 16 Spares (W) 15 15 Salvage (%) * Utilization (Hr./Yr.) 2900 * Resize the 1985 I0C propfan airplane for the new propfan data " Resize the 1985 IOC propfan airplane for a 2000 nautical mile design range to allow wider usage of the airplane " Resize the 1985 IOC propfan airplane for a cruise speed of Mach 0.75 to take advantage of additional fuel savings and potential reduced acoustic treatment requirements.

* Incorporation of an alternate turboshaft engine (PD 370-22) with component technology and overall pressure ratio comparable to the JT10D-2 turbofan.

" Incorporation of the Pratt and Whitney STS 487 turboshaft engine, representative of 1990 IOC technology and comparison with an equivalent 1990 IOC technology turbofan - STF 477.

The fuel conservation merits and the advantages in direct operating cost of the propfan powered aircraft was evaluated by utilizing an equal tech nology turbofan powered aircraft and comparing the two at identical design and mission conditions.

The new propfan data supplied by Hamilton Standard reflects the results of their wind tunnel tests of an 8 bladed propfan model and includes their predictions for a 10 bladed propfan. The effect on the new data was a slight increase in propulsive efficiency accompanied by a slight increase in gener ated sound pressure level for the 8 bladed propfan at 800 fps tip speed.

Directivity of the generated noise was re-defined, allowing more efficient utilization of the acoustic treatment material in the aircraft fuselage.

The near field noise generated by the propfan necessitates acoustic cabin interior noise at a level con treatment of the fuselage to maintain the sistent with current wide body turbofan aircraft. For this study, the Hamilton Standard supplied acoustic characteristics for the propfan (Appendix A) were utilized in conjunction with a "Double Limp Wall" concept to establish loss through the fuselage wall and the required mass treat the transmission levels. A discussion of the ment needed to attain acceptable interior noise is included as method utilized and the results obtained acoustic analysis concept utilized for this study is the Section 3. The acoustic treatment to the fuselage mass, such as lead vinyl, addition of "limp" (non structural)

xix ORIGINAL PAGE IS

OF

POOR QUALI1Y

to obtain the required transmission loss. This concept results'in approximately 5000 pounds of additional fuselage weight for the baseline propfan airplane (additional weight over that required for the baseline turbofan airplane). Since experimental verification of this acoustic analy sis concept (double limp wall) has not been accomplished, uncertainties exist concerning the magnitude of fuselage treatment required. To compensate for this uncertainty, Lockheed has included the conservatism of assuming that treatment of the entire fuselage diameter and cabin length will be required.

As subsequently discussed, additional acoustic assessment (both analytical and experimental) is required.

This study shows that an advanced propfan powered aircraft, utilizing the Hamilton Standard 8 bladed propfan, is a viable alternate to the turbofan powered aircraft and offers significant savings in fuel and operating costs without compromising passenger comfort. Additionally, the reduction in cruise speed to 0.75 Mach, consistent with current operation of short and medium range transports, offers further significant savings in fuel and operating costs.

Assessment of the Hamilton Standard data for a 10 bladed propfan indicates a further potential advantage in fuel and DOC savings since the projected sound pressure levels are reduced. This reduction in SPL along with an increase in blade passage frequency results in a significent reduction in acoustic treatment weight.

To realize the potential fuel and operating cost savings available with the advanced propfan powered aircraft, and to further enhance its viability, the following actions must be implemented: * Further investigation of ten and twelve bladed propfans to assess their characteristics (performance, acoustics, mechanical design, and economics) and the impact on aircraft performance * Investigate propfan aircraft acoustic treatment concepts and in to further assess noise transmission mechanism configurations conjunction with aircraft fuselage wall structure/requirements * Investigate further advances in turboshaft engine technology for additional improvements in fuel consumption and engine economic characteristics xx * Investigate alternate engine/aircraft installation configuration to minimize the effect of propfan exterior noise transmission to the fuselage interior * Investigate the maintenance characteristics and costs of thrust reverser mechanism and aircraft tires/brakes for the turbofan and propfan aircraft to enhance the maintenance cost data, used for operating cost comparisons.

xxi INTRODUCTION The energy restrictions imposed in late 1973 by the oil embargo and the compelling need for energy conservation in all sectors of our national trans portation system led to a concerted effort by the air transportation industry to conserve fuel. Resolvement of the oil embargo, though alleviating the energy crisis of 1973, did not negate the need for fuel conservation. The escalation in fuel prices which have resulted combined with those prices projected for the future indicate a severe economic impact which must be off set by advancements in aircraft technology and operating procedures. Fore casts of the demand for air transportation shows a doubling or tripling by the year 1990 in all our major metropolitan areas. These projections along with the economic impact of estimated fuel prices for 1990 dictate a con certed effort to provide aircraft which are significantly more fuel efficient.

The previous RECAT study, part of the Aircraft Energy Efficient (ACEE) program, investigated practical means of achieving reduced fuel consumption in commercial air transportation in the following areas: * Current aircraft types * Revised operational procedures " Modifications to current aircraft " Derivatives of current aircraft * New near-term fuel conservative aircraft.

Results of the previous RECAT study showed that significant potential savings in fuel and operating costs are available by utilizing the propfan propulsion system. The Hamilton Standard propfan is a multibladed, highly loaded, variable pitch propeller utilized in conjunction with an advanced turboshaft engine. Advanced aerodynamic characteristics of the propfan, which include xxii thin blades with swept tips and advanced airfoils, produce a significantly efficiency, than that attained with a standard propeller de higher propulsive with turbofan powered aircraft.

sign, and operation at Mach numbers competitive In the previous study, a turbofan and a propfan airplane were designed and optimized for minimum fuel and operating cost for a 1500 nautical mile fuel usage range, Mach 0.80 cruise speed design mission. Comparisons of and operating costs for the design mission showed that the propfai aircraft (at 60W/gal. fuel cost) of results in a savings in fuel and operating costs 17.8, percent and 8.2 percent respectively. The turbofan aircraft employed and Whitney JT10D turbofan engine. A Pratt a scaled version of the Pratt and Whitney STS 476 turboshaft engine with the Hamilton Standard 8 bladed 800 feet per second tip speed, was utilized for the propfan propfan, aircraft.

by this document examined the further potential for The study reported fuel and operating cost savings of the advanced propfan aircraft for the following conditions: a New propfan data * Revised design mission profiles " Additional engine performance characteristics.

from wind Propfan performance and acoustic characteristics, resulting tunnel testing by Hamilton Standard of their propfan model, were supplied data for assessment of the impact on aircraft performance. The new propfan and in generated sound results in a slight increase in propulsive efficiency 0.8 with the 8 bladed, pressure level at the design point of 30,000 feet, Mach of acoustic directivity.

800 fps propfan. Also included was a redefinition The previously used design range of 1500 nautical miles limited accep studies. A design range of 2000 nauti tance of the aircraft in airline fleet cal miles, equivalent to the B727-200, could provide a much wider potential additional use of the propfan aircraft. Also, preliminary analysis indicates fuel savings may be available with the propfan propulsion by reducing the' cruise speed to a value consistent with current operating experience for short to medium range transports. Mach 0.75 was selected as the reduced cruise speed.

xxiii ORIGINAL PAGE IS

OP POOR QUALf

The turbofan and turboshaft engines used equal component technology but the loss of fan supercharging in the STS 476 turboshaft engine resulted in a lower overall pressure ratio than the JT10D-2 turbofan. An alternate turbo shaft engine, Allison PD 370-22, with both component technology and overall pressure ratio comparable to the JT10D-2, was incorporated. Studies of unconventional engine cycles conducted under NASA-Lewis Research Center con tract have identified two comparable advanced technology engines which could be available for a 1990 IOC. These engines, identified as the Pratt and Whitney STF 477 turbofan and STS 487 turboshaft, were incorporated.

The mission profile used for all performance calculations is included as Figure 1 and the study ground rules are presented in Table 5.

xxiv CRUISE AT ALTITUDE STEP DESCEND TO CRUISE AT ft ALTITUDE 45 MINUTE SLIB10000 HOLD AT INITIAL CLIMB /ALTITUDE CRUISE DESCEND TO CLIMB DECELERATE ACCELERATE - ,-' $ft ACCELERATE DECELERATE ="\ TO CLIMB TO .DESCEND 10 000 ft /10 000 ft 10.000 ft J 1500 ft r.. 10 G0O ft' - DESCEND TO AIR MANEUVER CLIMB TO 100E'-1500T AT 15D0 ft FOR 10000 ft "<ST TAKEOFF 3V MIUES.APPROACH TO 1500 ft

APPROACH ,.-"N MISSED

START, TAXI TAXI, STOP APPROACH TO 1500 ft AND GROUND AND HOLD SHUTDOWN 9 MINUTES 3 MINUTES * LANDING RAMP LANDING RAMP RAMP TAKEOFF n.m'.

FLIGHT DISTANCE - n.mi. j200 FUEL RESERVE FLIGHT TIME

I

BLOCK FUEL AND BLOCK TIME ALTERNATE DESTINATION ORIGIN • FUEL FROM RESERVE Domestic Mission Flight Profile Figure 1.

RULES TABLE 5. STUDY GROUND Economic Parameters Dollars * 1973 Fuel * 60/Gallon With 10% Residual * Depreciation Period = 16 Years * Spares = 15% of Flyaway Cost * Insurance Rate = 1% * Production Quantity = 250 Aircraft * Inflation = 5% 8% * Discount Rate = Configuration * 200 Passengers * Wide Body Fuselage Engines * Four Mission * M 0.80 and M 0.75 Cruise * 1500 n.mi. and 2000 n.mi. range Cruise Altitude 30,000 feet * Initial * Field Length 7000 feet knots " Approach Speed 135 Technologies Advanced * Supercritical wing " Active controls * Advanced composites xxvi ABBREVIATIONS/SYNBOLS/CONVERSIONS Abbreviations ASM Airplane Seat Nautical Mile ASSET Advanced System Synthesis and Evaluation Technique (Lockheed Computer Program) blk-hr Block-hour BPF Blade passage frequency DOC Direct operating cost EPNdB Equivalent perceived noise level, decibels EPR Engine overall pressure ratio FAR Federal Air Regulation ft Feet gal Gallon in. Inch kt Knot lb Pound LF Load factor LFL Landing field length, ft.

MAC Mean Aerodynamic Chord MEW Manufacturer's empty weight, lb.

min Minutes n.mi. Nautical mile xxvii OEW Operating empty weight, lb Pax Passenger SFC Specific fuel consumption, lb fuel/hr/lb thrusl shp Shaft horsepower SL. Sea Level SLS Sea level static TOFL Takeoff field length, ft TOGW Takeoff gross weight, lb Symbols AR Aspect ratio, b 2/S b Wing span, ft c Wing Chord, ft c Propeller blade chord, ft b CD Drag coefficient CL Lift coefficient d Distance between inner and outer fuselage walls D Drag force, lb D_ Propeller diameter, ft p dB Decibel FN Net thrust force, lb f frequency, Hz fn natural frequency, Hz n f rRing frequency, Hz M Mach number MCR Cruise Mach Number MH Helical tip Mach number xxviii ft area, Wing S t/c Thickness ratio T/W Thrust to weight ratio lb/ft loading, Wing W/S 1Propeller efficiency x Wing sweep angle, degrees Conversions To Convert From To Multiply By Fahrenheit Celsius T (5/9)(T -32) c F foot meter 0.3048 2 2 foot meter 0.09290304 3 3 foot meter 0.028316846592 foot/second meter/second 0.3048 gallon meter 0.003785411784 horsepower (550 ft-lb/sec) watt 745.69987 inch meter 0.0254 knot meter/second 0.5144444444 nautical mile meter 1852 pound (force) Newton 4.4482216152605 pound (mass) kilogram 0.45359237 xxix

SECTION 1 AIRCRAFT DESIGN EVALUATION Table 6 provides a matrix of the design and mission characteristics utilized to evaluate the aircraft investigated during this study. Evaluation of the aircraft was accomplished for both a 1985 and 1990 IOC and included the following: o New Propfan data " Revised mission profiles * Additional engine performance characteristics The propfan and turbofan aircraft designed during Task 7 of the previous RECAT study were utilized as the baseline configurations for this study.

Each of the study conditions depicted in Table 6 resulted in re-sizing of the baselines to obtain the optimum point design characteristics.

The criterion utilized to select optimum point design characteristics was mini mum direct operating cost, at 60/gallon fuel cost. This criterion is identical to that utilized for the previous RECAT study.

1.1 TURBOFAN AIRCRAFT 1.1.1 Baseline The baseline turbofan powered aircraft, CL1320-11 is shown in the general arrangement drawing, Figure 2, and the general characteristics are shown in Table 7.- As previously documented in NASA Report CR 137926, the airframe technology levels include a supercritical wing, active controls, and advanced composite secondary structure.

ORIGINAL PAGE IS OF POOR QUALITY, 1-1 TABLE 6. STUDY MATRIX FOR AIRCRAFT CONFIGURATIONS 1985 IOC 1990 1OC DECREASED ALTERNATE OFF DESIGN 1990 TECH CRUISE SPEED ENGINE CRUISE SPEED ENGINE BASELINE INCREASED RANGE T/F P/F T/F P/F T/F P/F T/F P/F T/F P/F T/F P/F 1500 2000 2000 1500 1500 1500 N/A 1500 N/A 1500 1500 Range nm) 1500 0.8 0.75 0.75 0.8 0.75 0.8 0.8 Cruise Speed (m) 0.8 o.8 0.8 Pax 200 200 200 200 200 200 200 200 200 Fuel Cost (0/gal) 30/60 30/60 30/60 30/60 30/60 30/60 30/60 30/60 30/60 30/60 30K 30K 30K 30K 30K 30K 30K Cruise Alt (ft) 30K 30K 30K Field Length (ft) 7K 7K 7K 7K 7K 7K 7K 7K 7K 7K 135 135 135 App. Speed (kts) 135 135 135 135 135 135 135 Powerplant STS 476 JT1OD-2 STS 476 J.1OD-2 STS 476 JT1OD-2 Po370-22 STS 476 STS 487 STF 477 WING CHARACTERISTICS BASIC TOTAL HORIZ VERT

o . AREA (ft ) 1955 2209 275 253

10 - 5 1.6 ASPECT RATIO SPAN (1t) 139.8 37 20.1 t 303/1 137 232 ROOT CHORD (in.) 258 41 70 (in.) 77 TIP CHORD 0.3 0.3 TAPER RATIO 0.3 - 184 97.5 165.6 MAC (in., 25 25 30 SWEEP (DEG) TIC ROOT (%) 14A 10 10 TIC TIP m ) i

A ATBI. 117.5

PLANT: PRATT & WHITNEY JT1O0D-2 POWER SCALEDSLS THRUST 14 672 Ib ea 140 ft- 2 in.

H LA- ,

ln fn, nn

45 INT

E~00of

* FOUR TURBOFANS * 2DOPAX 000000 a 0 0 oo 000 4I * MACH 0.8 * 1500 n~mi.

n no 155 ft - 10 in.

Baseline Aircraft Figure 2. General Arrangement-Turbofan TABLE 7. TURBOFAN BASELINE AIRCRAFT CHARACTERISTICS (CL 1320-11) WEIGHTS Max. Takeoff Gross Wt (ib) 217015 Max. Landing Gross Wt (lb) 205000 Operational Empty Wt (ib) 138402 Max. Fuel Capacity (ib) 50000 POWER PLANTS Number and Type -JTlOD-2 (S caled) Bypass Ratio 5.4 SLS Thrust/Engine (lb) 14672 BODY Length (ft) 155.8 Max. Diameter (in) 235 - Accommodations 200 (10/90) 8 Abrea st WING AND EMPENNAGE WING HORIZONTAL TAIL VE RTICAL TAIL Area (sq. ft) 275 253 Aspect Ratio 10 5 i.6 Span (ft) 139.8 37 20.1 Sweep (deg) 25 25 30 Mac (in) 97.5 165.6 The selected

design was a 4 engine, wide body aircraft

with a design range of 1500 nautical miles, 0.8 Mach cruise speed, and 200 passengers. Additional mission constraints were an initial cruise altitude of at least 30,000 feet, takeoff field length of 7000 feet maximum, and a maxiuum approach speed of 135 knots. The results of the parametric study used to size the baseline turbofan aircraft are shown in Figure 3. Sizing of the aircraft for minimum DOC, at 600/gallon fuel, resulted in a wing AR of 10 and a t/c of 12%.

Design and performance characteristics of the baseline turbofan air craft are included in Table 8.

The supercritical wing has an aspect ratio of 10 and a sweep (.25C) of 25 degrees. Active controls, allowing smaller, lighter airframes, are incorporated into the airframe design for the turbofan aircraft.

A 3 percent reduction in wing weight is obtained by employing active ailerons to provide maneuver and gust load alleviation. Relaxation of static stability margins through use of an active horizontal tail results in a reduction in tail size and a corresponding 30 percent reduction in.tail weight. The net reduction in empty weight, due to incorporation of active controls, is 1.2 percent. Secondary structure employing advanced composite materials in cludes the fixed wing leading edge, fuel tank baffles, floor supports, interior doors, and-dividers.

The reduction in empty weight, attributed to composite structure, is 3.3 percent. Incorporation of advance composites and active controls results in a total empty weight reduction of 4.5 percent.

Included in the baseline configuration is the JT10D turbofan engine, scaled to the aircraft performance and mission requirements.

The features of the engine, designated JT10D-2, are included in Table 9.

1.1.2 1985 Technology Assessment The 1985 technology assessment of the turbofan powered aircraft consisted of re-sizing the baseline configuration to assess the impact on mission fuel 1-5

ORIGINAL PAGn is

P Pod QU GEI

2.0

®/ N110

1.9 0.40 DOC* d/ASM 1.8 / ~MINDOC = 0.28 ( f 0.25W/S ill = T/W 0.25 SPEED = 135 KTS

0 APPROACH

1.7 30 000 ft CRUISE ALTITUDE * 60d/gal fuel cost Figure 3. Selection of Baseline Turobfan Airplane Design 1-6 AIRCRAFT PERFORMANCE CHARACTERISTICS OF TURBOFAN TABLE 8. DESIGN AND 1985 IOC 1991 IOC 1990 ENGINE INCREASED RANGE REDUCED CRUISE BASELINE CONFIG JTIOD-2 JTIOD-2 STF Engine Identification J1OD-2 0.8M 0.8M 0.8M 0.75M Cruise Speed 2000 1500 Range (nrm) 1500 Design 200 200 200 No. Passengers 112 109.8 ill 115 W/S (ib/ft ) 0.24 0.26 0.28 0.28 T/W 10 10 10 AR 12 12 12 t/c (%) 212365 206212 217015 230386 TOGW (ib) 134648 132298 138402 141697 OEW (lb) 13249 16127' 12582 (SLSIb) 15191 Thrust/Engine 2003 1896 1878 Wing Area (ft ) 1955 139.8 ihi.5 137.7 137.0 Wing Span (ft) 1.294 1.364 1.381 1.381 DOC @30/gal 1500 nm (CIASM) 1.737 1.627 1.737 1.715 DOC @300/ga- nm (C/ASM) 1.674 1.809 1.793 1.810 Doc @604/gal 1500 m C¢ASM) 6 2.069 2.236 2.216 2.236 DOC @ 0/gal 475 nm CASM) 25418 38276 27962 - 1500 am (Cb) 28673 Block Fuel 9717 8832 13303 - 475 m (ib) 9965 Block Fuel 0.0847 0.0957 0.0932 Fuel Efficiency (Ib/ASM) 0.0956 0.588 o.656 o.641 Cruise SFC (lb/hr/lb) o.656 37000 31000 32000 (ft) 37000 Initial Cruise Alt.

5577 5787 6845 TOFL (ft) 6138 6174 LFL (ft) 135 135 135 135 Approach Speed (Kt) 11217 10015 13436 14379 Propulsion Weight (ib) IS OIGINAL PAGE

OF POOR QUALITY

1-7 FEATURES FOR JT1OD-2 (SCALED) TURBOFAN TABLE 9. ENGINE -Twin Spool - Design fan press. ratio _e Description of 1.69- bypass press. ratio of 5.4.

stage fan, 12 stage compressor, Single 2 stage HP turbine, 4 stage LP turbine 0.618 " Scaling Factor 14672 • Installed Thrust (SLS - lb) Press. Ratio 28:1 * Overall (30,000 ft, 0.8 mach) Inlet 2400 * Max. Turbine Temp. (OF) * Engine Length (in) 97.8 52.6 " Engine Diameter (in) Engine Maintenance Cost ($/Flt hr)

122.6 I

cost of changes in design range (2000 naut requirements and direct operating speed (0.75 Mach in ical miles in lieu of 1500 nautical miles) and cruise again minimum fieu of 0.80). Re-sizing criteria for each turbofan design was fuel cost.

DOC at 600/gallon range provides a potential wider usage of the An increase in the design range.

aircraft in current fleet operations equivalent to the B727-200 design the original cruise speed of Mach 0.80 may have unduly compromised Likewise, the propfan aircraft and preliminary analysis indicates that additional with a reduced cruise speed of Mach 0.75.

fuel savings are available Range Increase 1.1.2.1 Design nautical miles necessitated a re Increasing the design range to 2000 type of parametric analysis conducted sizing of the aircraft utilizing the same analysis, wing AR, -for the previous baseline design. For this parametric and T/W were varied as shown in Table 10. A total of 144 designs were t/ci W/S the ASSET, parametric analysis program. Plots of DOC accomplished using drawn so that optimum values for minimum DOC could versus t/c for each A were at 600/gallon versus t/c for the range of be selected. Figure 4, depicting DOC 1-8 TABLE 10. PARAMETRIC STUDY MATRIX TURBOFAN - JTIOD-2 0.8 MACH, 200 PAX/2000 NM AR '/c

W/S. T/W

8 9 100 110 120 130 0.24 0.26 0.28 0.30 8 12 100 110 120 130 0.26 0.28 0.30 0.32

8 14

100 110 120 130 0.27 0.29 0.31 0.33 10 9 100 100 120 130 0.22 0.26 0.30 0.34 12 100 110 120 130 0.24 0.28 0.32 0.36 10 14 100 110 120 130 0.24 0.28 0.32 0.36 12 9 100 110 120 130 0.22 0.26 0.30 0.34 12 12 100 110 120 130 0.24 0.28 0.32 0.36

12 14 100

110 120 130 0.24 0.28 0.32 0.36

-AR's considered, shows the basis for selection. The selected values for AR and t/c of 10 and 11.5% (rounded to 12) are consistent with the results of wing optimization studies previously accomplished for the RECAT study for 00/gallon fuel cost. ASSET carpet plots are utilized to select W/S and T/W values for minimum DOC and the mission constraints (i.e., field length, approach speed, cruise altitude, etc.).

Figure 5 is the ASSET carpet plot for minimum DOC (600/gallon) for the turbofan aircraft at the 2000 nautical mile design mission, and depicts the selection of the point design para meters. Point design parameters selected were AR = 10,: t/c = 12%, W/S = 115, and T/W = 0.28.

Design and performance characteristics of the turbofan aircraft sized for the 0.8 Mach, 2000 nautical mile design mission are shown in Table 8.

The effect of re-sizing the baseline turbofan aircraft for the 2000 nautical mile, Mach 0.8 mission is an increase in block fuel of approximately 26 percent which is consistent with the 25 percent increase in range and the

1-9 PAGE IS

1-9 o ULr 1.88 - DOC VERSUS THICKNESS-RATIO FOR AR 8, 10, & 12 1.87 1.86 1.85 1.84 91.83 4 _jt8 AR ~ n C1.81

81.80

1.79 1.77 TURBOFAN - JT100-2 2000 NM. RANGE PAX 1.76 -200 0.8 MACH 1.7511 9 10 11 12 13 14 THICKNESS RATIO (tlcV-%

Figure 4.Turbofan DOC versus t/c - 2000 N~M. Range

1_-jo

RECAT TURBOFAN - JT1OD-2 MACH 0.80,200 PAXI2000 NM. ARlO, t/c 12 1.96 - 0.24 W/S 1.92 - 036 7A . 1.88 0.32 APPROACHSPEED, 135 KNOTS o 1.84 0.28 1.80 .

.

W/S = T/W = 0.28 L 1.76 Figure 5. Asset Crossplot - Turbofan DOC (2000 NvM. Range) 1-11 increased and climb due to the fuel required for take-off small increases in miles does to 2000 nautical the design range weight. Increasing aircraft fuel cost.

in DOC at 60¢/gallon a small savings effect Cruise Speed Reduction 1.1.2.2 being consistent with current 0.75 was selected as A cruise speed of Mac of the transports. Re-sizing for short to medium range operator practice a basis for to provide for this mission was accomplished turbofan aircraft Mach 0.75' the propfan aircraft at fuel savings available'with comparison of the speed.

cruise analysis.

using the ASSET parametric baseline turbofan was resized The T/W respectively and W/S and held constant at 10 and 12 Wing AR and t/c were were varied as follows: T/W W/S 100 .22 .24 .26 .28 fuel cost for plot of minimum DOC at 600/gallon 6, the ASSET carpet Figure selection of aircraft,-depicts the PAX/1500 n.mi. turbofan the Mach 0.75, 200 the point design parameters.

W/S and T/W utilized as optimum values for selected were: Point design parameters AR = 10 = 12% tic WIs = 112 T/W = .24 aircraft, sized of the turbofan powered and performance characteristics Design included in Table 8.

1500 nautical mile range, are for Mach 0.75 cruise and nautical mile, aircraft for the 1500 The effect of re-sizing the turbofan 2.5 percent with no measurable is a savings in block fuel of Mach 0.75 mission 1-12

lb

ORIGINAL PAGE

OF POOR QUALITY

RECAT TURBOFAN - JT10 D-2 NM. AR10, t/c 12 MACH 015, 200 PAX/1500 1.92 0.22 100 1.90 APPROACH SPEED, 135 KNOTS 1.88 TM\ L -110

1.86 - 0.28

40.24 1.84 0 120 1.82 = 112= 112 130 1.80 TAV = 0.24 1.78 - Turbofan DOC (0,75 Mach) Figure 6. Asset Crossplot 1-13 change in DOC at 600/gallon fuel cost. The small decrease in block fuel at the 1500 nautical mile design range is attributed to an approximate 2 percent improvement in average cruise SFC.

1.1.3 1990 Technology Assessment Previous studies'accomplished under contract to NASA-Lewis Research Center of unconventional engine cycles have identified two comparable advanced technology engines for the turbofan and turboprop powered aircraft.

These engines, the Pratt and Whitney STF 477 turbofan and STS 487 turboshaft, are representative of those which could be available for a 1990 IOC aircraft.

The 1990 technology assessment of the turbofan powered aircraft consisted of re-sizing the baseline turbofan for incorporation of the STF 477 engine at the 1500 nautical mile, Mach 0.8 mission and ascertaining the fuel savings and operating cost advantages. Baseline airframe technology levels remained unchanged. The re-sizing criteria was minimum DOC at 600/gallon fuel cost.

1.1.3.1 Advanced Technology Turbofan Engine The Pratt and Whitney STF 477 turbofan engine was selected as representa tive of the best configuration for conserving fuel while presenting a practical configuration, attractive economic factors, and reasonable availability (1990 IOC) for advanced/technology transport aircraft. The STF 477 engine is a two spool design with an overall pressure ratio of 45:1 and maximum turbine inlet temperature of 2600°F as compared to 28:1 and 2400 F for the JT1OD-2 -turbofan engine. A description of the engine parameters is shown in Table 11.

Performance data for the STF 477 engine, along with engine and nacelle dimensions, engine weight, and appropriate scaling factor was provided by Pratt and Whitney for adaptation to the RECAT turbofan aircraft design mission requirements.

1.1.3.2 Aircraft Optimization Optimization of the 1990 turbofan aircraf', was accomplished by re-sizing the baseline to incorporate the STF 477 turbofan engine in lieu of the JT10D-2.

Utilization of the STF 477 engine necessitated alterations of the ASSET sub routines for configuration, weight, drag, and engine performance consistent

1-14

OU1OM4AL

PAGE IS

OF POOR 'QUAI4TY

TABLE 11. STF 477 ENGINE PARAMETERS PARAMETRIC DESCRIPTION Base Size, Thrust, N(lbf)* 118100(26550) 71200-178000(16000-40000) Scaling Range, Thrust, N(lbf)* Nominal Cruise Design Cycle at Mn 0.83 and 10,058m(33,000 ft) Fan Pressure Ratio 1.70:1 Bypass Ratio 8.0 a1 45:1 Overall Pressure Ratio (2600) Maximum Combustor Exit Temperature, OC(%F) 1427 472(1040) Inlet Flow (Corrected), kg/sec(lbm/sec) FAR 36 minus 10 EPNdB Acoustics (Engine Plus Nacelle) PERFORMANCE (Representative Conditions) TSFC Mach No. Net Thrust Condition Altitude N (lbf) kg/hr/N (lbm/hr/lbf) km (ft) (0.351) Take-off* 0 0 0.147 93635 (21050) 0.0358 Max. Climb** 9.14 (30000) 0.8 32912 (7399) 0.0588 (0.577) Max. Cruise** 9.14 (30000) 0.8 29910 (6724) 0.0586 (0.575) WEIGHTS AND DIMENSIONS Base Engine Weight, kg (lbm) 1787(3940) Dimensions Maximum Diameter, m(in.) 1.92(75.6) Overall Length, m(in.) 2.88(113.2) Nozzle Throat Areas 2 2 Duct, m (in. ) 1.150(1783) 2 2 Primary, m (in, ) 0,303(470) 118.28 Engine Maintenance Cost ($/Flt-Hr) U.S. Standard Atmosphere, 1962; *Sea level static take-off, 28.9%C (84 F) ambient temperature; recovery; no customer bleed or power extraction; representative nozzle thrust 100% ram coefficient.

performance calculated on basis of: U.S. Standard Atmosphere, 1962; 100 percent ram **Estimated duct bleed; recovery; 1.04 kg/sec (2.3 lbm/sec) mid-compressor bleed; 1.01 kg/sec(2.4 ibm/sec) day; representative nozzle thrust coefficients.

112 kw (150 hp) extraction; standard with the performance and dimensional data supplied for the engine.

For the parametric analysis, Wing AR and t/c were maintained at 10 and 12 % respectively and values of W/S and T/W were varied as follows: W/S = 100, 110, 120, and 130 T/W = .24, .26, .28, and .30 Figure 7, the ASSET carpet plot, depicts the selection of W/S and T/W values to be utilized for the turbofan aircraft point design.

Point design parameters selected are AR = 10, t/c = 12, W/S = 109.8, and T/W = .26. The performance and design characteristics of the turbofan aircraft with the STF 477 engine at 1500 nautical mile, Mach 0.8 mission, are shown in Table 8.

Utilization of the STF 477 turbofan engine results in a 1990 IOC turbofan aircraft at the 1500 nautical mile design mission with a block fuel savings of 11.4 percent and a DOC savings .of 7.5 percent, at 0C/gallon fuel cost, when compared to the 1985 I0C baseline turbofan aircraft. These savings are the result of an improvement in engine SFC characteristics, at cruise, of approximately 10 percent and a reduction in propulsion system installed weight of approximately 25 percent.

1.2 PROPFAN AIRCRAFT The propfan powered aircraft designed during Task 7 of the previous RECAT study was utilized as the baseline configuration for this study. Re-sizing of the the baseline configuration was accomplished for each of the study conditions to obtain the best point design consistent with minimum DOC at the 600/gallon fuel cost.

Installed propfan engine performance is based on engine manufacturers uninstalled engine data corrected for 100 hp per engine power extraction, 100 percent engine air inlet total pressure recovery, zero bleed flow rate, and 99 percent gearbox efficiency as noted in Table 12. The 200 percent recovery is based on the assumption that inlet duct losses are equal and opposite to the pressure rise across the propfan. Cabin pressurization and environmental control are provided by an engine driven compressor to avoid the potentially large losses associated with bleeding the turboshaft engines.

1-16 STF TURBOFAN RECAT CENTS/GAL FUEL- S M 0.80 1500 NMI, 200 PASS, 1.780 0.240 100.

1.760 W 1.740

Wis

110.

KNOTS SPEED, 135 APPROACH m 1.720 I-I 21i003 'I TW 0 .26 UU- m1.7400 1.640 - 1.620 (STF 477 Engine) - Turbofan DOC Crossplot 7, Asset Figure PAGOE'jo, 0oR QUALIY OF 1-17 TABLE 12. ENGINE INSTALLATION LOSSES Cruise M = 0.80 Propfan Turbofan Inlet recovery, PT /P0 1.00 0.998 IP compressor bleed, % 0 2.0 Horespower extraction 100 50 Fan Duct loss % APT/P 0 T - 0.99 Gear Efficiency Core cowl drag %AnFN/FN - 1.6 Notes (1) Exhaust nozzle thrust and airflow coefficients included in uninstalled engine performance (2) Nacelle drag included in aircraft drag All propfan engine performance is based on Hamilton Standard projected levels of efficiency for a propfan design having eight blades, each with 0.12 integrated lift coefficient and 200 activity factor, and operating at 800 fps tip speed. Initial point designs for the prcpfan aircraft were computed using the propfan performance and acoustic characteristics and installation considerations incorporated in the previous RECAT. Subsequent data, supplied by Hamilton Standard as a result of their continuing propfan testing, indicates an increase in the propfan induced external sound pressure levels as well as a slight improvement in performance. This data necessitated a re-sizing of the propfan aircraft and generation of new point designs con sistent with new performance values and the increase in required acoustic treatment.

Unlike the turbofan engine, the output of the turboshaft engine is shaft power which is transmitted through a gearbox and converted to useful thrust by the propfan.

As discussed in Report No. CR137926 for the previous RECAT study, selection of the propfan disk loading (diameter), geometry ('blade type and number), and tip speed is dependent on a tradeoff between net efficiency (fuel consumption) and installation weight (including acoustic treatment weight) and their impact on aircraft performance (DOC).

For example, Figure 8 shows that decreased cruise point design disk loading (increased prop-diameter) results in improved propfan efficiency. While this results in lower pro pulsion system specific fuel consumption, the increased installation weight associated with the larger propfan diameter and increased near field sound pressure levels may potentially counteract the benefits of higher fuel efficiency. The increased wight results in a heavier aircraft, larger pro pulsion system, and higher fuel flow rates.

Figure 8 indicates that while peak efficiency for the baseline STS h76 powered aircraft is achieved at a disk loading of 25 Shp/D , the optimum value for minimization of DOC and TOGW is 37.1 Shp/D , as shown in Figure 9.

1-19

0.90 30000 FT NET EFFICIENCY 0.8MACHNO.

800 FPS TIP SPEED -- PEAK >- 0.85

/

z /SELECTED SHP/D - FOR MIN D0C LL LU IL aJ 0.80 0.75 40,000 + GEARBOX + ENGINE + PROP WEIGHTS W -ACOUSTIC TREATMENT D 30,00 LB THRUST, 30,000 - J _j WEIGHTS FOR 3860 " rPT, 0.8 MACH NO., 800 FPS cc TIP SPEED Z k 20,000 0z ow fL-I 10,000 40 50 60 70 20 30 - HP/FT DISK LOADING SHP/D - System Figure 8. Propfan Efficiency and Propulsion Weight Trends with Disk Loading

1-20 @ jxx v

0 iv

~~1.6 42 44 210 28 30 32 34 36 38 40 ~ SHP/D DISK LOADING PROPELLER Figure 9, TOGW DOC, and Block Fuel vs. Propeller Disk Loading Similar propfan sizing studies for the other aircraft resulted in the loading and propfan diameter: following design disk Design Prop Design Diameter Mach No. Range Engine Disk Loading 0.8 1500 STS476 37.1 12.6 0.75 1500 STS476 35.9 11.4 0.8 2000 STS476 37.1 12.6 11.5

0.8 1500 PD370-22 42

0.8 1500 STS487 46 11.0 cost advantages of Assessment of the fuel conservation and operating accomplished for both a 1985 and 1990 IOC configura the propfan aircraft was propfan baseline was resized to -reflect updated propfan and tion. The 1985 of the propfan consisted of acoustic characteristics. Subsequent assessment re-sizing to 1) increase the design range to 2000 nautical miles to provide aircraft; 2) decrease the cruise potential as a replacement for the B727-200 speed to Mach 0.75 to obtain added fuel savings for the turboshaft engine at slower speed; and 3) incorporate an alternate turboshaft engine which has an overall pressure ratio and component technology comparable to the JT10D-2 configuration consisted of re-sizing turbofan. Assessment of a 1990 propfan turboshaft engine, representative the 1985 baseline to incorporate an advanced for a 1990 IOC. All final of the engine technology expected to be available Hamilton aircraft point designs generated during this study incorporate the Standard updated propfan data, presented in Appendix A.

1.2.1 RECAT Baseline The baseline propfan aircraft of the previous RECAT study, is shown in the general arrangement drawing, Figure 10, and the general characteristics are shown in Table 13. The airframe technology levels are identical to those a rematched the turbofan baseline. The propulsion system is utilized for the Hamilton

version of the Pratt and Whitney STS 476 turboshaft engine using

per second.

bladed propfan operating at a tip speed of 800 feet

Standard 8 Features of the rematched STS 476 turboshaft engine are shown in Table 14.

1-22 CHARACTERISTICS WING VERT BASIC I TOTAL HORIZ 2 ) AREA (fm 1995 2250 284 1.6 RATIO 10 - 5 ASPECT 37.7 20.4 SPAN (ft) 141.26 ROOT CHORD (in.) 261 306/& 139 236 (i 78 42 71 TIP CHORD 0.3 TAPER RATIO 0.3' 0.3 186 99.2 168 MAC fin.)

SWEEP (deg) 25 - 25 14& 10 10 TIC ROOT (%) 11 8 8 TIC TIP M)

A AT BL 117.5

C PRATT & WHITNEY STS 476 POWER PLANT: I 12.6 ft HAMILTON TURBOSHAFT ENGINE REMATCH STANDARD PROPELLER <!

SLS THRUST 15135 Ib (8863 SHP) EACH 0I 141 ft 7 in.

46 ft * 4PROPFANS, 0nn m's-.fooof .4 in.

* 200 PAX flocoMoro 01600 n.ni.

MAC 00.801 C) r) 120 155 ft- 10 in.

L3J Figure 10. General Arrangement-Baseline RECAT Propfan Aircraft TABLE 13. BASELINE RECAT PROPFAN Weights Maximum takeoff gross weight (ib) 217 466 Maximum landing gross weight (lb) 205 000 (lb) -146 417 Operational empty weight Maximum fuel capacity (lb) 50 000 Powerplants 4 STS 476 rematch Number & Type Propeller 12.6 ft/8 bladed SLS thrust/engine (lb) 14135 (8863 shp) Body H Length (ft) 155.8 ro Maximum diameter (in.) 235 (10/90%) Accommodations (No. Pax) 200 8 abreast Wing and Empenage Wing Horizontal Tail Vertical Tail Area (sq ft) 1995 284 Aspect ratio 10 5 1.6 Span (ft) 141.3 37.7 20.4 Sweep (deg) 25 25 32 165.6 MAC (in.) 186 97.5 TABLE 14. BASELINE BECAT ENGINE CHARACTERISTICS P&W STS 476 Rematch (Scaled) * Description Turboshaft Engine of Comparable Technology to JTlOD-2.

New Compressor and LP Turbine. Engine Rescheduled to Meet LCC Requirements * Scaling Factor 0.964 * Installed Rating Thrust (SLS, STD.)

- lb 14 shp (SLS, STD.) - hp 8 863 \.n Max shp (250 KEAS, SL, + 18 F) - hp l0 488 * Overall Pressure Ratio 20:1 36 000 ft M = 0.80 Cruise * Max Combustor Exit Temp 'F * Engine Length - in. 84.3 * Engine Diameter - in. 21.8 The baseline RECAT propfan airplane was optimized for minimum direct operating cost, at 00/gallon fuel cost, for a design range of 1500 nautical miles, 0.8 Mach cruise speed, and 200 passengers. Additional constraints imposed were an initial cruise altitude of 30,000 feet minimum, takeoff field length of 7,000 feet maximum, and a maximum approach speed of 135 knots.

The propfan aircraft, developed for the previous RECAT study, utilized performance, weight, and acoustic data for the 8 bladed propfan at 800 fps.

tip speed as supplied by Hamilton Standard, per Report SP 02A76 and SP 05A76.

Propeller disk loading and diameter, along with the magnitude of acoustic treatment in the aircraft fuselage, was determined using this data as described in Section 7.2 of Report No. CR 137926. A propeller disk loading of 37.1 Shp/D was selected by considering the tradeoffs between propeller efficiency and installation weights and the impact on aircraft performance. At the selected'disk loading for the turboprop baseline, 3089 lbs. of acoustic treat ment in the aircraft fuselage is required to attain interior (cabin) noise levels of 90 dB or less.

1.2.1.1 -Revised Baseline As part of this study effort, the propfan performance and acoustic characteristics were updated by Hamilton Standard, as shown in Appendix A, to incorporate their latest wind tunnel test results. Data supplied includes both an 8 bladed and 10 bladed propfan, each operating at tip speeds of 600, 700, and 800 fps, at Mach numbers of 0.7, 0.75, and 0.8. Figure 11 shows a comparison of revised to original efficiencies at 0.8 Mach number. A detailed discussion of the propfan acoustic characteristics and the effect on fuselage treatment methods and results is included in Section 3 of this report.

Propeller disk loading was maintained at 37.1 Shp/D for the revised baseline propfan aircraft. Because of the increased propfan SPL (Section 3), the weight of acoustic treatment for the revised baseline increases by approximately 2130 lbs from 3089 (previous RECAT) to 5220 lbs.

The acoustic treatment weight of 5220 lbs. is obtained by interpolating between the values 25 and 26, Section 3, for the point (at 37.1 Shp/D ) depicted in Figures design cruise thrust of approximately 3200 lbs required for the revised base line propfan airplane.

1-26 8 BLADES, FPS VTIP M = 0.8, 30,000 FT REVISED PROPFAN 82 - DATA I 78 U PER SP02A 76 0 10 20 30 40 50 60 70 80 2 2 DISK LOADING (SHP/D ) - HP/FT Ct Figure 11. Effect of Disk Loading on Propfan Efficiency The design and performance characteristics for the revised propfan baseline are shown in Table 15. The effect of the revised propfan data on the baseline configuration is an increase in block fuel -of 269 lbs and an increase in DOC at 60C/gallon fuel of 0.01¢/ASM for the Mach 0.8, 1500 nautical mile design Mission.

The sensitivities of changes in propfan efficiency and acoustic material weight on the baseline aircraft fuel and DOC savings are shown in Figures 12 through 15. These data indicate that a 1 percent decrease in propfan effi ciency affects DOC (0.5%) the same as a 1000 lb. increase in acoustic treat ment material.

At the baseline disk loading of 37.1 Shp/D , net efficiency increases by approximately 1.7 percent while installed weight increases by approximately 2130 lbs due to the increased acoustic treatment required.

1.2.2 1985 Technology Assessment For the propfan powered aircraft, the 1985 technology assessment con sisted of re-sizing the baseline configuration, based on minimum direct operating cost,''at 00/gallon fuel cost, for a design range of 2000 nautical miles in lieu of the 1500 nautical mile range; a cruise speed of 0.75 Mach in lieu of 0.80;'and incorporation of an alternate turboshaft engine with an overall pressure ratio comparable to the JTl0D-2 turbofan. In addition to re-sizing the baseline configuration, as stated above, an assessment of the performance impact was made when the cruise speed of the baseline configuration was reduced to Mach 0.75 with no re-sizing.

As previously discussed in Section 1.2 of this report, the change in design range was accomplished to provide a potentially wider use of the prop fan airplane in current fleet operations and the change in cruise speed was accomplished to obtain the potential additional fuel savings available with the turboshaft engine.

.1-28 TABLE 15. DESIGN AND PERFORMANCE CHARACTERISTICS OF PBOPFAN AIRCRAFT 1985 IOC 1990 IOC Cruise Opt. Alter. Engine 1990 Engine Revised Baseline Design Range

Engine Identification STS 476 STS 476 STS 476 PD 370-22 STS 487

Speed 0.8M 0.8M 0.75M 0.8M 0.8M Cruise Design Range (NM.) 1500 2000 1500 1500 1500 200 200 200 200 No. Passengers 200

W/S (Cb/ft ) 109 112 108 108 107.5

0.18 0.26 0.25 0.22 0.22 T/W 10 10 10 10 10 AR

E/C (%) 12 12 12 12 12

205749 TOWG (lb) 220572 231282 211264 211034 (lb) i49124 151223 142711 14!o8a 138513 OEW ib) 13785 14455 11613 11607 9257 Thrust/Ang (SLS, (ft ) 2042 2068 1992 1959 Wing Area

142.9 144.5 141.2 14o.4 139.0

Wing Span (ft) 1.283 1.282 1.228 DOC @300/Ga1-l500NM.(0/ASM) 1.314 1.294 1.584 1.603 1.602 1.503

DOC @300/Gal-h75NM.(0/ASM) 1.641

1.627 1.543 DOC @60¢/Ga1-1500NM.(¢/ASM) 1.667 1.649 1.629 2.008 2.006 1.901 DOC @60€/Gal-475NM.(/ASM) 2.055 2.033 Block Fuel - 150ONM. (lb) 23625 31970 22086 23559 21072

Block Fuel - 475NM. (ib) 8012 10840 7487 7987 7143

Fuel Efficiency (Ib/ASM) 0.0788 0.0799 0.0736 0.0785 0.0702

0.536 0.489

(lb/br/lb) 0.528 0.531 0.504

Cruise SFC 30000 30000 30000 Initial Cruise Alt (ft) 31000 30000 TOFL (ft) 4650 5009 5415 4555 4645

5994 6024

LFL (ft) 6056 6018 6033

Approach Speed (Kt) 135 135 135 135 135 16652 13332 11675 10882 Q Propulsion Weight (lb) 16471 TURBOFAN BASELINE 200 PAX/1500 NM.

M 0.8 PROP!AN H 8 - 01 4 SI II +10 +5 0 -5 -10 A EFFICIENCY - % Figure 12. DOC Sensitivity to Propfan Efficiency 200 PAX/1500 NM.

M=0.8 TURBOFAN BASELINE t~r 0 cn16

i.RO.,AN

U.

D 24' wJ " 28' +=5 0 -5 -10 +10 A. EFFICIENCY % Figure 13. Block Fuel Sensitivity to Propfan Efficiency TURBOFAN BASELINE 200 PAX/1500 NM.

M =0.8 n 6 N) CD tL)

a I I 1I

4 5 6 7 0 1 2 3 WEIGHT TREATMENT 1000 LB ACOUSTIC 14. Effect of Acoustic Weight on DOC Savings Figure TURBOFAN BASELINE 200 PAX/1500 NM.

M -0.8 - z > W l12 (A I, w U.

.14 • PROPFAN -'J 0 ::J" -- BASELINE 0 1 2 3 4 5 6 7 ACOUSTIC WEIGHT TREATMENT 1000 LB Figure 15. Effect of Acoustic Weight on Block Fuel Savings 1.8 DOC VERSUS THICKNESS RATIO FOR AR 8, 10,& 12 PROPFAN STS 476 2000 NM. RANGE 200 PAX 0.8 MACH AR 1- -12 CD L) 1.6 10 12 14 THICKNESS RATIO (t/c) " % NM. Range Figure 16. Propfan DOC Versus t/c -2000 1-34 STS 476 RECAT PROPFAN = 10.

tic = 12.0./AR FUEL = 60 CENTS/GAL M = 0.80 200 PASS, 2000 NMI, 1.850 SPEED, APPROACH 135 KNOTS 0.36 100 1.800 ~0.32 1.750 0.24 - 1.700 a- LU 1.650 W/S 112 / =/ 0.25 1.600 17. AE SETCrossplot - Propfan DOC (2000 NM. Rduge) Figure PAGE IS ORIGINAL

OF POOR QUALITY1

1-35 EFFECT OF MACH NUMBER ON ENGINE SFC 1.5 1.4 -8 BLADED PROPFAN 800 FPS TIP SPEED FT

/ 30,000

uL.° z 1.3 1.2 0.6 0.5 - 0.4 0.3 0.9 0.5 0.6 0.7 0.8 MACH NUMBER Figure 18. Effect of Mach Number on Engine SFC 1-36 RECAT PROPFAN STS SWEEP = 25 DEG, t/c = 12, AR = 10 200 PASS, 1500 NMI, M = 0.75, FUEL = 60 CENTS/GAL 1.700 0.28 100. APPROACH SPEED, 1.680 W 1.660 LU .. J0.22 I 0.2 I-.

Q C.

APPROACH 0O SPEED 1.620 .6 T/W = 0.22/ 1.580 Figure 19. ASSET Crossplot - Propfan DOC (075 Mach) 1-37 RECAT PROPFAN PD 370-22 200 PASS, 1500 NMI, M = 0.80 FUEL = 60 CENTS/GAL 1.670 -0.20' T/W 1.660 - 0.22 0.26 -0.24 1.650 APPROACH SPEED 135 KNOTS .. 1.640 .uj W/S ,J , 1.630 a.

Z 1.620

°

-

0 W/S =108 0.22 T/W= 1.610 1.600 1.590 1.580 Figure 20. ASSET Crossplot - Propfan DOC (PD 370-22 Engine) 1-38 RECAT PROPFAN 200 PASS, 1500 NMI, M=.80 FUEL=60 CENTS/GAL 1.580 0.18 0.240 1.570 -CRUISE FT 30,000 ALTITUDE T/W - 1.560 0.22. C 1.550 0 1.540 ' APPROACH SPEED, 135 KNOTS " • W/S u 1.530 C.

'110 LU - 1.520 W/S = 107.5 O TW = 0.18 1.510 1.500 1.490 1.480 ASSET Crossplot - Propfan DOC (STS 487 Engine) Figure 21.

1-39 Design Range Increase 1.2.2.1 design range of of the propfan aircraft at an increased Assessment the aircraft using the ASSET miles resulted in a resizing of 2000 nautical analysis, wing AR, t/c, W/S, and parametric analysis. For this parametric design combinations optimum varied as shown in Table 16. From these T/W were Figure 16, depicting DOC AR and t/c for minimum DOC were selected.

values of the basis for for the range of AR's considered, shows (60/gal.) versus t/c ASSET carpet plots were then uti selection of an AR of 10 and t/c of 12%.

and T/W for minimum DOC consistent to select the optimum values of W/S lized length, approach speed, and cruise with the mission constraints (field plot for minimum DOC ( 00/gal.)

altitude). Figure 17, is the ASSET carpet and depicts aircraft at the 2000 nautical mile mission, for the propfan selected design parameters. Point design parameters selection of the point 112, and T/W = .25. Design and performance were AR = 10, t/c = 12%, W/S = are shown in the Mach 0.8, 2000 nautical mile mission characteristics for Table 15.

aircraft for a 2000 nautical mile, The effect of resizing the propfan 26 percent.

is an increase in block fuel of approximately Mach 0.8 mission increase in range and a in block fuel results from a 25 percent This increase follows: slight decrease in fuel efficiency as NM. = 0.0788 lb/ASM Fuel efficiency @ 1500 2000 NM. = 0.0799 lb/ASM Fuel efficiency @ the increase in for the 2000 NM. range is due to Decrease in fuel efficiency of airplane and the requirement for 225 lbs.

thrust required for the heavier Additional weight for this higher cruise thrust.

additional acoustic treatment realized as a result of the increased savings of approximately 1 percent is DOC range.

1RIGINJP I

TABLE 6. PARAMETRIC STUDY MATRIX PROPFAN - STS 476, 0.8 MACH, 2000 N.M. RANGE AR t/c W/S T/W

8 9 100 110 120 130 .22 .26

.30 .34

8 12 .2~4 .28

.32 .36 8 14 .24 .28 .32 .36

10 9

.24 .28 .32 .36 12 .24 .28 .32 .36

10 14

.24 .28 .32 .36

12 9 .22 .26 .30 .34 12 12 .24

.28 .32 .36

12 14

100 10 120 130 .24 .28 .32 .36

ORGINAL PAGE

IS OF POOR QUX

.j

1-41 1.2.2.2 Cruise Speed Reduction As previously stated, reduction of the cruise speed for the propfan air craft has the potential of additional fuel savings due to the improvement in the fuel consumption characteristics of the turboshaft engine at the reduced thrust level. Figure 18 depicts the trend of fuel consumption versus Mach number for the turboshaft engine. Mach 0.75 was selected as the reduced cruise .speed since this value is consistent with current practice on short to medium range transports. Reducing the cruise speed to Mach 0.75 effects additional fuel savings (additional to engine fuel consumption characteristics) since the propfan noise and the required thrust are both reduced which in turn allow a reduction in acoustic treatment weight.

Resizing of the baseline propfan aircraft was completed with minimum DOC at 600/gallon fuel cost as the criteria for optimization.

Values of W/S and T/W were varied as follows for the parametric analysis: W/S = 100, 110, 120, and 130 T/W = .22, .24, .26, and .28 Figure 19 depicts the selection of optimum W/S and T/W values consistent with minimum DOC at 60/gallon fuel cost for the turboprop point design at Mach 0.75 cruise. Design and performance characteristics of the propfan aircraft, for the 1500 nautical mile, Mach 0.75 design mission, are shown in Table 15.

The effect of reducing the design cruise speed to Mach 0.75 is a savings in block fuel and DOC, at 600/gallon fuel cost, of 21.1 percent and 10 percent, respectively, when compared to the baseline (1500 nautical mile, Mach 0.8) turbofan aircraft.

-1.2.2.3 Alternate Turboshaft Engine Subsequent to the previous RECAT study, an alternate turboshaft engine, Detroit Allison Diesel PD 370-22, was identified which offers an overall pressure ratio and component technology comparable to the JT10D-2 turbofan.

Utilization of this engine in the propfan aircraft offers the potential of 1-42 additional fuel savings.

The PD 370-22 engine has an overall pressure ratio of 25:1 and maximum turbine inlet temperature of 2500°F as compared to 28:1 and 2400°F for the JTlOD-2 turbofan.

A description of the PD 370-22 engine parameters is shown in Table 17.

Installed performance data, along with engine and nacelle dimensions, engine weight, and appropriate scaling factor, was provided by Allison for adaptation to the RECAT design mission. The Hamilton Standard 8 bladed prop fan, operating at 800 fps tip speed, was utilized.

Installation guidelines previously supplied by Hamilton Standard were applied, where appropriate.

A propfan disk loading of 42 Shp/D was selected from the propeller sizing study which resulted in a requirement of 4720 lbs of acoustic treatment in the air craft fuselage.

The propfan baseline aircraft (revised) was resized to incorporate the PD 370-22 turboshaft engine for a 1500 nautical mile, Mach 0.8 design mission.

Resizing of the aircraft was accomplished using the ASSET parametric analysis with the following variations in W/S and T/W: W/S T/W 0.20 0.22 0.24 0.26 The ASSET carpet plot, shown as Figure 20, depicts the selection of values for W/S and T/W, based on minimum DOC at 600/gal.

fuel cost, for the point design.

Point design parameters selected were AR = 10, t/c = 12%, W/S = 108, and T/W = 0.22. Design and performance characteristics of the aircraft are included in Table 15.

The effect of incorporating the PD 370-22 engine for the 1500 nautical mile, Mach 0.8 design mission is a savings in mission fuel of 1.1 percent and a savings in DOC, at 60/gal.

fuel cost, of 2.0 percent due to a decrease in installed propulsion system weight relative to the revised STS 476 baseline.

OIGINAL PAGE IS

OF POOR QUALITY

1-43

TABLE 17. ENGINE DATA COMPARISON JT10D-2 STF 477 STS 476 PD 370-22 STS 487,

A

'A

* CYCLE TURBOFAN TURBOFAN TURBOSHAFT TURBOSHAFT TURBOSHAFT * MANUFACTURER P&W ,P&W P&W DDAD P&W * IOC (yr) 1981/82 1990+ 1983 1985 1990+' * RATING (SLS) THRUST (lb) 24500 26550 NA NA NA HORSEPOWER (hp) NA NA 9294 12328 20624 * TIT (OF) 2470 2600 2400 2500 * PRESSURE RATIO 27.3 45 20 25 4o.

* BYPASS RATIO 5.6 8.0 NA NA NA

* WEIGHT (ib) A

48oo 3940

2180 1566 2134

* UNINSTALLED PERFORMANCE

A

M = 0.8 35000 FT NRP THRUST (LB) 5683 6530 3363 3832 548o SFC (LB/HR/LB) 0.638 0.542 .515 0.509 0.444

NOTES: A Engines

for initial RECAT study, NASA CR-137926

A Turboshaft engine weights

for gas generator only (does not include gearbox or prop weights)

A Performance calculation

assumes the following: 'Turbofan: uninstalled with 18400 BTU/LB fuel heating value Turboshaft: uninstalled with 82% propeller efficiency, 99% gearbox efficiency and 18400 BTU/LB fuel heating value.

1.2.2.4 Off Design Cruise Speed Effects Included as a part of this study, an assessment of the baseline propfan aircraft design, flying at cruise speed of Mach 0.75, was accomplished.

Re sizing for this mission condition was not accomplished so that the effect of operation of the aircraft, sized for a specific design mission and operated at an off-design condition (consistent with current operator experience) could be assessed. The propfan aircraft baseline configuration, CL 1320-15, was subjected to the same mission profile utilized throughout the study with the cruise speed

reduced from Mach 0.8 to Mach 0.75. Take-off

gross weight of the aircraft was maintained at 217,466 pounds (identical to the baseline) and the effect of reduced cruise speed on aircraft performince and economics was determined.

For the design mission range of 1500 nautical miles, flying the propfan aircraft at a cruise speed of Mach 0.75 effects a savings in mission fuel of approximately 2.4 percent and a savings in DOC, at 600/gal. fuel cost, of approximately 1.1 percent relative to the same aircraft at Mach 0.8. For the same mission fuel as the baseline aircraft, the design range can be increased to approximately 1600 nautical miles (approximately 6.2 percent increase).

1.2.3 1990 Technology Assessment The 1990 technology assessment consisted of incorporating an advanced technology turboshaft engine, Pratt and Whitney STS 487, representative of that which could be available for a 1990 IOC aircraft.

The airframe technology levels (supercritical wing, advanced composites, and active controls) utilized for the 1985 IOC aircraft were retained as was the 8 bladed, 800 fps tip speed propfan.

Resizing of the propfan powered aircraft was accomplished during this assessment for the 1500 nautical mile, Mach 0.8 design mission with the STS 487 turboshaft engine. Resizing criteria was minimum DOC at 60/gal.

fuel cost.

ORIGINAL PAGE

IS

POOR QUALITY

1-45 OF

1.2.3.1 Advanced Technology Turboshaft Engine The Pratt and Whitney STS 487 turboshaft engine resulted from a Pratt and Whitney -contract -with- NASA-Lew-is Research -Center to study unconventional engines designed for low energy consumption for medium and long range transport appli cation.' The STS 487 engine employs the same advanced technology features as the STF 477 turbofan. A description of the engine design parameters is included ,in Table 17.

Performance data for the STS 487 engine, along with engine and nacelle dimensions, engine weight, and appropriate scaling factor was provided by Pratt and Whitney for adaptation to the RECAT design mission requirements.

2 2 l. .3. Aircraft Optimization Optimization of the 1990 technology propfan aircraft was accomplished by resizing the baseline configuration to incorporate the STS 487 turboshaft engine and the 8 bladed, 800 fps tip speed propfan. Utilization of the STS 487 engine necessitated alterations of the ASSET sub-routines for con figuration, weight, drag, and engine performance consistent with the perfor mance and dimensional data supplied by Pratt and Whitney for the engine and by Hamilton Standard for the propfan. The propfan baseline configuration was resized using the ASSET parametric analysis with minimum DOC at 60C/gal. fuel cost as the optimization criteria. For the parametric analysis, wing AR and t/c were maintained at 10 and 12% respectively with values of W/S and T/W varied as follows: W/S T/W 0.18 110 0.20 120 0.22 130 0.24 Figure 21, the ASSET carpet plot, depicts the selection of W/S and T/W values to be utilized for the aircraft point design.

1-46 Point design parameters selected were AR = 10, t/C = 12%, W/S = 107.5, and T/W = 0.18. The performance and design characteristics of the aircraft with the STS 487 engine at the 1500 nautical mile, Mach 0.8 design mission are shown in Table 15.

Incorporation of the STS 487 turboshaft engine results in a 1990 I0C aircraft at the 1500 nautical mile design mission with a block fuel savings of 10.8 percent and a DOC savings of 7.4 percent, at 600/gal. fuel cost, when compared to the 1985 IOC baseline aircraft. These savings are the result of an improvement in engine SFC at cruise, of approximately 7.6 percent and a reduction in propulsion system installed weight of approximately 34 percent.

ORIGINAL PAGE IS

OF POOR QTAL J

1-47

SECTION PROPFAN TECHNOLOGY BASE Initial assessment of the fuel conservation potential of the turboprop aircraft, Task 7 of the previous RECAT study (Report CR 137926), utilized the 8 bladed propfan operating at a tip speed of 800 fps. Performance and acoustic characteristics of the propfan were supplied by Hamilton Standard per their reports SP02A76, dated 27 February 1976, and SPO9A76, dated March l976.

Subsequent to the initiation of this study, updated performance and acoustic characteristics, for the propfans were supplied by Hamilton Standard as a result of their ongoing propfan wind tunnel tests. This data, supplied on 13 June 1976, is included as Appendix A of this report. The effect of the revised propfan data is a slight increase in efficiency accompanied by a slight increase in induced SPL for the 8 bladed, 800 fps tip speed propfan.

Also included in the revised data package, is definition of the directivity of propfan induced SPL (directivity of impingement on fuselage wall). Incor poration of the revised acoustic characteristics and directivity pattern re sulted in a revision to the acoustic treatment methodology utilized for the aircraft fuselage wall.

This revision in methodology and resultant acoustic treatment weights is discussed in Section 3 of this report.

The propeller sizing study, for the various turboprop aircraft designs, was reviewed with the foll6wing disk loadings and acoustic treatment weights established: 2-1 REVISED ACOUSTIC 0RIG. ACOUSTIC CONFIGURATION DISK LOADING WEIGHT (lb) WEIGHT (lb) STs 476, 1500 NMI, o.8M 37.1 SHP/D 5220 STS 476, 2000 NMI, 0.8M 37.1 SHP/D 5445 3089 STS 476, 1500 NMI, 0.75M 35.9 SHP/D 4405 PD 370-22, 1500 NMI, 0.8M 42 SHP/D 4720 STS 487, 1500 NMI, O.8M 46 SHP/D 4390 The above listed acoustic treatment weights (revised) and attendant propeller efficiencies at the selected disk loadings were incorporated into the turboprop aircraft designs. The net effect of this data on aircraft performance is a slight increase in block fuel of approximately 0.1 percent and an increase in DOC, at 0C/gal.

fuel cost, of approximately 0.6 percent for the 1500 nauti'cal mile, Mach 0.8 design mission.

ORIGINAL PAGE IS OF POOR QUALITY 2-2

SECTION 3 ACOUSTIC TREATMENT METHOD 3.1 TYPICAL TURBOFAN CABIN NOISE ENVIRONMENT For current widebody turbofan aircraft, the maximum interior noise level at high speed cruise conditions is dominated by turbulent boundary layer induced vibrations of the cabin wall. The cabin wall vibrations cause acous tic radiation to the interior in a manner similar to a loudspeaker.

The transmitted boundary layer noise is broadband in its frequency content, excit ing many structural vibration modes, the listener perceives an innocuous "whooshing" sound.

The peak boundary layer excitation frequency is of the order of Uo/6 where U is the free stream velocity and, 6 , is the boundary layer thickness. Since the boundary layer external pressure fluctuation spectrum varies slowly with frequency, near the peak frequency, the maximum interior sound-pressure will occur at frequencies near U o/6, but within a frequency band where also a condition of coincidence exists (phase velocity equality) between the boundary layer turbulent pressure fluctuation pattern and the flexural waves in the fuselage.

The boundary layer thickness on a typical fuselage can be estimated at percent of the equivalent flat plate boundary layer thickness.

The 20 percent factor allows for roughness and adverse pressure gradients. At Reynolds Numbers above 100 million (typical full scale flight) the boundary layer velocity profile varies approximately as the one seventh power of dis tance normal to the surface (Ref 1 p 536). In this case the boundary thick ness is 5.14 times skin friction coefficient, C., times L, the distance from the nose. At a Mach number of 0.85 and at 30,000 ft, the Reynolds Number is 2.42 x 106 per foot. For a representative aft cabin point, L = 150 ft, Re = 364 x l0 , CF = 0.00180, the flat plate boundary I layer thickness would be, 6 = 0.00924L.

3-1 Increasing this by 20 percent yield 6/, = 0.0109 at a point L = 150 from the = nose, hence'6 1.66 ft. The external free stream velocity is 827 ft/sec, and -the typical peak frequency for aft cabin noise (At L = 150 ft) would be 90 Hz.1 In the forward cabin the BL excitation frequency would be 3 times higher and perhaps twice as high in the mid cabin. Typical measured maximum interior noise levels at window seats for these turbofan aircraft range from 90 to 95 dB for OASPL and the A weighted SPL's are from 80 to 85 dBA.

3.2 PROPFAN PASSENGER COMFORT CRITERION For turboprops a preliminary interior noise comfort criterion has been selected at 90 dB SPL for the transmitted blade passage frequency harmonic tone. For a pure tone at a blade passage frequency of 160 Hz the 90 dB tone SPL value would correspond to an "A weighted" SPL of 75 dBA (Ref.

2, p 16-13).

The second harmonic tone in this example is at 320 Hz and a 90 dB tone would correspond to a value of 83 dBA. In order that the sum of the first two har monics should not exceed 75 dBA, it would be necessary that each tone could contribute only 72 dBA. This means that if the fundamental tone level was allowed to be 87 dB then the second harmonic could be 79 dB.

The transmission loss concept utilized for this study is a heavily damped, massive double wall construction separated by an airspace, as shown in Figure 22.

This concept produces 18 dB of added transmission loss for each doubling of frequency above the "mass-air-mass frequency" (air space stiffness resonance frequency of the double wall masses). The double "limp wall" mass law theory has important consequences in that higher harmonics are rapidly suppressed.

The theory is discussed more fully in Section 3.4.1 below. Data received from Hamilton Standard for the current 8-bladed propfan shows that the external tone level'SPL values for the first four harmonics, relative to the OASPL, are -1, -9, -15 and -20 dB, at a tip speed of 800 ft/sec. From these data it is clear that one would expect the interior tone level SPL for the second harmonic (2 times fBP) to be lower by 26 dB than the blade passage frequency tone level.

3-2 TRANSMITTED TRIM PANEL (INNER WALL) SOUND WAVE VIRATION I O A O SV I BR A T I O N IISOLATORS TRINGS R SAIRSPACE VISCOELASTICX SOUND WAVE INPUT EXTERNAL

IS

PAGE

ORIGINAL

QUALITY

pOOR

OF

for Acoustic Treatment "Limp Wall" Concept Double Figure 22.

Cabin Walls of 3-3 In conclusion, one can say that the blade passage frequency tone will dominate the interior noise level. The selected design level 90 dB corre sponds to 75 dBA at 160 Hz and 83 dBA at 320 Hz which is the upper range of blade passage frequencies for the lO-bladed propeller at 800 fps. These values compare favorably with the

80 to 85 dBA range

for current turbofans.

Therefore, the passenger comfort criterion selected may even be slightly con servative at the lower blade passage frequencies 3.3 PROCEDURE FOR DEFINING ACOUSTIC TREATMENT WEIGHT The required procedure is as follows: (1) Define the external sound pressure distribution on the cabin surface in terms of circumferential variation, and axial distance from the propeller disc plane.

.(2) Define the required noise transmission loss-(NTL) between the exterior SPL and the design goal interior SPL. The NTL is defined by: STL=SPL (X, e ) SPesign E = SPLE - 90 dB (3) Compute the total weight per unit area required (including practical design constraints) to achieve the specified NTL. The acoustic penalty is the increment above the reference turbofan weight per unit area.

(4) Integrate the excess acoustic treatment weight per unit area over the cabin wall.

The next subsection on acoustics discusses the cabin wall transmission loss aspects. The remaining acoustics subsections will (1) compare the cur rent external near field noise data with the predictions of the previous RECAT Study (Ref.

4) and, (2) will show weight penalty results 6f the current study, and their comparison with previous RECAT results.

3.4 CABIN

WALL TRANSMISSION LOSS PREDICTION 3.4.1 Transmission Loss Assumptions The noise transmission loss, NTL, for the cabin wall is predicted on the basis of double "limp wall" mass law theory as described in pp 187-189 of 3-4 herein as Ref. 4 and in Table 126, p 223 of Ref. 4, which is reproduced the structural response is dominated by Table 18. This theory assumes that frequencies, the sum of a large number of vibration modes at non-resonant rather than a few resonant modes. This theory is plausible if one assumes that part of the mass on each of the walls is a suitable viscoelastic damping material. The outer wall mass consists of the outer skin plus the rings and stringer masses are added to the skin at frequencies stringers. The ring and below the ring frequency (about 288 Hz for a 19.58 ft diameter aluminum fuse lage (Ref. 4, p 189)) because the flexural wave lengths are much longer than the structural bay lengths.

The double "limp wall" theory is presented in approximate form by Cremar, in Refer Heckl, and Ungar (Equation 79a, pg. 505 of Reference 3). Also shown ence 3 is an alternate expression for the double wall increment of noise trans mission loss (NTL) due to vibrations transmitted through the vibration isolators interior. One (see Figure 22) which provide a possible "flanking path" to the limitation of double wall theory is that the lower value of NTL should be described in Table 18 or that through the chosen (either that of the air path "flanking path" afforded by the trim panel vibration isolator): 18 lOgl0=2 2) ANTLisolator = 0 0- In this equation a trim panel bay of area S is attached to the outer wall via n vibration isolators, The Velocities, V1 and V , represent the vibration isolator attach points.

velocities of the outer and inner walls at the kc = C/fe, represents the critical wave length, C is the speed of sound in the cabin air, and fc the critical frequency above which a vibrating skin panel achieves maximum acoustic radiation efficiency (Reference 3, pg. 482 and 492).

For the case of air at 700F and aluminum

IS

PAGE

ODYGINAL

OF, poop.

QUA

3-5

TABLE 18. DOUBLE WALL MASS LAW - "LIMP WALL" THEORY Noise Transmission

Loss (NTL) = 20 LOG 7'

+ 20 LOG - 1 F K Total Wall Double Wall a M1 2 Mass Law Increment n 2r + M )

6 dB Per Octave 12 dB

Per Octave When M = M 1 Total Increase in NTL Per Octave is 18 dB.

2 1 fn T d (-)w 1.8 = hw c C air = hw (1.8) (16380) = 26.13 hw where hw is the outer wall skin thickness.

example, for hw = 0.060 inches, the critical frequency becomes For C . C air _ air 26.13 hw c = Hz (0.00567) (26.13) = 1.78 inches and k = (26.13) (0.068) = as those depicted in Figure 22) For well designed vibration isolators (such airpath ratio V /V , of 1/10 or less, then the which could provide a velocity 2 1 would be the critical path-and the mass law theory is valid. In general, the a mass mounted on soft springs, and if the trim panel can be considered as the air stiffness between the is sufficiently softer than isolator spring and the isolators dominates the noise transmission, double wall, then the airpath circuit" path for noise transmission.

will not cause an undesired "short for preliminary design The double wall mass law theory is convenient reflect the realities of cylindrical shell purposes; however, it does not single with Lockheed laboratory research tests on dynamics. Some comparisons above about one-half that mass law theory is toe optimistic wall cylinders show at lower frequencies.

frequency, but is somewhat conservative of the ring sensitivity to incidence angle.

Also, test data show some 3.4.2 Structural Design Constraints the total wall NTL equations (Table 18) are solved to find The double a specified NTL. This weight wall weight per unit area required to achieve

3-7 ORIGINAL PAGE is

POOR QUALITY

OF

increased because of a number of constraints beyond those used is, however, prediction modifications in the previous Study (Ref. 4). The treatment weight for the current study are as follows: 61.5 sq ft of (1) The treatment now covers the entire circumference 16 sq. ft of side wall treatment per ft of cabin cabin length versus length previously used.

(2) The design NTL is decreased stepwise by 10 dB with distance from external SPL

the propeller disk plane in 5 steps, according to new

directivity data. The required treatment segment lengths vary with relative tip clearance according to Figures 23 and 24. An example of the longitudinal distribution of treatment material is shown in .8.

Figure 25 for a relative tip clearance Ay/D = 22 shows plots of the required ratios of segment treatment Figure In the first length to prop-fan diameter for each of five segments.

loss is based on the maximum segment the required noise transmission exterior SPL at blade passage frequency (see Section 3.3 item (2)).

segment k = 1, Thus, for = NTL (1) SPL - 9O dB = NTLma E x max loss is For segments k = 2 to 5 the required transmission =10 d.B NTL(k) = (SPLE - 90 dB) - (k-1) max We define the increment of required transmission loss for each segment (k = 1 to 5) as follows ANTL (k) = NTL(k) - NTLma max = (k-1) X 10 dB It is noted that the mathematical model of AL /D is quite conserva (5) below, however).

under item (see remarks k = 2 tive for segment

3-8

SOURCE KEY NO BLADES SYMBOL 8 H.S. DATA DATA ------ 10 H.S.

MATH MODEL OF LENGTH 1ST SEGMENT TREATMENT 1) (k = = = : NTL SPLEMAx -9 dB NTLMA Z Lt MODEL \1 x = 0 ANTL(1) = NTL - NTLMAX D REQUIRED TRANSMISSION LOSS 1.0 NTL= 1.6 2.0 0.4 0.8 1.2 _ _ _ _ 2.0 2nd SEGMENT LENGTH -10 dB (k = 2) NTL = NTLMAX A Lt D 1.0NTL=-0dB 2.0 1.6 0.4 0.8 1.2 DIAMETER (Ay/D) PROPELLER TIP OF FUSELAGE CLEARANCE/PROPELLER Relative Tip Clearance Treatment Length With Figure 23, Required PAGE IS ORIGINAL 3-9 QUALITY OF POOR SYMBOL NO BLADES MATH MODEL LENGTH 3RD SEGMENT 1.0 = - 20 dB (k3) NTL MAX . .

.L A .

t

t(ANTL -20 do)

-~ I .A~h~~4

D .1 o . srrt 0 I 1.2 1.6 2.0 0.4 0.8 - 4TH SEGMENT LENGTH 1.0 _ NTL = MAX -30 dB _ - ._ (k4) .O AL t dB) (ANTL=-30 D 1.6 2.0 0.8 1.2 0.4 SEGMENT LENGTH .o - - -TH (k'5) NTL - MAX - 40 dB _ t (ANTL=--40 dB) ALt D 2.0 0.4 0.8 1.2 1.6 DIAMETER ( y/D) .PROPELLER TIP TO FUSELAGE CLEARANCE/PROPELLER With Relative Tip Clearance Figure 24. Required Treatment Length 3-10 D +10 ANTL

M Ii'

dB X/D, AXIAL DISTANCE (FWD OF DISC PLANE) AV/D 0.47 - 1.23 k=1, ANTL=0 dB I I I

t I I I

2.0 1.0 -11 -2 -3 -4 ..

-10 L . 9.

k- 2, ANTL =-10dB 7- 3 k=3, ANTL = -2 dB 0.3 0.3 -30 - k4ANTL=-30OdB 0.62 0DATA 0.6 FROM HAM STD, qk=5,A TLLL -40 dB -40 JUNE, 1977 END OF -50END OF TREATMENT TRA TENTZ Figure 25.

Example of 5 Step Acoustic Treatment for a Relative Tip Clearance of N-A /D = 0.8

OpIIA 'PAGE

IS

&IGINAL3-11

QUAJ1TY

pCOR ()F ,(3) Minimum inner and outer wall weight structural design constraints are imposed for each of the 5 treatment segments. These constraints are summarized in Table 19. The most notable feature is the choice of double the reference outer wall weight to 2.4 psf for the minimum outer wall weight for the first two segments nearest the propeller disk plane. At a relative tip clearance, A y/D, of 0.8 these two segments cover Lt = 3.6D, where, D is the propeller diameter. For the segments 3 and 4 the minimum weights are increased by 25 percent (to 1.5 psf). Finally, segment 5 has a minimum increase of 10 per cent to the outer wall. The minimum weight increases are provided to allow for viscoelastic damping material which will force the outer wall to behave in accordance with the limp wall mass theory.

(4) Minimum inner wall trim panel weight constraints (see Table 19) are set at 75 percent of the reference turbofan trim panel weight (0.33 psf).

In some cases because of the heavy minimum values of outer wall weight (due to constraints stipulated above) the double wall theory would require even less weight for the trim panel; therefore, this constraint is considered as a structural constraint.

(5) The total treatment length now varies according to diameter and propeller tip clearance as discussed above.

In the previous study (Ref. 4) the total cabin length was treated*(even though confined to the side walls), because of uncertainty concerning the axial location of the maximum external SPL signature. In the present study it was decided to adopt the external SPL levels and directivity data of Appendix A (measured at M = 0.3), with the understanding that these data are subject to future revision when new external SPL is avail able at flight Mach numbers of 0.70 to 0.80. Lockheed believes the shockwave position uncertainty discussions of Ref. 4, pp 187-188, Figures 70 to 73 pp 207 and 208, and Table 125, p 222 are still relevant to the external SPL environment at cruise Mach numbers of 0.7 to 0.8. There it is noted that at the high flight Mach numbers, the shockwave pattern defining the external sound pressure would be moved farther aft of the'disk plane than would be the case for the test data at a tunnel flight Mach number of 0.3. It is possible that the total signature length would not be greatly different, even though the axial location of peak intensity might vary from the pattern shown in Figure 25; The extra treatment length is an attempt to provide a margin of safety due to the anticipated variability of signature due to shockwave position change with flight Mach number, as discussed in Ref. 4.

(6) The current study is restricted to an airspace depth of 4.8 inches, as was used in the previous study. Increased airspace depth would be beneficial and should be considered in future design studies.

(7) In this study acoustic treatment weight variation with propeller diameter is considered during the selection of optimum propeller 3-12 TABLE 19. WALL WEIGHT CONSTRAINTS FOB 5 STEP DOUBLE WALL TREATMNT IaA (Lt\ ANTL (W /A) (W /A) SEGMENT DJK dB PSF AF 1 0.2 1.1 0 2.4 0.25 0.4 1.5 o.8

1.7

1.2 2.1 1.6 2.5 0 2.4 0.25 2 0.2 1.50 -10 2.4 0.25

o.4

1.63

0.8 1.90

1.2 2.17 -10 2.4 0.25 3 0.2 o.4o -20 1.5 0.25 o.4 o.47

o.8 0.60

1.2 0.73 -20 1.5 0.25 4 0.2 o.4o -30 1.5 0.25 o.4 0.47

o.8

o.6o

1.2 0.73 -30 1.5 0.25 5 0.2 0.40 -40 1.32 0.25 0.4 o.68 0.8 1.24 1.2 1.80 -4o 1.5 0.25

IS

PAGE

OBIGINAL

(QUAILMY

POOR

OF

3-13 diameter, blade count, tip speed, and cruise Mach number, for a range of net thrust values appropriate to each point design aircraft (payload, range and cruise speed combination).

(8) The propeller shaft axis is held at a fixed spanwise location as propeller diameter is varied in this study. This maintains constancy of nonacoustic weight of the wing and empennage structure even though propeller tip to fuselage clearance now varies also as the propeller diameter is changed.

3.5 EXTERIOR NOISE DATA Appendix A contains the exterior noise data supplied by Hamilton Standard.

Appendix A shows OASPL data versus tip speed'at cruise Mach numbers of 0.7, 0.75 and 0.8 for both eight and ten bladed propellers at 30,000 ft altitude, at-various propeller efficiency values, for a relative tip clearance,Ay/D, of 0.8. Appendix A also shows the estimated directivity data for the eight and ten bladed propfan designs, taking into account the most recent test data.

Also shown are the increments in SPL levels (HL's) of the blade passage fre quency harmonics, relative to the OASPL. These are the data used in the present study. The acoustic data of Appendix A differs somewhat from the preliminary Lockheed predictions used in the previous study (Ref. 4, pp 187-188). Table 20 provides a comparison of data used for the previous study and this assessment.

3.5.1 Previous and Current Prediction Results The first noticeable difference is the external SPL at blade passage fre quency.

Lockheed estimated the values shown in Table 21, which is a reproduc tion of Table 125 p 222 of Ref. 4. It is noted in Table 21that Lockheed and Hamilton Standard prediction methods were apparently in fairly close agreement as to the blade passage frequency harmonic (n = 1) SPL value (124 dB vs 126 dB); both of these values are much lower than the current prediction of 134.5 dB.

It is noted, however, that Hamilton Standard originally predicted (unpublished Ref. 5) an OASPL of 136 dB with a -10 dB correction for each of the first data, 10 harmonics. Furthermore, the altitude correction used in Ref. 5 and also, presumably, for the current data was only -4.3 dB re sea level. By contrast, the Lockheed original altitude was evaluated for 35,000 ft, and correction included in Table 20 was -12.5 dB, based on 20 Logl (Pamb/PSL). This repre sents dynamic pressure scaling at constant helical tip Mach number.

3-14 TABLE 20. COMPARISON OF PBEVIOUS RECAT VS. CURRENT CABIN NOISE TREATMENT METHODOLOGY AND DATA CURRENT PREVIOUS HAM STD JUNE 1977 External SPL Level SPL, @ BPF (M = 0.8, VL = 800 124 dB (134.5 dB @ same o conditions) = 0.8) D = 12.8,Ay/D Clearance, Ay/D 0.8 Varied in prop size trade

Tip Speed V (ft/sec) 8oo 800, 700, 6c0

t

Point Design Cruise Mach No, M 0.80 0.80, 0.75, 0.70

2 2 Tradeoff variable SHP/D re (HP/ft ) 37.1 f Net Thrust (lb/engine) 3860 3860, Diameter (ft) (@ 37.1 SHP/D ) 12.8 12.55/traded Prop Efficiency 0.82 0.83/traded Number of Blades 8 8, 10 1568 f (Ay/D,D) per Ham Std Treated Areas (sq. ft.)

Data June, 1977 per Ham Std Cabin Diameter (ft) 19.58 f (Ay/D,D) Data June, 1977 Cabin Length (ft) 98 f CAy/D,D) per Ham Std Data June, 1977 Std Total Cabin Surface 6028 f (Ay/D,D) per Ham Area (sq. ft) Data June, 1977 26.0% 100% % of Circumference Treated Constraints None New minimum wall weights near prop plane Defined Treatment Method Damped Damped Double Wall Double wall ft 30,000 ft Altitude 30,000 Blade Passae Freq (Hz) 159.2 163.2/traded @ Ref SHP D 3-15

IS

OhIGINAL PAGE

QUALITZ

oF POOR

TABLE 21. EXTERNAL SPL RESULTS

M = 0.8 8 Blades, D =

12.6 ft V = 8oo ft/s

rLE/cb = 0.0015 Clearance: 0.8 Dp m = l.06 H SPL HARMONIC

FREQUENCY LOCKHEED ©

HAM STD.© Blade Passage 156 Hz 124 dB 126 dB Second Harmonic 313 Hz 121 dB 126 dB Third Harmonic 470 Hz 116 dB 126 dB Fourth Harmonic 626 Hz 104 dB 126 dB * Pulse time/blade passage period = 0.330 * Cosinusoidal pulse -- ry_...

( Note

Lockheed calculations performed for 35,000 ft, including an altitude correction of -12.5 dB. At 30,000 ft the correction is -10.5 dB.

@ HS data includes an altitude correction of -4.3 dB at 30,000 ft altitude.

It is noted that altitude corrections used by Lockheed in Reference 4 and the current Hamilton Standard theory (Reference 6) both utilize a correction of 20 log (Pamb/PSL) which corresponds to the near-field noise of 1 0 a single blade or fixed blade area at fixed values of rotational and forward Mach number, and at a fixed relative blade tip clearance distance. The original Hamilton Standard altitude correction of Reference 5 is exactly equal to 10 log (PPSL) . The method of Reference 5 is 1 0 a preliminary design type method.

It would represent the variation with altitude of the noise of a loading (SHP/D of prop-fan power source for a fixed value dipole acoustic and for fixed values of tip speed, helical tip blade number, and relative blade tip to fuselage clearance. The "old" and "new" altitude corrections are essentially consistent because a factor of 10 log (Pamb/PSL) is absorbed into the SHP/D factor of the "old" (Reference 5) method.

Notice in Table 22 that Lockheed's prediction of harmonic level variation given in Ref. 4, Table 125, was more realistic than Hamilton Standard's origi nal prediction. A larger altitude correction may be more accurate, which means the new acoustic data could be too pessimistic in this respect by 5 to TABLE 22. COMPARISON OF HARMONIC LEVELS OF EXTERNAL SPL DATA AT 800 FT/SEC TIP SPEED, MCR = 0.80 Definition: HL(n) = SPL(n) MINUS OASPL, dB PREVIOUS CURRENT DATA (REF. 4, TABLE 125) APPENDIX A SOURCE LOCKHEED HAM STANDARD HAM STANDARD n HL(n) HL(n) HL(n) dB dB db -2.2 -10 -1 2 -5.2 -10 -9 -10.2 -10 -15 4 -22.2 -10 -20 3-17

ORIGINAL

PAqn Is

OF POOR QUAJTYI'

predicted data are based 30,000 ft. On the other hand, the currently

6 dB at

provides the wrong on tests at a tunnel (flight) Mach number of 0.3; this helical tip-Mach angle, even if the resultant, supersonic propeller advance The and therefore the directivity may be questioned.

number is matched, is con impingement in Ref. 4, pp 187-188, Lockheed discussion of shockwave of these this respect. It is possible that many sidered still pertinent to errors estimates may produce cancelling differences in exterior sound pressure when all in acoustic treatment weight penalties and lead to small differences corrections are taken together.

Prediction 3.5.2 Exterior Near Field Noise analysis of external SPL contained a Note in Table 21, that Lockheed's about what may be possible with number of blade shape oriented assumptions leading edge radius, and the effective regard especially to achieving a small in reducing the effective helical tip Mach number. These ness of blade sweep detachment. These estimates variables all affect the estimates of shockwave geometry of any of detailed knowledge of the exact were made without benefit or contemplated as a basis for the of the fan blades which were actually tested in Appendix A. It is believed that the shockwave new data package contained to better blade concepts analysis approach used by Lockheed in Ref. 4 could lead methods and deserves further development. There are other noise prediction which are reviewed here briefly.

6) has recently published a more elaborate It is noted that Hanson (Ref.

scheme based on Ffowcs-Williams and Hawkings solution (Ref. 7)

analytical equations of fluid motion. Hanson's analysis is based of the acoustic analogy using fluid fixed coordinates, which cannot account for on linearized theory field. Hanson the shockwave effects which are present in the near geometric and Regier's static data (Ref. 8) claims to have good agreement with Hubbard

of 4 and 8 inches on a 4 ft

at tip Mach numbers up to 1.0, and for clearances of the first two harmonics diameter prop. Hanson also states that calculations P-51 Mustang test data at a helical tip Mach num were in good agreement with a dive (Ref. 9). Hanson shows agreement with ber, M = 1.07, obtained in H method

at a distance of 5 rotor diameters. His

results by Farassat (Ref. 10) suprising that it field prediction; however, it is appears logical for far 3-18 small clearances. More test data is predicts the Hubbard-Regier data at Mach numbers speeds at high subsonic forward needed for supersonic helical edge and for various leading blade thickness distributions for various to or loading effects appear ratios. Blade angle of attack radius to chord operated near proposed thus far, when the propfan designs be secondary for peak efficiency PENALTY DATA 3.6 ACOUSTIC TREATMENT WEIGHT differences of parameter and methodology Table 20 shows a summary current RECAT studies.

the previous and penalties used for between weight the previous subsections.

have been discussed in differences The methodology disc power effects of propeller are parametric to be considered What remains level, propeller design cruise thrust design Mach number, loading, cruise and blade count.

tip speed, is , it SHP/D disk power loading, effect of propeller evaluating the In This remains constant.

shaft centerline position that the propeller assumed propeller Ay/D, changes with blade tip clearance, means that the relative near clearance is still that the relative It turns out, however, diameter.

diameters).

(large propeller very low disk'loadings value, except for optimum the relative by optimizing weight could be achieved cases, a lower Fore these tip clearance of the propeller.

Penalty Data Treatment Weight of Acoustic Presentation 3.6.1 800 ft/sec, the treatment a tip speed of 26 and 27 show, for Figures propeller at 30,000 ft for an 8-bladed penalties versus SHP/D weight 3,000 lb respectively.

of 3,860 lb and thrust levels for propeller altitude Mach numbers.

at various cruise thrust required levels bracket the These thrust design cruise Mach for different point shown are curves Parametrically data for a 29 show the same Figures 28 and 0.75 and 0.80.

numbers of 0.7, clearly much penalties are These weight speed of 700 ft/sec.

propeller tip the effect 30 and 31 display speed. Figures for 800 ft/sec tip worse than and also treatment weight required acoustic speed upon the of propeller tip 3-19 PAGE IS ORIGINAL OF POOR QUALITY 18,000 16,000 14,000 C, ; 12,000 I.

,U Cj U' U 10,000 LU Q ORIGINAL RECAT 3089 LB ] • 37.1 HP/SQFT 0 MACH 0.8 AMACH 0.75 0.7 -MACH = 20 00,VTIP 800 FPS 0 1 7. HPSF NO BLADES - 8 F = 3860 LB N ALTITUDE = 30,N FT 50 60 70 0 20 30 40 HP/FT , SHP/D Weight, 8 Blades, 800 fps, 3860 lb Thrust Figure 26. Acoustic Treatment 3-20 14,000 12,000 ,-J 10,000 I z LU < 8000 w S v T6=8000P O - MACH 0.80 0- MACH 0.70 NO. BLADES = 8 ALTITUDE = 30,000 FT 20 30 40 50 60' 70 2 , HP/FT SHP/D Figure 27. Acoustic Treatment Weight, 8 Blades, 800 fps, 3000 lb Thrust 3-21

ORIGINAL PAGE IS

OF POOR QUALITY

16,000 1 MACH 0.8 AZ MACH 0.75 MACH 0.7 14,000 - FPS VTIP= 700 NO BLADES = 8 F - 3860 LB N ALTITUDE = 30,000 FT 12,000 10,00t111 z LU LU e.)

- 800 .J I-6 LI

z

______ 15 20 30 40 50 60 2 , HP/FT SHP/D Figure 28. Acoustic Treatment Weight.

8 Blades. 700 fPs. 3860 lb Thrust 3-22 16;000 14,000 -j 12,000 _ _ 2w 10,000

° --

LU i- C* ____ ____ - I 0 MACH 0.8

4000 EOl

400AGMACH .UA5T MACH 0OOI

0.75 - = 700 FPS VTIp NO BLADES = 8 FN = 3000 LB 2000 ALTITUDE = 30,000 FT_________ 0 15 20 30 so 60 2 2 SHP/D . HP/FT Figure 29. Acoustic Treatment Weight, 8 Blades, 700 fps, 3000 lb Thrust 3-23

ORIGINAL

PAGE IS

OF

POOR

QUALITY

30,000 FN = 3800 LB/ENGINE 8 BLADES OPERATION 4 M = 0.8 30,000 FT ALTITUDE O ~AT PEAK DOUBLE WALL ACOUSTIC EFFATIPEAK . 20,000 TREATMENT I 2 1 83% EFFICIENCY I LU w 10,000 < SHP =37.1 HP/FT (VARIABLE EFFICIENCY) D 700 800 TIP SPEED FT/SEC Figure 30. Acoustic Treatment Weight Required vs, Tip Speed 0.85 EFFCEA CONSTANT EFFICIENCY , ACHIEVABLE HP lHP37.1 z 0.80 FN = 3860 LB/ENGINE = 0.80" M O z 8 BLADES.

<L 0.

FT ALTITUDE 30,000 0.

x 0.75 0.70 600 700 800 TIP SPEED FT/SEC Figure 31, Propfan Efficiency vs, Tip Speed 3-24 prop-fan efficiency. The lower curve of Figure 30 shows the weight penalty if the same power loading SPD/D = 37.1 HP/sq ft was maintained. It is seen from Figure 31, however, that a significant loss of prop-fan efficiency would occur at the low tip speeds, if the same power loading was maintained.

The upper curve of Figure 31 shows the improvement in prop-fan efficiencies which could be achieved by reducing the power loading below 37.1 HP/sq ft; however, Figure 30 shows that the acoustic treatment penalty must be increased significantly to achieve these higher efficiency levels because the much larger propeller size and the corresponding lower blade passage of penalties frequencies. It is clear from Figure 30 that the acoustic treatment become prohibitive, if required for operation at a tip speed of 600 ft/sec a reasonable prop-fan efficiency is to be maintained. The absolute weight used with some caution at these low tip speeds, since penalties must be some other form of acoustic treatment may be more suitable. It is also other than the acoustic treatment noted that power plant weight components material requirements increase rapidly with prop fan diameter, and therefore, would further penalize a design based upon 600 ft/sec tip speed.

Figures 32 and 33 show results for a 10-bladed propeller at 800 ft/sec at the same thrust levels and cruise Mach numbers. These tip speed, show a clear advantage compared to the 8-bladed prop fan.

results It is evident that high disk loadings, high tip speeds, and high blade counts are desirable from the exclusive standpoint of minimum acoustic treat the current study, however, these data were used as inputs ment weight. In optimize the

the aercacoustic versus propulsion trade off studies Lo

to selection of propeller diameter. Interpolation of these data, for the selected thrust levels, results in the acoustic treatment weights reported The next section discusses the underlying in Section 2 of this report.

basis of the trends presented in these figures.

Effects on Acoustic Treatment 3.6.2 Disk Loading Figure 25 shows the effect of disk, loading for an 8-bladed profan at at 3860 lb thrust. Also shown is the original RECAT 800 ft/sec tip speed data point at the design disk loading, SHP/D = 37.1 HP/sq ft for 0.8 cruise 3-25 12,000 o MACH 0.8 A MACH 0.5 0 MACH 0.7 10,000 FPS VTIP= NO BLADES = 10 FN = 3860 LB ALTITUDE = 30,000 FT -U 20 3000D56 3-2 0.8 0 -MACH ai" 0.75 A MACH w MACH 0.7 \0 FP' 6000 VTIP = 800 I- = 10 BLADES NO.

= 3000 LB I-FN FT =30,000 ALTITUDE w _____6 24000 z

0 I20 30 40 50 60

, HP/FT SHPID 3-j lb Thrust

800 f'ps, 3000

10 Blades, Weight, Acoustic Treatment Figure 33.

3-27

IS

PAGE

O-RIGINAL

QUAITUY

OF POOR

<C 0. -j 9L z cc 'C w .

o2 F I Q <j -4 0.2 0.4 0.6 0,8 1.0 1.2 A Y/D Figure 34. External SPL Correction vs. Relative Tip Clearance 3-28 8 BLADES F =3860 LB n V =800 FT/SEC B=8 t N LT=-30,000 FT z Ui o 180 LU c- LU 0 SYMBO M " C Lu 140 0 .80 a 0.75

LL

El 0.70

10 20 30 40 50 60 SHP/D2. HP/SQ FT ICI a.

Figure 355. Blade Passage Frequency vs. Disk PowTer Loading ii . El 0.70 16 800 FT/SEC TIP SPEED, 3860 LB THRUST- Mo SYMBOL - _____ ,.j 13 -J L~U .12 10 20 30 40 50 60 Figure 36. Propeller Diameter vs.

Disk Power Loading 3-29

OR IGINAL

PAGE IS

OF POOP, QUALITY

EXTERNAL SOUND PRESSURE LEVEL .AT BLADE PASSAGE FREQUENCY, SPLH VS DISC POWER LOADING VTIP -800FPS NO. BLADES =8 AYr_=16.4 FT.

ALTITUDE=30,000 FT.

M -j

C

a: -.4 C.

a uLJ SYMBOL Moo

0.80

< 0

..

I= a 0.75 10 20 30 40 50 60 S'HP/D2, HP/SQ.FT.

Figure 37. External SPL at Blade Passage Frequency vs. Disk Power Loading 3-30 38 60 30,000 FT, FN= LB, Vt=800FT/SEC, 8 BLADES '*I'TOTAL'CABIN SURFACE AREA 6028.21 FT w SYMBOL MCR CC <.0 0.80 0.75 - z A w0 0.70 N U' LM 3000 I 'C ORIGINAL RECAT 1568 SQ FT 1000.

10 20 30 40 50 60 SHP/D , HP/SQ FT Figure 38. Total Acoustic Treatment Area Required vs. Disk Loading 3-31 Mach number. It is seen that for this study the acoustic treatment weight in the weight rela penalty is now 6500 lb, representing a-3400 lb increase tive to the previous study. Note also that increasing thei disk loading to by about 1500 lb to about 50 HP/sq ft would reduce the weight penalty consider the acoustic treatment 5100 lb. This empahsizes the need to of propeller size. Note that weight as a significant input in the selection are sized for the required weights are further reduced if the propellers Mach numbers of 0.75 or 0.70.

lower point design cruise The external SPL data of Ref. 3 (Appendix A) show-no difference between M = 0.8 and M = 0.75, at equal values of relative blade tip clearance.

er cr This is so, despite the reduction of helical tip Mach number by 0.08. The since the relative data in Figure 25 are affected by external SPL changes, blade tip clearance fy/D varies with the propeller diameter because .the ft. This assumption propeller shaft center line is maintained at fyc = 16.4 L constancy of wing and empennage weight (which would change if -the maintains to maintain a constant relative tip clearance).

engines were moved spanwise the present study the relative blade tip clearance, therefore, is varia In ble with propeller diameter, according to the following equation.

-Ay/D (16.4 - 1/2) Figure 34 shows the effect of relative tip clearance on the external SPL as determined from the data of Appendix A.

variable affecting the acoustic treatment weight re The-strongest quirement is the blade passage frequency, according to the double wall transmission loss loss theory. As described earlier, the transmission increases at 18 dB per octave increase of the blade passage frequency 17). Figure 35 shows the variation of the (Ref. 4 pp 187-189, and Table blade passage frequency with disk loading for the 800 ft/sec tip speed case.

Figure 36 shows the corresponding propeller diameter requirements which are determined by the thrust and propeller efficiency data of Ref. 3, given in Appendix A. The blade passage frequency is easily calculated, given the tip 3-32 speed, blade count, and propeller diameter as

vt

BVt BPF = lTD with disk loading for the Figure 37 shows the variation of external SPL tip speed, F = 3860 lb and M = 0.7, .75, and conditions of 8000 ft/sec n cr from the effect of external SPL, the interior noise is governed 0.80. Apart and propeller diameter.

by the blade passage frequency acoustic treatment area according Figure 38 shows the variation of total propeller diameter to the 5 segment treatment scheme. This increases with shows the and clearance as shown in Figure 23 and Table 17. Figure 39 = 800 ft/sec, total treatment length versus SHP/D for the same conditions (V t 40 shows the lengths of M = 0.80, 30,000 ft, F = 3850 lb/engine). Figure cr n wall weight the first two treatment segments, and Figure 41 shows the total the reference turbofan value, 1.53 psf) versus per unit area (including for each of the first two segments. Figure 42 shows the treatment area SHP/D and for the various segments. It is noted that the total treatment areas, even the treatment areas for the first two segments are considerably larger study.

than the fixed value of 1568 sq ft, used for the previous RECAT to treat the entire circumfer This is due to the decision in this study ft/ft of side wall ence of the cabin wall (61.5 sq ft/ft), rather than 16 sq important conservatism used in the present only. This represents the most study to offset the risk where many technological uncertainties exist. It is anticipated that the treatment weight per unit area could be reduced circumferential near the top and bottom of the fuselage, if reliable available for the external SPL.

distribution data were is the The second most important conservatism in the present study in Table 19. This sets schedule of minimum wall weights per unit area shown above the reference weight (1.53 psf) of 1.12 for minimum weight penalties treatment segments 1 and 2, 0.22 psf for treatment segments 3 and 4, and study air 0.04 psf for segment 5." For the disk loading of the previous and plane (SHP/D = 37.1) a 12.8 ft diameter propeller would be required,

3-33 ORIGINAL

pAGE IS

QUALM

pOOR

OF

100 - 10 - 30,000 FT, F = 3860 LB, Vt =800 FT/SEC n 8 BLADES -80 0 8 LLU I z 60 6 I- a.

-J - 2-0 - 562o _ z P: 40 w 4 -j SYMBOL MCR 0 3 0 0.0 Lii ~1 ci 0.70 20 30 40 50 60 SHP/D , HP/SQ FT Figure 39. Total Acoustic Treatment Length vs. Disk Loading 3-34 - - - --- ___ BLADES, Vt = 800 FT/SEC F = 3860 LB/ENGINE SYMBOL n K2

" K3 [

K4 AL

J e3 "KS I CL u S z ,-1 LU

020 30 40 50 60

HP/FT SHPID ORIGINAL PAGE IS

POOR ejFJALIT l

OF

Figure, 40. Seginent Length/Diameter vs, Disk Loading 3-35 SEGMENT TOTAL WALL WEIGHT UNIT AREA _PER VS SHP/D VT = 800 FT/SEC MCR - 0.80 F 3860 LB/ENGINE n Uj- 8 BLADES SYMBOL K1 CD R: 4

K2 A

: 31 K4 KS I-- AVG AA 1 WREF1=1.53 PSF - AVG AWT-AT .IV 20 30 50 60 . HP/FT SHP/D Figure 41.

Segment Total Wall Weight vs.

Disk Loading SYMBOL K1 0 I~l,.CIK2 LI wK13

0]

K4 z K5 u 20 30 40 50 , HP/FT SHP/D 2 2 Figure 42. SHP/D , HF/FT vs, Disk Loading 3-36 about 0.8D. In this case the first two the relative tip clearance would be be 3.6D = 46.08 ft, covering a treated area of segment lengths would area 2834.5 sq ft. Applying the minimum wall weight penalty per unit penalty of 3174.6 lb, for segments 1 and 2. Segments (1.12 psf) yields a cover 1.2D = 15.36 ft and 944.83 sq ft. Applying the mandatory 3 and 4 penalty yields a minimum weight increment of 207.9 lb.

0.22 psf unit area segment length is 1.24D = 15.87 ft covering 976.32 sq ft, and The fifth weight penalty adding 39.1 lb requiring a mandatory 0.04 psf unit area the minimum total weight penalty would additional weight. Altogether, for the selected example 12.8 ft propeller diameter with a be 3421.6 lb penalty for the outer relative tip clearance of 0.8. The minimum weight lb.

3 segments alone is 247 described above has a tendency to The minimum weight penalty procedure SPL reductions and diminish the weight reduction benefits of external the minimum unit area higher blade passage frequencies. Without considering weight penalties for the lowest cruise Mach numbers weight constraints, the compared to the data of and higher disk loadings could be further reduced disk Figures 24 to 29. It is thus possible that the optimum propeller on the basis of the loading could be even higher than would be selected 24 to 29. The constraint procedure does not affect the data of Figures data at low SHP/D . This is so, because acoustic treatment weight penalty per unit area for each segment is well above the minimum the minimum weight values given in Table 19.

allowances imposed by the In order to reduce'the weight penalty for each segment, it is necessary requirement for minimum unit area weights wall transmission loss predictions from double to experimentally verify the goal, since theory. This would appear to be an urgent technology development gained from increases in blade it would allow further advantage to be field SPL. Figure h3 passage frequency and/or reductions of external near weight penalties plotted against shows a correlation of acoustic treatment = and 0.80 where frequency. This data shows that at M 0.75 blade passage cr nearly into a SPL data are the same, all of the data collapse the external PAGE IS 3-37 ORIGINAL

OF POOR QUALITYi

18,000 16,000 14,000 -J 12,000 F 3860 LB/ENGINE n 2 SYMBOL B Moo -VT Boo

a 8 0.80

L- 10,000 - 1 "0.75

0 0.o0

V 10 0.80 D0.70 r, V ~0.75 LU z 6.000 z 4,000 -'ORIGINAL RECAT DESIGN FUSELAGE 2,000 RING FREQUENCY fr = Hz,.

100 140 ISO 220 260 300 340 BLADE PASSAGE FREQUENCY, BPF (Hz) Figure 43. Acoustic Treatment Weight Increment vs. Blade Passage Frequency 3-38 single curve when plotted against blade passage frequency, for both 8 and 10 ]blades.

3.7 POSSIBILITIES FOR FUTURE TREATMENT WEIGHT REDUCTIONS The trends of Figure 43 invite the development of propfans with more blades, higher tip speeds, and higher diskloadings, in order to increase blade passage frequency. The minimum constrained weight at M = 0.80 for cr 10 blades at SHP/D = 50, is about 3650 lb which is within 700 lb for the previous RECAT weight penalty, despite the higher exterior noise levels, and more conservative design philosophy employed in the current study. It is believed that the data of Figure 43 might be reducible by 40 percent by eliminating some of the current conservatisms with respect to the large amount of treated surface area and the mandatory minimum weight per unit area stipulated in Table 17 for the various treatment segments. These conservatisms have been injected in this study to offset uncertainties concerning the exterior SPL distribution, and the validity of the simplified double wall transmission loss theory.

With regards to the transmission loss theory, notice in Figure 43 that the structural ring frequency is 288 Hz for a 19.58 ft diameter aluminum fuselage. This value is higher by factors of 111 percent to 191 percent of the typical range of-propfan blade passage frequencies .(150 to 260 Hz) for 8 and 10 blades at 800 ft/sec.

Figure 44 is'a reproduction of Figure 74 of Ref. 4. This shows the modal density parameter for single wall cylin drical shell vibration modes which are "acoustically fast" (efficient noise radiators), as a function of the ratio of excitation frequency to ring frequency. Lockheed is working on the development of data and a theory for counterpart to this curve for the proposed double limp wall damped treatment shown schematically in Figure 22. When such a curve is available, it will be possible to correct for loss of transmission loss at frequencies near the ring frequency according to ANTL = A + B log (YMDP) 1 0 where A and B are empirical constants to be determined from transmission loss tests on Lockheed's double wall concept, and, YMDP" is the modal 3-39 3.0 BLADE 2nd HARMONIC PASSAGE OF BPF FREQUENCY 1.0 LU I- LU a- I .1 -J 0.1 1.0 10 FREQUENCY RING FREQUENCY Acoustically Fast Modes Figure 1414 Modal Density. for

3-h o

density parameter for the particular double wall construction. It is a goal of the double wall technology development to minimize the response of these acoustically fast modes to the external excitation.

3.8 CONCLUDING REMARKS CONCERNING ACOUSTIC TREATMENT WEIGHT PENALTIES * New exterior near field SPL and propeller performance data (Ref. 3) have been evaluated with respect to acoustic treatment weight penalties. The new data are estimated to increase the weight penalty by 3400 lb to 6500 lb, compared to the 3100 lb requirement estimated for the original RECAT study (Ref. 4) at the same disk loading SHP/D = 37.1. For this disk loading a 12.8 ft diameter propeller is required for a net thrust of 3860 lb, at a cruise point design Mach number of 0.8 at 30,000 ft. Parametric studies have been conducted of the effects, of disk loading, upon acoustic treatment weight, cruise Mach number, blade count and thrust level.

* The parametric studies include a more conservative prediction method ology which is partly responsible for the higher weight penalties.

The more conservative approach has been employed to reduce the risk associated with technology uncertainties. In this sense, the attain ment of the interior noise goals with the current weight estimates in this study have a higher probability of achievement through development than the estimates in the previous RECAT study.

* It appears, by increasing the blade passage frequency, that the weight penalties could still be reduced to about 3600 lb, even with the currently more conservative methodology. The range of weight penalties contained in these studies is apparently small enough to make turboprop aircraft remain attractive, based on the weight versus DOC sensitivity data of Ref. 4.

* Achievement of certain goals of technology development, outlined penalty reductions, of the herein, could provide further weight order of 40 percent, through the elimination of conservatisms which are imposed on this methodology in order to offset technological uncertainty. In particular, it is believed that the total treat ment area assumptions used herein are definitely conservative, often requiring three times the 1568 sq ft treated in the original RECAT.

A large part of the treatment area increase comes from the treatment of the full cabin circumference in the current study.

ORIGNA PAGE Is

OF POOR QUALF

3-41

SECTION 4 COST/BENEFIT COMPARISONS At the conclusion of the previous EECAT study both the turbofan and propfan powered aircraft had been designed using 1985 levels of technology and the same payload/range requirements and mission constraints. The base line aircraft established during the previous study were competitive in terms of cruise speed, cruise altitude, block time, and passenger comfort. At the design range of 1500 nautical miles for a Mach 0.80 mission, comparison of the fuel and cost to operate these baseline aircraft showed an advantage of the propfan over the turbofan of 17.8 percent less fuel and 8.2 percent DOC savings at a 60/gal. fuel cost. Comparison of these baseline aircraft at a range of 475 nautical miles with a 58 percent load factor (L.F.) shows an advantage of the propfan over the turbofan of 20.4 percent less fuel and 8.5 percent savings for 0/gal. fuel. These comparisons are shown in Figure 45.

For this study, the competitive baseline design concept was retained so that direct comparison between the turbofan and propfan propulsion could be determined. The original baseline propfan powered aircraft was revised to reflect the latst propfan performance and acoustic data supplied by Hamilton Standard as a result of their propfan wind tunnel test program. The effect of the new propfan data, and a revised Lockheed analysis, is added weight required in the fuselage to accommodate the increase in propfan acoustic noise level. The comparisons of the revised turboprop baseline with the turbofan baseline are shown in Figure 46.

For each design/mission change investigated for the turboprop aircraft, a similar change was incorporated into the turbofan aircraft with each design ORIGINAL PAGE IS 4-1

OF POOR QUALITY

3.0 -8.5% -8.2%

2 I

$2.0 1.0 Z z < z z 0 u. .

o o U. 0 m ao 0 c 0 1500 NM. 475 NM (100% L.F) (58% LF.)

-17.8% I -J 16 -- 20.4% z o 0 U.0 cc I- 0 1500 NM. 475 NM.

(100% L.F) (58% L.F) Figure 45. Fuel and Cost Comparison for Baseline Aircraft of Previous Recat Study 4-2 3.0 -8.1% -7.8%1

--J

1 1.0 z o <c z <c z o S 0 0-I 475 NM.

1500 NM.

L.F.)

(58% L.F.)

(100% -17.6% j -19.6% -j 16

IS

PAGE

ORIGINAL

OF POOR QUALITY

- 820 0

Ca

o a 475 NM.

1500 NM.

L.F.)

(58% (100% L.F.)

Aircraft Turboprop Baseline for Revised Comparison and Cost 46. Fuel Figure 4-3 optimized for the desired design/mission characteristic.

All subsequent comparisons of the baseline aircraft are the origoinal turbofan baseline and the revised turboprop baseline (incorporating revised propfan data).

4.1 PERFORMANCE COMPARISON The major differences in fuel and operating costs between the turbofan and turboprop aircraft in this study are caused by differences in engine specific fuel consumption and aircraft weight.

The most significant dif ference in performance is in the propulsion system and its fuel consumption characteristics at the cruise condition for the 1500 nautical mile design mission.

Figure 47 indicates the improvement in average cruise SFC obtained with the turboshaft engine for the design/mission conditions investigated.

The turboshaft propulsion offers a 19 percent decrease in average cruise fuel consumption at the 1500 nautical mile, -Mach 0.8 design mission and addition ally offers another 3 percent decrease for the 1500 nautical mile, Mach 0.75 design mission.

As indicated in the previous study, the empty weight of the turboprop exceeds that of the turbofan.

Figure 48 depicts the differences in aircraft empty weight between the turbofan and turboprop aircraft 1985 IOC and the IOC designs.

The empty weight ofthe turboprop baseline design is approx imately 6.4 percent greater than the turbofanbaseline with the major dif ferences being in the wing and propulsion system weight and the amount of acoustic treatment required.

For the 1990 IOC aircraft, the turboprop empty weight exceeds the turbofan empty weight by approximately 3.3 percent due to decreases in the propulsion system weight (which is reflected in wing weight) and the amount of acoustic treatment required due to the reduction in induced sound level with the smaller diameter propfan.

The largest single weight increment between the turboprop and turbofan aircraft is the amount of acoust treatment required to maintain the cabin interior SPL at 90 dB with the propfan.

The amount of acoustic treatment required for each of the turboprop design/mission conditions is as follows: Revised Baseline 2000 N.Mi.

0.75M PD 370-22 STS 4E7

4-4

-18% -19% -19% -21.8% 0.8

-18%

j j

j

0.6 LL S w W L 0.4 2Z U N Sz z 2 2 N v CO . j Nl mi o N 0 0 coCF <w N I E0 .2 z I z Iz 0.22 .-.

4 z < 2 4 < z a4 IL 0 IL 0 L 0 0 LLL 0 m mm a,. m m.

9 > 0 cc 0 cc CC 0 ~ w 0 U C i af C.

sa.i - C Ci F r . Efe o

47.' Effect of Design/Mission Characteristics on Aircraft Cruise SFC

Figure +5.1% +6.4%

j +3.8% -0% +3.3%

,7- 1 1 j

- I ~100 U.

N aIM 2 z zj S Uj N Lu CO CV co NL LL .< LUL. LL < . L <1 z < z 4 <z << S0 0 0 0 - - . - - . -, 0,, 0-0 Fu 48. E. ra Figure F8. Effect of Design/Mission Characteristics on Aircraft Empty Weight 4.2 ECONOMIC COMPARISON For the economic comparison of the turbofan and turboprop aircraft, all point designs were compared using the 1500 nautical mile design mission with a 100% L.F. as well as a "typical" mission of 475 nautical miles with a 58 per cent L.F.

DOC values for Ot/gal.

fuel cost were calculated for each mission.

Figures 49 and 50 present the results of the effects of design/mission charac teristics on turboprop DOC savings at the two sthge lengths for 0W/gal. fuel cost. Comparison of the 1985 IOC revised baseline propfan and turbofan aircraft indicates a 7.8 percent DOC advantage, at 60C/gal.

fuel cost, for the turboprop at the design mission of 1500 nautical miles and Mach 0.8. An additional advantage in turboprop DOC of 2.2 percent is attained by reducing the cruise speed to Mach 0.75, due to the greater advantage in fuel consump tion characteristics of the turboshaft engine at reduced speed.

For the 1990 IOC aircraft, the 1990 propfan design shows an advantage in DOC, at 0W/gal. fuel cost, of 7.8 percent over the 1990 turbofan design at the 1500 nautical mile, Mach 0.8 mission.

Incorporation of the alternate turboshaft engine, PD 370-22, results in an 10.1 percent advantage in turboprop DOC, at 600/gal. fuel cost, over the baseline turbofan at the 1500 nautical mile, Mach 0.8 mission. This addi tional decrease in DOC is due to a significant decrease in installed pro pulsion system weight for this engine (approximately 40 percent) and the resultant effect on aircraft weight.

4.3 MISSION FUEL COMPARISON Figures 51 and 52 present the results of the effects of design/mission characteristics on mission fuel requirements, at both the design and "typical" (475 N.Mi.) range, for the propfan and turbofan powered aircraft. The advantage in mission'fuel of the baseline propfan over the baseline turbofan is 17.6 percent. Comparison of the mission fuel requirements indicate that the largest percentage of fuel saving (21 percent) is attained by reducing the cruise speed of the turboprop to Mach 0.75.

Incorporation of the 1990 engine technology in both the propfan and turbofan powered aircraft results

SOIGIOAL PAGL

IS

4-7

OF

POOR QVALFV

PAX /1500 NM.

100% L.F.

-10.1% -10.0% -8.0% -7.8% 2.0

7X

-

.4 Co<

"<

<; z U,

LU z

L1

, 2 La 20I C co (4 Co CLC 14CI 0; a.

z I 2 0 2 o00 F--F 50NM iso DO1% onAic-f hrceitc einMiso Fiu u, feto

1500 WM. Mission

Aircraft DC on Characteristics Effect of Design/Mission Figure 4-9.

200 PAXJ475 NM.

58% L.F.

cI0 2.5 -10.2% -10.3% -8.1% -8.3% -8.1% 2.0 $1.5 \ u 0" w ~0 5 n in UU w z m N oz z 2!: z 0 M U 4 KM U) a CL _. mi_ CL.

0 z 0 2 0 Z 0 z 0 z

m

o to w m °m m

Fg - - o Dsi-- C n A Figure 50. Effect of Design/Mission Characteristics on Aircraft DOC - h75 NM. Mission PAX/1500 NM.

35 - 200 100% L.F.

-17.6% , -17.8/ aoJ ~~-21.0%-18%1.%

3 _

L S w

a

x 10= U. .. j4 LO < 4 <

5uZ 2 o 0 0

0 N

o 0 0 0 0 0 0

cawmm In m CO Figure 51. Effect of Design/Mission Characteristics on Block Fuel - 1500 NM. Mission /475 NM, 200 PAX -18.5% 58% L. F.

-19.8% -22.9% -19.6% 19.1% --- -J U J LU Li, iI?

.

a.CNC ' - I- '.- .

- Li.LLL C2 - a. 0.L-a 0.

I 0 0 00I 0 0 0 o o Mission NM. -475 Block Fuel on Characteristics of Design/Mission Effect 52.

Figure in a 17.1 percent advantage in mission fuel for the turboprop since the improvement in fuel consumption characteristics is similar for both the. turbo shaft and turbofan engines. This advantage in fuel savings is similar to that shown for the baseline aircraft, however, the fuel savings available by incorporating the 1990 technology engine into the 1985 IOC turboprop aircraft is an additional 10.8 percent.

Incorporation of the alternate turboshaft engine, PD 370-22, into the 1985 IOC aircraft results in a additional small savings in mission fuel of 0.2 percent due to the decrease in installed propulsion system approximately weight.

Figure 53 depicts the potential fuel savings available for the turboprop aircraft,-with the design/mission characteristics, investigated in this study, oveir the baseline 1985 IOC turbofan aircraft. A potential of approximately 32 percent fuel savings and approximately 17 percent DOC savings, shown in Figure 54, are available over the 1985 IOC turbofan aircraft by utilization of a 1990 teclnology turboshaft engine with the 8 bladed propfan flying at a cruise speed of Mach 0.75.

4-12 200 PAX 1500 NM. MISSION -17.6% -17.8% -23.0% -26.5% -31.9%

30 -_V

Co -J L Co D w € '4 j 15 1 r... P, .

=.u inCo6C n I 1 I to .

LL m1 0Z Co u. <Z cc =- =H- I= i-u 10 000 00mL LnL ~ Ln LL versus 1985 IOC Turbofan Aircraft Figure 53. Fuel Savings oftPropfan h-i 3 200 PAX 1500 NM. MISSION -7.8% -10.1% -100/ -14.7% -16.9%

1. L) l -e

C; C cmz N d 01. <~ 8l < c < cc I z 0I F- cc CC < c - 0 oX 0 - o C. CL C.. C <<

co, 8" ,o w m= 0=

Figure 54. DOC Savings of Propfan versus 1985 IOC Turbofan Aircraft 4-14

SECTION CONCLUSIONS AND RECOMMENDATIONS The results obtained from this study show that the advanced propfan Powered transport aircraft, with the 8 bladed propfan, is a viable alternative to the turbofan powered aircraft and offers significant savings in fuel and operating costs without compromising passenger comfort. The advantage in fuel and operating costs of the propfan over the turbofan continues to be significant for the 1990 IOC time frame. Additionally, further fuel and the reduced cost advantages are shown for the propfan aircraft at operating cruise speed consistent with current operator experience for the design mission range.

Propfan data supplied by Hamilton Standard, as a result of their ongoing propfan test program, shows that the performance goals are attainable.

The noise generated by the propfan continues to be somewhat of a problem in that weight penalties required to damp the noise transmission into the aircraft cabin detract somewhat from fuel and operating cost savings. Estimates of the performance and acoustic characteristics of a 10 bladed propfan indicate potential for reducing the weight penalty required for the propfan aircraft.

To realize the potential fuel and operating cost advantages with the advanced turboprop aircraft, as identified during this study, the following research and technology items should be accomplished.

5.1 PROPFAN DESIGN Data from the 8 bladed Hamilton Standard propfan, as a result of wind tunnel tests conducted on a 2 foot diameter model; indicate that propfan efficiency goals can be attained or exceeded.

Acoustic measurements taken in other testing indicate that induced sound pressure levels are higher 5-1 conducted at different tip speeds and Mach numbers than estimated. Testing con 8 bladed propfan and a projection of this data to a 10 bladed for the in acoustic noise, figuration indicates a significant -potential reduction while maintaining efficiency goals.

the performance, acoustics, Further design studies are required to assess of 10 bladed and 12 bladed economics, and mechanical design characteristics be supplemented with propfan configurations. These design studies should and testing to provide a viable, demonstrated propfan component development design for utilization in a 1990 IOC aircraft.

5.2 AIRCRAFT ACOUSTIC TREATMENT advanced turboprop aircraft, one of the major design considera For the interior.

is the reduction of excessive noise transmitted to the cabin tions goal of 80 percent fan efficiency and a cruise Mach number At the design fan with a 0.80 at 30,000 feet, the tip noise generated by an 8 bladed of speed of 800 fps is approximately 138 dB at the fuselage wall.

tip dB the cabin interior noise levels at a maximum of 90 Maintaining in acoustic transmission of some 48 dB. Conventional requires a reduction in this study) requires that the burden of wing mounted engines (as utilized cabin walls.

noise reduction be obtained by structural design of the the cabin walls as well as skin The mechanism of noise transmission through by blade tip passage, is not well understood. The initiated in the fuselage with wing mounted propfans will design of an advanced turboprop aircraft through the fuse probably be paced by the noise transmission losses required cabin.

walls and the acceptable sound pressure level inside the lage of damping the noise through The approach taken during this study is one con using limp wall mass treatment. Using a double wall the cabin walls air space alleviates the cabin noise struction with the maximum possible increased but results in increased fuselage diameter along with attenuation to be the use of aircraft weight. The best potential solution appears as much structural damping in the affected double wall construction providing areas as possible.

PAGE IS

ORIGINAL

QUALITY

POOR

OR 5-2 Design studies, in conjunction with development testing should be conducted to provide the necessary information regarding the mechanism of noise transmission and damping and structure excitation when utilizing the propfan configuration. Also, further studies should be conducted, and followed by development testing, to examine fuselage wall structural and damping concepts, optimized for reduction in noise levels, weight, produci bility, maintainability, and economics.

5.3 AIRCRAFT CONFIGURATIONS Accomplishment of this study, and the previous RECAT study, utilized a 4 engine (conventional wing mounted engines) turboprop aircraft. Locating the propfans away from the cabin area would greatly reduce the amount of noise transmitted to the cabin. To accomplishithis a configuration study, including a 3 engine design, could be conducted to investigate alternate engine/aircraft installation configurations.

Another purpose of a 3 engine configuration would be to enhance utilization of the propfan concept for a complete range of aircraft sizes.

The study approach would be to evaluate a series of fin positions for the third engine considering fan diameter, tip to fuselage clearance, weight and balance effects, stability and control, potential acoustic fatigue and noise transmitted to the cabin. These results could then be extended to pylon or stabilizer positions and an evaluation could be made for a variety of 2, 3, 4 wing and tail engined configurations.

5.4 ADVANCED TECHNOLOGY ENGINES Significant design and technology studies, along with component develop ment testing, are currently in process for advanced technology, energy efficient turbofan engines. Similar studies and component tests should be conducted for the turboshaft engine utilizing those technology areas, where applicable, which are being developed for the turbofan. The test program now in process on the propfan configuration should be supplemented with a similar program to develop an advanced technology, energy efficient, economically viable turboshaft engine.

ORIGINAL- PAGE IS

OF

pOOR

SQUA

fl, 5-3 - REFERENCES 1.

Schlichting, H., "Boundary Layer Theory," 4th Ed., pp 536-543, McGraw Hill, 1960 2. Peterson, A., and Bruel, P.V., "Instruments for Noise Measurements" Chapter 16 Handbook of Noise Control, McGraw Hill, 3. Letter, B. Gatzen (HS) to L.K. Carson (Lockheed), June 4. Hopkins, J.P., and Wharton, H.E., "Study of the Cost/Benefit Trade Offs for Reducing the Energy Consumption of the Commercial Air Transportation System," Lockheed California Company LR 27769-2, NASA CR 137926, August, 1976 5. Anon., "Preliminary Prop Fan Noise Estimation," Hamilton Standard, SP09A76 (Unpublished) March, 1976 6. Hanson, D.B., "Near Field Noise of High Tip Speed Propellers in Forward Flight," AIAA Paper 76-565 presented at 3rd AIAA Aero-acoustics Conference, Palo Alto July 20 to 22, 1976 7. Ffowcs-Williams, J.E., and Hawkings, D.L., "Sound Generated by Turbulence and Surfaces in Arbitrary Motion," Phil. Trans Roy Soc. London, Series A, Vol. 264, 1969 8. Hubbard, H.H., and Regier, A.A., "Free Space Oscillating Pressures Near the Tips of Rotating Propellers," NACA TN 1870, 1949 9. Kubjun, M.C., and Vogeley, A.W., "Measurement of Free-Space Oscillating Pressures Near Propellers at Flight Mach Numbers to 0.72," NACA Report 1377, 1958 10. Farassat, F., "Theory of Noise Generation From Wing Helicopter Blades With an Application to Helicopter Rotors," NASA TR-R-451, December,

I

PAGE

ORIGNAL

QUAIX'

pOOR

OF.

APFENDIX A HAMILTON STANDARD PROPFAN DATA FOR UTILIZATION IN RECAT FOLLOW-ON STUDIES A-i

HAMILTON STANDARD

Windsor Locks, Connecticut 06096 Please address answer to Mail Stop No. 1A-3-1 June 13, 1977 Lockheed-California Company Lockheed Aircraft Corporation 2555 North Hollywood Way - Box 551 Burbank, California 91503 Attention: Kit Carson - Bldg. 63, Plant A-i, Dept. 75-21 Subject: Prop-Fans for RECAT II Reference: HS/LCC technical discussions at LCC on 5-3-77 Dear Kit: An aero/acoustic parametric study has been conducted as was agreed to in the reference discussions. Curves presenting the results are enclosed. The performance and noise data are updated by the latest test results and reflect future Prop-Fan designs. The study covers 8 and lT-aTes, 600 to 800 feet per second. tip speed, 0.7 to 0.8 Mach number, and a range of efficiency (SHP/D ).

The first three curves are the generalized efficiency maps for eight blades at 0.8, 0.75, and 0.7 Mn and 30,000 feet. The fourth thru sixth curves represent Prop-Fans sized for 3860 pounds, Tnet + Tjet, for each Mn. These were generated the efficiency maps for each respective Mn. The seventh curve provides using the engine power information for each Mn. Curves eight thru thirteen provide the same information for ten blades.

Curves fourteen and fifteen provide the parametric overall SPL which complement curves one thru three and eight thru ten, respectively. Curve sixteen shows the spectrum shapes at 600, 700, and 800 feet per second. Although labeled for 0.7 Mn, it should be considered representative for the entire Mn range under consideration here. Curves seventeen and eighteen show the directivities with varying tip clearance to the fuselage for 8 and 10 blades. Again these curves can be used over the Mn range.

Both the performance and noise curves are generalized based on your need to resize for a lower thrust at 0.75 and 0.7 Mn. Using the generalized curves, LCC can accomplish the same results as shown on curves four thru seven and eleven thru thirteen for any thrust level.

HS has selected an increased number of blades in addition to the 8 LCC requested based on acoustic considerations. Increasing the number of blades while keeping total solidity about the same will lower the overall SPL (hence, the level of first blade passing frequency also reduces), will increase the frequency at which Division of UNITED 0 TECHNOLOGIEST.

Telephone (203) 623-1621 - Telex 9-9288 TWX 710-420-0584 HAMILTON STANDARD Lockheed-California Company -2- June 13, 1977 the tones occur, and will improve efficiency slightly.

It is estimated that a 10 blade Prop-Fan will have a rotor weight which is 10% higher than the 8 blade weight provided by data package SPO5A76 dated 2-27-76. Acquisition and maintenance costs for a 10 blade Prop-Fan will also increase slightly over the 8 bladed rotor.

While not much change is expected in acquisition cost of the baseline 8 blade Prop-Fan, the results of the recently completed NASA funded maintenance study indicate that the maintenance cost information supplied for RECAT I is conservative.

Since LCC will use the enclosed data to optimize the propulsion system, it would be best to estimate the costs (both acquisition and maintenance) after the configuration matrix has been narrowed somewhat.

Please provide the.selected Prop-Fan diameter, horsepower, tip speed, and number of blades when available for this task.

If any questions come up, please contact me.

Very truly yours, HAMILTON STANDARD Division of United Technologies Corp.

New Produ evelopment BSG/csd Enclosures cc: Messrs. B. Miller (NASA-Lewis) L. Williams (NASA-Ames) J. Dupak (LCC) bcc: Messrs. C. Rohrbach (2) F. Metzger W. Adamson R.

Levintan/R. Bussolari R. Baum (Los Angeles) File 2.3.3 IS PAGE ORIGsNAL QUALITY POOR OF A-3 t~ TT ~ ii yr T H f f .

rrf flr wn.

rr~ ,--,l-thii-r-m-T m fTrm-r Tn ff r.

..

j I II p2: !7 ; II .= ..

.7 ....

VIL

, L

,, . f" , , rI ,'.t;' I I I .

hI1

1~ ...

hir t1t pp g i Ii . .

Sii

a]I j[1

i:Jt l

I LIl

Iii

Ifl i14 T 7 i .

l

a )I

1b r ::77w0 TiI

1-

PAG.,

lhl

I i

Y

IT 1

A L

Q U

O R

P O

u

GRAPH PAPER DIETZGEN CORPORATION NO. 340-2D OILTZOEN 20 X 2G1 PER INCH #4 N

I.. . ... W . ,~

I :T7 V. I

t I i ii l i

'

it i

!,Ji

IW, tillS

fl I'M.7

Q Liti Jill'l

n r r

.I.r

I.IrIiii;I

~~~~~~~~~ I

I nr q - - it 1 1..

iII 1I

p , . ..7. . .. .

-: --- - .- - - - - rt----- -- - - - -=_=--=z.. -r----- =T.ZZ-= .

- - -.... - . ..... -----

.. _ F-_-.- -- - - .----

....., ww ....... 2 ...

-- -- -- - - - .

.a

- -- -.... .- .... t ---- --- -- __---__-_-± ,

f:-t---- :-::: --- :ZzE--2

-rt.fr --- -

-~ -- z_

.

_- --- ... ---- - - .

...... ___-__--____ - -_- - - - -- -...

... . . . ... . . .

. .

----- 4------" - -- - - .- --...

___-_.---

-- -- I -- ----- _--- __ . __

'--

----- '-_--_ :---__ _ --__ :.

t

....-- -,. -...

........- - 2 .:"Z . . .. Z ", - -:: .. .. .. .. .. .. . . . . .. .. . :' . .. .. . .. .

.... . . . . .. . . . . . . ....

. .. .~ 4 .. . . . .

:,z-=-:-:- -. - -: - :_-') ?: 1' -' ...

. : ::::::::::::::::::::::::::: .

i- : f -4.... . . . . . - -... .. .... . . . .. .

.. . .. . . .. . . . . . . . . .

__ . .

_ _ _......

_ = = _-- _ _ _ _ :, ,-r ... . . ..... .___.. .__ := .........- ..

pti :K ~%-a -- ::7__________-_

-111 NVO. zly( r?")

z Cc IFE69 T z k6 * 9 Ik

i _ I_ I -- 7 I " i _ __F__17

-. . i . 7i 7 77 1 TV

t___---- __ - __-___ __ ___________ _______s- ------- . -- -- + _ ....

.. ...

..

-.

- . ...

I _ __ _ _ _ ........ .. .... _ _ , . . I - -.. ........ --- --- -- ------ - : .. . .. II .. -lV l S.. .. .. . . _ J f--rl -- I-nP% Zo : t -- --- o ~ ~ _________.- - 20 _-_ - .i--2___2_ -__-I_ __ : f... . .. ... . . i ... . . . .I { . .. . . . . .. . - - . . .I. ..

o - .. - - ... - .. . . .. ...

- ------ - - ------ - ,....... L _ ----- -- - .......-.. -- - ----- - - ~ -~~ ---------- I7

Jj __A-7

• .C.O tt t, u 2t2, .

.... .. . : .....

LII I'~ R~ I

-#-:r.:: tt7A-rr"+,p:

wt:rcr:fr:7vcy

......... ~ ~ ~ ~ ~ ~ ~~.. .. .. ... '::"... .... ... "

....... .....

. . . . . .

.. t- -' -- - ---- --- ----- - - - - - - - - - -: . ... - - - - - - - - _ :

t- _u "'4_ _--_

.. . .. ... . - ------

t ---- C --: ----

-. ........ ----- . ..

--- t ---- _ If _ _ _ _ _ 2b_ -/ - ~ I - -- t-:7 -::-- ---- 7 ----- A-8

C,..qe.. 4*'.iz C r)

.

t.

/ S i I '..".....".".."...................................... .............. .......

_ _-_.- . ... .. ,: . : .

-- -I . . . . .. - 1 _ - - -- - : - - - -:-- - - - 7_ .--- 7= .= - - --- . ,

. .. -,--

-- -_--- - --- -- - -- -= :

n 11 .t .... N ~~~~ ...) .. . 1 - I

-4 .. ...... ~~~~~~ - ...

-_ -_-- -. : ----- - --- - - --- __- - --- .n n,-------------

. ... .... _ __... t.... . ... _ _

... . .. . . . .

_.--- -- -L---- . ------.... .--- : .

_ - -_- - - - ' .. . . ... . L- - - - - - : . . . . . . . . . : - ' : t:: -- . .

-- - _ __..... .. .- . .. ... .. . . .. . . . _ . . ___--_ .

.. . . ... ..... .

f.. ---. ----- - ... . H-% -- - -- - . .

. .. ._. .

--- -. -;:__. . :,t: _ -- . . .f . . , - . . . tO_ . ,.,_ - -_---i-_.,-.

' -- - .. . -. . . . + . . . .. _ .. . 2- . .. -- -.

. --- -f - : = : :Z L''T''' 4::= - -

. .. . ... . . __4 - -- =-.. . . .~ .. .. .. .- ---

-' _ _--_, U..a. - '-_----: .-- -- t------ .:--.- - .. 4.. -- ... ...

- :=_ r / --- ___. I - -.

5, ... .. ..- . x--

.. . r. ..- .. . . - - _ --... . - :~~1~ ... : - -" _------- - - _. . . : - _ _ ...- -- - -- . . .

I-- S--''------- - : -- - . . ... . . . . . . . - tt___._ _. _ __ ___ .. .--.. . . .. . .. . . -- .. . . . . ..- . . .-...

- -i...~ -.. - ,____

____-_ _ . .. .

... . ... .. . . ..-..

. .... .... . ...--.-. ..---- -- - __

--......

-==- . .a-= ,--T___ - i-;--=-- --- -- --- _ .. .... ....

L-a€= -= / -- =: - -. .... :..... . . " -=

.. . . . . -- A - -._ _

S --

W- 7W'"

L

F:3,T:

,: ,, . T , '- , -

.'

- - __1___,,_ _I__ __:'n I.-

I

~ il , ,.. 1 , ,.,,,'......,,.

IHs .. .. ;ili r''k

'I , ...

.i: .

, i '"i"K..

LI ', H; " ,1 "'

i:[" H 'I

-, ,, .I 1 1 ; ! '' I " 9' , ; I .

'.I,,' , :244' k.

' CC , 'Al

T, ",i

,7. 00 7.

Ii Jill :" H :

-,K,7 f- (

'i.tH''''IiI

,, ....

, iIi

I15?! i .1 I'i'

I N 'I'IN'

!I

ji":ji

i" I I ' '

"

' .

I~h H

...

!

..

I..1

4 71 i , , .. .. , ,'l, . ' " ' . . .

t'uh~v' 4 111

I ti

"lGINAL PAGE

IS

@P DOOR~ QUAITY

OWITGEN GRAPH PAPLR DIETZGCN CORPORATION MCO. 34Q-203 20 X ZO F'LR INCH Pr, A , 0i-f 1: it ri fijf i it, .. h'I UL; ,~ i:i 'J t

I t- il li

i IiI I,,,I[f ' T rI 1Il H.......l

Ii i- j

I ....

illl ::7 ;,;TI ITI 11,7

k f LVIi

ii~jjfTjTiF

Ji jit[i

h :I..*

''LwE'> (Wi t.1 b

!2....

.L.

'

1. i :

!

I

1! .

i INI tit l

t mi:i

L I

f[

bt4;1U}

fb

INht I.

Hit M I ,.i "b !i F I I :;Ii

l:1

DIV !!

t!;

i

'

iur I

i !N l

b i

.'

IL

1 ,I

I:

*

:Itm

lit, <...FI

,, ithl I : ki ., :

[/ !It

ji: " , F, I: :' : h 1''1 I.

• C) : ,Ii' :!i.

;::~~~i

a,!

7 .

,H

.RGI ... . .. ....

Fl

7A

F

4O1 "."

GT is-

pj!

I

WI

OP~~~ DIETZSEN CORPORATION DIETZC3EN G1RAPH PAPER NI. 340-20 MAFE UIC A 20 X 20 I'rR R INCH -'

I: l L

LnbL

H,!II'r'. Pr,, ISA .K

T S ocxx

E.i AT ::.

IdtA

777ii 20 ,,j

: ,. :.

t~3t it:

liii; BLt

ill

f~lh"Li

11!'

t!'0

Fil I , iij i.,t ....

I

' lw '

'u '

.4-

11-VLII

it i i ; ' i I I ;: . 1 I I ~;

id!

i,,i

7= fV L ,

lI

J

SU

RE.CAT/PD 370-22! EN N. OP-P41 SL- c.-.Cn

:LC-

I:-::- ,-: -. r i

-- '' I '-I '- :=--L = -

......... T EFHCIN i/..DIAMET- . .T1PCD_.. ___

....... I ~~~~~ ':..F ..

- ~~~~ . .............

.. ... -. ---- --.

... ......

1.. . . . . . . . 4 . EFE .... 4 TO.

-- -- .... - -

+

- - -- - - *----- ---- -- - "----- - I . . .. . " .....-- ....- ' "h - " --

- - t- - ------- r . . - -------

-- --- - --- _ -_-* . _ _--___._=- _ liio J 1&:..3 ... == ___ _ _ -.- _ - &: -- ..... -.

e .,-- ,---,' -- ---- = --F:. .=ay,: .-.- ,).-

-.---------- ' " ------ ------ . --- -- . - -- - . - -- - - -. .. .--- - -

V::.::: ' I :::::: :::'ZZ:: ZZZ- .::: -. : : : :::.. .:--

-- -- --- - :- -- --- -=- :-. ---------- ------ --- 7- 1 ----- -. : : ..... -. -.-.-- -- -- - .. ..-. =J--- -: - _ -- - - --... - -.. --- -....

- - - - - -- .---- - - ---- ----- -- - - '-' ,----.- *--------.-- ----- T : :

L --- = I-: =:: ::- -- = = - --- 1 - - -- L _

_ _- _ :

---- _L - = . . . . : = r .. ._-

- L -- :-= - - = : _- - S1- -14____-_ - -- /"- f / - - :-----,- -- :-- ---------. - - ')== = = :-----= ... . . .. ,- .. ..--- - - - -" -' -" -• -2 : _ . _- - - - . . .

. .

. . . . . . .l . . .. . . . . . . . . . A -.. U.. . . .. . . . .I. . . . . . .. . . .

NO.

340-20 DIETZQE,4 GRAPH PAPER QIETZOEN CORPORATION t 2.0 X 20oER IN20,X20 INCH .. ,PEu.

I ;"L ijH i l l ! 1 " ;:f "'nH

, 'ii . rI j .1 !7 li i ! jI

... ..

T- . 7 11.

h ....

'' < __ I . W*V . . . 4: i 4 r, I I.*'.. __ [1 . l , J J ,l ,

It Ji l

i ll L

. . "

r I -... n ' i.,,, F~ 'r , ,I

H K .:I J.,, 'P flN o. :.

.1IJ 0..,

in

,i I Jill

I;..II I

I V

I' ll I

Jill

Ii I i, ,,[, I I.

I .

, ', | ..Tr , i'

h'l 'jI

I i , a ) j ,t ,. I C

. . . . "

I i I

H .. I4 :><

. . . . . . 'r I.II[ .

S'O .. ii hp

.- ,I

ll P

. . . . .I . I . Il i

7.

1- 1: Itfhlip ''ihdfl "nI ' U i 'li' /;iI 'H i ,',

, ' 1

N!' l l7 :/.;, .il lh ,h , Ulli il ','l ~ ll ~ l i''l't, ... I l,/ , 1, r.

. ic..

r .

A'

I4Iit

j; t , tl, -....

. .. .. .. . ."CALC ,,

LCC EEc4T/yD ,70-22

-P-N.,- P -t

'jf.

. ..... ...

. .. . . ..

-'a-. t , S , - --

-

F--.---:~~~~~~~ = C - ,I.N , I N = ~~~~~ = ., P.------ ------- . ----

.. .....

- .... - - ... - . .. "..

. . . --- .. . . . .. . -- - . --------

-.......-- _- .....

....... .... . .

_ _ _~---

- t-n-----P

n ~

- ------ -------- ... . !. .. .. _ I ... c ..... -. . . . ---- . .

--- -- ----- -- ---- - .. .

---------- - - ------ ------- -- ...... .----- _ - -:..... _----_- S==--L-=:. - .- . - .-- ....... - - ---.....

-....... . ... .. _ .... . . _ - -- -- ._ .4 --------- - -

: ~ ~ ~ ~ - -- _.--------,----

- -- _-- - .. - ---. -.- - -- .+ ...... !::- ..... .. . .- 1..6 ............

-- - - ---- --- - .-----------

.... w-- - Lr ---- -- . . --------

.-- ------- . -

--- - --

---- -

460782 SKEUFF"£L & ESSER CO PAGE. IS 46 0782 S0 X 10 TO THE INCH I X 10 INCHES

.XOLORI

NAL PAGE

IS

OF POOR QUAMT

-!-A

)' 00 I

I I

71 IL2 ltI

I'0, F I I

I:i

[ - .

c , : I -,

I., - * -. -

"- 4.I

E["

U,)V

Ito

4 11w

u. ' . jT. * ;, t i :

K:I 1 -

-F L 'H I If - l-tI-.ell iiI Iu

I -

t I .I

I IIlI-.- IT Ii.a, , '' 12 in l 11 .

I ,-

i

Itlu ii

"__ --- --- -- -1ARmA d :--S-flA&fll=t ,/.4J .,lI - ...........

- - - . -- .'... .,. . . ... .. .

Z Q W rtS------L-t >.---- --

- -- - -- - -

- - -- - --

.- _- ---

- - ------- --- . .

-z--_ri: 22: .. _2: 2 -__.___.___....

_.

- - --

-t - ... - .L

.. .... . . ~..- ....

: ::"-::-:" - ----- Z - -4- -: l-"-.-- -- ..- -- .. - i -' --. "--" " --- E-:: -= z::_ :::p...

- - -Th-- ------ ----

-- : - - --.-

- - 7 :- --- --- --- - . .. -- +....

- oZ-L N 0-:::-::_ _ K_ _ _ _..

--- -:=: : _-_

-:-:4 - -... zv- _ - F---- -

___ ___- _-

o

.--- -- - :--- - .

x- - a ..-- ---.- ----- -- - . D. . .

, . .. - .-- _ - ........- -- ----.

-- ---

---

r

_----- 7 77- ... .. -- - ------ Z-4

-z ___ uuI E~E

_ _____ -f--- - - --- - -- .

: . .. --

um u-, H9- L.,f-.

k--fl I

i--I ---

J

F: I"-AT--+ _

i

1 -t--- --- - I

1i

it

-4 1 -

F ~ ~ ~ ~ ~ -- - ------.

.

.v . .- - -----

WE

IIt [V1,

.111

HPal

CORPORATION DgETZSEN PAPER GRAPH NO. 340-10 DIETZGEN PER INCH 10OX 10

Vii

7~Tb

§

-j.%-ii-.:-IV-ii-ii- A Ir

L -I -11 -j - -i

j~ ~~