Energy efficient engine: Propulsion system-aircraft integration evaluation

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V i^^11' TABLE OF CONTENTS Title Page Section 1.0 SUMMARY 1 2.0 INTRODUCTION EVALUATION PROCEDURE 10 3.0 FLIGHT PROPULSION SYSTEM CHARACTERISTICS 13 4.0 4.1 Configuration Description 4.2 Cycle Selection 4.3 Cycles and Performance 4.4 Engine Weight, Price, and Maintenance Cost AIRFRAME EVALUATION 29 5.0 5.1 Aircraft and Mission Selection 29 5.2 Engine Integration Considerations 45 3.3 Airplane/Engine Peformance Comparison 52 6.0 ECONOMIC EVALUATION 6.1 Economic Model Description 6.2 Economic Evaluation Results NOISE EVAj UATION 17 .0 7.1 Introduction 81 Acoustic Configuration 81 7.2 7.3 Noise Prediction Methodology 84 Predicted Noise Levels 86 7.4 EMISSIONS EVALUATION 91

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8.1 Predicted Emission Levels 91 8.2 Emission Prediction Methodology 91 Margins 93 8.3 EVALUATION OF THE FPS GROWTH POTENTIAL 95' 9.0 Thrust Growth Approach 95 9.1 Performance Effects 95 9.2 Noise Effects 99' 9.3 9.4 Emission Effects GOAL ACHIEVEMENT PROBABILITY ANALYSIS 101 10.0 10.1 Introduction 101 TSFC Probability Assessment 101 10.2 10.3 DOC Probability Assessment Noise Probability Assessment 105 10.4 Emissions Probability Assessment 106 10.5 V

TABLE OF CONTENTS (Cont'd) Section Title Page 11.0 CONCLUSIONS 109 12.0 REFERENCES 110 13.0 SYMBOLS-TERMINOLOGY 112 APPENDICES 113 A. Boeing 115 B. Douglas 197 C. Lockheed 235 D. Distribution List 305 F LIST OF ILLUSTRATIONS Fi gure No.

Title Page Figure 1 Energy Efficient Engine Fuel Savings Fi gure 2 Energy Efficient Engine DOC Improvement Figure 3 Energy Efficient Engine Predicted Noise Levels Figure 4 Propulsion System/Aircraft Integration Evaluation Procedure Figure 5 Energy Efficient Engine Cross Section Figure 6 Installation Sketch of STF505M-7D Figure 7 The JT9D-7A/747-200 Installation Sketch Figure 8 Boeing Advanced Airframe Structures Figure 9 Boeing Airplane 32 Figure 10 Douglas Domestic Airplane Figure 11 Douglas Intercontinental Airplane t Figure 12 Lockheed Traffic Forecast Figure 13 Lockheed Total Long-Haul Aircraft Requirements in Year 2000 Figure 14 Lockheed Comparison of Refined Supercritical Wing of Study Airplane and of Current-Tech- nology Wing of L1011 Airplane Figure 15 Effect of Composite Structure on Empty Weight of Lockheed Aircraft 41' Figure 16 Lockheed Domestic Airplane Figure 17 Lockheed Intercontinental Airplane Figure 18 Wing-Engine Placement Summary 47 Figure 19 Douglas Trijet (Typical) Reverse Flow, Directivity Requirements (Lockheed Trijet Requirements Are Similar) Vii LIST OF ILLUSTRATIONS (Con'd) Figure No. Title Page Boeing Twinjet Reverse Flow Directivity Figure 20 Requirements 50 Figure 21 STF505M-7D Reverse Thrust Capability (Sea Level Std + 14 0C, 10% Leakage) 51 Figure 22 Customer Bleed Schedule, STF505M-7 Data Pack 52 Figure 23 Normal Customer Bleed Requirements 53 Figure 24 Boeing Airplane Weight Breakdown 55 Figure 25 Douglas Airplane Weight Breakdown 56 Lockheed Airplane Weight Breakdown 57 Figure 26 P&WA Aircraft Weight Breakdown Figure 27 58 Thrust Size Required for Each Energy Efficient Figure 28 Engine Powered Aircraft 60 Fuel Savings Relative to the JT9D-7A Figure 29 Powered Aircraft (Design Mission) Fuel Savings Relative to the JT9D-7A Figure 30 Powered Aircraft (Typical Mission) 61 Comparison of Direct Operating Costs 66 Figure 31 Utilization 66 Figure 32 Figure 33 Crew Cost Figure 34 Relative Fuel Cost Figure 35 Airframe Maintenance Costs 68 Engine Maintenance Costs 68 Figure 36 Figure 37 Depreciation Costs Figure 38 Insurance Costs Figure 39 Indirect Operating Cost Breakdown 71 STF505M-7D Savings in DOC Relative to the Figure 40 JT9D-7A for Typical Missions Viii LIST OF ILLUSTRATIONS (Con'd) Figure Title Page No.

Figure 41 STF505M-7D Savings in DOC Relative to the JT9D-7A for Design Missions DOC Breakdown (Boeing Domestic Twinjet, Figure 42 Typical Mission: 1,850 km) 74 Figure 43 DOC Breakdown (Douglas Domestic Trifjet, L Typical Mission: 1,850 kn 75 Figure 44 DOC Breakdown (Douglas Intercontinental Trijet, Typical Mission) 76 Figure 45 DOC Breakdown (Lockheed Domestic Trijet, Typical Mission: 2,600 km) 77 Figure 46 DOC Breakdown (Lockheed Intercontinental Quadjet, Typical Mission: 5,560 km.) 77 Figure 47 DOC Breakdown (P&WA Domestic Trijet, Typical Mission: 5,560 km) 78 Figure 48 DOC Breakdown (P&WA Intercontinental Quadjet, Typical Mission: 3,700 km) 78 Figure 49 Acoustic Configuration of STF505M-7D 82 Fiaure 50 STF505M-7D Nacelle Acoustic Treatment Design Features 83 Figure 51 Component Noise Level Predictions (P&WA Trijet) 87 Figure 52 Predicted Noise Levels at FAR Part 36 (1978) Certification Conditions 88 Figure 53 Thrust Growth Engine Noise 98 Figure 54 Thrust Growth Engine Design 99 Figure 55 TSFC and DOC Probability Assessment Process 101 Figl ,.re 56 Probability of Achievin g TSFC Goal 102 Figure 57 DOC Probability Buildup (P&WA Domestic Trijet, Typical Mission) 104 Figure 58 Probability of Achievin g DOC Goal 105 ix LIST OF ILLUSTRATIONS (Con'd) Figure No.

Title Page Figure 59 Probability of Achieving Emission Goals 107 Figure 60 Probability of Achieving NOx and Smoke Emission Goals LIST OF TABLES Table No. Title Page Cycle and Performance Comparison of Energy Efficient Engine and JT9D-7A Engine 2 Airplane Definitions 3 Estimated Emissions and Smoke Characteristics 7 4a EEE Performance Summary (STF50511-7D) (SI Units) 18 4b EEE Performance Summary (STF505M-7D) (English Units) 19 5 Summary Comparison of EEE and JT9D-7A Engine 20 6 Performance Comparison of the EEE and JT9D-7A Fully Installed With Bleed and Power Extraction Standard Day 21 7 Energy Efficient En g ine Weight Breakdown 23 EEE (ST'F505M-7D) Price and Maintenance Cost Component Breakdown 25 9 EEE Mean Time Betweend Repair By Module 27 Comparison of JT9D-7A and EEE 28 11 Boeing Aircraft Mission and Sizing Criteria 30 12 Douglas Aircraft Mission and Sizing Criteria 34 13 Lockheed Aircraft Mission and Sizing Criteria Effects of Advanced Active Control System on Lockheed Aircraft 40 15 Pratt & Whitney, Aircraft Mission and Sizing Criteria Domestic Airplane Definitions Summary 46 17 Intercontinental Airplane_ Definitions Summary 46 18 Accessory Location Study Summary 48 19 P&WA Airplane Weight Breakdowns 59 X1 - ^{ LIST OF TABLES (Con'd) Table No. Title Page 20 Energy Efficient Engine Fuel Burned Advantage Breakdown 62 Airline Operating Cost Model Elements 64 Economic Model Assumptions 65 Indirect Operating Cost Savings of EEE Relative to JT9D Reference Engine 79 24 Percent Return on Investment Advantage of EEE Over JT9D Reference Engine 80 90EPNdB Takeoff and Approach Noise Footprint Areas 89 Estimated Emissions and Smoke Characteristics 91 27 Landing and Takeoff Cycle 92 28 Summary of Margins 94 Thrust Growth Potential Summary Relative to Base Energy Efficient Engine 96 30 Effects of Growth Steps on Exhaust Emissions 100

of Achieving Noise Goal 106

31 Probability k 1.0 SUMMARY NASA is sponsoring an Energy Efficient Engine Program that is intended to develop and demonstrate an advanced technology base for a new generation of more fuel-conservative engines for commercial transport use. This report summarizes a portion of the effort conducted under this program. The purposes of the Propulsion System-Aircraft Integration Evaluation (PS-AIE) portion of the Energy Efficient Engine program are to estimate the flight performance and operating economics of future commercial transports utilizing the Energy Efficient Engine and to assess the probability of meeting the NASA goals of at least 12% reduction in thrust specific fuel consumption (TSFC) and at least 5% reduction in direct operating costs (DOC), relative to the JT9D-7A reference engine, while meeting FAR Part 36 (1978) noise requirements and the proposed 1981 EPA exhaust emissions standards.

This report presents the results of the initial PS-AIE, in which Pratt & Whitney Aircraft was assisted by Boeing, Dou g las, and Lockheed.

o ENGINE CYCLE AND PERFORMANCE/WEIGHT/COST CHARACTERISTICS The cycle and performance differences between Energy Efficient Engine (EEE) used in this study and the JT9D-7A reference engine are summarized in Table 1. The JT9D-7A, installed in the Boeing 747-200 short-duct nacelle, was chosen as the reference point because it is the most widely used Pratt & Whitney Aircraft high bypass ratio turbofan.

The cycle changes shown in Table 1 were combined with advanced component technologies to produce a predicted thrust specific fuel consumption advantage of 14.9 percent at maximum cruise thrust, 10,670 meters, Mach 0.8. Performance values shown in the table include installation effects and isolated nacelle drag, but no customer bleed or horsepower extraction. This predicted TSFC advantage clearly surpasses the NASA goal of a 12 percent minimum reduction.

Total flight propulsion system weight of the Energy Efficient Engine is estimated to be 1.7 percent heavier than the JT9D-7A/-200 reference (scaled to equal cruise thrust). Price is predicted to be 0.7 percent higher than the reference, while maintenance cost is expected to be 3.2 percent lower. The predicted increases in weight and price are due primarily to the change from the very short duct -200 nacelle of the reference engine to the long duct, mixed flow, Energy Efficient Engine nacelle.

o AIRPLANE PERFORMANCE AND ECONOMICS The advanced airplane configurations used by Boeing, Douglas, Lockheed (P&WA) in the evaluation of the Energy and Pratt & Whitney Aircr^^, ft TABLE 1 CYCLE AND PERFORMANCE COMPARISON OF ENERGY EFFICIENT ENGINE. AND JT9D-7A ENGINE SUMMARY EEE JT9D-7A Takeoff Thrust, N (lbf) 182,880 (41,115) 204,720 (46,025) Turbine Rotor Inlet Temp, o C (OF) Takeoff, +14 0 C (+25 0 F) Day 1369 (2495) 1260 (2300) Max Climb, +10 0 C (18 0 F) Day 1321 (2410) 1169 (2135) Max Cruise, Std. Day 1206 (2205) 1088 (1990) Overall Pressure Ratio 38.6a 25.4b Fan Bypass Ratio 6.51a 5.1b Fan Pressure Ratio 1.74a 1.58b Exhaust Type Mixed Separate Max Cruise Installed Performance (10,670 m (35,000 ft) Mn = 0.8) Thrust, N (lbf) 43,260 (9726) 44,320 (9964) TS:FC, kg/hr/N (lbm/hr/lbf) 0.05874 (0.576) 0.06904 (0.677) -,14.9% Notes: (a) Aerodynamic Design Point, 10,670 m (35,000 ft), MN = 0.8 (b) 10,670 m (35,000 ft) MN = 0.8 Max. Cruise Efficient Engine are defined in Table 2. These study airplanes were chosen by each company as representative of the missions, technolog nd likely to be required for early 1990's

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Each of the airplane manufacturers and Pratt & Whitney Aircraft evaluated both the Energy Efficient Engine and the JT9D-7A reference* engine performance on design and typical missions Figure 1 summarizes the fuel burned advantage shown by Energy Efficient Engine over the JT9D-7A in each airplane. The individual bars cover fuel burned for both design and typical- missions. These results correlate well with *Douglas used the JT9D-20 engine for reference. This -engine is the same basic engine as the -7A, except adapted to the DC10-40 airplane.

TABLE 2 AIRPLANE DEFINITIONS For Domestic Missions BOEING DOUGLAS LOCKHEED P&WA Type Twinjet Trijet Trijet Trijet In Service Date 1990's 1990's 1990's 1990's Design Range - km 3700 5560 5560 5560 (n.mi.)

(2000) (3000) (3000) (3000) Passengers 196 458 500 440 Cruise Speed - Mach No.

0.8 0.3 0.3 0.3 For Intercontinental Missions DOUGLAS LOCKHEED P&WA Type

Trijet Quadjet Quadjet

In Service Date 1990's 1990's 1990's Design Range - km 10190 12040 10190

(n.mi.) (5500) (6500) (5500)

Passengers 438 500 510 Cruise Speed - Mach No. 0.8 0.8 0.8 the total design fuel load (mission design fuel fraction, which is reserves) divided by design takeoff gross weight (TOGW). Average fuel saving for the Energy Efficient Engine over JT9D-7A is 16.6 percent on typical missions and 17.3 percent on design missions._- Airline operating economics were evaluated by Pratt & Whitney Aircraft, using the NASA approved economic model for all engine/airplane combinations. This model used 1977 dollars and assumed fuel prices of 10.6/liter (40t /gal) domestic and 11.9/liter (45C/gal) international. Direct operating cost reductions for the in Figure 2.

Energy Efficient Engine relative to JT9D-9A are shown Since the primary advantage of the Energy Efficient Engine is reduced to the fuel fuel consumption, the trends in DOC are quite similar R w (5500-6500 N.M.)

(3000) 10200.12050 KM 20 5560 DESIGN RANGE DAC_ (2000) _ _ P&WA ' LCC P&WA I_ 10, ~ .101-CC DAC 15 /' / PERCENT FUEL BOAC BURNED REDUCTION FROM JT9D-7A 10 5L 0.1 0.2 0.3 0.4 FUEL WEIGHT FRACTION — DESIGN FUEL WEIGHT TOGW Figure 1 Energy Efficient Engine Fuel Savings -- The average fuel and missions considered is savings for all airplanes 17.0 percent.

than the airplanes show a DOC advantage greater burned trends. All on design missions, and all but NASA goal of five percent (minimum) five percent on typical missions. Average the Boeing twinjet surpass 7.6 percent on DOC reduction was 9.7 percent on design missions and typical missions.

o NOISE, • Douglas, and Lockheed supplied Pratt & Whitney Aircraft with Boeing, for 36 noise flight conditions and airframe noise estimates FAR Part their Energy Efficient Engine powered study airplanes. Pratt & Whitney evaluated total an acoustic configuration and Aircraft defined airplane noise for each of these study airplanes and for the two Pratt & Whitney study airplanes. The results shown in Figure 3 indicate that is predicted to meet FAR Part 36 (1978) the Energy Efficient Engine noise limits in all study airplanes except the Boeing twinjet. In the relatively small Boeing twinjet it is cixrrently predicted to slightly exceed the FAR Part 36 limit at takeoff, while meeting the approach and sideline limits.

(5000.6500 N.M.)

(3000) 10200-12050 KM 12 DESIGN RANGE DAC ^- P&WA f -^ (2000) PERCENT 3700 /^ ' , ► — ._... , -.._ ...^ D iC' DOC REDUCTION g LCC P&WA FROM JT9D-7A LCC PROGRAM GOAL (> 5%) ,` N \ ^BCAC 0.1 0.2 0.3 0.4 DESIGN FUEL WEIGHT FUEL WEIGHT FRACTION ^ - TOGW Figure 2 Energy Efficient Engine DCC Improvement -- The average DOC improvement for all airplanes and missions considered is 8.5 percent, well above the NASA goal of 5 percent improvement o EMISSIONS The gaseous emissions estimates for the Energy Efficient Engine, shown in the Table 3, include allowances for engine-to-engine variations as well as deterioration and development margins. A comparison with the NASA goal of meeting the proposed 1981 EPA exhaust emissions standards shows that total unburned hydrocarbons (THC) and carbon monoxide (CO) emissions are well below goal limits,the smoke number meets the goal limit, and the oxides of nitrogen (NOx) emissions do not meet the goal limit.

o GROWTH POTENTIAL The Energy Efficient Engine was designed to have the potential for thrust growth. To evaluate this potential, two specific growth steps were defined. One was a 15% thrust increase accomplished by increasing .

APPROACH FAR 36 (1978) O P&WA p sCAC TAKEOFF O DAC m LCC z 100 FAR 36 (1978) a.

W SIDELINE m Z a W m 100 Z A W 180 220 360 (400) (600) (800) (200) E TAKEOFF GROSS WEIGHT (400) (600) (200) (800) — THOUSANDS OF kg— TAKEOFF GROSS WEIGHT (THOUSANDS OF Ibm) THOUSANDS OF kg — (THOUSANDS OF Ibm) Figure 3 Energy Efficient Engine Predicted Noise Levels -- All study airplanes achieved the NASA goal of meeting FAR Part 36 (1978) noise certification requirements.

overall and fan pressure ratio and rotor inlet temperature, while preserving existing external nacelle lines. The second step provided a 25% thrust increase by increasing overall pressure ratio, rotor inlet temperature and total airflow, requiring a larger nacelle. Evaluation of these thrust growth strategies showed that growth can be achieved with small impact on the performance and environmental goals of the engine.

The larger growth step produces a 1.1% improvement in cruise TSFC over the base Energy Efficient Engine, while increasing fan and jet noise less than 1 dB each. CO and THC exhaust emissions are decreased by 0.2 and .05 EPAP, respectively, while NO x emissions are increased by 1.0 EPAP.

-6 TABLE 3 ESTIMATED EMISSIONS AND SMOKE CHARACTERISTICS EEE GOAL (1981 EPA) 3.0 EPAP* CO 2.0 0.4 EPAP* THC 0.2 3.0 EPAP* NO x 4.3 20 (Max) 20 Smoke No.

*Environmental Protection Agency Parameter, lbm pollutant/1000 lbf thrust/hr/cycle ' 2.0 INTRODUCTION The National Aeronautics and Space Administration has the responsibility for advancing technology to improve the energy efficiency of future commercial transport aircraft. One element of the plan for meeting this responsibility is the Energy Efficient Engine Program. The objective of this program is to develop and demonstrate advanced turbofan engine component technologies for achieving the NASA goals of at least a twelve percent reduction in thrust specific fuel consumption (TSFC) and at least a five percent reduction in direct operating cost,(DOC) compared with current commercial commercial engines while meeting FAR Part 36 (1978) noise requirements and proposed 1981 EPA exhaust emission standards. Pratt & Whitney Aircraft is a major participant in the Energy Efficient Engine Program through NASA Contract NAS3-20646, which covers a six-year duration (1978-83) and is intended to develop an initial engine design and advance the technology level for an engine that could be introduced into commercial service in the early 1990's.

This report presents the results of the initial Propulsion System-Aircraft Evaluation (PS-AIE) portion of the Energy Efficient Engine program. The purposes of this evaluation, which took place during the Energy Efficient Engine preliminary design task (Reference 1) in 1978, were to provide flight and economic performance estimates of future commercial transports using the current design of the Energy Efficient Engine propulsion systems and to assess the probabilities of meeting the NASA established goals for TSFC, DOC, noise, and emissions. Three airframe manufacturers--Boeing Commercial Airplane- Company (BCAC), Douglas Aircraft Company (DAC), and Lockheed the evaluation through in California Company (LCC)--assisted will be updated periodically during subcontracted efforts. The PS-AIE the program as the results of the component development and testing become available.

Specific items evaluated in the initial PS-AIE and covered in this report include: • Flight propulsion system (FPS) performance predictions (TSFC) • FPS weight, price, maintenance cost estimates o Definition of possible 1990's airplanes integration effects and airplane fuel burned o FPS/airplane estimates* as a reference.

*With the Energy Efficient Engine and the JT9D-7A 8_ • Predicted airline operating economics (DOC and ROI) • Total engine and airplane noise predictions • Exhaust emissions predictions • Growth potential of FPS and effects on performance • Probability of program goals achievement 3.0 EVALUATION PROCEDURE The procedure followed in the initial PS-AIE is shown in Figure 4.

At the start of the evaluation Pratt & Whitney Aircraft and each of the three airframe manufacturers individually defined aircraft that would be suitable for introduction into commercial service in the early 1990's, reflecting their projections of the market conditions and technology levels that will prevail at that time. The definitions included design and typical mission range, number of passengers, design Mach number, configuration, and types and levels of advanced technologies.

Concurrently, Pratt & Whitney Aircraft in consultation with NASA defined an airline economic model for use in the evaluation. This model included methods for calculating direct and indirect operating costs (IOC), revenues, and return on investment. Fuel prices, labor rates, year-dollars, and airplane pricing formulas were specified in the model. The airplane definitions and economic model were approved by NASA.

PROPULSION SYSTEM ENERGY EFFICIENT AIRCRAFT INTEGRATION ENGINE BOEING EVALUATION •PRELIMINARY DOUGLAS DESIGN & INTEGRATION LOCKHEED PRWA STUDIES NASA I EVALUATION AIRCRAFT MISSION GROUND DEFINITIONS RULES (NAS3- 20628) ENERGY EFFICIENT I APPROVAL TRANSPORT I O /AIRF AMC I INTEGRA N R • PROPULSION AIRCRAFT & INTEGRA TION INSTALLATLON SYSTEM EVALUATION • AERO TECHNOLOGY EVALUATION I •ACTIVE CONTROLS I I INTEGRATED I ENGINE - AIRCRAFT ACEE/ EVALUATION I COMPOSITES PROGRAM I I APPROVAL REPORT Propulsion System/Aircraft Integration Evaluation

Figure 4

Procedure -- This procedure provided reliable estimates of the flight and economic performance of representative the Energy aircraft employing

1990's commercial

Efficient Engine as their flight propulsion system, Pratt & Whitney Aircraft, using computer simulations, produced installed engine performances that covered the flight envelope for both the Energy Efficient Engine and the reference JT9D-7A engine. The thrust and fuel flow simulations included the effects of isolated nacelle drag, customer bleed, and horsepower extraction. Other propulsion system characteristics--engine and nacelle weights, dimensions, and costs--were calculated in a consistent manner for both engines.

Since the PS-AIE and the engine preliminary design task were performed simultaneously, engine characteristics were changing up to the end of the PS-AIE. Because of the time required by the airframe manufacturers to size their airplanes and to perform mission analyses, as well as the time required to prepare a comprehensive engine performance data package, the engine performance and characteristics Pratt & Whitney Aircraft' provided to the airframers had to be of an early version of the engine.

Data for this early version--the STF505M-7C presented in the May 1978 data pack--and for the JT9D-7A engine were used by the airframe manufacturers t-) size and evaluate the performance of their advanced airplanes and r ,.iraish Pratt & Whitney Aircraft with airplane data: weights and dimensions, aerodynamics, fuel burned on design and typical missions, engine size, and flight conditions and airframe noise at FAR 36 noise measuring points. Design and typical mission sensitivities to TSFC, propulsion system weight, and nacelle drag were also provided.

In addition, the airframe manufacturers assisted in the preliminary design of the nacelle and in establishing installation requirements for the propulsion system.

Pratt & Whitney Aircraft combined flight propulsion system costs with the airframe evaluations to determine airline operating costs--DOC, IOC, ROI--for all study airplanes for both the Energy Efficient Engine and the JT9D-7A. The NASA approved economic models were used for these evaluations.

Pratt & Whitney Aircraft also calculated exhaust emissions and noise for the Energy Efficient Engine and combined these with the aircraft-alone noise supplied by the airframe manufacturer to determine the total airplane noise for all study airplanes.

Using the data from the performance and environmental evaluations, the overall probabilities of the flight propulsion system meeting NASA program goals were assessed for TSFC, DOC, noise, and emissions.

The flight propulsion system described in this report represents the status at the end of the initial engine preliminary design efforts -- engine model STF505M-7D. Since the earlier STF505M-7C version was used in the evaluations by the airframers, it was necessary to adjust the airplane performance results to reflect the difference in performance and weight between the two models.

The cycles of the two engine versions were essentially the same. The had difference between the engines was that the earlier STF505M-7C assumed levels of component performance, duct losses, secondary flows, weights, etc., while the STF505M-7D had values representing the status at the end of the initial preliminary design effort. In some cases the latter version had not achieved the values assumed in the earlier version.

The net performance loss from the STF505M-7C to the STF505M-7D was about +0.6 percent in cruise TSFC and -4 percent in takeoff and cruise thrust. To determine the adjustment required to account for these differences, Pratt & Whitney Aircraft ran airplane performance analyses of both the STF505M-7C and the STF505M-7D engines in the study airplanes used by Pratt & Whitney Aircraft for the domestic and intercontinental missions at a variety of design and typical mission distances, covering the range of missions considered. Adjustments to such parameters as takeoff gross weight, fuel burned, and direct operating cost for each airplane-mission combination were assessed in this manner. The results shown in the main body of this report represent STF505M-7D characteristics; however, the results in the airframe manufacturer's reports in the appendices represent STF505M-7C characteristics.

4.0 FLIGHT PROPULSION SYSTEM CHARACTERISTICS 4.1 CONFIGURATION DESCRIPTION The Energy Efficient Engine Design used in this study is an advanced, high bypass ratio, two spool turbofan engine with a full length nacelle and a mixed exhaust. The design and principal features of this engine are described in detail in Reference 1. A number of special features of this engine are shown in the cross section in Figure 5.

Starting from the front, pressure ratio, 2.7 aspect ratio the 1.79 (AR), 26 blade, shroudless, hollow titanium fan, is followed by 30 integral strut-fan exit guide vanes (10 structural, 20 non- structural). Fan blade containment is provided by a Kelvar wrap. The compressor counter-rotates with the 14:1 four stage, low-pressure overall pressure ratio, ten stage high-pressure compressor. The pressure ratio is 38.6:1.

stages of the Variable geometry is employed in the first four clearance control is high-pressure compressor, and external active employed on the remaining stages. The combustor is a two-stage annular configuration which was designed for low emissions. A single-stage, high, rim speed high-pressure turbine is a key feature of the Energy Efficient Engine design. This single-stage turbine, employing single-crystal alloy blades, was a major contributor to the 40 percent overall reduction in the total number of engine airfoils, relative to the JT9D-7A. Both the high-pressure turbine and the four stage counter-rotating low-pressure turbine have active clearance control for tighter tip clearances at cruise.

twelve-lobe design with a 0.5 'L/D The mixer is a short, scalloped, mixing length. A full authority digital electronic control is used.

installation sketch the Energy Efficient Engine is shown in An of Figure 6. Comparing this sketch with the JT9D-7A/200 installation in of the Figure 7 shows the difference between the short duct nacelle the Energy_ reference engine and the long duct mixed flow nacelle of Efficient Engine.

Key 'features of the long duct nacelle are an integrated engine/nacelle structure (which reduces rotor clearances and improves load sharing engine performance retention by reducing engine deflections), honeycomb materials, acoustic treatment lightweight composite and reverser with twelve replaceable throughout, a fan-stream thrust by the mixer and cascade racks, and core stream thrust spoiling primary nozzle in reverse thrust operation.

The overall engine design emphasizes mechanical simplicity, and a reasonable development performance retention, maintainability, risk, which, combined with the large fuel consumption benefit, results in a commercially acceptable energy efficient engine.

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^ ' 1 I x dl l .^ I 1 91 a 1 4r----_T•_-may---------- '^•w -''--^^_ II 1 1 I 1 -- II ,.J.______--L^I_---- • 319.3 (125.7) 659.9 (259.8) DIMENSIONS IN CENTIMETERS (INCHES) Figure 6 Installation Sketch of STF505M-7D 4.2 CYCLE SELECTION Selection of the Energy Efficient Engine cycle was based on extensive performed during the earlier Preliminary cycle/configuration studies 3. completed under Design and Integration Studies, Reference Studies initial configuration screening to ask I of that contract were for study.

determine the most promising engine types for further Four configurations were evaluated: two separate-exhaust configurations, one _direct drive and one geared; and two mixed-exhaust configurations, one direct drive and one geared. The cycle for each configuration was based on work performed during an earlier NASA-sponsored program, the Low Energy Consumption Program (Reference 3). Performance, DOC, noise, and emissions estimates were obtained for each configuration.

Each of the four configurations were further studied over a wide range of cycle variations during Task II of Contract NAS3-20628. The range of the cycle parameters studied was 33 to overall compression ratio (OPR), 11200 to 1290 0 C (2050 0 to 2360 0F) rotor inlet temperature (RIT), and 6 to 11 bypass ratio (BPR). Boeing, Douglas, burned and DOC of and Lockheed assisted in evaluating the fuel rt N 123 (49) DIA (128) (287) DIMENSIONS IN CENTIMETERS (INCHES) Figure 7 The JT9D-7A/747-200 Installation Sketch -- This engine was chosen as the reference for the Energy Efficient Engine because it is the most widely used Pratt & Whitney Aircraft high bypass ratio turbofan.

selected engines from this range--Pratt & Whitneyircraft made a fuel

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all engines studied. The results of burned and DOC analysis for the the selection of two mixed-flow engine configurations Task II led to for further evaluation in Task III.

During Task III, key areas of technical concern were investigated for drive and geared engines.

the mixed-exhaust configuration in direct Both single and two stage high pressure turbine versions were studied.

configuration chosen for Engine cycles representative of each were 38.6 these studies. All configurations were OPR and 12040C (2200 0F) RIT. The direct-drive engines had a 1.74 fan pressure ratio (FPR) and a 6.5 BPR; the geared engines, a 1.52 FPR and a 9.1 BPR.

Growth studies and performance/DOC risk studies were conducted for each of the four configurations; sufficient preliminary design work was completed to permit a feasibility evaluation and detailed comparison of the various configurations.

This study effort resulted in the selection of the following cycle for the Energy Efficient Engine: 38.6 BPR, 1204 0C (2200 0F) RIT, 1.74 single-stage HPT, mixed exhaust. The FPR, and 6.55 BPR direct drive, selected OPR was the highest currently considered to be feasible from material and mechanical standpoints to obtain the lowest TSFC and fuel burned, consistent with the need for OPR increase to 45:1 for thrust growth. RIT was the optimum level for TSFC and slightly below the optimum for fuel burned. Higher RIT resulted in a significant increase in estimated DOC. The need for increases of up to 940 C (170 0F) in RIT for thrust growth was also considered. The FPR/BPR was selected based on the optimum compromise between TSFC/fuel-burned and DOC.

Reducing FPR (increasing BPR) reduced TSFC and mission fuel burned, but increased DOC.

The choice of direct drive was consistent with the FPR/BPR selection.

The single-stage high-pressure turbine (HPT) configuration offered a simplified cooling system, requiring only two cooled rows for lower maintenance cost and improved DOC, relative to a two-stage HPT configuration. The selected mixed-exhaust system offered improvements in installed performance and DOC, relative to a separate flow configuration.

4.3 CYCLES AND PERFORMANCE Engine cycle parameters, component performance levels, and engine overall performance are shown in Table 4a and Table 4b. Except for the last two flight conditions in the table, all the engine- performance levels are for the uninstalled STF505M-7D engine (ideal inlet, zero nacelle drag, no customer air bleed or power extraction); the last two flight conditions in the table represent typical noise points and are fully installed with flight inlet, isolated nacelle drag, and typical customer bleed and power extraction. The aerodynamic design point (Table 4a) is used for component design and represents a typical cruise altitude and Mach number at a power setting between maximum cruise and maximum climb power. All cycle parameters are based on this condition. Table 4a uses Standard International (SI) units, while 4b is in English units.

A comparison of the Energy Efficient Engine (STF505M-7D) and the JT9D-7A is presented in Table 5. The takeoff thrust size (uninstalled) of the Energy Efficient Engine is somewhat less than that of the JT9D-7A, and the cycle pressure ratio, RIT, BPR, and FPR are higher.

These cycle differences combined with the superior component Energy technology and-mixed-flow nacelle of the Efficient Engine resulted in an estimated 14.9 percent improvement in installed (flight inlet, nacelle drag, customer bleed power) but no or cruise TSFC at the 10,670 m (35,000 ft), Mach number of 0.8, maximum cruise condition., The Energy Efficient Engine and the JT9D-7A engine are further compared in Table 6. Net thrust,, thrust specific fuel consumption, and fan total corrected airflow are included in the comparison. The performance levels in this table are fully installed including customer bleed and power extraction. The performance data is TABLE 4 a EEE PERFORMANCE SUMMARY (STF505M-7D) (SI Units) 85% Aerodynamic Maximum Maximum Maximum Ma x imum Typica l Typ ica l .

t Desion- Poin Cru Cl imb Takeoff Cl imb Cr u s Takeoff Approach ise Altitude (m) 10,568 10,663 10,668 Sea Level 6, 705 10,668 365.8 120 0. 80 0.80 Static 0.^oo 0.800 0.245 0.208 Ma ch Number o.io .

blant Temp. Std. Day Std. Day Std. Day Std. Day+13.9 0C Std, Day+100 0 Std. Day Std. Oay+IOO C Std. Day+100C IGO ioo 100 100 100 too 99.65 99.65 inic Recovery t%) Extraction (kW) 0 0 0 0 0 a 1 12. 6 112.6 Powa y Cu 0 a 0 0 0 1.13 1.202 stemer Steed (kg/sec) 0 ( , d /drag) ft;) 47,372 45,419 50,672 182,880 70,216 38,504 131,973 37,208 Thrust Fu el Flow 'kg/hr) 2658 2855 6133 4104 2179 5857 1,91 0.056 0.056 0.056 0.034 0.058 o.C56 0.044 0.048 (kglhrli) TS'C Rot o r Inlet Tamp. ( d C) 1225 1206 1257 136 1341 1.138 1348 9139 Overall Pressure Ratio 38.6 . 37 . 4 40. 5 30. 2 34. 5 33.5 29.1 11.9 Byp ass Ratio 6151 6.59 6-7 7.6 6,88 61i2 7.iO 9.39 Fan 00 bind/4 kg/sec 513 607 624 541.6 589 583. 5 538 350 PR 1,74 1.71 1.79 1.57 1.65 1.61 1.54 i.16 a{",) 37,3 87,1 88.2 87.5 8519 83.0 87.0 RPM 3660 3614 3741 3626 3710 3444 3565 221 1 9 Fan 10 and LPC W @ / a ' 4 9/sec 94.2 92,1 97.9 77.1 66 84,5 75.8 37.3 - PR 2.755 2.70 2.85 2.36 2.59 2.51 2.31 1.43 92.2 94.2 90.2 91.5 90.2 89.8 90.5 85.6 1(%) JPC W% r E) 14 Kg/SGC 40.0 390 40,3 37.2 38 33.7 37.2 27.5 PR 14.0 13.85 14.20 12.80 13.3 0.3 12.6 8.33 88.2 58 -1 88.0 85.9 88.4 88.7 82.3 96.4 t%) ' RPM 12,j62 12,291 12,473 13,006 12,801 12,C46 12,905 11"251 Burner W Nrel, ka/sec 3.56 3.58 3.58 3.58 3.6 3.6 3.6 3.77 P/P 0.055 0.055 0.055 0.055 0.055 0.056 0.055 0.061 99,95 99.95 99. 9 5 99,95 99.95 99.95 99.95 99.95 HPT 0.94 W \rTl P - ko/ • ec 0. 94 0.94 0.94 0.94 0.94 0.94 0.945 ( OK/042m^) PR 4, 0 4. 0 4. 0 4. 0 4. 03 4.05 4.03 3.95 8^.2 SE 2 8^.2 si.2 BK 3 88.2 87.3 87.6 W' V T/P - kg/;ec 3.7 3.7 3.7 3.7 3.7 3.8 3.77 3.70 OK/kti/ml) PR 5.6 5.5 5 7 a ,. 9 5.37 5.33 4 86 2.65 9115 9i.4 9i .6 96.2 -1 9i. 0 96. 0 84.5 A P/P Trans Duct 0.015 0.015 0.015 0.015 0.015 0.015 0.015 0.015 . 1 P/P TEGV 0.009 0.009 0.0094 0.0066 0.0083 C.0080 0.0065 0.0016 AP/P Duct 0.006 0.006 0.0059 0.0054 0.0061 0.0062 0.0056 0.0034 - S P/P Core Mixer 0.0024 0.0024 0.0025 0.0019 0.0022 0.0022 0.0018 0.0005 4 P/P Duct Mixer 0.0018 0.0018 0.0018 0.0017 0.0018 0.0013 0.0017 0.0012 •11 ! /P Tailpipe 0.0034 0.0034 0.0034 0.0029 0.0032 0.0031 0.0030 0.0015 % Mixing 85 85 85 85 85 85 85 85 Nozzle Gross Thrust Coefficient 0.9958 0.9960 0.9955 0.990 0.996 0.996 0.9906 0.9935 Drag (N) 0 0 0 0 0 0 565 Turbine Cooling and Leakage Air (% WA Core) 15.95 15.95 15.95 15.45 15.45 15.95 15.45 15.45 oRIG PA

OFPVV

A 06i Qtr-4 b TABLE 4 SUMMARY EEE PERFORMANCE (STF505M-7D) (English Units) 85% Maximum Aerodynamic Naximum Maximum Ma x imum Typical Typical Design Point Cruise Climb Takeoff Climb Cruise Takeoff Ap p roach Altitude (ft) 35,000 35,000 35,000 Sea Level 22,000 35,000 1200 394.0 MachNumber 0.80 0.80 0.80 Static 0.70 0.800 0.2455 0.208 Ambi en t Temp. Std. Day Std. Day Std. Gay Std. Day +25 O F Std. Day+18 0F Std. Coy Std, Day+IS O F Std. Day-18OF Inlet R ecovery (%) 100 too 100 100 100 100 99.65 99.65 Power Extraction (hp) 0 0 0 0 0 a 151 151 Customer Bleed (ibm/sec) 0 0 0 0 0 0 2.63 2.65 Thrust (w/drag) (lbf) 10,650 10,211 11,392 41,115 15,786 8679 29,670 8365 Fuel Flow (Ibm/hr) 5860 5620 6295 13,521 9047 4803 12,916 3959 TSFC (lbi n /hr/lbf) 0.55 0.55 0.55 0.18 0.57 0.55 0.44 0,47 Rotor Inlet Temp. ( O F) 2238 2203 1 1 295 2495 2446 2081 2457 1683 Overall Pressure Ratio 38.6 37.4 40.5 30.2 34.5 33.5 29.1 11,9 Bypass Ratio 6.51 6.59 6.37 7.0 6.88 6.91 7.10 9.39 Fan 00 W CE) /j lbm/ser 35Z 1338 1375 1194 1297.5 1286'5 1187 772.2 PR 1 74 1.71 1.79 1.57 1.65 1.61 1. 54 1.16 7 (w) si 3 8i 3 ai I 8i 2 8^ B^ 9 8^ 0 ai 0 3 7i, 0 RPM 36 00 3 4 3 1 3 626 6i 7i 444 356 * 5 22i9 to Fan and LPC WVO/h lbm/sec 207.7 203.1 215.8 170 188.6 186.3 167.2 82.2 PR 2. 75 2.70 2 . 85 2.36 2.59 2.51 2.31 1.43 8;,3 89.5 89.0 91.5 90.2 S9.8 90.5 25.6 M HPC W ^e - /' Ibm/sec 88.1 87.5 88.9 82.0 84 85.3 82.0 60.7 PR 14.0 13.85 14.2 12.8 13.3 13.3 12.6 8.33 r 88.2 88.3 88.0 88.9 88.4 88.7 88.8 86.4 q (%) RPM 12,362 12,291 12,473 13,006 12,801 12,048 12,905 11,251 Burner W\ / E)/, Ibm/sec 7.9 7.9 7.9 7.9 7.9 8.0 7.9 8.32 a P/P 0.055 0.055 0.055 0.055 0.055 0.056 0.055 0.061 1(%) 99.95 99.95 99.95 99.95 99.95 99.95 99.95 99.95 HPT W^P - lbmlsa^ 19.1 19.1 19.1 19.1 19.1 19.1 19.1 19.2 ( OR/Ibf/in.

4.0 4.0 4.0 4.0 4.03 4.05 4.03 3.95 PR *(%) 88.2 88.2 88.2 87.3 87.3 88.2 87.3 87.6 LPT lbm/seC 76.3 76.3 76.1 76.4 76.4 76.6 76.5 75.1 W \ 1 71P - ( OR/Ibf/in. 4 ) 5. 6 5. 5 5. 7 4 . 9 5. 3 7 5. 3 3 4. 8 6 2. 6 5 PR (%) 91 . 5 91 . 4 91 . 6 9 0, 2 90 . 3 91 . 0 90 . 0 84 . 5 . P/P Trans Duct 0.015 0.015 0.015 0.015 0.015 0.015 0.015 0.015 I TEGV P/P 0.009 0.009 0.0094 0.0066 0.0083 0.0080 0.0065 0.0016 a 0.006 0.006 0.0059 0.0054 0.0061 0.0062 0.0056 0.0034 AP/P Duct 0.0024 0.0024 0.0025 0.0019 0.0022 0.0022 0.0018 0.0005 aP/P Core Mixer 0.0018 0.0018 0.0018 0.0017 0.0018 0.0018 0.0017 0.0012 4P/P Duct Mixer 0.0034 0.0034 0.0034 0.0029 0.0032 0.0031 0.0030 0.0015 , % P/P Tailpipe as 85 85 85 85 85 85 85 % Mixing Nozzle Gross Thrust 0.9958 0,9960 0.9955 0.990 0.996 0.996 0.9906 0,9935 Coefficient (ibf) 0 0 0 0 0 0 166 127 Drag Turbine Coolina and Leakage Air (% ' WA Care) 15.95 15.95 15.95 15.45 15,45 15.95 15.45 15,45 TABLE 5 OF EEE AND JT9D-7A ENGINE SUMMARY COMPARISON EEE JT9D-7A (STF505M-7D) (46,025) Takeoff Thrust, N (lbf) 182,880 (41,115) 204,720 Turbine Rotor Inlat Iemp, o C (OF) 1260 (2300) Takeoff, +14 1 C (+25 0 F) Day 1369 (2495) Max Climb, +10 0 C (180F) Day 1321 (2410) 1169 (2135) Max Cruise, Std. Day 1206 (2205) 1088 (1990) Overall Pressure Ratio 38.6* 15.4** Fan Bypass Ratio 6.5:1* 5.1** Fan Pressure Ratio 1.74* 1.58** Mixed Separate Exhaust Type (B747-200) Max Cruise Installed 10,670m (35,000 ft) Mn = 0.8 0.06904 (0.677) 0.05874 (0.576) TSFC, kg/hr/N (lbm/hr/lbf) -14.9% ---^ *Aero Design Point Max. Cruise **10,670 m (35,000 ft) Mn = 0.8, representative of airplane mission and economic studies. As the comparison shows, the STF505M-7D has a significantly better specific fuel consumption than the JT9D-7A engine for a wide range of altitudes at higher power settings. However, because of the different off-design characteristics of mixed flow and separated-flow engines, this advantage diminishes at idle descent and becomes a penalty at the lower altitudes.

The thrust comparison shows an even greater difference. The STF505M-7D has power ratings that result in a greater maximum climb and maximum cruise thrust, relative to takeoff thrust, than the JT9D-7A. When both engines are scaled to the same takeoff thrust, the STF505M-7D has about four percent more climb thrust than the JT9D-7A at 6100 m (20,000 ft) at a Mach number of 0.7 and about 17 percent higher climb thrust at 10,670 m (35,000) at a Mach number of 0.8. The cruise thrust of the STF505M-7D is about seven percent higher than that of the - JT9D-7A at 10,670 m/Mn 0.8, and about three percent higher at 13,720 m TABLE 6 PERFORMANCE COMPARISON OF THE BASE SIZE EEE AND JT9D-7A FULLY INSTALLED WITH BLEED AND POWER EXTRACTION STANDARD DAY EEE (STF505M-7D) Relative to Altitude JT9D-7A % Difference M M (ft) Mn Power Setting rr^nt TSFC W / total -- Mn ---- (0) 0.0 Takeoff -12.0 -10.2 -12.3 6100 (20000) 0.7 Max Climb - 8.7 -12.2 -12.5 0.8, Max Climb + 3.4 -14.9 - 8.3 10670 (35000) Max Cruise - 5.7 -14.1 -10.9 10670 (35000) 0.8!

0.8 Max Cruise - 8.3 -12.8 -11.8 13720 (45000) (20000) Idle +74.1 -64.4 -12.8 6100^' 0.7 + 0.2 Idle -55.2 +51.2 1.3 0 (0) *Fnt and TSFC are negative at 0.7 idle condition; STF 505M-7D has 74.1 more negative thrust and 64.4% less negative TSFC at this condition.

These rating differences improve (45,000 ft) and a Mach number of 0.8.

the mission performance of the STF505M-7D engine. For example, if the for climb or cruise, the required engine size and engines were sized the weight would be reduced. If the engines were sized for takeoff, STF505M-7D would have had an improved time-to-climb and would have used less climb fuel.

Idle descent is another area of improved performance for the STF505M-7D engine. The off-design operating characteristics of the mixed-flow configuration combined with the idle ratings of the STF505M-7D result in a significant improvement in descent and taxi fuel consumed, even at those flight conditions where the idle descent TSFC of the STF505M-7D is worse.

4.4 ENGINE WEIGHT, PRICE, AND MAINTENANCE COST 4.4.1 Methodology The methods employed for estimating price, weight, and maintenance all study cost are the same as used by Pratt & Whitney Aircraft on engines. These methods utilize detailed of each analyses component rather than simplistic statistical regressions.

V 21.

Weight and price scaling techniques were used extensively. The scaling techniques are the result of past studies in which engines and nacelles were evaluated in several thrust sizes. The scaling studies were performed in sufficient detail in determine the effects of size on individual major parts and components. Total weight or cost represents the summation of the individual components.

4.4.1.1 Wei ght Estimatin Engine weight was estimated by analytical techniques that utilized computer programs for an accurate weight analysis; statistical procedures were not employed. The weight of each component was estimated in detail as the layout evolved.

Trade studies were also carried out to ensure that minimum weight configurations were considered and incorporated into the final engine definition where practical. Advanced technology items that influenced weight in such programs as the JT10D and Variable Stream Control Engine were evaluated for the Energy Efficient Engine and were incorporated into the design and weight estimate procedure, thus providing aggressive weight estimates that have a high degree of credibility and technical substantiation.

Nacelle weight was estimated by scaling similar components of existing nacelles to Energy Efficient Engine size and then adjusting the scaled weights to account for the use of advanced materials. Aggressive use of composites and titanium in the nacelle allowed weight reductions of 18 percent in the inlet, 27 percent in the fan cowl, and ten percent in the fan reverser and core cowl, compared with conventional metal construction.

Component weight breakdown of the STF505M-7D is shown in Table 7.

Included for comparison are the bare engine, nacelle, and total weights assumed for the STF505M-7C at the beginning of the preliminary design--these weights were used by airplane manufacturers in their evaluations.

4.4.1.2 Price Estimating The engine and spare part prices of the STF505M-7D model were established based on production cost-estimating procedures. A bill-of-materials was generated from design layouts of major parts that represented over 90 percent of the total engine cost.

Cost was first approximated by computer programs that operate on a library of components of reference engines. Similar components were scaled to the STF505M-7D flowpath and adjusted for number of parts, the material and labor cost of each element being analyzed separately.

The many unique features of the Energy Efficient Engine design required an in-depth evaluation of many parts.

22, TABLE 7 WEIGHT BREAKDOWN ENERGY EFFICIENT ENGINE Component STF505M-7D Weight kg (lbm)

Fan 82 1 (1810)

LPC 218 (480) Intermediate/Fan Discharge 354 (780) HPC 345 (760) Diffuser/Burner 304 (670) HPT 376 (830) LPT 898 (1980) Mixer/Plug 118 (260) Controls/Accessories 300 (660) Total Bare Engine* 3734 (8230) 277 (610) Inlet Fan Cowl 88 (195) (1755) Fan Reverser/Core Cowl 796 Tailpipe 125 (275) Total Nacelle* 1286 (2835) Total Flight Propulsion System* 5020 (11.065) *STF505M-7C levels were: (8050) Bare Engine 3652 (2785) Nacelle 1265 Flight Propulsion System 4917 (10835) An extensive data base was available for the detail evaluations, including both production and development engine part costs, material cost correlations and scaling relationships, and trade studies from other programs. Materials and manufacturing specialists from Pratt & Whitney Aircraft and from vendors were consulted when determining costs of unique processes and configurations, such as superplastic forming and diffusion bonding of the hollow fan blades and intermediate case struts. Advanced materials and manufacturing methods were evaluated by utilizing the Pratt & Whitney Aircraft Materials Research Laboratory and Manufacturing Research and En g ineering Development Groups.

The engine cost estimates were reviewed by making detail comparisons with other en g ines. Differences between Energy Efficient Engine configurations and JT10D and JT9D configurations were analyzed to ensure that costs were reasonable.

The total estimated cost of the Energy Efficient Engine is a realistic manufacturing cost based on JT9D-7A production quantities. The changes in configuration evolving during the subsequent detailed design and development of the engine will be adjusted for by a design and development allowance based on trends of past programs.

Direct Operating Cost and Maintenance Cost required selling price rather than production cost levels. Therefore, the Financial Department generated the 1977 budgetary and planning prices. The pricing method used for the STF505-7D was consistent with that used for the JT9'D-7A. A breakdown of the price of the Energy Efficient Engine by major component is shown in Table 8.

Nacelle price was estimated on a constant price per unit weight for both engines.

In addition to estimates for determining engine status, many trade studies were performed to ensure that applicable cost reduction proposals were incorporated.

4.4.1.3 Maintenance Cost Estimating A comprehensive analysis was performed to estimate maintenance cost of the STF505M-7D engine. Maintenance cost includes cost of maintenance material, labor, and outside repair.

Outside repair cost is the cost for repairs not normally accomplished in the airline's repair shop; in this analysis these charges are the labor costs. All costs are for a mature engine included in expressed in 1977 dollars and represent a 15 year cumulative average for a fleet of aircraft introduced into service at the same time was (block feed). A full burdened labor rate of $29.00 per manhour used to convert manhours to dollars. Estimates were made for both intercontinental and domestic missions.

To establish a "hardpoint" base for maintenance cost studies, the maintenance cost- of a mature JT9D-7A engine was established; the maintenance cost of the Energy Efficient Engine was evaluated against this base. A breakdown of maintenance costs of Energy Efficient Engine is shown in Table 8.

TABLE 8 EEE (STF505M-7D) PRICE AND MAINTENANCE COST COMPONENT BREAKDOWN Price Maintenance Cost Fan 14.5% 10.5% LPC 5.56 4.0% Intermediate/Fan Discharge 9.0% 1.5% HPC 11.5% 11.0% Diffuser/Burner 7.5% 10.0% HPT 10.5% 21.0% LPT 25.5% 26.0% Mixer/Plug 3.0% 0.5% Controls/Accessories 10.5% 5.5% Assembly, Test, and Line Maint. 2.5% 10.0% 100.0% Total 100.0% A preliminary analysis of nacelle maintenance cost performed in support of the accessory location study indicated that an Energy Efficient Engine type nacelle design could have an advantage over a JT9D-7A/-200 type installation in labor costs. This advantage would be partially offset by the lower mean time between repair of the Energy Efficient Engine relative to the JT9D-7A. For the purposes of airplane economics comparisons, the Energy Efficient Engine and JT9D-7A/-200 nacelle maintenance costs were assumed to be equal.

Maintenance Material Cost The maintenance material cost (MMC) estimates were obtained from a computer program that simulated the operation and maintenance of a fleet of engines over a 15 year period. Approximately fifty groups of parts (highest MMC contributors) were modelled interactively by means of a Monte Carlo simulation. Mature scrap lives, module mean time between repair, and part prices provided the major input from which a 15 year cumulative average MMC estimate- was derived. In addition to the fifty groups of parts, the MMC of disks (prorated over 15 years) and of miscellaneous parts (those not modelled individually) were added to obtain the total MMC estimate.

Mature JT9D-7A part scrap and repair lives were obtained by extrapolating field experience. The current experience positions were derived from the analysis of spare-part sales records, data obtained from various JT9D operators, and other sources within Pratt & Whitney Aircraft. The JT9D positions were used as a base for extrapolation to the Energy Efficient Engine lives, considering changes in design, advanced technologies, operating environment, and part repairability.

Spare part prices for both the mature JT9D-7A and STF505M-7 models, provided by the P&WA Financial Department, were based on production cost estimates. The pricing method was consistent for both engines and was based on 1977 dollars.

Maintenance Labor Cost Model The maintenance labor cost (MLC) analysis covered maintenance performed on the flight line, in the operators shop, and on parts sent to a repair vendor. Manhours per repair and mean time between repair were estimated for each of the major sections (modules) of the engine.

The maintenance labor cost in terms of manhours per engine flight hour was calculated by dividing the manhours per repair for each section by its mean time between repairs. The manhours per engine flight hour for the complete engine was obtained by summing the manhours per engine flight hour for the individual sections and adding estimated value for line maintenance separately.

The manhours per repair were derived through a comparative analysis of the projected mature JT9D-7A engine manhours. The manhours per repair for the JT9D-7A base was a detailed model in which manhours for module repairs was broken down into module disassembly/inspect/assembly and repair of major part types such as blades, vanes, outer air seals, The manhours per Energy Efficient Engine repair estimates were etc.

generated from the mature JT9D model. 'Differences in engine design, operating environment, materials, size, etc. were taken into account in estimating the mahours.

Estimates of mean time between repairs were also made at the module level. These estimates (Table 9) reflected the total module repair rate independent of what module or part caused the shop visit. As was the case for the manhours per repair, a comparative analysis was performed using the mature JT9D-7A as a base. Differences in the engine design and operating temperatures, pressures, and speeds were taken into account in estimating the module mean time between repairs for the Energy Efficient Engine. - TABLE 9 EEE MEAN TIME BETWEEN REPAIR BY MODULE Module STF505M-7D MTBR (hr) Fan 4400 LPC 6100 HPC Diffuser 7500 Burner 2750 HPT LPT 5000 All Causes 2300 4.4.2 Weight, Cost, and Maintenance Cost Comparison With the Reference Engine The status of the Energy Efficient Engine (STF505M-7D) weight, price, and maintenance cost estimates at the end of the preliminary design phase is compared with the JT9D-7A reference engine in Table 10. The Energy Efficient Engine configuration provides reductions in bare-engine weight, price, and maintenance cost as a result of the mechanical simplicity achieved with advanced materials and manufacturing technology and with increased rotor speeds, high aerodynamic loadings, and advanced engine controls technology. Prime contributors to these benefits are the single-stage high-pressure turbine, the forty percent reduction in the number of engine airfoils, light-weight composite fan containment, full authority electronic control system, and the five bearing rotor support system located in two bearing compartments with two support frames. Although the nacelle design makes aggressive use of advanced materials, the change from the very short-duct, -200 nacelle to the long-duct mixed-flow nacelle increased total nacelle weight and price. Previous studies (NASA CR-135396, Reference 1 ) indicated, however, that these nacelle weight and price. increases are more than offset in terms of fuel burned and DOC by the performance improvements -offered by forced mixing of the exhausts. Nacelle maintenance costs were assumed to be equal for the purposes of this evaluation.

2 7 TABLE 10 COMPARISON OF JT9D-7A AND EEE JT9D-7A* EEE (STF505M-7D) {%) Weight Bare Engine Base - 3.1 Nacelle Base +18.5 Total + Base 1.7 Price Bare Engine Base - 3.4 Nacelle Base +18.5 Total + 0.7 Base Maintenance Cost Base - 3.2 *Scaled to STF505M-7D max. cruise thrust @ 10,670 m, M 0.8 (43.26 kN (9726 lbf)) 5.0 AIRFRAME EVALUATION 5.1 AIRCRAFT AND MISSION SELECTION Each of the airframe manufacturers--Boeing, Douglas, Lockheed-- as well as Pratt & Whitney Aircraft, recommended one domestic and one intercontinental range aircraft for use in the Propulsion System- Aircraft Evaluation (PS-AIE), each aircraft being suitable for early 1990's introduction into service. These aircraft and their mission definitions are described in the following sections, including the technology assumptions and marketing rationale used in their selection.

5.1.1 Boeing Aircraft & Mission Selection 5.1.1.1 Market Considerations and Desi gn Constraints Examination of the market situation indicated that airline requirements in the 1990's will be similar to those existing today.

This prediction assumes that the air traveling community of the 1990's will. be approximately the same percentage of the total population as today, with a small annual growth rate of four to six percent. Air cargo growth should be similar unless a large dedicated air freighter is developed, which might increase the growth rate.

The major airlines probably will retire many of the current narrow-body aircraft by the late 1980's. These aircraft include about 750 intercontinental range 707 and DC8 series airplanes, and over one thousand 727 domestic range airplanes. Therefore, barring unforeseen developments, a market should exist in the late 1980's for a large number of 180-220 passenger aircraft with domestic or intercontinental range capability. Accordingly, the design mission and sizing constraint selected for the Energy Efficient Engine study are shown in Table 11.

Boeing chose to concentrate on one airplane, examining only a domestic twinjet with wing mounted engines.

The takeoff field length of 1830- m (6000 ft) at sea level, 290C (84 0 F), was chosen to approximate a hot day, reduced-range mission takeoff from Denver.

Since passengers have shown a preference for double-aisle seating, a wide body with a seven abreast, two aisle seating arrangement was chosen. The fuselage determined by this seating arrangement accommodates 17 LD-13 containers side by side in the cargo compartment.

TABLE I BOEING AIRCRAFT MISSION AND SIZING CRITERIA 3706 (2000) .

Design Range, km (n.mi.)

,196 Passengers, 15/85 split 0.80 Cruise Mach No.

2286 (7500) Takeoff Field Length 64 (125) Approach Velocity, m/s (kts) 10,058 (33,000) Initial Cruise Alt., m (ft) Typical Mission for Economic Evaluation: 1853 (1000) Range, km (n.mi.)

Passengers, (556 load factor) 0.80 Cruise Mach No.

5.1.1.2 Advanced Technoloav Features Each technology area was reviewed and advanced technology features were identified as being available for a 1986 program start and an early 1990's inservice date. These features are discussed below.

Advanced aerodynamic features included: improved wing/airfoil design, wing/nacelle/strut design for minimum interference, and tailoring of x, the empennage to the wing-body flowfield. These features reduced cruise drag by two percent. In addition, the low speed (takeoff and landing) lift/drag was increased five percent through incorporation of sealed leading edge flaps, seals between nacelle struts and lateral edges of leading edge flaps, and aileron droop.

Advanced structural features depended on use of advanced aluminum alloys, high strength titanium, and composites, as shown in Figure 8.

Advanced flight control technology features incorporated in the airplane design were all-axes handling qualities stability augmentation systems, all flying tail, and double-hinged control surfaces Systems technology advancements applied to the design consisted of

cabin air reconstitution and recirculation, integration of anti-icing

with environmental control system, carbon brakes, and limited

slip

braking system.

30t

CURRENT NEW TECHNOLOGY TECHNOLOGY STRUCTURAL WEIGHT SAVING MATERIAL MATERIAL COMPONENT % COMPONENT WEIGHT 6% ADVANCED *.WING BOX STANDARD 4% ALUMINUM *FUSELAGE ALUMINUM ALLOYS *EMPENNAGE 6% ALLOYS BOX (CURRENT 747) *CONTROL 25% ADVANCED CONVENTIONAL SURFACES COMPOSITE ALUMINUM *LANDING GEAR STRUCTURE CONSTRUCTION DOORS (GRAPHITE) 40 '0 *MAIN LANDING CARBON GEAR BRAKES *LANDING GEAR TITANIUM SUPPORT 20% FITTINGS *SIDE OF BODY RIB •EMPENNAGE BODY ATTACH *ENGINE STRUT ATTACH *FLAP SUPPORT Figure 8 Boeing Advanced Airframe Structures -- Advanced structural features in the Boeing airplane include advanced aluminum alloys, high strength titanium, and composites.

5.1.1.3 Aircraft Design Point Selection Wing loading and thrust loading were chosen to minimize takeoff gross weight and block fuel, with the 1830 m (6000 ft) takeoff field length requirement limiting both parameters.

A configuration drawing of the Boeing aircraft is shown in Figure 9.

5.1.2 Douglas Aircraft and Mission Selection 5.1.2.1 Market Considerations and Design Constraints Normal development and traffic growth trends indicate that a growth airplane (450 to 500 passengers) program is the most likely new W N

O`

LI Figure 9 Boeing Airplane -- .Boeing chose to concentrate on one airplane, a domestic twinjet with wing-mounted engines.

program to be initiated in the 1990's. Although an all new airplane is possible, an improved-technology derivative of a current wide-body transport is more likely. This growth airplane would probably have both a domestic and an intercontinental version, similar to the DC10 series.

A new aircraft is clearly needed to replace DC8's and 707's in domestic operations. But this need is expected to be fulfilled by the DCX-200 and/or 7X7 aircraft,,which should be at their production peak in the mid to late 1980's. A replacement for these aircraft would not, therefore, be required in the early 1990's.

Similarly, the current wide-body transports should continue in production through the 1980's, with stretched versions based on the current wing being introduced in the early 1980's. If there were to be

a technical breakthrough in laminar flow technology in the early

1980's, an airplane sized to replace the existing wide-body fleet would be a logical development since this fleet would be the largest user of aviation fuel. Such a high technology airplane would then be available for the 1990-1995's.

Assuming these marketing projections and that aircraft development will proceed along normal lines without a major technical breakthrough, Douglas based its airplanes on a DC10 trijet derivative with a stretched fuselage and an all new wing. Aircraft sizing and mission criteria are presented in Table 12.

Although the domestic and intercontinental airplanes have different thrust requirements, they are externally similar, having the same wing, fuselage, and empennage. However, the interior arrangements are different. The domestic airplane has a lower galley, allowing more seating but less cargo space (40 vs. 50 LD-3 containers) than the upper galley interior configuration of the intercontinental airplane.

Fuselage diameter is the same as on current DC10 series aircraft, but the length of the fuselaqe has been stretched 18.4 m (60 ft).

5.1.2.2 Advanced Technology Features A review of technology areas indicated that the advanced technology features described below would be available for 1990's application.

Advanced aerodynamic features included a thick supercritical high aspect ratio wing, winglets, and an advanced high lift system. The advantages of a supercritical wing of increased thickness, such as now flying on the YC-15, include lower weight, reduced drag, and improved CL buffet. Increasing wing aspect ratio from the current levels of 6 to 8-1/2 to 10 to 12 reduces induced drag, leading to reductions in engine size and fuel consumption. Winglet design technology, while not ready for the next generation of aircraft, should be sufficiently - 33€.

^h a • TABLE 12 DOUGLAS AIRCRAFT MISSION AND SIZING CRITERIA Domestic Intercontinental Airplane Airplane Design Range, km (n.mi.) 5560 (3000) 10190 (5500) Passengers, 10/90 split 458 438 Cruise Mach No. 0.80 0.80 Takeoff Field Length, m (ft) 2440 (8000) 3350 (11000) Approach Velocity, m/sec (kts) 67 (130) 69 (135) Initial Cruise Alt., m (ft) 10060 (33000) 9450 (31000) Typical Mission for Economic Evaluation: Range, km (n.mi.)

1850 (1000) 2780 (1500) Passengers (60% load factor) 275 263 Cruise Mach No.

0.80 0.80 advanced for inclusion in an early 1990's airplane. The advanced high lift system, consisting of a variable camber Krueger leading edge flap and a translating two segment trailing edge flap, will provide improved CL max and lift/drag. These improvements permit reductions in wing and engine size, and reduce approach noise.

Composite materials should be ready for application in the next; generation of transport aircraft, and would be used in such areas as doors.

control surfaces, floor beams, fairings, and landing gear Design, fabrication and repair techniques should have advanced by the early 1990's to allow applications to be expanded to provide essentially fully composite wings and empennage. The fuselage pressure should still be of metal construction. The advantages shell, however, of composites include significant weight reductions and the potential of reducing airplane price.

the stability In area of advanced controls, a longitudinal augmentation system would be ,incorporated to reduce empennage area and trim drag. Active controls would also be used for- wing load alleviation.

;' ^ features would include digital avionics, reduced Advanced systems advanced APD, advanced cockpit bleed requirement air conditioning, displays, and flight performance management systems.

5.1.2.3 Aircraft Design Point Selection airplanes, was set the 1.3g buffet Wing area, common to the two by cruise altitude requirement margin at the 9450 oo (31,000 ft) initial .

` of the intercontinental aizcraft. Throat luudio8a for both airplanes were determined by the design takeoff field requirements.

arrangement drawings for the two Douglas airplanes are General presented in Figure lO and Figure ll.

^ Lockheed Aircraft and Mission Selection 5.1.3 5.1.3.1Market Considerationsand Design i ` projections to the year 2000 to establish total Lockheed used market world-traffic demand (Figure 12), Runge requirements for both domestic intercontinental airplane designs were established from studies of and ' Combining the world traffic forecast traffic distribution patterns.

distribution provided an estimate of the number of with traffic aircraft that would be required to accu^^o6ate the market (Figure 13 ^ tbo a^rcraft design range was established as 5600 km . From these data, and km (6508 o.zoi. ) for the (3000 n.mi. ) for domestic mission 12r000 encompass all mission. These design ranges iotercontioento!

total range traffic and percent of the Ipug domestic routes 93 .

projected for -he year 2000.

domestic and intercontinental The payload capability for both the all tourist oaiaai000 was selected as 500 passengers in u nine abreast, The was based on considerations of seat mile configuration. choice frequency of congestion, scheduling flexibility, costs, airport service, and number of aircraft required.

` | ' speed of Mar ket projections and airline preference indicated a cruise ' range mission. Previous studies Mach 0.85, especially for the longer showed that in a high fuel cost environment the by Lockheed, however, lowest o9eratioQ costs and optimum fuel utilization are attained at a with the fuel conservation cruise speed of Mach 0.8. In keeping .

aspects of Energy Efficient Engine program, a urulao speed of Mach 0.8 ^ was oelected. A aoouuary of the mission and aircraft design criteria is - presented in Table 13.

^ chosen for the A'tbree-engioe configuration similar to the L1011 was ne s was domestic aizplane. J\ configuration with tour wing-mounted- engi W ON 1243 FTI 4221 FTI.

674M1

1 16, f 11

_ _

_ 1R f>^t -- VOL `f III n ihdI on w ^- W IQa — Figure 10 Douglas Domestic Airplane -- Both Douglas airplanes, the intercontinental as well as the domestic, are based on a DC10 trijet derivative with a stretched fuselage and an all new wing.

(245 FT) 74.7 M ti 1r (221 FT) # 67A (y 18.6M.

_ FT) - (lsl i7 na 1mn _ n.n n c ft DC70 13049

Figure 11 Douglas Intercontinental Airplane -- Although the

domestic and intercontinental airplanes have different w thrust requirements, they are externally similar, having v the same wing, fuselage, and empennage.

k 1990-2000 1975-1990 AVG. ANNUAL AVG,ANNUAL GROWTH RATE 1990 GROWTH RATE 2000 1975• MARKET 4,01; 4.9% 51,777 16,986 34,979 NO. AMERICA-EUROPE 6.0 15,864 9.5 28,410 4,071 EUROPE-ASIA/OCEANIA 28,850 7.0 10.2 3,425 14,668 NO. AMERICA-ASIA/OCEANIA 11.1 22,078 7.0 2,328 11,222 EUROPE-AFRICA 7.0 11.0 15,787 1,644 8,025 EUROPE-SO. AMERICA 12,398 7.0 10.0 1,506 6,303 NO. AMERICA-LATIN/SO. AMERICA 7.7% 159,394 29,960 91,061 GROUPTOTAL ORIGINAL PAGE , IS • LOCKHEED ESTIMATE OF POOR QUALITY Figure 12 Lockheed Traffic Forecast -- Lockheed made projections major longhaul of one-wa y daily passenger demand on markets (over 4800 km).

AIRCRAFT SEATING CAPACITY 200 300 500 600 MILEAGE BLOCK 400 KIM NMI 5560-7410 (3,000 - 4,000) 448 298 222 176 146 7410-9260 (4,000 - 5,000) 395 259 195 152 126 374 246 186 145 120 9260-11110 (5,000 - 6,000) 245 161 93 78 11110 - 12970 (6,000 - 7,000) 119 26 15 7 7 6 12970 - 14820 (7,000 - 8,000) 6 5 14820 • 16670 (8,000 - 9,000) 21 13 8 16670 - 18520 (9,000 - 10,000) 27 17 13 10 18520 - 20370 (10,000 - 11,000) 20 13 10 8 7 1,556 1,022 760 597 496 TOTAL Figure 13 Lockheed Total Long-Haul Aircraft Requirements in Year 2000 -- Lockheed's projected requirements are based on daily service and a sixty percent load factor.

TABLE 13 LOCKHEED AIRCRAFT :FISSION AND SIZING CRITERIA Domestic Intercontinental Airplane - Airplane 12,050 (6500) Design Range, km (n.mi.) 5560 (3000) 500 Passengers 0/100 split 0.80 Cruise Mach No. 0.80 2900 (9500) Takeoff Field Length, m (ft) 2130 (7000) 69 (135) Approach Velocity, m/sec (kts) 69 (135) Initial Cruise Altitude, m (ft) 10,670 (33,000) 10,670 (33,000) Typical mission for economic evaluation: Range, km (n.mi.) 2590 (1400) 5560 (3000) Passengers (55% load factor) 275 275 0.80 Cruise Mach No. 0.80 chosen for the intercontinental airplane. Passenger seating is nine abreast throughout, with a fuselage diameter similar to the L1011 and the length stretched to accommodate the additional passengers.

5.1.3.2 Advanced Technology Features The levels of advanced technology appropriate for incorporation into in the following paragraphs.

the 1990's airframe design are discussed The primary advanced aerodynamic technology feature incorporated was a high aspect ratio supercritical wing. Figure 14 shows a comparison of the refined supercritical airfoil used in this wing and an airfoil of a current L1011 wing.

An advancedactive controls system was incorporated in the aircraft design, providing load relief and relaxed static stability. This system is currently under development for the L1011. Wing load relief is accomplished by means of computer-controlled active ailerons which redistribute wing loadings, resulting in reduced bending moments and, wing and body structural weights. Relaxation of static hence, reduced The effects of stability results in a smaller horizontal tail size.

shown in these active controls on the aircraft configurations are Table 14.

ADVANCED SUPERCRITICAL TODAY (1-1011 TYPE) SUPERCRITICAL AIRFOILS FEATURE: MORE ROUNDED NOSE MORE CAMBERED TRAILING EDGE PERMIT: HIGHER CRUISE SPEEDS REDUCED WING SWEEP THICKER AIRFOILS Figure 14 Lockheed Comparison of Refined Supercritical Wing of Study Airplane and of Current Technology wing of L1011 Airplane -- The primary advanced aerodynamic technology feature incorporated by Lockheed is a high aspect ratio supercritical wing.

TABLE 14 EFFECTS OF ADVANCED ACTIVE CONTROL SYSTEM ON LOCKHEED AIRCRAFT Load Relief Reduction (°G) Wing Weight 5.5 Body Weight 1.0 Relaxed Stability Tail Size 28 Advanced composites are used for the internal and external secondary structures and for a significant portion of the primary structure. The specific applications are .y External Secondary Structure Flaps, slats, spoilers, gear doors Internal Secondary Structure Floor supports, beams, posts, dividers, doors fuel tank baffles Primary Structure Vertical fin, horizontal stabilizers, wing, fuselage, engine nacelle The effect of composite structure on aircraft empty weight is shown in Figure 15.

% WEIGHT REDUCTION COMPONENT WING 23 TAIL 20° BODY 7 LANDING GEAR 4 NACELLES 19 AIR INDUCTION 19 SURFACE CONTROLS 5 FURNISHINGS 0 DOMESTIC INTERCONTINENTAL 8.7% TOTAL REDUCTION IN MANUFACTURING EMPTY WEIGHT 9.2% Figure 15 Effect of Composite Structure on Empty Weight of reduced the Lockheed Aircraft -- Advanced Composites empty weight of the domestic airplane by 8.7 percent and the intercontinental airplane by 9.2 percent.

Aircraft Design Point Selection 5.1.3.3 The design points for the domestic and intercontinental airplanes were established by means of parametric studies based on the Lockheed Asset minimize Synthesis Program. Wing and thrust loadings were chosen to direct operating cost and mission fuel consumption. Takeoff distance was the limiting factor for both. of the Energy Efficient Engine powered airplanes and for the JT9D= 7A powered domestic airplane; the JT9D-7A powered intercontinental

cruise altitude was limiting for

airplane._ Lockheed's domestic and intercontinental airplane designs are shown in Figure 16 and Figure 17.

IN METERS (FEET) DIMENSIONS

T+--- i~

12.20 (40.02) 19.1024 (62.672 Figure 16 Lockheed Domestic Airplane -- Lockheed chose a three-engine configuration similar to the L1011 for its domestic airplane.

5.1.4 Pratt and Whitney Aircraft Study Airplane and Mission Selection 5.1.4.1 Market Considerations and Design Constraints Studies conducted by P&WA in connection with the Energy Efficient Engine Preliminary Design and Integration Studies (Reference 2) indicated that there should be a very substantial market for large, wide-bodied aircraft in the 1990's. The existing first-generation wide-body transports (747, DC10, L1011) will have been in production for 20 years by the early 1990's. Traditionally, successful aircraft are replaced by newer designs at approximately 20 year intervals. The large, wide-body application, therefore, appears to be a natural one in which to introduce the Energy Efficient Engine.

I F DIMENSIONS IN METERS (FEET) I 66.63 (218.6) t uur u^ ala

b

I -^

11.915 (39.09) +^

r

^b

Lockheed Intercontinental Airplane -- Lockheed chose a Figure 17 w for its wing-mounted-engine configuration four intercontinental airplane.

Our studies also showed a very large future market for smaller, shorter range airplanes (200 - 250 passengers, less than 5560 km (3000 n.mi.) range), but these airplanes, of which the 767 is a prime example, are scheduled to enter service in the early 1980's. Since these. aircraft will not be ready for replacement in the 1990's, they are not considered to represent a practical first application for the Energy Efficient Engine. Later advanced versions of these aircraft will of course, use engines with Energy Efficient Engine technology.

When replacement aircraft have entered the market, they have been consistently larger than their predecessors, reflecting the natural growth in market demand. Based on this. trend, two aircraft, with passenger capacities larger than their present day counterparts, were chosen for the study. The first, a long range, four-engine transport with a nominal capacity of 510 passengers, is envisioned as a 747 (nominal 385 passengers) replacement. The second, a medium range, three-engine transport with a nominal capacity of 440 passengers, is designed for the DC10/L1011 market.

Both airplanes have fuselage widths similar to the Boeing 747. The passenger capacities assume nine abreast seating in tourist and six abreast in first class. The medium range domestic airplane has 15/85% first class/tourist split, while the intercontinental range airplane has a 10/90% split.

A design cruise speed of Mach 0.8 was chosen as the best compromise between minimizing operating costs and conserving fuel. A summary of the mission and aircraft design criteria is presented in Table 15.

5.1.4.2 Design Features Both airplane designs incorporate a number of advanced technology features. The chief aerodynamic design feature is an advanced, high aspect ratio, supercritical wing. Use of a supercritical airfoi allows wing thickness to be increased, which in turn reduces the wing weight penalty associated with increased aspect ratio. The wing al features an advanced leading and trailing edge flap system for improved low speed performance.

The structure weights of both airplanes reflect the assumption that composite materials will be used extensively in 1990's airplane designs. Composite materials are assumed- for primary and secondary structures.

Active controls are used in the designs to reduce empennage area and to reduce wing loads.

TABLE 15 PRATT & WHITNEY AIRCRAFT MISSION AND SIZING CRITERIA Intercontinental Domestic Airplane Airplane 5560 (3000) 10190 (5500) Design Range, km (n.mi.)

440 510 Passengers (15/85 - 10/90 split) 0.80 0.80 Cruise Mach No.

2440 (8000) 3350 (11000) Takeoff Field Length, m (ft) 69(135) 69 (135) Approach Speed, m/sec (kts) 10670 (35000) 10060 (33000) Initial Cruise Altitude, m (ft) Typical mission for economic evaluation: 3700 (2000) Range, km (n.mi.) 1300 (700) Passengers (55% load factor) 242 0.80 Cruise Mach No. 0.80 5.1.4.3 Aircraft Design Point Selection The design points for domestic and intercontinental airplanes were based on parametric studies conducted during an earlier NASA study, "Turbofan Engines Designed for Low Energy Consumption" (NAS3-19132).

Initial cruise altitude determined the engine size required for both JT9D-7A and Energy Efficient Engine powered airplanes, domestic and intercontinental.

5.1.5 Study Aircraft Summary Summaries of the domestic and intercontinental airplane, configurations are presented in Table 16 and Table 17 for each of the three airframe manufacturers and Pratt & Whitney Aircraft.

5.2 ENGINE INTEGRATION CONSIDERATIONS 5.2. 1 Installation Geometry Ground Rules The three airframe manufacturers (Boeing, Douglas, Lockheed) each developed installation ground rules for mounting a mixed-flow, Tong-duct engine on their advanced airplanes. These ground rules incorporate each airframe manufacturer's best compromise among a number of such conflicting considerations as interference aerodynamics, wing flutter, jet wake impingement, pylon-weight, and ground clearance: TABLE 16 DOMESTIC AIRPLANE DEFINITIONS SUMMARY P&WA BOEING DOUGLAS LOCKHEED 'irRIJET TRIJET. TRIJET TYPE TWIN 1990'S 19901S IN SERVICE DATE 1990'S 19901S 5560 (3000) 5560 (3000) 3700 (2000) 5560 (3000) DESIGN RANGE — km (NM) 500 440 196 458 PASSENGERS 0.8 0.8 0.8 0.8 CRUISE SPEED MACH NO, 2130 (7000) 2440 (8000) 1830 (6000) 2440 (8000) FIELD LENGTH m (FT) 10060 (33000) 10670 (35000) '10670 (35000) 10670 (35000) CRUISE ALTITUDE — rn (FT) kgLBM 569,4 (116.6) 439.4 (90.0) 522.4 (107.0) 560.5 (114.8) WING LOADING ^ 2 Z^ ( M FT 10 12 10,24 9.83 ASPECT RATIO 1850 (1000) 2590 (1400) 1300 (700) TYPICAL RANGE — km (NM) 1850 1 1000) 55 55 % 55 60 TYPICAL PAYLOAD — TABLE 17 INTERCONTINENTAL AIRPLANE DEFINITIONS SUMMARY P&WA DOUGLAS LOCKHEED TRIJET QUADJET QUADJET TYPE 19901S 19901S 19901S IN SERVICE DATE 10190 (5500) 12040 (6500) DESIGN RANGE — km (NM) 10190 (5500) 500 510 PASSENGERS 0.8 0.8 MACH NO. 0.8 CRUISE SPEED 3350 (11000) 3050(10000) m (FT) 3350 (11000) FIELD LENGTH 10060 (33000) 10360 (34000) m (FT) 9450 (31000) CRUISE ALTITUDE k9 LBM .

644.6 (132.0) 673.8 (138.0) WING LOADING ^ m 670.9 (137.4) 2 FT2) 9.83 10 ASPECT RATIO 5560 (3000) 3700 (2000) 2780 0500) TYPICAL RANGE — km IN M) 55 55 TYPICAL PAYLOAD % wing-engine placements envisaged by each of the A composite of the Efficient Engine is presented in airframe manufacturers for the Energy The variations in engine location from company to company Figure 18.

of factors, the primary one being the different are due 4o -a number.

installation considerations (d-rag, weight, relati;-aahips among the Interference drag and wing flutter, etc.) for different airplanes.

Y _ ^ B C A B WING CHORD LENGTH 0.1 C BOEING 0.4 C WING CHORD LENGTH 0,290 0.13C DOUGLAS WING CHORD LENGTH 0.17C LOCKHEED 0.17C Figure 18 Wing-Engine Placement Summary -- The variations in engine location are due to differences in such considerations as drag, weight, and flutter for the different airplanes.

flutter are especially sensitive to individual design details.

Airplane size and configuration are also factors: a smaller airplane has less wing-ground clearance, requiring the engine to be mounted closer to the wing. Comparing the 200-passenger Boeing twinjet installation with the 400/500-passenger Douglas and Lockheed trijet installations illustrates this point.

Since Douglas and Lockheed each had trijet airplanes, they also evaluated tail installations. These installations followed DC10 and L1011 practice, respectively, as can be seen in their configuration drawings (Figures 10, 11, 16, 17).

5.2.2 Engine and Airframe Accessory Location Assessments Each of the airframe manufacturers assisted in the design of the nacelle during the Energy Efficient Engine Preliminary Design task ('described in detail in Propulsion System Preliminary Design and Analysis Report, Reference 1). As part of this study, the airplane companies each performed a qualitative analysis of the merits of various accessory locations. These analyses are summarized in Table 18.

Results of a Pratt & Whitney Aircraft study on accessory locations, including some special concerns relating to the shroudless- fan design, are also shown on this table. Preliminary design studies have TABLE 18 ACCESSORY LOCATION STUDY SUMMARY ,.C:.FZC;s " .rY r ULL DUTY GEARBOX i.

CORE MOUNTED IRFPAME ENGINE ACCESSORIES SPLIT GEARBOX AIRFRAME ACCESSORIES PYLON MOUNTED AIRFRAME ENGINE ACCESSSORIES SPLIT GEARBOX-FAN CASE MOUNTED v ENGINE ACCESSORI— ENGINE ACCESSORIES FAN CASE MoUNTEO SPLIT GEARBOX AIRFRAME ACCESSORIES.

PYLON MOUNTED—^7 AIRFRAME AC RtE R E ESSO c :I r AC.ESSOFIE;^^ L ^...'r. : W- T ORIGINAL PAGE IS OF POOR QUALITY Type Accts Bury typo rs cat ion S.-6 L.—..t, Douglas Cn—.r,:s ed Comments PL,.A C .mments 6 I.nf.3[I^n. L..'.II •: Fui; Duty on c Pass..bt Fuel Spilt .a Perlarrud Wcatinn L ?rob,em o Goad Access.bt!Lty a Pan 7 Problem Fan rase dut tom o 'end Acceas.bilicy u ACceptab:c Location Poor Accessibility.

Sp Lic Spit[ Aeresao: iev c Very S Good Aerc Shape o Fan. 2E Prool.m LL - a Unacceptable co Some ,Nacelle Airlines o Unacceptable o Airicame - Pylon Requires Addit,.aal o -gin - Fan Case '.ark Stands Bottom o. Good A— Soape Aece?tabie o Split. Accessories o Very Poor Accessibility Ilacel Le III Split - Unacceptable cc Some Airlines o Requires Additional ALriraore -Pylon o Unacceptable Work stands o Possibi iicy Engine - Co..

o Hoc Environment o Poor Accessib:lity o Hot Environment. o Redu c es ttteasibility IV Full. Duty on Gore - Good Performance to Engine u o Acceptable o Hot Environment o Poor Accessibility o .Acceptable o.. Good Environment o Good Accessibility o Requires Tailored E- c V Split - Guide Vanes in Fan Case o Good Accessibility o Possible Drag and Weight J o Acceptable A- Lramo - Fan Penafcics Case May be AcceptablL o Possible Drag and o - Eng:ne - Fan base {1200 .Apart) Weight Penalties o Split Accessories Unacceptable to some Airlines - y indicated that the fan may be sensitive to 2E (twice per revolution) disturbances, such as would be caused by two flow blockages 180 degrees apart and located a short distance downstream of the fan.

Since the engine mounting system requires a wide (about 20 cm) strut at the top (or 12:00 position), it is desirable to avoid another wide strut, as would be required by a towershaft, at the bottom (or 6:00 position). Hence, accessory locations that require a towershaft through the fan duct at the 6:00 position are indicated as potential fan problems in Tabl . 2 18.

These assessments indicate at least three accessory locations are acceptable to one or more of the airplane companies and also satisfy the fan 2E considerations: (III) pylon-mounted airframe accessories with core-mounted engine accessories, (IV) full-duty core-mounted accessories, (V) airframe and engine accessories :counted 120 degrees apart on fan case. The full-duty core-mounted accessory configuration was chosen for use in this study in order to nave a common configuration for comparison purposes. In practice, however, accessory location would be determined by each airframe manufacturer to match each particular engine application.

5.2.3 Reverse Thrust Requirements Reverse thrust level and directivity requirements are dependent on the airplane configuration. Figure 19 shows the Douglas trijet requirements (Lockheed trijet requirement is similar), and Figure 20 shows Boeing twinjet requirements. Reverse thrust directivity is necessary to prevent reingestion of engine exhaust, to avoid interference with control surfaces, and to prevent impingement of exhaust on airplane. The nacelle was designed with 12 replaceable fan Figure 19 Douglas Trijet (Typical) Reverse Flow, Directivity Requirements (Lockheed Trijet Requirements are similar) -- Reverse thrust level and directivity requirements are dependent on airplane configuration.

I PREVENT IMPINGEMENT Figure 20 Boeing Twinjet Reverse Flow Directivity Requirements -- Energy Efficient Engine Nacelle was designed with twelve sections, replaceable fan duct reverser cascade permitting flow to be matched to the application.

duct reverser cascade sections in order to allow matching the reverse flow to the application.

The required levels of reverse thrust are more difficult to predict, since they can vary according to airline practice. Lockheed and Douglas have indicated that a reverse thrust level of about 35 to 40 percent of forward thrust would be appropriate. Figure 21 shows that the reverse thrust capability of the Energy Efficient Engine exceeds 35 percent to speeds less than 26 m/sec (50 knots). This performance is achieved without reversing the primary stream and without overspeeding the low pressure rotor or violating low pressure compressor surge margin requirements. Since the Energy Efficient Engine is a mixed — flow engine and the reverse blocker doors are upstream of the mixer, the effective nozzle area seen by the primary stream in reverse mode is uncertain. The performance shown in Figure 22 represents the most pessimistic case, where the primary flow fills

only the primary mixer area and therefore actually provides some

forward thrust. A ten percent leakage of duct flow past the blocker

doors, was assumed for this figure.

N S LL LL W Y Q Q O LL N J N 40 60 80 0 20 M /S 100 150 200 0 50 KNOTS FLIGHT VELOCITY Figure 21 STF505M-7D Thrust Std + Reverse Capability (Sea Level 10% Leakage) -- Reverse thrust capability 14 0 C, exceeds 35 percent down to speeds close to twenty meters per second.

5.2.4 Customer Bleed and Horsepower Extraction Bleed and horsepower extraction requirements were evaluated by each of the three airplane companies and by Pratt & Whitney Aircraft. Energy Efficient Engine and the JT9D-7A engine data packs provided to each airplane company included the bleed schedule shown in Figure 22 and a power extraction of 113 kW (151 hp) at all conditions. Bleed and horsepower influence coefficients for thrust and specific fuel consumption were also included in the data packs, permitting the engine performance to be modified to reflect specified requirements.

The bleed requirements assessed by each airplane company are shown in Figure 23. Typical cruise power extraction requirements were assessed as 67 kW (90 hp) per engine by Boeing and 79 kW (106 hp) by Douglas.

For the purposes of this study, Lockheed chose to use the bleed and horsepower levels provided by Pratt & Whitney Aircraft, and Boeing and Douglas chose- to modify the engine performance to -reflect the schedules shown in Figure 23.

2.0 1,5 U Y W W N 1.0 LIJ J 2 m CG 1 0.5.0 12 14 16 8 10 4 6 1000 m 1 } 30 40 10 20 1000 ft ALTITUDE Figure 22 Customer Bleed Schedule, STF505M-7 Data Pack -- Bleed and power extraction influence coefficients for thrust and specific fuel consumption were included in the data pack to permit the performance to be modified for specific requirements.

5.3 AIRPLANE-ENGINE PERFORMANCE COMPARISON The three airplane companies and Pratt & Whitney Aircraft evaluated the mission performance of both the Energy Efficient Engine and the JT9D-7A reference engine. The evaluation was based on the airplanes and missions described in Section 5.1 and the Pratt & Whitney Aircraft provided isolated nacelle engine data, modified and installed as described in Section 5.2.

5.3.1 Airplane Performance Evaluation The following procedure is reasonably typical of that used by Pratt & Whitney Aircraft and the airplane companies to evaluate airplane performance for this study. The airplanes and missions are defined first. Next, the aerodynamic and weight methods, including scaling functions, are chosen consistent with the technology levels assumed for each airplane--these methods are unique to each company. The engine data provided by Pratt & Whitney Aircraft is then translated to a form suitable to each company's mission analysis program, including changes to customer bleed and/or horsepower extraction levels. Weight and aerodynamic penalties unique to each engine, such as interference 52:.

2.0 1.5 U 3 DOUGLAS m Uj Cn uj Uj _j 1.0, ^ ► 2 DATA PACK BOEING 8 14 16 4 6 10 12 1 050 2 1000 m I I 1 I 1 30 40 50 0 10 20 1000 ft ALTITUDE Customer Bleed Requirements (anti-icing not Figure 23 Normal included) -- Bleed and power requirements assessed by the airplane companies were similar to those assumed by Pratt & Whitney Aircraft.

drag and pylon weight are assessed, and engine and nacelle weights are is assumed and airplane included. Next, a takeoff gross weight (TOGW) component and engine sizes and weights determined from the sizing minimum cruise altitude, wing conditions (e.g., takeoff field length, loading). Airplane operating empty weight OEW can then be obtained and available fuel load determined by subtracting the OEW and payload from the TOGW.

design mission is The ability of this size airplane to perform the through the simulation of the then assessed: The airplane is "flown" design mission profile in order to determine if there is sufficient fuel, including reserves, at the assumed TOGW to fly the design range.

If the range that can be flown with the available fuel is greater or less than the design range, a new TOGW assumed, the airplane and is engine are resized, new OEW and fuel available are calculated, and the mission is "reflown". This process is repeated until the range flown with available fuel (minus reserves) exactly matches the design range, determining the design TOGW and engine size, OEW, and airplane component weights and sizes.

Once sized for the design mission, the airplane can be - "flown" on a in typical mission. The typical mission is of primary importance assessing the merits of an engine or airplane because it represents the average mission an airplane of this passenger capacity and range would fly in actual airline operation. Thus, the typical mission performance (fuel burned, operating costs) of the airplane more closely simulates the experience of an airline operating a fleet of these airplanes than does design mission performance. Definitions of typical missions for each airplane were presented in Section 5.1.

If any of the airplane design parameters, such as wing loading (TOGW/wing area) or thrust loading (total thrust/TOGW), are to be optimized, the process described above is repeated many times, and minimums of the chosen figure-of-merit (usually DOC or fuel burned) are determined. Each of the airplanes used by the airplane companies in this evaluation represent minimum fuel burned and/or DOC designs.

5.3.2 Airplane Peformance Results Comparisons of the design mission takeoff gross weights for Energy Efficient Engine (STF505M-7D) and JT9D reference* engine powered airplanes are shown in Figure 24, Figure 25, Figure 26 and Figure 27.

In all cases the reduction in total fuel (mission plus reserves) is the primary contributor to the TOGW advantage of the Energy Efficient Engine. Thus aircraft that have small design fuel fractions (total fuel/TOGW) tend to demonstrate less TOGW advantage for the Energy Efficient Engine.

The Pratt & Whitney Aircraft domestic and intercontinental study airplanes are used in Table 19 to further illust:^ate the effects of the Energy Efficient Engine on aircraft weight. The improved performance of this engine causes reductions in most structural component weights.

Figure 28 shows the takeoff thrust size required by each of the Energy Efficient Engine powered aircraft. For the PS-AIE evaluations, the engine is treated as a "rubber" engine: one scalable to any size. The takeoff thrust rating of the base size STF505M-7D is 182.8 kN (41,000 lbf). Except for the Douglas intercontinental airplane, thrust requirements are clustered in a band from 164 kN to 182 kN (32,000 lbf to 41,000 lbf)--the Douglas airplane is the only trijet among the intercontinental airplanes, accounting for its much larger thrust.

Because of the advanced technologies included in both the airplanes and their flight propulsion systems, the levels of thrust required by these advanced airplanes are considerably lower than that of current airplanes of comparable payload and range.

*Lockheed, Boeing, and P&WA used the JT9D-7A engine with the 747-200 type :nacelle as reference engine, Douglas used the JT9D-20 with DC10-40 type nacelle. The JT9D-2O is essentially the same as the -7A, except for accessory location.

DOMESTIC JT9D-7A STF505M-7D c^ z to Z5 so o.

Q o.

L o' w!

Q,

z

W V Cr W a.

Figure 24 Boeing Airplane Weight Breakdown -- Boeing twinjet has lowest fuel fraction of the study airplanes.

The mission Efficient Engine fuel-burned advantage of the Energy shown in Figure 29 for compared with the JT9D-7A reference engine is missions. The results design missions and in Figure 30 for typical a mission correlate well with design fuel fraction. Table 20 shows segment breakdown of the Energy Efficient Engine fuel burned savings for P&WA study aircraft on typical missions. The large advantage of Energy Efficient Engine over the reference engine at off design flight conditions is evident in this Table. Overall, fuel-burned reductions vary from 13.5 percent for the Boeing twinjet on a typical mission to over 18 percent for the Lockheed and Pratt & Whitney Aircraft intercontinental airplanes and the Douglas domestic airplane on design missions. Average fuel-burned reduction is 16.6 percent on typical missions and 17.3 percent on design missions.

55:: INTERNATIONAL DOMESTIC JT9D-20 JT9D-20 STF505 M-71) STF505 M-71) TOTAL TOTAL r FUEL TOTAL FUEL FUEL C7 TOTAL I— ENGINES FUEL ENGINES N W 60 ENGINES 0 ENGINES N O.E.W.

O.E.W.

LESS LESS 0.

O.E.W.

O: E.W.

ENGINES ENGINES e 40 CW7 LESS LESS Q ENGINES r ENGINES W U cc W CL PAYLOAD PAYLOAD PAYLOAD PAYLOAD Figure 25 Douglas Airplane Weight Breakdown -- Douglas used the JT9D-20 engine with the DCIO-40 nacelle as the reference engine/nacelle. the JT9D-20 is essentially the same as the JT9D-7A, differing in accessory location and nacelle design.

DOMESTIC INTERNATIONAL a JT9D-7A JT9D-7A STF 505 M-713 STF505 0.9-7D TOTAL TOTAL FUEL FUEL TOTAL TOTAL FUEL ENGINES FUEL ENGINES F- Z a ENGINES n ENGINES O.E.W, O.E.W.

LL LESS LESS Uj ENGINE.- ENGINES O.E.W.

O.E.W.

a - F LESS LESS W U ^ ENGINES ENGINES W PAYLOAD PAYLOAD PAYLOAD PAYLOAD Figure 26 Lockheed Airplane Weight Breakdoxm -- The 12040 km design range intercontinental airplane has a much larger v fuel fraction than the 5560 km domestic airplane.

LJ7 co + 111 INT E RNATIONAL DO MESTIC STF505 M-713 STF505 M-7D JT9D-7A JT9D-7A TOTAL TOTAL TOTAL FUEL ^ FUEL FUEL TOTAL O O FUEL ENGINES ENGINES W 60 0 ENGINES ENGINES

a

LL O.E.W.

O.E.W.

O w LESS LESS O.E.W.

ENGINES O.E.W.

ENGINES LESS LESS - W U ENGINES ENGINES cc W <L PAYLOAD PAYLOAD PAYLOAD PAYLOAD Figure 27 P&WA Aircraft Weight Breakdown --Weight breakdowns of P&WA .airplanes are similar to those of Douglas airplanes, especial, in the case of the domestic trijet.

',CABLE 19 WEIGHT BREAKDOWNS P&WA AIRPLANE i DOMESTI C INTERCONTINENTAL C • 1-7D JT9D-7A STF505 11-7D J 9D-7A STF505 (lbm) kg lbm) kg ^(Ibm) kg (lbm) kg 48017 42866 (94503) Wing 34024 (75069) 31535 (69523) (105859) 28648- (63157) 28543 (62925) 31609 (69686) 31552 (69560) Fuselage 3309 (72916) 3103 (6840) 4354 (9599) 3921 (8645) Empennage 12159 (26806) 15852 (34947) 15149 (33398) Alighting Gear 12534 (27632') Propulsion 19594 (43198) 17522 (3862&) 26490 (58399) 23372 (51526) Fixed Equip. & 28499 (62828) 28206 •(62182) 34070 (75111) 33514 (73886) systems 160392 (353601) 150376 (331518) MEW 126608 (279120) 121067 (266904) 12618 (27818) 17000 (37478) 16870 (37192) Std. & Oper. 12662 (27915) Items OEW 139270 (307035) 133685 (294722) 177392 (391079) 167246 (368710) (90200) (108375) Passengers & 40914 (90200) 40914 49159 (108375) 49159 Baggage 55784 (275852) Fuel 67014 (147739) (122981) 153243 (337839) 125126 Ln TOGW 247199 (544974) 230383 (507903) 379794 (837293) 341530 (752937) DOMESTIC INTERCONTINENTAL 5: 2 Z LU _j LL.

LL 0 co Z LU Z 30) BOEING DOUGLAS LOCKHEED P&WA DOUGLAS LOCKHEED P&VVA *STATUS FPS Figure 28 Thrust Size Required for Each Energy Efficient Engine Powered Aircraft -- For the evaluation, the engine was assumed scaleable to any size.

10200-12050 (5500 - 6500) (3000), P & VVA DESIGN RANGE DAC km - 3700 / .^' DAC- 0 (NM - 2000) LCC Uj LCC x O / P&WA.

BCAC a/ /• !

w 0 ^ W LL FOR DESIGN MISSIONS 10 , -^ 0.15 0.20 0.25 0.30 0.35 0.40 DESIGN FUEL WEIGHT DESIGN FUEL WEIGHT FRACTION TOGW Figure 29 Fuel Savings Relative to the JT9D-7A Powered Aircraft (Design Mission) -- The advantage in fuel savings ter-ds to peak and then level off because the STF505M-7D engine has its biggest advantage during climb and descent.

10200-12v50 20 r (5500 - 6500) (3000} DESIGN RANGE DA ^ C _^.^ O ^" km - 3700 .L^^ P&WA 0 P&WA Q LCC W (NM - 2000) ^Q ^.

/ ^• DAC ^ Q m 15 LCC JV W ^ LL LL W FOR TYPICAL MISSIONS o ^ BOAC 0.35 0.40, 0.25 0.30 0.15 DESIGN FUEL WEIGHT DESIGN FUEL WEIGHT FRACTION TOGW Figure 30 Fuel Savings Relative to the JT9D -7A Powered Aircraft (Typical Mission) -- The advantage of the STF505M-7D tends to peak and then level off because- its biggest advantage is during climb and descent, not during cruise. As a result the effects of increased fuel fraction is somewhat offset at longer ranges.

TABLE 20 ENERGY EFFICIENT ENGINE FUEL BURNED ADVANTAGE BREAKDOWN 2000 n.mi. Airplane Mission - P&WA Intercontinental JT9D-7A STF505 ti1-7D Fuel /km Fuel/km Distance Fuel Fuel/Km Distance Fuel km kg kg/km km kg kg/km Taxi & - 1954 - - 1280 - Takeoff Climb 282 6512 23.09 317 5638 17.79 -23.0 Cruise 3213 30597 9.52 3197 25308 7.92 -16.8 Descent 209 751 3.59 190 431 2.27 -36.8 Total 3704 39814 10.75 3704 32656 8.82 -18.0 Mission Reserves - 13590 - - 11159 - 700 n.mi. Mission - P&WA Domestic Airplane JT9D-7A STF505 M7D %6 - i & - 1,225 - - - Climb 16.84 -22.6 295 4967 351 4578 13.04 Cruise 796 5938 7.46 757 4745 6.27 -16.0 -38.8 Descent 205 550 2.68 188 308 1.64 -17.6 Total 1296 12679 9.78 1296 10448 8.06 Mission Reserves - 10685' - - ` 9001 6.0 ECONOMIC EVALUATION Airline operating economics for the Energy Efficient Engine and the JT9D-7A reference engine were determined by combining the results of the airplane performance evaluation provided by the airframe manufacturers with the engine price and maintenance cost estimated by Pratt & Whitney Aircraft. The economic advantages of the Energy Efficient Engine were then determined by comparing the results obtained for the two engines. The NASA-approved economic model was used for this effort.

6.1 ECONOMIC MODEL DESCRIPTION The Air Transport Association's (ATA) operating cost method, its formulas modified to reflect current airplane technology and airline environments, was used for the economic evaluation. The ATA method was originally published in 1967. The formula modifications were based on a 1977 Boeing update.

The elements of which direct operating cost (DOC) and indirect operating cost (IOC) are composed are presented in Table 21; the parameters controlling these elements are also identified. DOC includes most elements of operating cost directly influenced by airplane and/or engine performance. All other airline operating costs are included in IOC. Important assumptions upon which the economic evaluation was based are shown in Table 22.

6.1.1 Direct O peratin g Cost Model This section discusses the effect of each DOC element on the Energy Efficient Engine/JT9D-7A comparison. The Pratt & Whitney Aircraft domestic trijet study airplane on a typical 1300 km (700 N.Mi.)

mission is used to illustrate each effect. The overall DOC of this airplane with either STF505M-7D or JT9D-7A engines is compared in Figure 31.

Flight crew cost, which includes both wages and fringe benefits, varies with airplane design speed, utilization, size of crew, and with takeoff gross weight. Since both STF505M-7D and JT9D-7A airplanes use three-man crews and have a design speed of Mach 0.8, crew size and speed does not affect the comparisons. Utilization, shown in Figure 32, is also essentially the same for both engines on the same airplane and trip distance. Crew cost is, therefore, a function of TOGW only.

Figure 33 shows the effect of TOGW on DOC; costs are per block hour and are shown as a percent of the total DOC of the JT9D-7A powered airplane.

Fuel cost reflects fuel burned on the mission (see Section 5.3) and fuel price. Fuel prices of 10.60/liter (40^/gal.) domestic and TABLE 21 AIRLINE OPERATING COST MODEL ELEMENTS ELEMENTS FUNCTION OF Direct Operating Costs Flight Crew TOGW, Speed, Utilization Fuel Block Fuel, Fuel Price Airframe Maintenance Material Airframe Weight, Flight Length Labor Airframe Weight, Flight Length Engine Maintenance Material Engine, Engine Size, Flight Length Labor Engine, Engine Size, Flight Length Maintenance Burden Airframe and Engine Maint. Labor Insurance Airplane and Engine Price, Utilization Depreciation Airplane and Engine Price, Utilization Indirect Operating Costs Ground Property & Equipment Max Landing Weight, Block Time Airplane Related Costs Max Landing Weight, No. of Seats, Block Time Passenger Related Costs No. of Passengers, Block Time Cargo Related Costs Tons of Cargo, Block Time General and Administrative Max Landing Weight, Operating Costs, Block Time TABLE 22 ECONOMIC MODEL ASSUMPTIONS • Dollars - 1977 • Flight Crew - 3 People • Fuel Price - 10.6(,-/litar (40q^/gul) Domestic, 11,91^/liter 45^./&al Intercontinental • Maintenance - Labor Bate = $9.70/br.

• Maintenance Burden - 200% of Labor Coat • Non-Revenue Flying - 2% Factor Added to Fuel and Maintenance • Ground Time - 15 min. Domestic, 20 min. International • Insurance - 0.5% Fly-Away Cost Per Year • Spares - 6% Airframe, 30% Engine • Depreciation - 15 Year Straight Line to 10% Residual Value • Utilization - (see Figure 32) 11.9,,^/liter (45q , /gal.)

international are used to represent 1985 prices expressed in 1977 dollars. Trip fuel and time used in economic calculations (referred to as block fuel and time) include standard ATA allowances and ground idle and taxi allowances derived from current airline experience (Table 22). Relative fuel costs for the STF505M-7D and JT9D-7A are shown in Figure 34. Comparing Figure 34 with other DOC component plots revealed that fuel cost was the primary difference between STF505M-7D and the JT9D-7A engines.

The airframe maintenance costs are presented in Figure 35; the costs include materials, labor, ,nd burden.

Flight length also affects airframe maintenance cost since many maintenance items such as brakes, are cycle dependent rather than time dependent (1 flight = 1 cycle).

For this comparison, flight length was assumed to be a constant, making airframe weight the onl^.y variable. Airframe weight is, in turn, a function of design TOGW, which reflects the performance-capabilities of the engine.

u JT9D-7A STF505 M-7D 100F FUEL FUEL

y

H O U Q n 0 ENGINE MAINT.

ENGINE MAINT.

U.

AIRFRAME MAINT.

W AIRFRAME MAINT.

Q' F., 40 Z W V DEPRECIATION ^ DEPRECIATION W C.

INSURANCE INSURANCE FLIGHT CREW FLIGHT CREW 0 1 Figure 31 Comparison of Direct Operating Costs -- The P&WA domestic trijet on a typical mission was used for this comparison of STF505M-7D with JT9D-7A reference engine costs.

Q 2000, W } 15001 H a 500- .6000 3000 4000 5000 1000 2000 0 0 km 2000 2500 0 500 1000 NM TRIP DISTANCE Figure 32 Utilization -- For a given airplane and trip distance,, utilization is essentially the same for both engines..

E3 JT9D-7A Q STF505 M-70 U O O J Q F Q F U.

O W Q H Z W U W CL D 100 150 200 250 300 350 400 TAKEOFF GROSS WEIGHT ~ 1000 kg Figure 33 Crew Cost (P&WA Domestic Trijet, Typical Mission) -- Crew costs, which are primarily a function of TOGW, are similar for both engines.

V O J 30 H O 20 STF505 M•70 JT9D•7A U.

O LL H W W a 0.

Figure 34 Relative Fuel Cost (P&WA Domestic Trij_et, Typical Mission) -- Fuel costs are the main difference in the direct operating costs of the two engines.

UU 30 C] JT90-7A O STF505 M•70 J Q H O Q 20 rn F- LL W 10 .r+'^ Q H W U a 0 100 110 120 130 1 80 90 AIRFRAME WEIGHT — 1000 kg Figure 35 Airframe Maintenance_ Costs (P&WA Domestic Trijet, Typical Mission) -- Costs, which are dependent on airframe weight, include materials, labor, and burden, and are similar for both engines.

Engine maintenance costs for base size (scale factor of 1.0) STF505M-7D and JT9D engines were calculated as described in Section 4.3. The engine maintenance costs shown in Figure 36 have been scaled to the engine size required to fly the design mission and adjusted to the proper flight length. Engine maintenance, like airframe maintenance, is dependent on flight length and requires adjustment to the actual flight length of each ai::plane/mission. Since both engines are evaluated for the same missions, their adjustment is the same.

Materials, labor, and burden were included in the maintenance cost.

W a 1Q ^, 0

Sao

W -^ 10 U H CC: L6 p a 0'~

_off

Figure 36 Engine Maintenance Costs (P&WA Domestic ,Trijet, Typical Mission) The STF505M-7D has a small advantage in engine maintenance costs over the JT9D-7A.

Depreciation, in dollars per block hour, is the total investment in the airplane (airframe and engine and spares) minus residual value (10% in this case) divided by.depreciation period (15 years) and hours flown per year. The airframe price equation, based on a Pratt & Whitney Aircraft correlation of present airplanes, is 0'7 /airframe weight 6 10.0 + furnishings + avionics Airframe Price = 0.5* I- 1000 1 where: 6 (furnishings) Domestic 1(0.008 * number of seats -0.284) * 10 (avionics) Airplane 1(0.0022 * number of seats +1.54) * 10 6 or International J(0.0089 * number of seats -0.31) * 10 (furnishings) Airplane 1,(0.0022 * number of seats +1.81) k 10 (avionics) The effect on depreciation of the domestic airframe price equation is shown in Figure 37, including six percent airframe spares. Base engine prices, calculated as described in Section 4-3, have been scaled to the engine size required for the design mission. The two points in the figure indicate the total effect of airframe and engine price on depreciation. The engine depreciation is for three engines and thirty percent spares.

TOTAL EFFECTS OF ENGINE & AIRFRAME PRICE U O

0 a

w J 20 EFFECT ON DEPRECIATION C7 OF DOMESTIC AIRFRAME PRICE, L a. O INCLUDING SIX PERCENT SPARE Z 1T9D-7A j n TOTAL q t W ^ 10 DEPRECIATION O STF505 M-7D ap F- 110 150 160 80 90 100 120 130 140 AIRFRAME WEIGHT 1000 kg Depreciation Costs (P&WA Domestic Trijet, Typical Figure 37 Mission) -- Depreciation, for a given airplane configuration, is a function of airframe weight and engine price.

The insurance rate used was 0.5 percent per year of fly-away price, which is airplane price without spares. Figure 38 indicates that insurance has only a small effect on the DOC comparison. The line in the figure indicates the contribution of airframe price to insurance.

The points indicate the sum of the airframe and engine insurance.

0 Q [3 JT9D-7A r a' ^ TOTALINSURANCE 1 0 Q STF 505 M-71D W'1- V " /AIRFRAME ONLY TOTAL ^ O -- i i - _ _ 1 I 0, 80 90 100 110 120 130 140 150 160 AIRFRAME WEIGHT ~ 1000 kg Figure 38 Insurance Costs (P&WA Domestic Trijet, Typical Mission) - The insurance rate used was 0.5 percent per year fly-away price.

6.1.2 IOC and ROI Models The IOC model was based on CAB data. For this model the costs are grouped into five categories: Airplane Related - 1- Aircraft handling, cabin crew, landing fees 2- Passenger Related Passenger and baggage handling, ticket sales, commissions, advertising, food 3- Cargo Related Cargo handling, insurances, sales, commissions, advertising Ground Property and Equipment 4- Depreciation and maintenance of ground property and equipment 5- General and Administrative The parameters upon which these categories are dependent are shown in Table 21. Indirect operating cost is determined primarily by speed of 70` the airplane, trip distance, and number of passengers and/or tons of cargo. Variations in engine performance and characteristics affect IOC only through landing weight, which determines landing fees and is a factor in the correlation of ground property and general and administrative costs.

The IOC components for the STF505M- 7D and JT9D-7A powered airplanes are compared in Figure 39. The Pratt & Whitney Aircraft domestic airplane was used as the basis for the comparison. The advantage of the STF505M-7D (1.6%) is primarily due to decreased landing fee costs.

(0.7% out of 1.6%), with ground properties and equipment (0.5%) and general and administrative (0.4%) accounting for the rest.

Since no cargo was assumed for this case, there were no cargo related costs. Cargo would not have changed the absolute IOC difference between engines because cargo related costs like passenger related costs are functions of the amount carried and are not influenced by engine performance or characteristics when identical missions are flown.

JT9D-7A STF505 M-7D AIRPLANE AIRPLANE RELATED RELATED H p Q ^ 60 n LL PASSENGER PASSENGER Q RELATED RELATED Z 40 W U W CL GROUNDPROP GROUND PROP + EQUIP.

+ EQUIP.

GEN 4. ADMIN GEN + ADMIN Figure 39 Indirect Operating Cost Breakdown (P&WA Domestic Trijet, Typical, Mission) -- IOC is much less sensitive to engine performance differences than_DOC is.

Return on investment (ROI) is calculated using a traditional discounted cash flow technique where the annual ROI is determined by zero present value of future cash flow benefits. Cash flow is defined as after-tax profits plus depreciation, where depreciation is a noncash expense.

Cash Flow = (Revenue - DOC - IOC -- Taxes) + Depreciation (where tax rate = 50%) As in DOC, a straight line depreciation over 15 years to a ten percent residual value was used for ROI.

Return on investment is sensitive to revenue and load factor assumptions. The following revenue functions were assumed for all airplanes: Domestic Passenger Yield: $20.88 + 0.0362$/km (0.0582 $/s.m.)

$23.42 + 0.0406$/km (0.0653 $/s.m.)

International Passenger Yield: Cargo Yield: $145.0/ton + 0.0972$/ton-km ($131.6/ton +0.142 $/ton s.m.)

Typical mission load factors (ROI is shown only for typical missions) were chosen by each airframe manufacturer--Boeing and Lockheed chose 55% and Douglas chose 60%. A 55% load factor was also used for the P&WA study airplanes.

6.2 ECONOMIC EVALUATION RESULTS The performance input from each of the three airframe manufacturers was used to compare the DOC, IOC, and ROI of the STF505M-7D with the reference JT9D-7A. The performance of the STF505M-7C powered airplanes was adjusted to reflect the STF505M-7D status characteristics, as explained in Section 3.3.

6.2.1 DOC Comparison Figure 40 shows the results of the DOC comparison for the design missions; and Figure 41, for the typical missions. Figure 41 should come closest to approximating actual airline experience. These two plots show trends similar to the fuel-burned trends in Figure 29 and Figure 30. As shown in Section 6.1.1, the primary reason for the DOC advantage of the STF505M-7D is reduced fuel consumption. In general, airplanes with higher fuel fractions tend to have a greater DOC advantage for the STF505M-7D. Company-to-company variations at similar fuel fractions are due to difference in design systems and modeling techniques; for example, the rate at which airplane structure weight increases with increasing TOGW is different in each company's airplane model. Even with these differences, the variations of DOC advantages fall within a +11 band.

10200 - 12050 km (5500 - 6500 N M ) 13000) DESIGN RANGE P&WAO Q DAC O 00.0 P&WA ----(D.

3700 km LCC p ` (2000 NM) .^^ riQ LCC

U 0 /

O 0/ D o j 3 BCAC FOR TYPICAL MISSION 0.40 0.35 0.25 0.30 0.15 0.20 DESIGN FUEL WEIGHT DESIGN FUEL WEIGHT FRACTION TOGW DOC Relative to the JT9D-7A for Figure 40 STF50.5M-7D Savings in Typical Missions -- The DOC trends for both the typical and design missions are similar to the fuel-burned trends. Average typical mission DOC savings for STF505M-7D is 7.6% well above the NASA goal of at least 5% On design missions, all airplanes with STF505-7D engines show a greater than five percent reduction in DOC. The average savings is 9.7 percent. When flown on typical missions, all STF505M-7D airplanes except the Boeing domestic twinjet show DOC savings greater than five percent (average savings = 7.6%). Comparison of Figure 42 with the other DOC pie charts (Figure 43, Figure 44, Figure 45, Figure 46, Figure 47, Figure 48) clearly show why the Boeing STF505M-7D airplane has less of a DOC advantage. The fuel cost portion of DOC for the Boeing airplane is significantly lower (29% vs. 35 to 41%) than for the other airplanes, and improved fuel consumption is the prime attribute of the STF505-7D engine. The reason for the lower fuel cost contribution to DOC in the Boeing airplane is that in smaller, shorter-range airplanes--Boeing twinjet carries 196 passengers 3700 km (2000 N.Mi.) vs. 400-500 passengers and 5600 to 12,000 km (3000 to 6500 N.Mi.) in the other airplanes--DOC tends to be dominated by costs less sensitive to engine performance (like crew cost and depreciation) In the case of the Boeing airplane, the sum of crew .

.

1s 10200'12050bn ( 5500 ~~ pmwA^^ _^~~ =~^ (^ ~=~~~=^ ~~ oAc ^-~~~~ ~~^~~~ ^^^ -/^^~^~ ' ~~^ ~+*~ - oAc' - ~^~~^' m ^^ Lco ^~ 4U IVIX NM) (2000 LU ^-Cc ~ FOR DESIGN MISSION o } ^^ o—c o "^" "~.

0.25 0.30 0.20 0.15 DESIGN FUEL WEIGHT DESIGN FUEL WEIGHT FRACTION -~ TOGW Figure 41 SZF505M-7D Savings in D0O Relative to the JT9D-7A for Design Missions -- The main DOO advantage of the STF505M-7D for both typical and design missions is fuel consumption. Average design mission DOC advantage is 9.7%.

STF 505wpo JT9o-7A BASE puoor | FUEL| onsw FUEL puoxr CREW 4.3% SAVINGS AIRFRAME MAINT.

12.5% DEPRECIATION AIRFRAME MAINT.

DEPRECIATION Figure 42 DOC Breakdown (Boeing Domestic Twinjet, Typical Mission: 1,850 km) -- The Boeing twinjet,- which shows the least DOC advantage for STF505M-7D, has the smallest fuel cost portion of DOC (25.5 percent) of the study STF 505 M•70 JT9D-20 BASE 30.2% - 37.0% 14.7% FLIGHT CREW 15.1% FUEL FUEL 1.9% 145, FLIGHT CREW DOC AIRFRAME IN S , SAVINGS MAINL DEPRECIATION 19% 9.4% 11.9% DEPRECIATION AIRFRAME ENG.

MAINT.

MAINL.

23.9% ENG.

MAINL 11.3% 24.8%.

8.7% a Figure 43 DOC Breakdown (Douglas Domestic Trijet, Typical Mission: 1,850 km) -- The Douglas trijet demonstrates the largest DOC advantage for domestic airplanes for the STF505M-7D (9.4%). Fuel cost of this airplane is 30.2% of DOC, STF 505 M•7D JT90.20 BASE 34.2% 41,0% 15.591 FLIGHT FUEL FUEL / CR EIN FLIGHT 16.1°6 CREW INS.

1.7% DOC AIRFRAME SAVINGS MAIN'(, 9,5% DEPRECIATION 10.4% DNS AIRFRAME 1 79' ENG. MAINT..

ENG.

MAINT. DEPRECIATION MAINT. 27,5% 9.8% 8.2%^ 7.8% 22.6% Figure 44 DOG Breakdown (Douglas Intercontinental Trijet, Typical Mission: 2 1 780 km) -- The fuel cost of the STF505M-7D airplane is 34 .

2 percent of DOG.

it JT9D•7A STF 505 WD BASE t zoa° FLIGHT

30.2/ FUEL

CREW INS.

1.8% MAINT.^v^c 6.6%/ DEPR ECIATON 11.4% AIRFRAME MAINT.

DEPRECIATION ENG, 23.4% MAINT.

ENG.

MAINT.

11.3% 10.2% 24.0;6.

Typical Domestic Trijet, Breakdown (Lockheed DOC Figure 45 STF505M-7D cost ol the 2,600 km) -- The fuel Mission: DOC.

` airplane is 30.2 percent of STF 505 M•713 M JT9DJA BASE 33.2% 39.3°:° 17.3% FUEL FLIGHT FUEL CREW FLIGHT 17.8% CREW INS.

1.6% DOC AIRFRAME SAVINGS MAINZ 8.5% 97% DEPRECIATION AIRFRAME INS MAINZ 117% ENG.

DEPRECIATION ENG.

MAINT.

21.0% MAINT.

9.6% - 21.9% 9.0% Quadjet, Intercontinental Breakdown (Lockheed Figure 46 DOC of the km) -- The fuel cost Mission: 5,560 Typical IOC.

STF505M-7D airplane is 33.2 percent of _ STF 505 M-7D JT9D-7A BASE 29.3% 15.2% 35.5% FLIGHT FUEL 15.5% FUEL CREW FLIGHT 1 .8% 1N5 CHEW INS. 1,8%

DOC

AIRFRAME SAVINGS 8.376 MAINT, DEPRECIATION 12.9% DEPRECIATION ' AIRFRAME 23.2% MAINT, ENG.

ENG, MAINT, MAINT, 24.0% 12.;96 10.3% 9.5% Figure 47 DOC Breakdown (P&WA Domestic Trijet, Typical Mission: 5,560 km) -- The DOC advantage of the STF505M-7D is between that of the Douglas and Lockheed domestic airplanes.

JT9D-7A STF 505 M-7D BASE 41,1% 33.8% 15.2% FUEL FUELP FLIGHT CREW FLIGHT 15.7% CREW INS. 1,6% DOC AIRFRAME / IAAINT. SAVINGSI 10.4% 9,8% DEPRECIATIONS iN S AIRFRAME ENG.

DEPRECIATION MAINT.

MAINL 21.0% M NT.

10.0% 9.3% 21.8% 8.6% Figure DOC Breakdown (P&WA Intercontinental Quadjet, 48 Typical Mission; 3,700 km) -- The of STF505M-7D fuel cost the airplane is 33.8 percent of DOC.

costs and depreciation is sixty percent greater than fuel cost, while in the Lockheed intercontinental airplane, for example, the sum is equal to fuel cost.

6.2.2 IOC Comparison The reductions in IOC for the STF505M-7D airplanes compared with the JT9D-7A airplanes are shown in Table 23. The advantage in IOC did not correlate with fuel fraction as well as DOC did because engine performance has little influence on IOC. Because the primary use of IOC was in calculating ROI, the results are only shown for the

economically relevant typical missions. Design mission ROI has no

significance as it does not reflect actual airline experience.

TABLE 23 INDIRECT OPERATING COST SAVINGS OF EEE RELATIVE TO JT9D REFERENCE ENGINE Domestic Intercontinental • Airplane Airplane Percent Savings Percent Savings Boein g 0.6 - Doug las 1.2 1.8 Lockheed 0.8 1.4 1.6 2.1 P&WA 6.2.3 ROI Comparison The return on investment advantages of Energy Efficient Engine powered study aircraft over JT9D reference engine powered study aircraft are shown in Table 24. Since the Energy Efficient Engine when sized for the aircraft application, generally combines a lower initial investment with improved total operating costs (relative to the JT9D-7A), the incremental ROI (or "hurdle rate") of Energy Efficient Engine is mathematically undefined. Incremental ROI is primarily useful for determining the desirability of modifications to existing systems; where the modification requires an initial additional investment, but lowers the future operating costs (or increases future revenues) of the system. When comparing the merits of two competing systems, as in the present study, the difference in their absolute ROI's (Table 24) is more useful.

TABLE 24 PERCENT RETURN ON INVESTMENT ADVANTAGE OF EEE OVER JT9D REFERENCE ENGINE Domestic Intercontinental Airplane Airplane Percent Percent 0.5 - Boeing 2.3 Douglas 2.4 2.4 Lockheed 1.4 2.6 P SHWA 1.9 7.0 NOISE EVALUATION 7.1 INTRODUCTION Predicted noise levels for study airplanes meet FAR Part 36-1978 certification requirements, generally by sufficient margin to provide a high probability of compliance. These noise levels were determined for all study airplanes by means of Pratt & Whitney Aircraft developed procedures to predict the noise characteristics of each source.

The major objecttve of ,'his effort was to assess the noise levels of the study airplanes relative to FAR Part 36-1978 requirements for new type airplanes. This version of Part 36 includes two amendments added in 1978. The primary purpose of the first amendment (Amendment 8) was to adjust noise level limits and measuring locations to align with international noise certification standards recently adopted by the International Civil Aviation Organization. The second amendment (Amendment 9) provided modifications to the measurement and analysis procedures for conducting aircraft noise certification tests to improve uniformity and repeatability.

A second objective of this study was to determine the area within the ° approach and takeoff 90 EPNdB footprint of each study airplane.

To meet the above noise objectives, an acoustic configuration was defined for the two Pratt & Whitney Aircraft study airplanes. The same configuration then was assumed for five other study airplanes defined by the airframe manufacturers. This configuration, defined in greater detail in Section 7.2, included a long common flow exhaust system extensively lined with acoustic treatment.

Noise levels were calculated for the seven study airplanes, using procedures developed by Pratt & Whitney Aircraft to predict the characteristics of each noise component: fan, core (combustor), turbine, jet exhaust, and airframe. Discussions of the prediction methodologies used for each component noise source are contained in Section 7.3, and the estimated noise levels and the calculated probabilities with certification requirements of compliance are presented in Section 7.4. Section 7.4 also includes the estimated area within the approach and takeoff 90 EPNdB footprint for each study airplane.

7.2 ACOUSTIC CONFIGURATION An overview of the acoustic configuration is shown in Figure 49.

Engine acoustic features included substantial spacing between the single-stage, shroudless, 2E-blade fan and the strut -stator (4.3 blade-chord gaps at the outer radii) to minimize the fundamental blade passing noise generated by the - rotor-strut stator interaction.

TURBINE TONE 4000 HZ PP/BRAZED HC FAN TONE FAN TONE FAN TONE 1600 HZ 3150 HZ 2000 HZ BUZZSAW WM/PP /HC WM/PP /HC PP/HC FAN TONE 630.800 HZ 2000 HZ PP/HC NOTE: WM/PP/HC PP = PFRFORATEDPLATE HC = HONEYCOMB WM = WOVEN MESH SCALLOPED MIXER SHROUDLESS STRUT STATORS 27` 8 LA D E FAN 4.3 BLADE —CHORD SPACING e Figure 49 Acoustic Configuration of STF505M-7D -- Acoustic features include substantial spacing between fan and strut stators, mixed, long duct nacelle, and extensive acoustic treatment.

Aerodynamic and structural constraints prevented the selection of a large enough number of strut-stator vanes to acoustically "cut-off" this fan noise source. A large number of core stator vanes ensures that the fundamental noise from the fan/core stator interaction will be cut off and not propagate to the far field.

Acoustic treatment was extensively employed in the inlet and fan discharge duct in order to suppress blade passing tones and buzzsaw noise generated by the fan and tones generated by the turbine.

Treatment requirements were determined from hardwall (untreated) noise estimates for both takeoff and approach conditions in order to define the dominant noise sources and their spectral characteristics. This information established the treatment-tuning requirements for obtaining approach and takeoff noise levels that meet, program' U 4 b' t'ves for the various segments of uning o sec i t T e ob J eCtives.

treatment are shown in Figure 49. Treatment design features, based on factors such as tuning objectives, duct Mach number, and temperature` are defined in Figure 50.

D M FLOW RESISTANCE HONEYCOMB OR HOLE TREATED GENERAL SKIN CELL-SIZE DEPTH DIA. THICKNESS AREA LOCATION CONSTRUCT.

(CM) (CM) POROSITY (CM) (CM) (SO. M) 6.953 A WM/PP/HC 3.378 79 RAYL — — 6.08 0.953 5.080 0.203 8 PP/BONDED HC 20% 0.127 2.10 C WM/PP/HC 0.953 2.286 60RAYL — — 3.28 24.24 D WM/PP/HC 0.953 2.667 60 RAYL — — E 1 524 PP/BONDED HC 0.953 BM 0.127 0.064/0.081 3.05 NOTE: PP = PERFORATED PLATE PP/BRAZED HC 0.953 1.397 13% F 0.203 0.064 0.83 HC = HONEYCOMB WM = WOVEN MESH G P P/BRAZED HC 0.953 1.118 8% 0.203 0.064 5.25 H PP/BRAZED HC 0.953 2.159 11% 0.239 0.064 3.14 -- — - — — — — —. — r^.

r IT —^

l f

I F

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C p O ^

B ` I j A E zr.rrrrrrr4 J20235-50 R7s1,12

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STF505M-7D Nacelle acoustic Treatment Design Features -- Figure 50 Treatment design features are based on such factors as

; 4

tuning objectives, duct Mach number, and temperature.

co LJ k The scalloped exhaust mixer, included in the engine configuration for performance purposes, is not expected to benefit exhaust noise significantly.

7.3 NOISE PREDICTION METHODOLOGY Airplane noise levels predicted in this study were performed on an individual component--fan, core, turbine, jet, airframe--basis. For most components the predictions were derived from a component data base established from engine noise measurements. Correlating parameters, developed from analytical procedures to condense data from several tests and configurations, were used to scale predicted noise levels from each component data base.

7.3.1 Fan Noise Fan noise, dominant during both takeoff and approach, is the most important noise source. Since JT9D-7A and the Energy Efficient Engine fan sections have similar acoustic designs, the JT9D data base was used for the fan noise predictions. The similarities included single-stage configuration, inlet vanes, similar - tip-speed/pressure-ratio relationships, and substantial fan-to-exit A vane spacing.

An important feature of the JT9D data base is that all data were obtained with `an Inflow Control Structure installed on the engine to simulate -flight inflow conditions. Without the Inflow Control Structure, significant levels of inflow distortion exist during static testing that do not exist in flight. Inflow characteristics affect noise generation,thus, only with the proper simulation of inflow conditions can the static data be reliably used for estimating inflight fan noise.

The data base provided noise levels in each 1/3-(ctave band that contained fan noise as a function of fan tip Mach number for each measurement angle around the engine. The data base levels were corrected for differences in diameter, pressure ratio, spacing, and blade number between the JT9D and study engines.

One significant difference between the acoustic designs of the JT9D and Energy Efficient Engines fan section is that while the JT9D contains a large enough number of fan exit vanes to acoustically "cut-off" the fundamental blade assin tone enerated b the P g g Y fan-stator interactions, the Energy Efficient Engine does not.

Structural and performance considerations for Energy Efficient Engine preclude the use of acoustically optimum numbers of vanes; therefore, the - interaction that generates blade passing tone is "cut-on".

Procedures do not exist to define analytically the impact on noise of these differences. It has been assumed that any adverse effect would be offset by the increased rotor-stator spacing of the Energy Efficient Engine, which is nearly double that of the JT9D.

The Energy Efficient Engine fan also differs from the JT9D fan in that it is shroudless. This difference is not expected to have a significant effect on noise. Any acoustic differences probably would favor the Energy Efficient Engine as lesser disturbances would be introduced into the flow that could interact with downstream stator vanes to generate noise.

7.3.2 Core Noise Predictions of core noise generated by combustor

burning

processes were based on data base from a variety of Pratt & Whitney Aircraft engines and combustors. Predicted values were scaled, from the data base by means of a correlating parameter that was developed analytically to collapse the data from the various sources. The correlating parameter included terms for fuel-air ratio, inlet temperature, a flow parameter, and number of fuel nozzles. The calculated values of the correlating parameter for the Energy Efficient Engine fell within the range of available data.

7.3.3 Turbine Noise Turbine noise levels were predicted from a data base obtained from a variety of low and high bypass ratio engines tests. Values of correlating parameters used to scale the predictions fell within the range of existing data. The correlating parameter included terms for loading, tip speed, a flow parameter, blade-vane spacing, and size.

7.3.4 Jet Noise Recently revised SAE procedures (SAE ARP 876, March 1978) that relate noise level primarily to the logarithm of the jet velocity were used for jet noise predictions. Mixing of 85 percent was assumed for the force mixed, common flow nozzle. This is consistent with the percent mixing assumed for performance calculations. To account for 85 percent mixing, the relationships in the SAE procedure were entered at a velocity that was 85 percent of the range between the nonmixed primary velocity and the fully mixed velocity.

7.3.5 Airframe Noise Airframe noise levels for the Pratt & Whitney Aircraft study airplanes were estimated using a relationship between airframe noise data and the airplane takeoff gross weight. The data were obtained from various published data (Reference 4, 5, 6). Airframe noise estimates for airframe company study airplanes were provided by _the airframe manufacturers, and the levels were in general agreement with the procedures used by Pratt & Whitney Aircraft.

85X:' 7.3.6 Acoustic Treatment - Separate procedures were used to predict the inlet and aft treatment effectiveness. The inputs for both procedures were the same: design frequency, duct parameters, and treatment type. Inlet attenuation spectra were obtained from a NASA-developed procedure based on mode cut-off ratio (reference 7).

Fan duct attenuation spectra were based on Pratt & Whitney Aircraft flow duct data. Predictions were obtained by applying to the predicted inlet and fan duct spectra: 1) a calibration factor obtained by comparing predicted and measured JT9D attenuation spectra and, 2) a correction for treated area.

7.3.7 Calculation of Effective Perceived Noise Level Using the previously described procedures, noise levels were predicted increments (5 0 increments at critical for each component in 10 0 angles). The predicted values were extrapolated to required values and tone corrected perceived noise levels (PNLT) we're calculated at each

angle from the 1/3-octave band levels for each source and the combined

sources (total noise). Using the airplane. altitude and airspeed for each case, angles were translated to time. From the total noise PNLT versus time relationships, effective perceived noise levels were calculated for comparison with noise certification requirements.

PREDICTED NOISE LEVELS 7.4 7.4.1 Component Noise Levels the dominant noise source at all conditions.

The fan, as expected, was

significant

and airframe also provided

At certain conditions, the jet

noise. Figure 51 shows the relative contributions to the total for the three noise certification importance of each noise component & and sideline -- for the Pratt conditions --- approach, takeoff, a typical example. At approach the Whitney Aircraft trijet, which is the inlet is the dominant source, fan noise propagating forward from the propagating fan noise is at the other two conditions aft and had the second highest noise level at dominant. Airframe noise at the second most important source approach, and jet noise was The figure also shows the reduction takeoff and sideline conditions.

total noise) that was predicted for the in fan and turbine noise (and acoustic treatment.

Predicted Noise Levels Vs. Objectives (FAR Part 36) 7.4.2 Based on predicted noise levels, all study airplanes could comply with 36-1975. Figure 52 certification requirements of FAR Part the noise airplanes and the FAR for the various study shows the noise levels •?^+""" HARIDWALL (NO ACOUSTIC TREATMENT) WITH ACOUSTIC TREATMENT J 95 Z CL_ A X R F R F F T A A U M A A T N F R I N F U E A R A R N N T f T I F e N C C O O N R J I A L O E 0 rj'g^

I J T T L

E M A E R E A E X R UH T E L T T L T H E 65 L APPROACH SIDELINE TAKEOFF Figure 51 (P&WA Trijet) -- The Component Noise Level Predictions fan was the dominant noise source at all conditions.

requirements for the three certification conditions. It can be noted that the requirements for each condition were a function of airplane takeoff gross weight. In addition, for the takeoff condition, the requirement was a function of the number of engines. Noise levels for all study airplanes were well below the limits with one exception: The Boeing twinjet by one-half an EPNdB.

at takeoff exceeded the limit This small exceedance would not prevent the airplane from meeting a certification requirements as the regulations permit trading of surplus at one condition for an exceedance of up to 2 EPNdB at another condition.

Also it should be noted that no attempt was made to refine for minimum noise the nacelle configuration of this airplane or any other study airplanes defined by the airframe manufacturers. The nacelle configuration optimized for the Pratt & Whitney Aircraft defined airplanes was used throughout this effort.

' APPROACH FAR 36 (1978)

E

O P&WA q BCAC TAKEOFF DAC m O 6 LCC 0.100 FAR 36 (1978) a w SIDELINE 105 100 CG FAR 36 (1978) Z a w ,n 100 Z a, w

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9095 O CJ 180 220 360 E3 (400) (600) (800) (200) 270 360 TAKEOFF GROSS WEIGHT (200) (400) (600) (800) -THOUSANDS OF k9- TAKEOFF GROSS WEIGHT (THOUSANDS OF Ibm) -THOUSANDS OF kg - (THOUSANDS OF Ibm) Figure 52 Predicted Noise Levels ar FAR Part 36 ((978) Certification Conditions - The STF505M-7D achieves the NASA goal of meeting FAR part 36 (1978) noise certification standards.

At approach, the two Douglas trijets had significantly more margin below the limit than the other study airplanes. The lower estimated noise levels resulted primarily from the lower values of approach thrust required for the Douglas airplanes because of higher lift/dra designs and lower flap setting requirements.

Optimization for minimum noise of the performance of all study airplanes and of the nacelle configurations of the airframe manufacturer designed airplanes would decrease the nominal noise levels.

} r 7.4.3 Noise Footprint Areas The area covered by 90 EPNdB contours during landing, approach, and takeoff was calculated for each study airplane powered by the Energy Efficient Engine. These estimates are presented in Table 25. The noise contour area produced by the JT9D-7A powered Pratt & Whitney Aircraft TABLE 25 90EPNdB TAKEOFF AND APPROACH NOISE FOOTPRINT AREAS Airplane Footprint Area - sq.. km Boeing Twinjet 25.4 Douglas Domestic Trijet 30.0 Douglas International Trijet 34.7 Lockheed Trijet 33.4 Lockheed Quadjet 47.7 P&WA Trijet 37.6 P&WA Quadjet 50.2 intercontinental quadjet was also calculated for reference purposc.^ The JT9D-7A engine was assumed to be installed in a nacelle with a short fan-discharge duct, typical of current configurations.

The footprint areas of the Energy Efficient Engine powered study airplanes are 22 percent to 60 percent lower than the JT9D powered reference airplane.

In terms of absolute values, the footprint areas of the study airplanes range. from 50.2 square kilometers (19.4 square

miles)

for the Pratt & Whitney Aircraft intercontinental quad jet to

25.4 square kilometers (9.8 square miles) for the Boeing domestic -twinjet.

It should be noted that values of footprint areas should be used for comparative purposes only, limited to the airplanes within this study.

Footprint areas calculated in other studies may not be comparable as there are no standard procedures, for calculating these areas. Also, absolute values of footprint areas should not be considered exact ^ because of the uncertainties associated with the extrapolation of airplane noise levels to the long distances required in footprint calenlaftinnra_ 8.0 EMISSIONS EVALUATION 8.1 PREDICTED EMISSIONS LEVELS Estimates of gaseous emissions and smoke levels for the Energy Efficient Engine cycle are presented is Table 26. These estimates are based on two-stage Experimental Clean Combustor Program (ECCP) TABLE 26 - ESTIMATED EMISSIONS AND SMOKE CHARACTERISTICS EPAP 2.0 CO 0.2 THC 4.3 NOx 20 (max.)

Smoke No.

*lbm pollutant/1000 lbf thrust/hr/cycle combustor engine data for carbon monoxide (CO) and total unburned hydrocarbon (THC) emissions and on single-stage carburetor tube combustor rig data for oxides of nitrogen (NOx) and smoke emissions.

T $.2 EMLSSION PREDICTION METHODOLOGY The emissions, reported in an EPAP (Environmental Protection Agency Parameter), represent a weighted average of Emissions Index (ED during a typical landing and takeoff (LTO) cycle within the airport environment. Emissions Index is composed of various engine power -settings for a length of time typical of a particular class of aircraft. The power settings and time blocks corresponding to the Energy Efficient Engine class is shown in Table 27. The equation for the EPA parameter can be expressed as cycle (EI)i (Wf)i(TIM)l EPAP cycle (TIM)i (FN)i where Wfi = lbm/hr of 'fuel flow EIi Ibm of pollutant/1000 lbm of fuel TIMi = time in mode(xin) FNi thrust (lbf) subscript i particular engine power setting or mode The CO EPAP was calculated from the ECCP data on a combustor inlet temperature (Tt3) basis, and involved correcting the engine data for the Energy Efficient Engine pressure levels and evaluating the EI's at the appropriate combustor inlet temperatures. The THC EI's were observed to be approximately 1/10 of the CO values. The pressure correction for these constituents is linear: Pt3 ref ; THC) 0.1 CO) EEE CO)EEE = CO)ref EEE P t3 EEE r TABLE 27 LANDINC AND TAKEOFF CYCLE Mode Mode Takeoff Thrust M Time in (min) Taxi/Idle (out) Assigned (mfg) 19 100 0.7 Takeoff Climb Out 85 2.2 4.0 Approach 30 Assigned (mfg)* 7 Taxi/Idle (in) is being _employed for *Installed idle thrust of 5.5 percent the Energy Efficient Engine.

'i 92 `j The NO, EPAP was calculated form rig data on a fuel-air ratio (f/a) The corrected for pressure and temperature and basis. data were correlated against measured fuel/air ratio. The NOx EI's corresponding to the Energy Efficient Engine fuel/air ratios were then, employed to calculate the EPAP. The pressure/ temperature correction utilized for the calculations was ]1/2 Pt3EEE - T _Tt3 EEE t3 ref NO exp x ) EEE NOOref ref 1 P t3 V ref exP [ 18 - 8 (Href- HEEE) VEEE where Pt3 combustor inlet pressure TO combustor inlet temperature (OK) H humidity (lbm H20/lbm air) V combustor reference velocity The humidity and reference velocity terms drop out of the equation since the rig and combustor reference velocities are approximately equal and all data are corrected to the standard 60 percent relative humidity.

The maximum smoke level anticipated during the LTO cycle product was Pt3 corresponding to estimated from the rig data at a value of (f/a) data from a takeoff conditions. The method of correlating smoke configuration has been employed particular combustor by Pratt & effects in Whitney Aircraft to account for variations in ambient engine tests. The method also enables estimation of smoke levels for high pressure ratio engines operating at comparable f/a ratios.

8.3 MARGINS The levels of gaseous emissions shown in Table 26 include allowances for engine-to-engine variations as well as deterioration and development margins. The breakdown of these margins are shown in Table 28.

The allowance for engine-to-engine variation was determined by means consisting of of a statistical analysis of a JT9D-7A pilot lot data, 19 engines. A 3o- level was chosen, which implies that all but 1.5 engines in 1000 will probably meet the requirement.

TABLE 23 SUMMARY OF MARGINS Engine-to-Engine Deterioration Development Total Emissions Variation (%) Margin (%) Margin (%) (%) i CO 22, 5 20 47 THC 46 5 20 71 NO 14 3 10 27 x a 4 # 9.0 EVALUATION OF GROWTH POTENTIAL 9.1 THRUST GROWTH APPROACH A thrust growth plan has been devised for the Energy Efficient Engine.

The approach to thrust growth was developed under the earlier study contract, NAS3-20628, and reported in , NASA CR-135396. Preliminary growth studies were conducted during the first two tasks of that contract, and detailed growth studies were conducted during Task III.

studies Task III, selected During the detailed growth of 1: investigated. This investigation included configurations were cooling-air increases required to maintain hot-section analyses of life, changes in the aerodynamic performance of components resulting from changes in pressure ratio or cooling flows, and evaluation of the structural impact of changes in rotor speeds, pressure, and + temperature levels. The investigation verified the feasibility of the chosen approach to engine thrust growth.

The selected growth path was defined in two growth steps: f, A) An initial step of about 15 percent increased thrust, without changes in nacelle geometry. This increase is to be ` major accomplished by means, of ' an increased overall pressure ratio, fan pressure ratio, a small increase in fan an increased airflow, and an increased rotor inlet temperature.

B) A final of about increased- thrust that step 25 percent include an requires a new nacelle. The engine changes increased overall pressure ratio, an increased fan diameter, and an increased.rotdr inlet temperature.

in Neither of the growth steps will require additional advances technology (i.e., no improvements in materials, coatings, or aerodynamic technology were assumed): 9.2 PERFORMANCE EFFECTS Changes in engine cycle, performance, airflow and diameter are Thrust growth will f summarized in Table 29 for the two growth steps.

be, by a slight improvement in installed (flight 1 TSFC for step (A)and a inlet, nacelle drag, no bleed or power extraction) significant improvement for (B). The improved TSFC will be a step inlet result of increased overall compression ratio and rotor temperature, which partially offsets the fuel consumption penalty of increased fan pressure ratio. In addition, the nacelle drag/thrust of step (A) is lower than the base because the greater thrust is achieved step (B), the without an increase in nacelle diameter. For improvement in TSFC is the efficiency a direct result of thermal • TABLE, 29 GROWTH CYCLE PERFORMANCE AT AERODYNAMIC DESIGN POINT RELATIVE TO BASE ENERGY EFFICIENT ENGINE Step A 15% Step B 25% Thrust Increase Thrust Increase Change in Corrected Airflow - kg/sec (lbm/sec) +36 (+78) +186 (+410) +0.07 0 Change in Fan Pressure Ratio -0.59 +0.82 Change in Bypass Ratio +6.4 +6.;4 Change in Overall Pressure Ratio OC ( OF) +37 (+67) +95 (+172) Change in Rotor Inlet Temperature +3.3 Change in Turbine Cooling Air (% Core Flow) +5.7 +3 (+1.1) Change in Fan Diameter cm. (in) +27 (+10.7) -0.2 Change in TSFC (Percent) -1.1 associated with - the OPR/RIT increases, since fan pressure improvement ratio is constant (propulsive efficiency is nearly constant).

Drag/thrust for step (B) is nearly the same as for the base engine.

The effect of thrust growth on direct operating cost is difficult to assess, since the additional thrust can be used for many purposes, from reducing takeoff distance to increasing payload. The cycle changes used to achieve thrust growth produce DOC penalties, relative to the base EEE, of 0.6% for Step A and 1.2% for Step B, assuming no DOC benefit for increased thrust. These penalties are primarily due to the RIT increases. If the thrust increases are used to increase payload, then these penalties are more than offset, on a seat-kilometer cost basis.

Two stretched versions of the P&WA domestic study airplane were thrust growth conceived to quantify the possible DOC value of these steps. Passenger capacities were solved for to match both steps by holding design range and wing area constant while TOGW, OEW, fuel load and thrust were increased as required. The results show that, on a seat-kilometer cost basis, the DOC of the 15% growth airplane is 4.5% lower than that of the basic EEE powered airplane and the DOC of the 25% growth airplane is 8.7% lower than the base EEE. Both advantages include the penalty for growth cycle changes The changes to the engine required to achieve growth steps are: Step (A): Base single-stage fan is replaced with a single-stage design of higher pressure ratio, increased tip speed, and slightly greater specific flow and diameter. Low-pressure compressor has an additional. supercharging stage. High-pressure compressor is aerodynamically unchanged, but operates at higher physical speed with increased gaspath temperature levels as a result of increased inlet temperature; pressure ratio and corrected airflow are unchanged from base. Combustor operates at higher exit temperature and higher vane cooling flow than base. High-pressure turbine cooling flows are increased to maintain life at the elevated rotor inlet temperature, and the turbine is rebladed (same annulus) because of changes in cooling air and expansion ratio. The low-pressure turbine has a cooled first vane row for hot section life, but no aerodynamic changes are required. The forced mixer/plug is revised to furnish the mixing plane area adjustments required by the cycle revisions. No significant changes in total mixing plane area or jet nozzle are required. No changes to the inlet throat geometry or the external forward nacelle lines are required, with the possible exception of an added section to extend the length.

Step (B): More extensive changes to the engine are required, some of which are illustrated on Figure 53. The base engine fan is replaced with a scaled-up fan having the same aerodynamic parameters. The low-pressure compressor has an added supercharging stage, as in Step but A. The high-pressure compressor is unchanged aerodynamically, operates at elevated pressures, temperatures, and rotor speed -- pressure ratio corrected from base. Burner and flow are unchanged cooling is j exit temperature is increased, and first turbine vane increased by turbines beyond that required Step (A). High-pressure and the turbine is cooling flows are increased to maintain life, rebladed (same annulus) as in Step (A). Since the low-pressure rotor fan diameter at the speed is decreased as a result of an increased same corrected tip speed, the low-pressure turbine requires more extensive changes for Step (B) than are required for Step (A). The low.-pressure turbine has increased elevation, requiring a new transition section between the turbines, and has an additional (fifth) stage. Both the low-pressure turbine first blade and first vane are cooled. Because the increased airflow and cycle changes, mixing of plane geometry differs from the base engine, requiring mixer/plug revisions. The entire nacelle for this growth engine is new.

9.3; NOISE EFFECTS Thrust growth can be achieved with only a small impact on noise. The effects on engine noise of the two approaches for growth discussed in Section 8.1 are summarized in Figure 54. The increases in fan pressure ratio (Step A) and fan size (Step B) result only in a small BASE ENGINE +g 50G'(+172^F) TURBINE ROTOR INLET ODIFIED TURBINE TRANSISTION DUCT TEMPERATURE NEW LARGER FAN EW 5-STAGE GE E N LP TURBINE SUPERCHARGING STAGE ULTIMATE GROWTH _ENGINE Figure 53 Thrust Growth Engine Design — Thrust growth will be accomplished in two ;steps: an initial step of about 15 percent increased thrust without major changes in the nacelle, and a final step of 25 percent, requiring a new nacelle.

increase in fan noise: less than 1 dB. The increase in jet velocity associated with Step A resulted in an approximately 2 dB increase in jet noise. Extending this approach beyond a _15 percent thrust growth would result in additional jet noise which probably would be unacceptable. By increasing bypass ratio (Step B), the jet velocities can be reduced and a more ,substantial (25%) thrust growth can be obtained with small ( l dB) increases in jet noise.

98.

c i thrust growth on airplane noise is more difficult to The impact of growth i3 usually accompanied by changes in the assess; thrust configuration. However; if a constant thrust loading is airplane increases the same rate as engine assumed (takeoff gross weight at

thrust) the margin below the FAR Part 36-1978 noise limit would be

retained as the noise limit increaseswith takeoff gross weight. Very heavy airplanes are possible exceptions at the approach condition because the approach limit is constant at takeoff gross weights above 280,000 kg (617,300 lbm).

9.4 EMISSIONS EFFECTS The primary cause of the changes in exhaust emissions shown below is the base engine to

the increase in overall pressure ratio from 38.6 in

This increase -causes carbon monoxide and 45 in both growth engines.

emissions decrease and oxides of nitrogen unburned hydrocarbon to emissions to increase.

Projected Change in Exhaust Emissions Due to Growth Steps is shown in

Table 30.

TABLE 30 10.0 GOAL ACHIEVEMENT PROBABILITY ANALYSIS 10.1 INTRODUCTION The probability of achieving NASA-.specified goals for the flight propulsion system in TSFC, DOC, noise, and emissions was evaluated during the PS-AIE. Evaluation of the DOC and noise goals required input from the airframe manufacturers.

10.2 TSFC PROBABILITY ASSESSMENT The procedure by which the probability of meeting TSFC goals was determined is outlined in Figure 55. This process was performed in detail during the initial study phase of the Energy Efficient Engine TSFC TURBINE COMBUSTOR TSFC POLL EXPERTS PROS PROS I \ INFLUENCE "BEST" "WORST" 7 COEFFICIENTS "MOST LIKELY" 7 PROS COMPONENT

PROS I "/

ETC.

PERFORMANCE TSFC NACELLE 4 DRAG DOC INFLUENCE WEIGHT, COST, MAINTENANCE COST COEFFICIENTS WT TSFC POLL EXPERTS LIST OF • PROBABILITY DESIGN OPTIONS PROS OF INCLUSION OF OPTIONS • IMPROVEMENTS O DOC Dc a • CONCERNS ^ COST MC • ESTABLISH WEIGHT, COST, MC PROS PROBABILITIES FOR I EACH ITEM • ESTABLISH G DOC G DOC I NTER-RELATIONSHIPS COMPARE NOMINAL E3 ENGINE PERFORMANCE TO JT90-7A PROS % G DOC RE L TO JT9D-7A Figure 55 TSFC and DOC Probability Assessment Process -- Component performance and characteristics probabilities are statistically, combined to predict overall, probabilities of achieving TSFC and DOC goals.

Program (Contract NAS3-20628), and is described in the final report for that contract (Reference 2).

Experts for each component (fan, burner, turbine, ...) review their special area and estimate the "best possible", the "most likely", and the "worst possible" levels of component performance. These expert estimates for each component are then converted into component probability-of-achievement curves. The probability -of- achievement curves for all the components are combined statistically using TSFC influence coefficients to establish a curve of overall TSFC probability.

The TSFC probability for the current contract, NAS3-20646, was obtained by re-estimating the most likely component performance analyzing the overall performance of an engine composed of these most likely components, then shifting the overall TSFC probability curve to reflect the difference between this performance and that of the previous contract. This re-evaluation was based on the results of the Preliminary Design Analysis reported in Reference 1.

The TSFC values in Figure 56 are installed values and include flight inlet and nozzle effects, isolated and nacelle drag, but no bleed or horsepower extraction. The figure shows that the probability is very meeting the NASA goal of at least 12 percent TSFC high (99%) for improvement over the base JT9D-7A engine.

10.3 DOC PROBABILITY ASSESSMENT shown in Figure 55) The evaluation of the DOC probability (also probability and required estimating engine weight, included the TSFC These other parameters, price, and maintenance cost probabilities.

unlike TSFC, had not been evaluated in detail in the previous contract effort and had to be completely determined during the (NAS3-20628) present study.

and maintenance The assessment of these parameters (weight, price, cost) started with the development of a baseline engine configuration and the determination of the weights, prices, and maintenance costs of The design was reviewed and a list of the individual component areas.

improvements to reduce the weight, price, or maintenance cost possib l e was developed. A list of possible design concerns that might increase also compiled. Most items on these lists gave these parameters was an item might decrease weight but increase price mixed results (e.g., and/or maintenance cost).

of These lists, containing over 80 items, were shown to a number estimated the in the various aspects of engine design who experts probability of the item being included in the final flight propulsion cost system. Simultaneously, weight, price, and maintenance The relationships among probabilities were established for each item.

weight, price, and maintenance cost were also determined.

probability-of- estimates of the experts were used to establish The combined for each item. These curves were inclusion curves cost statistically, along with the weight, prices, and maintenance inter-relationships) to probabilities of each item (accounting for and maintenance cost probabilities for produce overall weight, price, These probabilities, expressed in terms the flight propulsion system.

are shown in of percent variation in DOC from the baseline design, was the Figure 57. The Pratt & Whitney Aircraft domestic airplane A similar basis for the influence coefficients used in this figure.

plot was prepared for the Pratt & Whitney Aircraft intercontinental airplane.

The TSFC probability curve, converted to DOC variation, is also shown analyzed as a separate component for in Figure 57. The nacelle was weight and cost and these effects are shown coupled in the figure. The the overall DOC probability, curves were then combined to produce still based on a nominal Energy Efficient Engine design.

10c s0 s0 t 70 z W W,j 60 V 4 50 LL O H J 40 C1 m a -1,z —1.0 —U.S —U.S —0.4 —0.2 0 0.2 0.4 0.6 0,8 1.0 1.2 1.4 1.6 %ADOC i Figure 57 DOC Probability Buildup (P&WA Domestic Trijet, Typical Mission) - TSFC, weight, price and maintenance cost probabilities were combined to predict the probability of meeting the DOC goal.

The airplane economic performance with the nominal engine was evaluated and compared with the JT9D-7A reference engine in both Pratt & Whitney Aircraft airplanes. On the basis of these comparisons, the overall DOC probability curves were converted to a JT9D-7A reference base, as shown in Figure 58. The relationships between DOC results for the Pratt & Whitney Aircraft and the airframe manufacture airplanes were used to establish probability curves the for airframe manufacturers' airplanes--also shown in Figure 58. shown The results are for typical missions. -F The probability of meeting the NASA goal of a five percent DOC F` advantage over the JT9D-7A reference engine is greater than 99 percent a on all airplanes except the Boeing- twinjet. The reasons for the relatively low advantage in the Boeing <twinjet are explained in Sections 5 and 6.

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o

I cc aD CL BCAC // I DOMESTIC DAC / DOM.

—12 —11 —10 -9 —8 -7 —6 _5 -4 —3 + % Q DOC FROM JT9D REFERENCE Figure 58 Probability of Achieving DOC Goal (Typical Mission) The average probability of meeting_ the NASA DOC goal of five percent is 86 percent.

10.4 NOISE PROBABILITY ASSESSMENT Although the predicted noise values shown in Section 7 are generally well below the FAR 36-1978 requirement, it must be recognized that the - predicted levels are nominal values. The uncertainties associated with these predictions must be taken into account when assessing the probability of an airplane complying with noise certification requirements. A manufacturer would want a relatively high probability development before considering the risk acceptable and launching a program. This probability can be calculated using statistical procedures. In this study the statistical methods took into account the predicted margin below each noise limit and the uncertainty or tolerance associated with the prediction. The standard deviations for the noise predictions were estimated to be 3 EPNdB for approach and .2.7 EPNdB for takeoff and sideline. The estimated probabilities for each study airplane are summarized in Table 31. As noted in Section 7, optimization for minimum noise of the performance of all study r ^.k airplanes ' and of the nacelle configurations of the airframe manufacturer designed airplanes would increase the compliance probabilities.

TABLE 31 PROBABILITY OF ACHIEVING NOISE GOAL Airplane Probability of Meeting Goal* 63% BOEING D014ESTIC TWIN DOUGLAS DOMESTIC TRI 95 DOUGLAS INTERNATIONAL TRI 93 LOCKHEED DOMESTIC TRI LOCKHEED INTERNATIONAL QUAD P&WA DOMESTIC TRI P&WA INTERNATINAL QUAD 79 *Goal is compliance with FAR Part 36 (1978) EMISSIONS PROBABILITY ASSESSMENT 10.5 probability margins applied to the gaseous exhaust emission The estimates are discussed in Section 8.0. Figure 59 and Figure 60 show how these probability margins compare with the NASA goals of meeting the proposed 1987 EPA Emissions Standards. Pratt & Whitney Aircraft The quoted estimated emissions are shown by the circles on each plot.

results indicate that the Energy Efficient Engine has a greater than 99 percent probability of meeting the CO and THC standards about a 10 the goal, and a 50 percent percent probability of meeting NOx probability of meeting the maximum smoke standard.

f z r S t n y f CO THC 100 ( I

I I

I ,

I

i

I

} I so > I ^ ^_.

NASA GOAL i J NASA GOAL m I m O I I I o

L

0 1 2 3 0 0.1 ' 0.2 0.3 0.4 ,' CO EPAP THC EPAP Figure Probability of Achieving Emission Goals The probability of meeting CO and THC standards is greater than 99 percent.

a

i

NO X SMOKE 100 ' I

I I

NASA GOAL t 60 I I cc QI m I I

a

I

I

NASA GOAL ) 1 ' -I 0 I 1 2 3 4 5 14 16 1s 20 22 24 26 28 NO X EPAP SMOKE NUMBER - g Figure 60 Probability of Achieving NOX and Smoke Emission Goals -- The probability of meeting the NOx _goal is ten percent. There probability is of fifty percent of achieving the maximum smoke standard.

} i 11.0 CONCLUSIONS o The technologies being developed during the Energy Efficient Engine program offer substantial fuel burned and economic payoffs in a wide range of advanced commerical transport airplanes.

o The current design of the Energy Efficient Engine installed in medium to long range trijet or quadjet aircraft offers a high probability of meeting all NASA program goals in performance, economics, and environmental factors except for nitrous oxide exhaust emissions. In Boeing's shorter-range twinjet application, the current - design, of the Energy Efficient Engine also ma not meet the NASA oals Y $ for either direct operating costs or noise.

o The Energy Efficient Engine design has the potential for thrust growth of up to 25% without significant impact on its ability to meet NASA program goals.

i r r z a ;a 12.0 SYMBOLS & ABBREVIATIONS FOR PS-AIE REPORT oC - degrees Celsius OF - degrees Farhenheit- OK - degrees Kelvin APU - Auxiliary Power Unit AR - aspect ratio BCAC - Boeing Commercial Airplane Company, Division of the Boeing Company BPR - bypass ratio CAB - Civil Aeronautics Board CL- lift coefficient - centimeter c m CO - carbon monoxide DAC - Douglas Aircraft Company, Division of McDonnell Douglas Corporation DOC - direct operating cost ECCP - Experimental Clean Combustor Program EEE - Energy Efficient Engine Emissions Index EI - EPA Environmental Protection Agency EPAP - EPA Parameter (measure of exhaust emissions) EPNdB - effective perceived noise in decibels FAR - Federal Airworthiness Regulations FPR - fan pressure ratio FPS - flight propulsion system ft - feet gal - gallons hp; - horsepower HPC - high pressue compressor HPT - high pressure turbine

hour

hr,': -

- inch

in

- INS insurance IOC - indirect operating cost _kNi -- kilonewtons kg - kilogram - kilometers km - kilowatt kW - pound force lbf - pound mass lbm Lockheed California Company, Division of Lockheed LCC - r Corporation LPC - low pressure compressor LPT - low pressure turbine LTO - landing and takeoff cycle _ m meters MEW = manufactuers empty weight manufacturer Mfg - r u 110 - q ;Y MI, Mach number n,mi. nautical miles N - newtons - NOx oxides of nitrogen - OEW operating empty weight - OPR overall pressure ratio - P pressure - P&WA Pratt & Whitney Aircraft - PR pressure ratio - Prob probability efficiency R RIT rotor inlet temperature ROI return on investment local total pressure/sea level standard pressure S.M. statute mile T total temperature TEGV turbine exit guide vane THC total unburned hydrocarbons TOGW takeoff gross weight - TSFC thrust specific fuel consumption, fuel flow/thrust - W mass flow - Wa airflow - S seconds - local total/sea level standard temperature REFERENCES 1. "Energy Efficient Engine Preliminary Design and Analysis Report" NASA CR-135396, (To be Publishad) "Energy 2. Gray,D.E.: Efficient Engine Preliminary,Design and Integration Studies", NASA CR-135396.

3. Gray D.E.: "The study of Turbofan Engines Designed for low energy Consumption, NASA CR-135002, PWA-5313 April 1976 Revelle, J.D.: 4. "The Calculation of Aerodynamic Noise, Generated by Large Aircraft at Landing Approach", Acoustic Society of America, 87th Meeting, 26 April 1974 5. Putnam, T.W. and Lasagna, P.L.: "Measur,ements, and Analysis of Aircraft Airframe Noise" AIAA Paper 75-510 March 1975 *V.L. al, 6. Blumenthal, et "Aircraft Environmental Problems", AIAA Paper 75-3, January 1973 7. Rice, E. J.: "Optimum Wall Impedance for Spinning Modes, a Correlation with Cutoff Ratio", AIAA Paper 78-193, Jan. 1978.

_

{ A i L APPENDICES The Appendices present the final reports prepared by each of the airframe subcontractors--Boeing, Douglas, and Lockheed--for Pratt & Whitney Aircraft. The reports provide the results of their performance comparisons and integration studies and summarize the subcontract effort in support of the preliminary nacelle design task.

During the nacelle design process--which included two design review/ coordination meetings with each subcontractor--the airframers raised a number of concerns about the design, Pratt & Whitney Aircraft endeavored to account for these concerns as the nacelle design evolved. In cases where there were conflicting opinion among the airframers on aspects of the design, Pratt & Whitney Aircraft chose the approach most in keeping with the objectives and philosophy of the Energy Efficient Engine Program. In other cases, where satisfactory resolution of the concern would have required a detailed design effort beyond that appropriate to the purposes of this program, the concern was left unresolved.

The preliminary_ nacelle design resulting from this coordinated effort the Energy Efficient Engine Preliminary Design and is described in Analysis Report, Reference 1.

r

D6-48027

DOCUMENT NO.

ENERGY EFFICIENT ENGINE AND AIRPLANE INTECROJION STUDY TITLE: MODEL ISSUE NO. TO: c JDATE) c Subcontract No. 20646-1 " Under NASA Contract NAS3-20646 For United Technologies Corporation Pratt and Whitney Aircraft Group Commercial Products Division East Hartford, Connecticut • _ -7 .PREPARED BY J SUPERVISED BY v ^____t_^!

r -. nson v_ APPROVED BY r W APPROVED BY a ° Y•q A.

_ o ay, f ii4 D6-48027 N,.

PAGE 17 1 REV SYM

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REVISIONS

a

REV DATE APPROVAL SYM DESCRIPTION A Corrected typographical errors on Pages vii, 9, 10, 33, 43 57, 65, and 68. Revised wording of second sentence on Page 3. Revised wording of first sentence, page 22, to to correct "...maximum cruise speed of 0.8 Mach ---."

"---- cruise thrust, 0.8 Mach ----." Revised wording line 4 'through 8, page 29, to change standard day to 84 0 F day.

Page 36 revised 500M point to 6500M point. Page 40 revised signs of TOGW and SLST sensitivity to &FN TO' t,.

TABLE OF CONTENTS Page 1.0 SUMMARY INTRODUCTION 2.0 INTRODUCTION 5 7' 3:0 SYMBOLS AND ABBREVIATIONS 4.0 9 AIRPLANE AND MISSION DEFINITION 4.1 Mission Selection 4.2 10 Advanced Technology Features 4.2.1 Aerodynamics 4.2.2 Weights and Structures 13 4.3 Airplane Geometry Guidelines 4.4 Engine Installation 16 4.4.1 Engine Placement 4.4.2 Nacelle Drag 16 4.4.3 Engine Bleed and Power Extraction 16 4.5 Preliminary Airplane Configuration 21 4.5.1 Airplane Description 4.5.2 Engine Description 21 4.6 Procedures for Determining Direct Operating Cost (DOC) and Return on Investment (ROI) 22 4.6.1 Direct Operating Cost 4.6.2 Return on Investment 27 AIRPLANE PERFORMANCE AND SENSITIVITY 5.0 5.1 Airplane Sizing 29_ 5,1.1 Airplane Performance and Characteristics 5.1.2 Airplane Weight 33 5.1.3 Airframe Noise and FAR 36 Flight Conditions 5.1.4 Engine and Airframe Noise '36 Sized Airplane 37 5.1.5 Airplane Drawings of Airplane Drag Pol ars 37 5.1.6

C

5.2 Airplane Sensitivity Factors 37 5.3 Takeoff Gross. Weight and Fuel Burned Comparison 5.4 Typical Mission DOC and ROI 42 6.0 AIRCRAFT ENGINE INTEGRATION 51 6.1 Nacelle Arrangement and Construction 6.2 Airframe Accessory Requirements and Location 51 6.3 Maintainability, Accessibility, and Safety Requirements 6.4 Nacelle Mount System 58 6.5 Nacelle Design •6.6 Nacelle Weight Evaluation 66 72' 7.0- CONCLUSIONS AND RECOMMENDATIONS REFERENCES LIST OF TABLES Page 1-1 Airplane Characteristics and Performance 1-2 Nominal Noise Estimates 3 LIST OF FIGURES Page Typical Mission Profile Energy Efficient Engine Configuration (General Arrangement)-- Model 768-865 Airplane Geometry Guidelines Page 5-8 768-867 Design Mission Flight Profile--Model 5-9 768-867 46 500 nmi Mission Flight Profile--Model 768-866 5-10 Design Mission Flight Profile--Model 5-11 500 nmi Mission Flight Profile--Model 768-866 48 6-1 P&WA Energy Efficient Nacelle - 6-2 Hydraulic Loads 55 6-3 - Electric Loads 6-4 Long Duct Nacelle Mount 6-5 Airl oads--Maximum Takeoff _ 63 6-6 Airloads--Maximum Dynamic Pressure 6-7 Airloads--0-deg Flap, 1.3 Vstall 6-8 65 Airloads--0-deg Flap, 1.3 Vstall 1.0 SUMMARY NASA objectives for the Energy Efficient Engine (E 3 ) program are to develop technology to achieve: (1) a 12% reduction in cruise specific fuel consumption, (2) 5% reduction in direct operating cost (DOC), and (3) reduction of engine performance deterioration common to current technology high-bypass-ratio engines. Future noise and emission requirements must also be met. Boeing's role in the E 3 program was to help determine if the P&WA E3 engine cycle met NASA goals. In this capacity, Boeing defined an advanced technology airplane and provided mission performance, economics, noise, and nacelle assessment data with E 3 and current technology engines installed.

No D6-48027.

REV SYM PAGE D3 42037500 REV. 5/76 Table 1-1. Airplane Characteristics and Performance Domestic Airplane 7A Engine STE505M-7C Table 1-2 shows that airport and community noise levels for the STF505M-7C airplane meets FAR 36, amendment 8, requirements for a twin-engine airplane.

A nominal noise estimate 3 EPNdB below FAR noise requirements is generally considered sufficient margin to ensure a certifiable engine installation.

Using this criterion, the approach noise is marginal; however, no attempt was made in this preliminary estimate to refine the nacelle treatment to lowest noise levels.

Table 1-2. Nominal Noise Estimate STF505M-7C FAR 36 (1978) EPNdB Requirement EPNdB z EPNdB o Boeing evaluation of the STF505M-7C nacelle weights indicated the nacelle weight to be 1310 lb over the P&WA estimated weight.

1% Boeing's weight estimate was based on methods reflecting low technical risk for commercial operation. This weight increase reduced fuel burned savings from 17.9% to 16.6% and reduced the DOC advantage from 6% to 5.7%.

o The STF505M-7C engine price supplied by P&WA is too low according to Boeing projections. Boeing's assessment indicated a price increase of $590,000 per engine. The Boeing estimated price reduced the DOC

below the NASA

advantage of the STF505M-7C from 6 to 4.2%, which is 5% goal.

o Nacelle assessment and evaluation requires continual review as the design evolves to ensure that the nacelle design meets airplane requirements and objectives, Boeing design practice, and airline and FAA certification requirements. During the Boeing assessment,— several versions of the STF505M-7C nacelle design were reviewed.

In P&WA's nacelle layouts, however, material callouts and construction details were too incomplete to conduct an indepth evaluation. Concerns based on a critique of the nacelle design were developed and coordinated with P&WA. Some nacelle design problems were identified. Much additional effort would be required to ensure a flight-acceptable-nacelle installation, but no work of this type is future programs.

being considered in

To ensure that the E 3 program results in an engine configuration that meets

the program goals and that can be installed in a nacelle acceptable to the airframes and airlines, it is important for the airframer to be actively, involved in the installation design and evaluation.

D6-48027 r1o.

REV SYM 128` PAGE FY. 5/76 2.0 INTRODUCTION The NASA Aircraft Energy Efficient program (ACEE) has the objective of improving the energy efficiency of future U.S. aircraft so that substantial fuel savings and economies can be achieved.

The "Energy Efficient Engine (E 3 ) Preliminary Design and Integration Study" is one of the elements of this program. The recommended advanced technology propulsion system resulting from this study is projected for use on airplanes introduced into service in the late 1980's or early 1990's. NASA goals for the E3 program are a 12% improvement in installed cruise specific fuel consumption, a 5% improvement in DOC, and performance retention of 50% or more as compared with a current technology high-bypass-ratio turbofan engine.

The present study is a follow-on to work performed for Pratt & Whitney Aircraft (P&WA) under subcontract No. 20528-1 in support of P&WA prime contract NAS3-20628. Objective of the P&WA prime contract, NAS3-20628, was to evaluate advanced technology engine cycles and to select an advanced cycle that best fulfilled the NASA E 3 program goals. Objective of the current study was to evaluate the advanced technology turbofan engine comparing it with a current technology reference engine to determine if NASA goals will be met when these engines are installed on commercial airplanes of the late 1980'x.

The tasks designed to accomplish this objective included: a. Aircraft and Mission Definition. Under this task an advanced technology transport aircraft was defined with a design range, performance passenger capacity, and mission appropriate for domestic use.

b. Aircraft Performance and Sensitivity. This task evaluated a current technology reference engine, the JT9D-7A (ref. 3) scaled to the airplane requirements and a similarly scaled advanced technology engine, the STF 505M-7C (ref. 4), as installed in the advanced technology airplane.

The aircraft size was optimized for each engine for the defined mission.

Aircraft performance and mission sensitivities were then generated for the

aircraft powered with the advanced engine.

c. Aircraft and Engine Integration.. Under this task a P&WA nacelle was

evaluated for nacelle construction, airframe accessory requirements and

location, maintainability, accessibility and safety requirements.

Section 4.0 of this report reviews and updates the mission selection and

airplane definition studies accomplished in earlier E3 studies reported in

reference 5. Mission definition differed from these earlier studies primarily

in its reduction of takeoff field length (TOFL) requirement from 7500 to 6000

ft; the major airplane-configuration change was an aft relocation of the engine exhaust plane to 40% wing chord. The latter change was made as a result of a flutter-weight penalty trade study.

Section 5.0 summarizes the sizing studies of the JT9D-7A- and STF505M-7C powered airplanes and compares the resulting performance, noise, and economics of the two airplanes. These studies were based on the P&WA-supplied engine performance, engine weight, engine noise, and engine economic data. DOC and

ROI sensitivity to fuel price was determined by using fuel prices of 35, 40

and 45t/gal. Also, an additional DOC and ROI calculation shows the impact of

a Boeing estimated engine price that was about 50% higher than P&WA's estimate.

Section 6.0 comments on the Boeing assessment and evaluation of the P&WA-designed nacelle installation. Design comments, accessory requirements and location, design loads, mount structure, and a weight assessment are

included in the criti

quue of the P&WA nacelle design.

t

-48027

po. D6 N REV SYM 130 AGE 6 G3 4203 7500 REV. 5/76 3.0 ABBREVIATIONS AND SYMBOLS A/P airplane AR aspect ratio BLKF block fuel, pounds BLKT block time, hours focal chord c wing lift coefficient, L/gSREF C L C LR CL ratio CD drag coefficient, D/gSREF

nacelle drag coefficient, DNAC/gSNAC

C DNAC

CET combustor exit temperature, of D airplane drag, pounds dB(A) weighted sound pressure level, decibels

No. D6 - 48027

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PAGE D3 4203 7500 REV. 5/76 SFC specific fuel consumption lb/hr-lb SLST sea level static thrust (uninstalled) wing reference area, ft2 SREF nacelle wetted area, ft2 SNAC t/c wing thickness-to,chord ratio, measured streamwise TE trailing edge TOGW takeoff gross weight, pounds TOR takeoff field length, feet WCP wing chord plane WR P wing reference plane VAPP approach speed, keas design dive speed V sweepback angle at wing quarter chord, degrees A - 0.25C 4.0 AIRPLANE AND MISSION DEFINITION Selection of the design mission and a corresponding design payload and range was based on a projection of the commercial airplane market of the 1990's.

Various design requirements, wing geometry, and advanced technology features were established for a 1990 domestic service airplane.

4.1 MISSION SELECTION Examination of the possible 1990 market suggested that the future airline market would be similar to the existing marketplace. This prediction was bused on the assumption that the air traveling community in the 1990's will constitute approximately the same percentage of the total population as today's air travelers, with a 4 to 6% annual growth. The air cargo market should experience similar growth.

Many of the current narrow body aircraft will be retired from active service by the major airlines in the late 1980's. These include the intercontinental range 707-3206 and -320C models, the DC-8 Sixty series airplanes, and some of the early 727-200 model domestic airplanes.

Hence, there should be a market in the late 1980's for a large number of replacement aircraft in the 180 to 220 passenger size range.. Accordingly, the design mission and sizing constraints selected for the E 3 study are: Domestic Airplane Design range, nmi 2000 Nominal payload, passengers (15/85% mix Cruise Mach number 0.8 _6000 TOFL, feet (max) VAPP, knots _(max) ICAO, feet (min) 33 000 Reserves ATA Domestic The following off-design missions were selected for economic assessments: Domestic Airplane ,665 Range, nmi Payload, 108 passengers (15/85% mix) Cruise Mach number _ 0.8 : 01 4100 7740 ORIG.1/71_..

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4.3 AIRPLANE GEOMETRY GUIDELINES The airplane geometry guidelines shown in figure 4-3 were adopted to ensure adequate ground clearance during taxi, takeoff, and landing. These are the same guidelines used in the earlier study under subcontract No. 20628-1.

CURRENT

NEW TECHNOLOGY

TECHNOLOGY

MATERIAL

MATERIAL STRUCTURAL WEIGHT SAVING

COMPONENT

% OF

COMPONENT

WEIGHT

STANDARD

ADVANCED WING BOX

6%

ALUMINUM

ALUMINUM FUSELAGE 4%

ALLOYS

EMPENNAGE

ALLOYS

6%

( CURRENT 747)

BOX

CONVENTIONAL ADVANCED

CONTROL

25%

COMPOSITE

ALUMINUM SURFACES

STRUCTURE

CONSTRUCTION LANDING GEAR

(GRAPHITE)

DOORS

MAIN LANDING 40%

CARBON

GEAR BRAKE

LANDING GEAR

TITANIUM

20%

SUPPORT

FITTINGS

SIDE OF BODY RIB

EMPENNAGE

BODY ATTACH

ENGINE STRUT

ATTACH

FLAP SUPPORT •-

TABLE 4-1

Advanced airframe structure for 0studies

J19-0.7

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y I f 119-0.7 D1 4 10 7740 ORic.3/ 71 ;7 rn N uj U JI a° 8° 2.3° v° ^--- N1CF' W R F T----- ---1-- X0.1 C • NACELLE PLACEMENT NACELLE PRIMARY NOZZLE AT 40% OF CHORD.

FAN COWL VERTICAL POSITION BELOW WING LOWER

SURFACE BY 10% CHORD OR GREATER.

• NO VORTEX SHEDDING OVER WING

FORWARD LIP OF COWL MUST BE BELOW AN 8-DEG LINE

MEASURED WITH RESPECT TO LOCAL CHORD PANE.

• NO JET WAKE IMPINGEMENT

JET WA KE BASED ON EQUIVALENT DIAMETER AT THE PLANE

OF PRIMARY NOZZLE AND EXPANDING 7 DEG MUST NOT

CONTACT LOWER WING SURFACE.

-^ Z > 0 m p

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FIGURE 4-6 Installation comparison

w

AIR CONDITIONING + WING & ENGINE INLET TAI

D6-48027

No.

REV SYM 144

PAGE

n

Engine power extraction for airplane off-design operation (e.g., operation in icing conditions) was not required for the airplane parametric studies.

System designs, however, considered off-design requirements.

4.5 PRELIMINARY AIRPLANE CONFIGURATION 4.5.1 Airplane Description For the preliminary airplane, this study selected a twin-engine wide-body configuration with double-aisle seven-abreast seating. Wing geometry -0.25C = 30 deg) was consistent with the cruise speed and (AR = 10, - A takeoff and landing characteristics. The lower lobe cargo space was configured to accommodate 17.LD-3 cos tainers side by side.

Main characteristics of the two engines at maximum cruise speed of 0.8 Mach and an altitude of 35 000 ft are: STF505M-7C JT9 D-7A 7.0 5.0 Bypass ratio Installed SFC - 0.56 0..68 1.74 1.58 Fan pressure ratio Overall pressure ratio 38.6 25.4 Maximum turbine rotor inlet temperature (SLS hot-day takeoff) 2450OF 2290OF O -1 Crew cost - f (TOGW, cruise speed, mission type) - fuel burn and fuel price specified • Fuel • Airframe maintenance - specified (Boeing) • Engine maintenance - specifi led (engine manufacturer) • Depreciation f (useful lifo-, residual value, utilization, Initial price, spares price) + Insurance f (initial flyaway price)- DOC per trip Utilization n f (block time) do its-047 0 4100 7740 OR iG. /71 s M M NU M ^i 1977 CREW PAY (S1BLK-HRI w + 2.838) F u + 19 80 (29 87 F 2-MAN CREW 1 iJei:nition of terms and units (33.54 F w + 3.483) F C + 2930 z 3-MAN CREW 1 FUEL (S/U.S- GAL) 036 TOGW Maximum takeoff gross weight--lb 1.02 ON FUEL AND MAINTENANCE girflem• price—S NONREVENUE FACTOR Ca Engine prica:angine—i Ce AIRFRAME MAINTENANCE—CYCLE (excluding reverwr) MATERIAL ($/CYC) No Number of •nglnes DIRECT LABOR (MH/CYC) MATURE LEVEL MAINTENANCE Sea level static thrust-4b T AIRFRAME MAINTENANCE—HOURLY M High speed crulae mwJi number MATERIAL (S/FH) BASED ON DETAILED VV a Airframe weight—lb DIRECT LABOR (MH /FH) Flight-hours FH ENGINE MAINTENANCE—CYCLE ANALYSIS MH Man-hours

MATERIAL /$ /CYC)

CYC Cycle DIRECT LABOR (MH/CYC) Tb Block time—hr ENGINE MAINTENANCE—HOURLY MATERIAL ( S /FH) Notes: DIRECT LABOR (MH/FH) 1. See attachments for F. and F u crew pay factors BURDEN ( MH/DIRECT LABOR MH) 2 0 MAINTENANCE LABOR RATE ( /M 9 70

S H)

2. For flight-hours <2 use; -- ----- INVESTMENT SPARES RATIO Cost at 2 hr — 0.73 (hourly cost) x (2 • flight-lows) AIRFRAME 0.08 ENGINE 003 For flight-hours >4 use: DEPRECIATION SCHEDULE Cori at 4 hr + 1.53 (hourly coat) x (flight-hours - 41 15/10 (YEARSM RESIDUAL) INSURANCE RATE 1 OF TOTAL PRICE/YEAR)

M 0.6

4,000 UTILIZATION U _ a z * 060 + ^ O (BLK-HR/YEAR) Tb+05 i^ ^7 (16 TRIPS/DAY MAXIMUM) N I A A Co O IV V Domestic direct operating cost formulas i e 04 0 4 10 7740 OR IQ. A/71 m CA

GW ))0.3

Fey - (SPEED' X (

100,000

E FW

"Speed - 7i5 (mach) --75 (mach)4 Cruise mach < 0.9

- 660 (mach) Cruise mach > 0.9

3,600 4,000 4,400

400 800 1,200 1,600 2,000 2,400 2,800 2 3,200

SPEECH ° X ( GW ) 100,000 O m p

A

O N V FW factor for crew pear FIGURE 4-8 It its-( X N

-c

U.S. DOMESTIC

CREW 560

UTILIZATION

U.S. INTERNATIONAL

(h/year)

•

Utilization factor, Fu

a

F Ft

Crew

utilization ( h /year)

Fu international

s

Crew utilization (h/year)

(F u J 1.0)

Z O

4 8 10 14

2 6 12

AIRPLANE UTILIZATION

(h/day)

co Crew utilization FIGURE 4-9 i ^t ROI is the discount rate at which the net present value of Definition: future cash inflows (cost savings) is equal to the initial cash outlay (investment) 0 useful life r Net present value (NPV)_ (NPV)CL -COUT + E n-I When NPV - 0, r - ROI = discount rate Calculations: 1. Before tax cash outflows (C0UT) • Incremental airplane price or modification cost • Additional spares inventory 2.

Before tax cash inflows (annual) (CW) • Cash operating cost savings • Fuel • Maintenance 3. After tax equivalence • Depreciation tax effects • Investment tax credit (if applicable) Return on investment method TABLE 4-5 u No. D6-48027 REV SYM 152 -- PAGE 28 I 5.0 AIRPLANE PERFORMANCE AND SENSITIVITY AIRPLANE SIZING 5.1 Both the JT9D -7A and the E 3 powered airplane were sized to meet the same design mission. Design selection charts for the two airplanes are shown in figures 5-1 and 5-2. The wing loading for these airplanes was chosen for BLKF and takeoff gross weight (TOGW) with an 84 0 F -day sea-level minimum takeoff field length (TOFL) constraint of 6000 ft determining the thrust loading. The takeoff constraint for the STF505M-7C required about 5% higher thrust-to-weight than the JT9D-7A. This was largely a result of relative increase in engine BPR and windmilling drag for the STF505-7C engine.

5.1.1 Airplane Performance and Characteristics -7A- and the STF505M -7C-powered Characteristics and performance of the JT9D 5 -1. Each airplane was designed to meet airplanes are compared in table airplane and mission requirements (sec. 4.1). The BLKF and TOGW shown in table 5 -1 are based on an airplane sizing program. A more detailed fuel burned comparison based on mission analysis is discussed in section 5.4.

01 .100 7740 ORIO.a/71 -C 047 m

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768-867 MUM) AIR PLANE . • i..... ; .:: ,...

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is

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'-RESERVES ; :: : ' A?A^OOMESTI^^'...

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FIGURE 5-1 Airplane design selection chart model 768-867 T 01 4100 7140 OR10.0/71 l 1. -047 .x m

768-866 (STF 505M-7C) AIRPLANE

:...

..

SAR `^ 2000 NMI !

_(kEAS) A !.PAYLOAD 196 'PASS (401®O L0) ...

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80 120:.;;:::..;: 130 .;..:.:140 1 :150 • W I NG' LOADING ` WS ..

^.. .......:.::....

G V FIGURE 5-2 Airplane design selection chart model 768 866 Table 5-1 Airplane Characteristics and Performance Domestic Airplane STF5 _ 05M-7C JT9D-7A Engine Engine 5.1.2 Airplane Weight Table 5-2 shows results of a weight analysis on domestic E 3 airplanes with the STF505M-7C and JT9D-7A engines. These weights reflect the advanced technology features discussed in section 4.2. The nacelle weights were supplied by P&WA and scaled to the appropriate thrust level. A preliminary balance analysis indicated acceptable loadability for both airplanes.

Table 5-2.

Weight Statement for P&WA E 3 Airplanes Weight (LB) Model 768-867 Model 7 68 -866 STF505M-7C)JT9D-7A Wing 33 590 37 890 Empennage 4790 4580 Body 33 620 33 730 Nacelle* Gear 12 870 12 760 Total structure (92 720) (90 660) Propulsion system (16 280) (16 290) Fixed equipment and options (42 110) (42 430) Standard and operational items 11 400 11 400 OEW 162 510 160 780 *P&WA provided nacelle weights used in above analysis.

i` Table 5 -4. Nominal Noise Estimates FAR 36-8 STF505M-1C* R_equirement Notes

Takeoff 91.0 dB 93.9 dB No cutback at

6500 m point p dB 98.3 d8 Sideline distance

Sideline 91.0

= 450m point 102.0 d6 102.0 dB 2000m from Approach threshold (two extended flap segments, 3 deg glide slope) Note Nominal noise estimates are shown--appropriate design and demonstration tolerances are required for certifiable/guarantee levels.

5.1.4 Engine and Airframe Noise In In the Boeing analysis, the acoustical design point was an 80% level of .

.

of certification. This goal could be achieved with current and

near-future _lining technology. The estimated noise

levels for the STF505M-1C

were based on a nominal acoustic treatment to the engine and nacelle, not on a fully iterated lining design study.

Because quiet operation was not the prime objective in configuring this airplane, no adjustments were made to the performance or flight configuration for the purpose of lowering noise levels.

Optimization of linings, flap settings, and thrust levels could improve the margin for the approach case.

The above Table 5-4 shows nominal noise estimates.

5.1.5 Airplane Drawings of Sized Airplanes Figures 5-3 and 5-4 show drawings of the JT9D-7A- and STF 505M-7C-powered airplanes.

5.1.6 Airplane Drag Pol ars The airplane drag polars were derived from wind tunnel test data obtained from a model closely resembling the study configurations. Beyond that drag optimism associated with advanced technology was incorporated as discussed in section 4.2. Estimated drag of isolated nacelles and drag caused by interference between the nacelles and the airframe were included in the airplane polars.

5.2 AIRPLANE SENSITIVITY FACTORS Sensitivities for airplanes are shown in tables 5-5 and 5-6. The airplanes are sized by TOFL and the sensitivity results are nonlinear for some parameters. In some cases, better airplane solutions (i.e., lower TOGW or BLKF) can be obtained by sizing to more stringent performance constraints.

This, however, requires additional diagnostic point designs that are time-consuming and costly. It is recommended that the sensitivities be used with caution and not outside the amount of change shown.

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ob- 4W27 OEEE 31A Y, ^ .

5% CR DRAG 5% OEW 5% FNTO BASE 5% F 5% SFC NCR +/- +/- +/- +/- CYCLED +/- +1.6/-1.5 +2.0/-1.7 +6.2/-5.7 -0.4/+0.8 TOGW 256600 -0,2/+0.4 +0.8/-0.8 -0.9/+1.2 OEW 160780 -0.1/+0.2 +1.1/-0.9 +8.51-7.9 +0.9/-0.9 -1.0/+1.3 MEW -0.11+0.2 +1.21-1.0 +8.8/-8.1 149380 +4.7/-4.6 +6.2/-5.2 BLKF +4.2l-3.7 +1.11+0,4 41410 -0.61+1.7 +5.71-5.3 -5.2/+6.-2 +1.41-1.4 +2.1/-1.9 SLST 37280 -0.11+0.4 BASE 5% FN CR 5% SFC 5% CR DRAG 5% OEW 5% FN TO +/- +/- +/- +/- CYCLE D +/- TOGW +1.3/-1.3 +6.2/-5.9 -0.51+0.8 248900 -0.2/+0.3 +1.9/-1.8 +0.71-0.7 -0.9/+1.2 OEW 162510 -0.11+0.2 +0.9/-0.8 +8.5/-7.9 +0.7/-0.7 T8.7/-8.2 -L0/+1.2 MEW -0.11+0,2 +1.0/-0.9 151110 +4.51-4.4 +0.9/0.0 BLKF 34290 -0.8/+1.6 +5.9/-5.3 +4.0/-3.7 +1.21-1.2 SLST 35820 -0.2/+0.3 +1.;9/-1.8 +5.81-5.4 -6.31+5.4 A comparison of design-mission fuel burned for the STF 505M-7C and JT9D-7A powered airplanes (fig. 5-6) shows cumulative fuel savings over the mission nearly constant. This was a result of minimal variation in SFC difference between the two engine-airframe combinations throughout the mission. A breakdown in fuel used during various mission segments is shown in figure 5-7. The large percentage of fuel burned 'during climb for typical stage lengths shows the importance of maintaining the advanced engine SFC improvement at climb power setting.

Figures 5-8 through 5-11 show the actual mission profiles (time and altitude versus distance) for both airplanes at mission ranges, of 500 and 2000 nmi with 100% payload. The engine thrust level at the beginning and end of cruise are noted for support of engine duty cycle studies.

f _lr

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^ N s D, ^^ ^^ ;r }1 a fi ^ O

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v `s t j t L°"' 3

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DESCENT + APPROACH + TAXI CRU I SE CLIMB + ACCELERATION TAXI + TAKEOFF

Q

u

a FIGURE 5-7 Percent block fuel by mission profile segment a D6-4 8027 No.

REV SYM 166 PAGE , t 2 ELAPSED TIME

x

r 500 N.MI. MISSION PROFILE MODEL 768-867 ( JT9D-7A ) to END CRUISE 93% FN MCR FN - 6595 LB LENG FN MCR - r ALT . 39000 FT W INITIAL CRUISE FN - 94% FN MCR 6595 LB I ENG a FN MCR - ALT - 39000 FT 300 400 0 100 200 500 DISTANCE - N. MI.

i ;i

500 nmi mission f/fight profile - mode/ 768-867

FIGURE 5-9

V Q Q f e f i x of 4100 »4o-_owia.s[ti

_

m

5-

ELAPSED TIME o 3'

DESIGN MISSION PROFILE

±

MODEL 768-866 ISTF 505M-7C - INITIAL CRUISEEND CRUISE

9696 FN MCR

FN

- FN •9096 FNMCR

4, LB 1 FN MCR - 7050 LB I ENG FN' MCR • 7050 ENG

o

39000 FT

ALT • 39000 FT ALT • _.

0 500 1000 1500 - 2000

o .

^ D I STANCE — N. MI.

FIGURE_ 5 -10

Design mission flight profile - mode / 768-866

-E.

i J U-047 'i ^^ & ^ 3 1' y END CRUISE FN MCR MCR % 7050 LB / ENG F N ALT n c A I r '. f'ot4s+^3!4^t3^1^3f1^4:.1N:8T.^+f{m^ n.r.

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505M7C 505M7C ENGINE JT90-7A (P&W ENGINE PRICE) (BOEING ENGINE PRICE) .45/GAL .35/GAL .40/GAL .45/GAL .35/GAL .40/GAL .45/GAL .35/GAL .40/GAL FUEL } c: 4.96 5.139 5.312 5.066 5.239 5.412 DOC: 5.259 5.467 5.675 FUEL 1.455 1.662 1.870 1.210 1.383 1.555 1.210 1.383 1.555 CREW - - 1.120 - - 1.130 - - 1.120 INSURANCE .1096 - - .1081 - - .1142 - - BURDEN .475 - - .482 - - .482 - - .0707 - - ENGINE LABOR .0671 - - .0707 - - ENGINE MATERIAL .273 - - .273 - - .299 - - .1703 - - AIRFRAME LABOR .1703 - - .1703 - - AIRFRAME MATERIAL .1239 - - .1239 - - .1239 - DEPRECIATION - - - 1.502 - - 1.430 - 1.403 it AIRPLANE AFTER TAX 11.0 11.3 10.8 10.2 10.9 10.3 9.6 12.1 11.6 ROI (X) INCREMENTAL AFTER TAX 13.3 15.4 17.4 ROI* (%) > o

^

A

A

v

;v r b.

Cash flows and their 'timing are considered as follows: Time prior to delivery Percent (%) of price paid 15 mo 20 12 mo 5 9 mo 6 mo 5 0 mo (delivery) 65 + spares 6.0 ENGINE/AIRPLANE 'INTEGRATION This section describes the Boeing assessment and evaluation of the P&WA designed STF 505M-7C engine/nacelle installation defined by P&W layout No.

L-109846. Comparison of nacelle features with Boeing standards and airline requirements is covered where appropriate.

6.1 NACELLE ARRANGEMENT AND CONSTRUCTION The inlet and major nacelle dimensions were generally consistent with Boeing practice. Nacelle lines were not evaluated, but both the afterbody leaving angle and plug leaving angles exceeded the Boeing recommended value of 12 deg.

Being preliminary, this layout lacked numerous construction details, and in-depth critique of detail construction was not possible. Comments were provided on areas such as the thrust reverser where some detail was shown.

Figure 6-1 represents the P&WA designed nacelle. Number codes on the figure have been keyed to the comments listed below.

1. Interference between fan air blocker door (from hinge line forward) and trapezoidal doorframe will occur durin g translation. This condition is characteristic of the configuration (fixed doorframe in cowling).

115-017 Do 4100 7710 'OA10J/771.

m N J i if a d e r y_- f P&WA LAYOUT NO.

>o

m o

!_-109846 SHEET 1 (REF)

N t

V

5. No, radial structural tie is shown for the outer translating sleeve when in the reverse mode. Radial loading could be substantial due to existing leak paths.

6. Outer translating sleeve appears to be the duct pressure wall. This condition is not compatible with existing latching or hinging means from the standpoint of sealing or strength.

Access to inside latches is not clear.

8. Roller and track translation mechanisms have very short service lives and have been replaced in existing thrust reverser sleeves with sliders.

Upper and lower bifurcation joint loading not clear, as load paths 9.

are interrupted.

10. Transfer - of cascade basket radial loads is not clear.

Honeycomb panel edge closeouts are not a practical design.

11..

12.

Cowl vent areas, drain means, and blow-out pane "I areas are not identified.

13.

Thrust reverser cowl hinges, as shown, allow interference between the nacelle and the strut structure.

{ 16. Bulb type seals around thrust reverser are not suitable for this type Y of service. is recommended.

A pressure-on-lip seal 17.

Longitudinal seals should be provided in thrust reverser cowl.

Drainage and Vents 18.

No routing for engine accessory and strut drains is shown.

i 19. Cooling air and duct burst venting are not provided.

20.

Acoustic surfaces shown should provide for fuel drainage in the area of the primary mixer.

6.2 AIRFRAME ACCESSORY REQUIREMENTS AND LOCATION r" Hydraulic and electric Loads are shown in figures 6-2 and 6-3. These loads can be handled by one hydraulic pump and one alternator on each engine gear box.

Gearbox and accessory location studies generally have shown the core mounting to have the least weight and best performance; however, accessibility, especially in a long duct nacelle, is not as good as for chin-mounted accesssori es.

'Table 6-1 presents a general study of accessory location. A numerical rating system, where 0 is unacceptable and 5 is the best or most acceptable, was u to obtain an overall figure of merit. Recent surveys of Boeing customers There also showed that chin mounting and core mounting had widest acceptance.

-appeared to be a strong feeling against split gearboxes.

Gearboxes apparently are high-maintenance items and airlines believe that splitting a gearbox increases its maintenance problems significantly. Another important j (DOT/FAA order 8110.19) that consideration was the fuel spill requirement specifies that no fuel may be spilled during a wheels-up-landing. The to to certify chin-mounted gearbox and engine fuel pump would be difficult this requirement.

D6-48027

No.

REV SYM PAGE 54 D3 4203 7500 REV. 5/76 J X15-047 01 4100 7740 e)RrG.fl71 m x I NOTE: USE OF THREE PROPERLY INTEGRATED HYDRAULIC SYSTEMS ENSURtS THAT LGSS OF ONE SYSTEM WILL NOT DEGRADE SAFETY OF FLIGHT OR CAUSE INFLIGHT TURN6ACK.

1 S0 • J GRND 100% CRUISE 1000,6 0 GEAR EMER SPOILER LATERAL ETRACT DESCENT PITCH 1000/o LDG FLAP TAXI GEAR FLAP NORMAL EXTEND YAW FLAIR EXTEN RETRACT DESCENT low Z GROUND FLf -7I1T > o cl m L f CT A O N V 6-2 Hydraulic loods FIGURE J J Ell 4I00 7/40 OAiO.!/71 jig-n4?

x rn IV J Q Y IOC G Q O J .J Q v SO OC 1-

U

LOAD EQUIPMENT THAT MUST BE W POWERED ELECTRICALLY T ",,-Zoa W.

91t co N V FIGURE 6 . 3 Electric loa(ls

L

N• .jam.

:10 dwwww a a i Table 6-1 E 3 Engine Gear Box Location Study i —I)l it Pump_ Split Fan S Fuel fa i Core Mount Fuel Pumo Bottom Frame 600 and 2700 T:^ Fuel Spill per 5 5 0 5 0 DOT/FAA order 8110.19 Accessibility to 3 3 5 5

accessories

Heat rejection 2 5 5 5

Accessibility to 2 5 5 5 5 } variable IGV A Compatibility with 5 5 5 5-

load. reduction

Compatibility with 2 5 5 5 5 zero moment mount Customer Acceptance 4 0 0 0 5 30/0 _ 24 28/0 23/0 35/0 a Note: Rating 0 to 5, with 5 most acceptable and 0 not acceptable i k h F k a I No, n6-48027 REV SYM A C PAGE D3 4203 7500 REV. 5/76 '^ , Table 6-1 reflects these considerations and shows the core-mounted gearbox to be the only acceptable location.

6.3 MAINTAINABILITY, ACCESSIBILITY, AND SAFETY Maintainability, accessibility, and safety provisions were reviewed and found to be generally acceptable. The reference layout did not contain sufficient detail, nor was it sufficiently complete, to warrant detailed study of these features.

6.4 MOUNTING SYSTEM Boeing's practice is to design mount systems so that the mount can accept all engine models that might be used on a given airplane. Since _P&WA's mount system does not have this flexibility, Boeing made the preliminary design mount system shown in figure 6 -4 as an alternative to the P&WA mount.

I-- f

FIGUR- 6 4 I(- - VIEW L!-WQLLL - IM til l m LL1 iO4

EEE W tO

- - E Id

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I - v

p

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^& 404-i- Table 6-2 Nacelle and Strut Design Load Factors strut and primary engine mounts shall be designed for the The nacelle, nacelle following inertia load conditions which are assumed to occur only once in the lifetime of the airplane; Condition Ultimate load factors Vertical 6.5 6.5 + 1.5 T(c) _3.5 -3.5 + T(c) Thrush 3.0 T(max) + 3.0 vertical 3.0 T(max) + 1.5 vertical 3.0 T(R) 3.0 T(R) + 3.0 vertical y Side + 3.0

Gyroscope + 2.25 rad/sec y aw + 1.5T (c) + 1.5 vertical

+ 2.25 rad/sec. ypitch + 1.5T(c) + 3.75 vertical

Engine seizure Torque equivalent to stopping rotating mass in approximately 0.60 sec T(max) = maximum takeoff thrust at sea level cruise thrust (maximum or minimum, whichever is Where: T(C)

critial)

T(R) = reverse thrust

Note: For design purposes, these ultimate factors shall be applied at the

nacelle and content weight and C.G. exclusive of thrust and contents.

6 - 48027

ZEV SYM 182 - -

y PAGE 60 EV, 5/76 DI 4 10 7 74 OR Ii.j/71

m

N 3r X (UP) SIGN RIGHT HAND RULE FOR CONVENTIONZ (AFT) NOTE: - - MOMENTS THE FOLLOWING RESULTANT AIRLOAD CONDITIONS ARE

VIY OUTBOARD)

OCCURING ONCE PER FLIGHT AND THE ASSOCIATED VERTICAL LOAD FACTOR IS 1.0 g.

My

ALT. Fy Fz Mz

HRUS Fx Mx

Ve

CONDITION: LBS LBS 10 3 IN. IN.

LBS IN. 10 3 10 3

FEET KNOTS LB/ENG LBS I. BS MAX. TAKE OFF1D 0 126.0 36000 3945 2890 -2960 162.3 -317.0 1 -11.0

MAX. Q 2 M :72.3 15600 -3270 -2370 -4550 -104.6 207.3 6.1 I

18500 -3350 184.8 -334.8 -12.7

L 3 -V 0 FLAPS 17000 161.2 6220 3800

STALL, 0

-10.2

1.3 17000 161.2 18500 3070 2140 -1550 150.0 -283.6

V STALL , 100 FLAPS

aW - 16 DEG. AND aINlFr . x 12 DEG.

BASED ON SLS THRUST F - 45,500 LBS

SCALE: LOADS BY FF

Z- M OM ENTS BY Or) 7. O n n ^ O^ O^

N

V TABLE 6-3 Engine nacelle airloads I

x

w 1111111 U1 4 10 1740 FRIG.!/?I J1m-061 .X m 114RUST= 36,000 LB/ENGINE \/C DT I/ • AI nA1 CA/'T/\D _: • n t .. I r O 2 > O n

n p

U 1 N t^ IV

V

A• r, m N s ii t 4.

} If .f tj Y\j tt { ' a N V c \ 8l ^ r OSEE QV N 1 008'tCC _ 8l OZZ9 0' t = U = NOlJdA ,

(IdOl IV:)liS3A

13N "WON3/81 OOS'8L isml cro—u r- 9 L19 . 9Ido o ► cc OOI r SO r J75-047 Of 4100 7740 ORLG.3/71.. ) m N

D

i j i

m

.k

Co J 6.5 NACELLE DESIGN With the exception of the tailpipe material, Boeing was in general agreement with the materials shown on the reference P&WA nacelle layout. Boeing had good results with Kevlar/aluminum containment structures in laboratory experiments and the fan containment concept shown appeared feasible. Boeing used Dyna Rohr in the inlet cowling of the 737 for about two years and experience was acceptable.

Graphite/Kevlar fabric skins, with a metal core on the exterior of the inlet cowl, would be particularly vunerable to lightning strikes unless a lightning Protective surface and possibly a nonmetallic core are used. Use of aluminum brazed titanium honeycomb for the core cowl structure is satisfactory provided cowl skin temperatures do not exceed 800 0 F. Because the tailpipe could be subjected to temperatures above 1000 0 F, aluminum brazed titanium honeycomb is not recommended. ' Inconel would be a logical material selection for the tailpipe.

In Boeing practice, new materials selected for application to flight structures are subjected to a rigorous time consuming test and evaluation program. This evaluation consists of laboratory tests of candidate materials, destructive tests to determine allowables, noncritical service testing of Lightly loaded structure, and noncritical service tests of loaded structure.

This evaluation process may take several years, the actual time depending on the severity of the intended application. Candidate materials may be dropped at any time during the evaluation process.

6.6 NACELLE WEIGHT EVALUATION Table 6-4.

STF505M-7C Nacelle Weight Evaluation and Comparison of Boeing and PWA Nacelle Weights Nacelle Nacelle weight (lb/pod) Weight difference component SLST - 42 200 lb (PWA minus Boeing) Boeing PWA .

lb % Estimate Estimate Inlet 984* 587 -397 -40.4.

Fan cowl 164 136 -28 -17.1 TABLE 6-5. Comparison of Boeing and P&WA Advanced Technology Weight Reduction Factors Weight Reduction Factor (%) Nacelle Component PWA Boeing J 10-0^7

M m

N' l ; a i t C c ^O Q N

V

F'

X10-047

M

Remarks N Outer surface weight was reduced 25% to account for composites. Relative to total fan-duct and -core cowl weight, a this is about 5.5% reduction. Design t complexities need further Investigation and refinement before the P&WA 10% i reduction can be realized.

E Fan reverser weight provisions for ^ ,Y Installation are included, with reverser.

P&WA data showed reverser cascades to be the only area of advanced tech- nology. Cascade weight saving of 15% due to use of composites is about Z 5% of total fan reverser weight. Combined fan-reverser and fan-duct weight reduction factor is 4.61k Fan-reverser design needs refinement to be acceptable.

o a V 31e_oA7 M M CA Table 6-6 Weight Analysis Summary (continued) Nacelle Component Substructure and Material Remarks Mixer, plug and Mixer lobes; single- Data base for mixer includes expert- tallpipe thickness titaniu m mental work on daisy-lobe mixers { Lobe support struts and and analytical studies for the JT8D _--t, ring: Iconei and JT9D long-duct mixers.

^f Lobe fairing: aluminum P&WA gave minimum design Plug: lconel, thickness Insufficient structural definition.

as required by depth for frames and for nozzle to ' minimum welding fan duct attachment were Boeing gage criteria, concerns.

Difference in design philosophy in this area accounts for significant part of weight difference.

f ^j NN

V

^O , r 7.0 CONCLUSIONS AND RECOMMENDATIONS I 1.

NASA's stated fuel consumption goal is a 12% reduction of cruise TSFC.

For the Boeing study, this was interpreted to mean a 12% reduction of airplane BLKF. Under this interpretation, the STF505M-7C as installed in the Boeing Model 768-866 surpasses the fuel consumption goal by 4 1/2 to 6%, depending on the propulsion system weight used in the airplane performance study.

2.

Boeing evaluation indicated the STF505M-7C nacelle to be about 1310 lb heavier than the P&WA weight estimate. Using the heavier nacelle increases the fuel burned by about 1.3%.

3. The NASA goal of 5% DOC reduction is bettered by 1% using P&WA supplied engine performance, weight, and economic data. However, Boeing considers the engine price quoted by P&WA unrealistically low for E 3 technology levels. When the $590,000 higher Boeing price estimate is applied, the DOC improvement drops from 6 to 4.2%. The DOC reduction due to the higher Boeing weight estimate would cut back the improvement to about 4%, which does not meet NASA's goal of 5% DOC reduction.

1. EEE Component Development and Integration Program Boeing, subcontract No.

20646-1.

2. EEE Component Development and Integration Program, Boeing subcontract No.

20628-1.

3.

Preliminary Performance and Installation Data for the JT9D-7A Turbofan Engine, Pratt and Whitney Aircraft, February 14, 1977.

STUDY TRANSPORTS POWERED BY ENERGY EFFICIENT ENGINES This report summarizes work done under Pratt & Whitney Purchase Order 20646-2 as part • of Pratt & Whitney's prime contract NAS 3-30646, Energy Efficient Engine (E3) Component Development and Integrated Program.

This report completes the requirements of paragraph_ 2.1.1 and 2.1.2 of Task I and paragraph 2.2.1 of Task II of the Purchase Order Statement of Work.

i

Prepared b Y R. T. Kawai ACEE Propulsion Systems Project Manager J. H. Lindley Advanced Engineering Project Engineer Approved by awt n W. J Advanced Engineering Chief Program Engineer Commercial Systems ti L. A. Wrig t, Director Design En ineering - Power Plant r.

M. Klotzsche Director, Progra Engineering ACEE.

aA PREFACE This report presents results of a study conducted by the Douglas Aircraft Company as a subcontractor to Pratt & Whitney Aircraft to investigate appli- cations of engines based on use of NASA supported Engergy Efficient Engine (E3) Technology. This work was done under Purchase Order 20646-2 as a part of the Pratt & Whitney prime contract NAS 3-30646.

The studies reported herein were conducted to identify commercial transport aircraft which could oossihl y use enainec haseH nn tarhnnlnnv frnm tho NASA 1.0 Introductic 2.0 Study Aircr 2.1 Advanced Technology Features 2.1.1 Advanced tiling Design 1 2.1.2 Advanced High Lift System 3 2.1.3 Longitudinal Stability Augmentation System 3

2.1.4 Wing Load Alleviation 4

2.1.5 Composite Structure 5 2.1.6 Systems 6 2.2 Aircraft Descriptions 6 2.2.1 Aircraft Characteristics 6 2.2.2 Airplane Drag 1>3 2.2.3 Weight 2.2.4 Sensitivity Factors 18 2.2.5 Noise 2.2.6 Secondary Power 18 2.2.7 Comparison Between E3 and JT9D 23 2.3 Airframe/Propulsion System Integration 26 2.-3.1 Study Installations 26 2.3.2 Preliminary 1990 Propulsion System Requirements 26 Appendix I 34 A i 1.0 INTRODUCTION This study is based on aircraft which are advanced technology derivatives of DC-10 aircraft. This selection was arrived at from a solicitation of the views of Douglas marketing and engineering personnel.

Taking into consideration traffic growth forecasts, airline fleet compositions and technology development activities, the logical transports to utilize engines based on NASA E3 technology in the early 1990 time period appeared to be aircraft with increased seating capacity relative to the DC-10 and design emphasis on reduced fuel consumption. The need to minimize new development costs resulted in the selection of stretched DC-10's-employing advanced technologies. A domestic and an international version incorporating a 65-foot were configured fuselage stretch and a common area advanced technology winq to acquire a_large market base.

2.0 STUDY AIRCRAFT 2.1- ADVANCED TECHNOLOGY FEATURES The selection of advanced technology features was based on results from recent studies and on-going technology development programs.

2.1.1 Advanced Wing Design One of the prominent features of the advanced airplane is the new high aspect ratio wing using supercritical airfoil sections and winglets. Fundamentally, the supercritical airfoil generates greater amounts of lift for a given thickness and drag than a conventional airfoil. The distinguishing geometric

characteristics are a slightly blunter nose, a flatter upper surface and a

highly cambered thin trailing edge relative to a conventional airfoil.

The benefits provided by the supercritical airfoil for wing design can be "utilized in several ways. From purelyaerodynamic considerations the cruise

speed and lifting capability (buffet boundary) could be increased for the

same wing sweep and thickness. Because of the emphasis on fuel efficiency,

the application of supercritical airfoil technology to the E3 aircraft has

been to increase wing thickness while still achieving some benefits in buffet boundary. The increased wing thickness provides a structural weight advan- tage as well as an increase in takeoff and landing The increased CLmax' CLmax' improved buffet boundary and weight reduction due to the thickness increase, result in a reduction in wing area (and thus further weight reduc- tion). Part of this weight reduction has been utilized to increase the wing aspect ratio to reduce induced drag. Winglets in conjunction with the moderately high wing aspect ratio will provide a large induced drag reduction without the excessive wing span and the consequent large airport gate space requirements that result from the use of very high aspect ratios.

The wing design incorporates airfoil shape and thickness variations across the span to counteract wing-fuselage interference and other three-dimensional planform effects and to maintain as much of the two-dimensional drag- divergence Mach number capability of the advanced airfoils as possible. The wing twist and taper ratio are selected to produce minimum induced drag, considering the tradeoffs in wing weight and stalling characteristics.

NASA has done exploratory development of these advanced airfoils including.

flight testing on an F-8 research airplane. Douglas has designed, developed and fli ht'tested supercritical airfoils on two different wins on the YC-15 P 9 AMST prototype aircraft. Results from recent EET wind tunnel programs have substantiated_ that these advanced airfoils will provide the desired charac- teristics for a high aspect ratio wing application.

The winglet concept as well as the supercritical wing were wind tunnel tested by Dr. Whitcomb of NASA Langley, and have been under study for a number of aircraft applications. A joint USAF/NASA program is currently pursuing winglet installation on a KC-135A aircraft. In preparation for thisactivity, extensive wind tunnel testing at cruise speed and low-speed high-lift con- ditions has been conducted.

A winglet development program for potential application to the Douglas DC-10 is currently active. The winglet design has taken into account the experi- mental results of .Dr. Whitcomb. This design, in various forms according to the specific model of DC-10, was successfully wind tunnel tested at cruise ' speed in the NASA Langley eight-foot 'wind tunnel _in 1978 as .part of the NASA { - r ACEE program, and demonstrated the performance potential compared to wing tip extensions The program will continue development through 1979 in the low- speed high-lift regime and will evaluate the stability and control characteri- stics. Other concurrent work at Douglas is investigating the structural and other facets of the winglet installation. Continuation of on-going efforts forms the basis for the advanced wing design in the 1990 E 3 airplanes.

Advanced High-Lift System 2.1.2 The high-lift system features two-segment trailing edge flaps in conjunction with a variable camber Krueger leading edge flap. The two-segment flap provides high extension capability and the large chord forward segment and smaller chord auxiliary flap provide an optimum camber distribution. The flap is continuous from the side of the fuselage to 80 percent of the wing span, avoiding the h 7lgh-speed (inboard) aileron cutout and the associated loss of lift and ip'crease in drag. The full-span leading edge Krueger flap will allow for tailoring to provide good stall characteristics and control stall progression across the span.

This high-lift system design will provide excellent capability and very CLmax high lift-to-drag ratios allowing the use of a small wing area and engine thrust size. Maximum flap deflection is limited to 30 degrees to reduce by approach noise minimizing both approach thrust and airframe generated noise. An additional benefit is reduced fuel consumption.

Development work on this high-lift system design is proceeding, leading to application in the next generation Douglas transport aircraft. Extensive two dimensional wind tunnel testing and analytical configuration studies have been conducted in the last few years. Based on these results, three dimen- sional development testing will be conducted shortly in conjunction with the NASA ACES program.

stability characteristics similar -to those of the DC-10.. The more aft center y of-gravity location reduces the aerodynamic balancing down load carried by the horizontal tail. This results in lower trim drag and a weight savings due to the smaller horizontal tail and wing required. The aLSAS system provides positive stability for all flight conditions, ensuring the proper sense for control column motions and forces required for maneuvering the aircraft. The system employs pitch rate, pitch attitude and normal accelera- tion as feedback parameters to independent augmentation computers which provide control inputs in series with pilot commands to the four elevator segments and the horizontal stabilizer.

In order to explore thoroughly the requirements and interrelationships of aircraft configuration, flying qualities, safety and reliab i lity, , y g q y i y, control..

g system design and economics, Douglas has embarked on a study utilizing an A advance derivative of the DC-10 transport. h substantial portion 01 t117s task is proceeding under the ACEE program.

During 1977 an extensive piloted simulation, to explore aircraft flying qualities on the Douglas six-degree of motion simulator, was conducted.

During 1978 a further piloted simulation, which includes the effect of control system characteristics includ-ing failure cases and transient phenomena, is being conducted..

2.1.4 Wing Load Alleviation The use of control surface movement to regulate the net load and its distri- bution on the wing structure can be used to reduce bending moments and therefore reduce weight.

An additional advantage is that ride quality will be improved.

principally,' the application of these functions will be applied to the control of maneuver loads and gust loads.

The use of active systems for flutter suppression, which alters the apparent

mass or stiffness, or aerodynamic damping, is expected to be employed to

-provide appropriate flutter speed margins.

Even in the extremely unlikely

event of complete .system -failure, the • aircraft-will not be flutter critical

. within the normal operating envelope.'

0; ,t The use of control devices to limit load are not uncommon.

However, the full application of wing load alleviation in a transport'aircraft involves careful consideration not only of the technical factors, but also the regulatory requirements and operating factors such as dispatch reliability.

Advanced techniques to improve the design processes are under development, for example, by NASA in the ACEE program. In this program, large-scale drones, using a high aspect ratio supercritical wing with active controls, will be tested to correlate design techniques.

A number of other applications are also under study or development. In the transport field, a significant interest has developed into applications for current transports or their derivatives.

The Lockheed L-1011 experimental development, conducted partly under the ACEE program, is now flying.

At Douglas, design is proceeding for a system related to the DC-10.

Activity in this field is also to be pursued in combination with the ACEE program.

2.1.5 Composite Structure A Major advanced composite technology development activities have been underway for several years.

Douglas composite programs, with major funding support from NASA,_are leading to widespread application of composites in future transport aircraft. Current NASA sponsored advanced composite programs at Douglas include development of the DC-10 rudder, vertical tail and a wing study.

Expected application areas for composite materials in the next generation of transport aircraft include control surfaces, floor beams, fairings, landing gear doors and carbon brakes.

If emphasis is planed on continued composite technology development, by the early 1990's, design, fabrication and repair techniques should have advanced to the point that application areas may be expanded to include wing and empennage primary structure.

Use of composites in primary structures for the E3 study aircraft is assumed.

The fuselage pressure shell will still be of metal construction and will not have changed noticeably from current DC-10 designs except for the increased use of bonded metal structure and improved alloys.- Composite advantages include significant structural weight reduction, and with=;the falling price of composite materials relative to metals, minimum price escalation due to inflation.

tems Systems 2.1 .6 Improvements in all aircraft systems are expected. Some of these are: Digital avionics - reduced weight and improved reliability and o capability.

Flight performance management - reduced aircraft operational o fuel consumption.

o Air conditioning - reduced engine bleed requirements, APU - reduced weight and fuel consumption.

o o Advanced cockpit displays - reduced weight and improved performance.

These improvements, relative to current aircraft systems, can be incorporated' into future aircraft designs and are assumed in the 1990 E3 study aircraft.

2.2 AIRCRAFT DESCRIPTIONS Using the advanced technologies described with results from on-going studies and technology development programs, aircraft sizing studies were conducted using Douglas computer programs. The design requirements for the two aircraft, shown in Table 1, are based to a great extent on the DC-10-10 transcontinental and DC-10-30 intercontinental range aircraft. Design cruise Mach number was reduced from the DC-10 levels to reduce fuel consumption. The domestic and international aircraft are shown in Figures The only major external 1 and 2.

difference tetween the two versions other than engine size is . that the two wheel centerline main landing gear on the domestic aircraft is replaced with a four wheel assembly to cope with the international aircraft's higher weights.

2.2.1 Aircraft Characteristics The aircraft characteristics are shown in Table 2. The aircraft incorporate a DC -10 fuselage stretched 65 feet, a new high aspect ratio wing with super- critical airfoils and winglets, a new empennage and advanced aircraft systems.

The basic mixed class seating capacity is 458 passengers in the domestic version with lower deck galley and 438 in the intercontinental version with upper galleys. Interior arrangements for the two configurations showing the main and lower deck layouts are presented in Figures 3 and 4. Oversize cargo doors permit the accommodation of pallets in both the forward and center cargo compartments. The aft bulk cargo compartment is the same size as in the 6 ' 208

AIRCRAFT DESIGN REOUIREMEiVTS

INTERNATIONAL

5,500

0.80

11,000

31,000

FIGURE 2

GEN ERAL ARRANG EMENT -

INTERNATIONAL VERSION

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r] TABLE 2 =` a AIRCRAFT CHARACTERISTICS P&W STF505M-7C ENGINES DOMESTIC INTERNATIONAL MIXED CLASS SEATS 458 438 (NAUTICAL. MILES) 3,000 5,500 DESIGN RANGE 36,930 47,570 ENGINE THRUST SIZE (LB/ENGINE) ADJUSTED WING AREA (SQUARE FEET) 4,640 4,640 WEIGHTS: 496,000 MAXIMUM TAKEOFF (LB) 638,000 MAXIMUM LANDING (LB) 456,000 506,000 OPERATOR'S EMPTY (LB) 286,820 309,170 PERFORMANCE 0.80 0.80 CRUISE MACH NUMBER TAKEOFF FIELD LENGTH, MTOGW, SL, 84° F (FT)- 8,000 11,000 APPROACH SPEEDS, PASSENGERS;, BAGS, RESERVES (KEAS) 129 THRUST' LIMITED INITIAL CRUISE ALTITUDE (FT) 34,400 33,300 x BUFFET LIMITED INITIAL CRUISE ALTITUDE (FT) 36,500 31,000 FUEL BURNED AT DESIGN RANGE (LB) 99,380 (100% PASSENGER LOAD FACTOR) 207,630 TYPICAL STAGE LENGTH (NAUTICAL MILES) 1,000 1,500 FUEL BURNED AT TYPICAL RANGE (LB) (60% PASSENGER LOAD FACTORS) 51,340 32,800 (30% CARGO • E i FIGURE 3

DOMESTIC AIRCRAFT INTERIOR

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DC-10-30. The flight crew consists of a three man cockpit

crew and 15 cabin attendants.

Wing area, common to the two versions, is set by the 1.3 g buffet margin at the 31,000 foot initial cruise altitude requirement of the intercontinental range aircraft. The wing design incorporates the results of the latest wind tunnel tests and analytical studies. Lateral control is provided by spoilers and the all-speed outboard aileron. This allows the flap to extend from the side of the body to 80 percent span without interruption, and with the limited flap deflection of 30 degrees, results in lower required thrust levels and less noise. Wing load alleviation consisting of maneuver and gust load alleviation is used to reduce wing weight.

Horizontal tail aspect ratio has been increased compared to the current DC-10 to reduce trim drag.

The wing, horizontal tail and vertical tails utilize composites in primary to minimize weight.

,.and secondary structures The scaled thrust sizes of the P&W STF505M-7C engines are set by the design The takeoff field length requirements for both versions of the aircraft.

cruise altitudes exceed requirements by 1400 and 2300 thrust limited initial feet respectively for the domestic and international aircraft-, indicating a small surplus of cruise thrust relative to the takeoff rating.

2.2.2 Airplane Drag

Nacelle drag is

are shown in Table 3.

The airplane parasite and induced drag

in

The compressibility drag increment is shown included in the engine data.

in Figure 6.

drag polars are presented

The takeoff and landing

Figure 5.

2.2.3 Weight The weights are based on Airframe weight breakdowns are shown in Table 4.

technology advancements including widespread use of advanced composites.

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NACELLE DRAG NOT _ _ / INCLUDED _ .

-- .04 .06 .03 .10 12 .114 .16 C D A 1 F TABLE 4 B AIRCRAFT WEIGHT BREAKDOWNS STF505M-7C ENGINES DOMESTIC INTERNATIONAL 59,200 WING 54,690 4,220 4,;700 HORIZONTAL TAIL 1,960 2,130 VF.R.T T LCAL TAIL 61,910 63,040 FUSELAGE 26,490 LANDING GEAR 20,230 34,010 449270 PROPULSION* 1,435 APU 1,435 2,130 FUEL SYSTEM 2,130 10,470 10,470 FLIGHT CONTROLS AND HYDRAULIC SYSTEM INSTRUMENTS 1,750 1,750 4,965 AIR CONDITIONING AND PNEUMATICS 4,965 6,460 6,460 ELECTRICAL 32060 AVIONICS 2,700 s FURNISHINGS. 53,290 51,980 ICE PROTECTION 650 f HANDLING GEAR 60 MANUFACTURER'S EMPTY WEIGHT 260,930 282,790 25,890 26,380 OPERATOR'S ITEMS OPERATOR'S EMPTY WEIGHT 286,820 309,170 *Includes lower vertical tail 2.2.4 Sensitivity Factors Sensitivity factors were generated and are shown in Table 5. These factors provide a means to assess the impact of perturbations in specific fuel consump- tion, engine weight and nacelle drag on aircraft weights, engine size and fuel burned for the study missions.

2.2.5 Noise The airframe or non-propulsive noise with flight conditions and engine power settings at the FAR noise measuring points are shown in Table 6 Tabulated noise spectral data are presented in Appendix I.

2.2.6 Secondary Power The secondary power requirements have been estimated and the mechanical power requirements are shown in Table 7. For hydraulic power, the time average cruise requirement in still air (without turbulence) is 31 horsepower per engine. This is based on hydraulic pumps in average condition with nominal aircraft hydraulic system leakage.. The maximum or sizing requirement for hydraulic power is for two pumps per engine operating at full capacity. One hundred seventy five horsepower per engine is required for pumps that have had considerable usage.

The time average accessory gearbox power required by the generators is 75 horsepower per engine. This is based on a survey made on power usage in the DC-10. The DC-10 average power usage was scaled up to provide for the 'increase in number of passengers in this study. The maximum or sizing requirement is 257 horsepower per engine.

The average pneumatic power required in the form of compressor bleed is shown Figure 7. .

in _ c The maximum bleed case is for one pneumatic system out and an engine out, under icing conditions. For this case, at a 15,000 foot hold condition, it is estimated that.one engine must provide 0.7 pounds/second inlet cowl anti-ice .

• ^'^ with a bleed temperature greater than 500°F plus' 5 pounds/second win g anti-ice flow at a temperature greater than 400°f plus '2.7-pounds/second to provide air - + 5% + 1000 LB + 20% ISOLATED NACELLE DRAG TSFC WEIGHT PER ENGINE DOMESTIC + 0.7% ' MAX TAKEOFF WEIGHT + 2.4% + 1_.2% + 0.5% i OPERATOR'S EMPTY WEIGHT + 1.7% + 1.8% + 1.3% 0.7% ENGINE THRUST SIZE + 2.1% r FUEL BURNED + 1.8% DESIGN MISSION + 6.1% + 0.9% ^ + 6.0% + 1.0% + 1.4% TYPICAL MISSION INTERNATIONAL + 0.9% + 4.0% + 1.2% MAX TAKEOFF WEIGHT + 0.6% OPERATOR'S EMPTY WEIGHT + 2.4% + 1.7% + 4.3% + 1.3% + 1.0% ENGINE THRUST SIZE FUEL BURNED + 1.5% + "7.5% + 1.0% s DESIGN MISSION + 5.8% + 0.9% + 1.0% — TYPICAL MISSION , {

TABLE 6

CONDITIONS AT FAR-36 MEASURING'' POINTS

N

SCALED STF505M-7C ENGINES Airframe Geometric True Installed Thrust Generated Altitude Airspeed per Engine Noise Aircraft Condition (FT) (KPJ) (% Takeoff) E( PNdB) Domestic Sideline 850 150 100 76.8 Takeoff 1500 151 100 78.1 Cutback 1412 151 69 78.5 Approach 394 146 18 91.6 `International Sideline 850 167 100 79.7 Takeoff 1152 167 100 83.1 C) Cutback 1072 167 70 83.7 Approach 394 153 18 92.7 L=; fj j t, a cj t - a.

TABLE 7 ACCESSORY GEARBOX POWER REQUIREMENTS SIZING TIME AVERAGE CRUISE REQUIREMENT POWER IN STILL AIR TYPE SOURCE .

175 HP/Engine Two 35 GPM Pumps 31 HP/Engine Hydraulic per Engine Power 257 HP/Engine 75 HP/Engine - Electric- One 120 KVA Power Generator per Engine tN c W C i c FIGURE 7 - t E STUDY AIRCRAFT..COMPRESSOR BLEED AIR REQUIRED 3 w rH ^^^ ^ r- • ^ ' ! 1 1 1 r ICE PROTECTION NOT INCLUDEDS I .

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cc • t r _^.._. t ..._ ...._ t 50 1 c 20— 40 0 10 a .c a ^ r ._ .„ 1 1= z ALTITUDE=(THOUSAND,FEET) Na 224 to drive one air conditioning pack. The sizing case therefore requires a '60% of climb thrust.

total of 8.4 pounds/second with the engine at 40 to The above values reflect preliminary analyses of a current test program to reduce bleed flow requirements for wing anti-icing.

Further evaluations may result in requirements to revise the wing anti-icing flow requirements. In addition, potential means to reduce bleed flow require- ments have been identified but sufficient work has not been done to reflect these reductions in this study.

And JT9D 2.2.7 Comparison Between E 3 In order to determine the fuel consumption benefits from E3 engine tech- nology, the advanced technology airplanes were sized using JT9D engines for TABLE 8 AIRCRAFT CHARACTERISTICS &W JT9D-20 ENGINES P DOMESTIC INTERNATIONAL MIXED CLASS SEATS 458 DESIGN RANGE (NAUTICAL MILES) 3,000 5,500 ENGINE THRUST SIZE (LB/ENGINE) 42,150 53,600

ADJUSTED WING AREA (SQUARE FEET) 5,260 5,260

WEIGHTS: MAXIMUM TAKEOFF (LB) 548,000 723,000 MAXIMUM LANDING (LB) 482,000 542,000 OPERATOR'S EMPTY (LB) 309,070 336,870 PERFORMANCE: CRUISE MACH NUMBER 0.80 0.80 TAKEOFF FIELD LENGTH, MTOGW, SL, 84°F (FT) 8,000 11,000 APPROACH SPEEDS, PASSENGERS, BAGS, RESERVES (KEAS) 120 126 THRUST LIMITED INITIAL CRUISE ALTITUDE (FT) 34,100 32,700 `} BUFFET LIMITED INITIAL CRUISE (FT) 37,100 31,000 ALTITUDE (LB) FUEL BURNED AT DESIGN RANGE (100% PASSENGER LOAD FACTOR) 124,260 256,880 TYPICAL STAGE LENGTH (NAUTICAL MILES) 1,000 1,500 FUEL BURNED AT TYPICAL RANGE (LB) (60% PASSENGER LOAD FACTORS) 41,210 63,160 } (30% CARGO 226 24 TABLE 9 i_ COMPARATIVE AIRCRAFT CHARACTERISTICS q, Domestic International Engine JT9D-2O STF505M-7C JT9D-20 STF505M-7C Maximum Takeoff Weight (LB,) 548,000 496,000 723,000 638,000 Operator's Empty Weight (LB) 309,070 286,820 336,870 309,170

Takeoff Thrust (LB) 42,150 36,930 53,600 47,570

Design Range (N MI) 3,000 3,000 5,500 5,500 256,880 Fuel Burned At Design Range (LB) 124,260 99,380 207,630 Relative Fuel Burned At Design Range (LB) --- -20% --- -19% Typical Range (N MI) 1,000 1,000 1,500 1,500 e Fuel Burned At Typical Range (LB) 41,210 32,800 63,160 51,340 t ' N Relative Fuel Burned At Typical Range (LB) --- -20% --- -19% fi N J E y .

2.3 AIRFRAME/PROPULSION SYSTEM INTEGRATION Preliminary propulsion system integration requirements were investigated.

Study engine installations provided by Pratt & Whitney were reviewed and requirements for installation in the E3 study aircraft were determined.

2.3.1 Study Installations Table 10 summarizes preliminary results of evaluations of the study engine installations provided by Pratt & Whitney. The evaluations were preliminary assessments of the aerodynamic fines and general arrangements.

The aerodynamic critique on P&W Drawings L-108594 dated 1/15/78 and L-109846 dated 6/8/78 indicates a potential_ for excessive nozzle afterbody angle.

These statements are based on a certain degree of uncertainty because the angles exceed those for which data is available.

The adequacy of flow directivity in the reverser configuration column refers to the ability to incorporate a directed flow reverser which will preclude debris ingestion due to reverse flow impingement onthe ground.

Pylon-mounted accessories are judged unacceptable because it would require special equipment in order to conduct servicing,_ inspection and maintenance, particularly on the tail engine. Airframe accessories must be easily removed and replaced without requiring the airplane to fly to a maintenance base. In addition, the pylon-mounted accessories can preclude simultaneous maintenance or servicing the engine and accessories because the open engine cowl door would interfere with access to the pylon. Further, the additional time required would result in additional flight dela s and cancellations.

y Experience with fan cowl-mounted accessories has demonstrated that this arrangement is satisfactory for maintainability. By comparison, the other arrangements.are judged to be poor.

Preliminary 1990 Propulsion System Requirements 2.3.2 New engines are introduced because they result in a major- improvement in economics, provide the thrust requirement for anew airplane size, or both.

In the 1990's, a new engine based on E 3 technology will be expected to improve economics because the thrust sizes of interest are expected to be d TABLE 10 INSTALLATION STUDIES SUMMARY STUDY INSTALLAT T ON 11 D u "r ^. ace a N :ensions (Inches) Dou las CL to Max R 1 Aerodynamic Pratt & Whitney -Interpretation i Critique Description Length Top Side ^8ottom Drawing Number •' DDig. No.

STF 505 M-7 Engine - wing instl. , Nozzle afterbody L-108594 Split accessory arrangement angle wa y be exces- dtd. 1-15-78 --- I fan mounted engine accessories 252.7 ' 55.0 55.3 55.0 sive. Outside of with airframe accessories in II data base.

pylon Wing Instl.

L-108594 J-112536 - STF505 M-7 Acceptable and i Split accessory arrangement L-108623 core mounted engine accessories 258.0 55.0 55.3 55.0 1 dtd. 5-5-78 with airframe accessories in F pylon - P&W design fan reverser I i^ Tail instl. - STF 505 M - 7 J-112539 Adaptation of J-112536 for 53.4 62.0 53.4 ! Acceptable different location i L-106594 J-112540 Wing instl. STF 505 M - 7 i v modified per Full duty core-mounted acces- 258.0 55.0 55.3 55.0 Acceptable L-108613 sory {p ackage DAC type fan re- dtd. 4-18-78 verser.

' i Wing Instl. - STF ,505 M-7 ' J-112541 Full duty core-mounted acces- 258.0 55.0 55,0 I Acceptable 1 51-1 sory pkg. DAC design fan re- verser-. Study dwg. for nacelle 1 i seals & latches STF 505 M-7 Engine - wing instl.f Nozzle afterbody L-109846 Split accessory pkg., core angle may be exces- dtd. 6-8-78 ---- mounted engine accessories with i 262.0 55.0 sive. Outside of airframe accessories in pylon I da^a base.

P&!d design fan reverser.

L-109846 and 5G.8 J-112545 Wing instl. - split accessories L-108620 @ 4 at 4 and 8 o'clock on fan case I 262.0 55.1 56.0(60 dtd. 5-5-78 } & 8 io'clock) 1 L-109846_ Wing instl.

marked 'up for ---- sl'im nacelle

N

^o line N s TABLE 10 (CONTINUED) INSTALLATION STUDIES SUMMARY

o

Accessory Structural Reverser Douglas Corranents Critique Configuration Configuration Maintainability I LntPrpra,tz,tipn Does not appear tc^ Pylon Mounted Poor Special ground stands for accessory Appears have adequate flow accessories unaccep- maintenance and replacement is reasonable directivity provisions table for tail unacceptable engine Appears Does not appear to Pylon mounted un- Poor See above J-112536 ' reasonable have adequate flow acceptable plus po- directivity provisions tential reduced re- liability for engine accessories F t` very poor See above Appears Pylon mounted J-112539 unacceptable reasonable ---- J-112540 Appears Appears reasonable Potential reduced Poor See above reasonable reliability for accessories E $ J-112541 Appears Appears reasonable Potential reduced Poor Study of nacelle seals and latches u reasonable reliability for accessories See above Appears Does not appear to Pylon mounted i Poor t) '--- reasonable have adequate flow unacceptable directivity provi- a sions i J-112545 Appears Appears reasonable Good May have weight and performance penalty reasonable ----- for addition paler take-off shaft --__ Insufficient information to evaluate t available from current and derivative versions of JT8D refan, JT9D, JT10D, CFM56, CF6 and RB211 engines. Since the E 3 goal is'to reduce specific fuel consumption by 12% and DOC by 5%, other cost components cannot increase, and may have to decrease to provide sufficient incentive for development of a new engine. It is therefore expected that other costs should improve, or at worse, remain the same. This needs to be accomplished while meeting more stringent - regulations and requirements.

Maintenance The installation maintainability goals should be comparable to today's standards.

This requires access to all borescope ports without removal of any component. Elapsed time goals are shown in Table 1'l.

Thrust Reversers Thrust reversers should be improved compared to current designs. Specific needs are listed below.'

• 1. Fan thrust reversers with efflux directivity that minimizes debris kickup while enabling routine use down to zero speed are desired. Directivity tailoring capability must exist to match airframe; requirements to main- tain airplane control and drag.

2. The overall reverser effectiveness goal is 40% for the primary plus fan on wing engines. Tail engine reversing effectiveness can be lower to prevent aircraft pitchup.

Current fail-safe design practice for ground only reversing will be 3.

maintained. The reversers will maintain their position in the event of an actuation system failure.

4.' A hydraulic actuation-system is preferred with reverser hydraulic fluid isolated from other airframe hydraulic fluid.

Ozone (r Consideration should be given to providing bleed air for cabin air conditioning r Since elevating the that has an ozone concentration of less than 0.1 ppm.

temperature of air containing ozone will destroy the ozone, heating and cooling the bleed air may be a viable_'•way to reduce the ozone concentration- in the cabin,

i

TABLE INSTALLATION °ELAPSED TIME GOALS DESCRIPTION ELAPSED TIME (Minutes) Engine Build Up Neutral QEC from Basic Engine 2000 Build UP Neutral QEC to Wing QEC Build UP Neutral QEC to Tail QEC Convert Wing QEC to Tail QEC 45 Convert Tail QEC to Wing QEC 45 Change, Wing Engine (Including Access Time and GSE) Engine 90 Change Tail Components/Accessories Remove and Replace Integrated Drive Generator 35 Hydraulic Pump 15- Fire Detector 15 Main Fuel Control 25 Fuel Pump 64 Fuel Heater 30 Primary Nozzle 90 Exhaust Plug with Primary Nozzle Removed 10 Exhaust-Plug with Primary Nozzle Installed 15 Fuel Heater Air Shutoff Solenoid Valve 10 Anti-Icing Air Shutoff Actuator Valve TABLE 11 (CONTINUED) INSTALLATION ELAPSED TIME GOALS DESCRIPTION ELAPSED TIME (Minutes) Components/Accessories Remove and Replace 'K Differential Pressure Switch 6 Nose Cowl Anti-Icing Pressure Regulator and Shutoff Valve Starter Starter Shutoff Valve Hydraulic Filters Fuel Flow Transmitter Ignition Exciter Ignitio p Plugs 7 Pressure Ratio Bleed Control B Compressor Stator Control 23 -Fan Air Case Cooling Shutoff Valves Bleed (Air/Fuel) Converter Valve Bleed Control Valves Pneumatic Pressure Regulating Valves Bleed Check Valves y Y Bleed Air Cleanliness Bleed ports must be designed to prevent the ingestion of solid particles that enter the engine inlet, or liquids (such as might be generated within the engine by fluid leakage), without unnecessarily sacrificing total pressure recovery.

' Because an engine compressor acts as a centrifugal separator, clean air may be extracted at the compressor inside diameter without significant loss of ram pressure. The associated disadvantages are the cost of making hollow stator vanes suitable for conducting this air to the outside diameter of the engine, and the pressure drop of the flow traversing these relatively small passages.

Outside diameter ports that are protected by locating them in a shadow zone sacrifice ram pressure but may be designed to provide clean air as Tong as the engine is running. When the engine is stopped, fluids can draw into such openings if they occur at a low point.

Desirable Stage Locations For Bleed Ports Bleed air must be available from the compressor discharge to accommodate operation at engine idle.

For economy reasons, bleed must be available at the lowest stage that will satisfy air conditioning system pressure requirements at maximum altitude with the lowest engine power useful for cruise. If the maximum altitude for the baseline airplane is 39,000 feet, a bleed pressure of 20 psig would

Pressures as low as 15 psig could be

permit using DC-10 type components.

considered if the associated economy improvement would justify the development of new and possibly more complicated air conditioning components.

An additional, lower stage port located so that the discharge temperature closely approached but did not exceed 450°F on a hot day sea level takeoff stage bleed, and would open the would eliminate the need for precooling low eliminating all precooliog. Complete elimination of pre- possibility" of Changing from cooling could only be justified by a thorough investigation.

DC-10 to DC-9 pneumatic system concepis for providing suitable ice protection bleed temperatures would probably be required. The investigation would have to include a study of the pressure suitability of the next lower stage pressure whenever high stage bleed exceeds 450°F at'idle power on a hot day, Any pressure above 25 psig at this lower stage would be satisfactory.

A completely independent port for engine inlet ice protection air supply is desired, located at compressor discharge, or preferably a lower stage if it would provide 400°F at engine idle power with ambient temperatures at the low limit of the FAA icing envelope.

Containment In addition to rotor blade containment requirements of FAR Part 33, any blade fragment exiting from the engine shall not have sufficient energy to penetrate nacelle structure or systems..

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LR 28664 23 JUNE 1978 rIwl w I nrnnn-r

LR 28664 TABLE OF CONTENTS Section Page LIST OF FIGURES v LIST OF TABLES vii 1 INTRODUCTION AND SUMMARY 1-1 2 STUDY EFFORT 2-1 MISSION AND DESIGN DEFINITION 2.1- 2-1 -2.2 PROPULSION SYSTEM - AIRCRAFT INTEGRATION 2-1 STF505M-7C Engine Evaluation 2.2.1 2_ -1 Sensitivity Analysis 2.2.2 2-3 Airframe Noise Estimates 2.2.3 2-4 2.2.4 Engine Bleed Requirements and Power Extraction 2-4 2.3 ENGINE INSTALLATION 2-11 2.3.1 Nacelle Configuration 2-11 2.3.2 Nacelle-Wing Interference 2`-12 2.3.3 Accessory Location 2-12 -13 2.3.4 Access 2.3.5 Thrust Reverser ns 2 -13 2.3.6 Center Engine Installation 2-16 2.4 PERFORMANCE AND ECONOMIC COMPARISONS 2-16 CONCLUSIONS AND RECOMMENDATIONS 3. 3-1 DIRECT OPERATING COST (DOC) CALCULATIONS - E 3 AIRCRAFT A-1 APPENDIX A x _ 238._, LR 28664 LIST OF FIGURES Figure Page 1 General Arrangement Domestic Aircraft (STF505M-7C Engine) 1-7 - 2 General Arrangement Intercontinental Aircraft (STF505M-7C Engine) 1-9 Wing Engine Installation Layout 1-11 Center Engine Installation 4 1-13 Wing Engine Location 5 1-15 6 ASSET Synthesis Cycle 2-4 Domestic Aircraft Sensitivity Factors 7 for ATSFC 2-5 8 Domestic Aircraft Sensitivity Factors for DEngine Weight 2-6 9 Domestic Aircraft Sensitivity Factors for ONacelle Drag 2-7 1.0 Intercontinental Aircraft Sensitivity Factors for aTSFC 2-8 11 Intercontinental Aircraft Sensitivity Factors for AEngine Weight 2-9 Intercontinental Aircraft Sensitivity Factors for ONacelle Drag 2-10 13 Thrust Reverser Flow Directivity 2-15 14 Block Fuel Advantage with STF505M-7C Engine 2-17 15 DOC Advantage with STF505M-7C Engine 2-18 Size Advantage with STF505M-7C Engine 2-20 v _ t LR 28664 "N LIST OF TABLES Page Table 1-3 Reference Aircraft Design and Performance Characteristics Aircraft Design and Performance Characteristics 2 E3 1-5 Design and Technology Features-1990's Transport Aircraft 2-2

LR 28664 SECTION 1 INTRODUCTION AND SUMMARY This study was accomplished by the Commercial Advanced Design Division of the Lockheed-California Company for the Pratt and Whitney Aircraft Group in support of their "Energy Efficient Engine Component Development and Integra- tion Program," The effort required was in accordance with Pratt and Whitney Subcontract 20646-3 and consisted of the initial. Propulsion System Aircraft Integration Evaluation as specified by Task 1. This initial evaluation was in support of Pratt and Whitney's engine preliminary design effort and two add j.- tional evaluations will be made by Lockheed during the program as follows: • Initiation of engine core manufacturing and testing - mid 1980 • Completion of integrated core/low spool testing - mid 1983

This evaluation is an update or follow -on to the previous Lockheed study

effort in su ort of the "Energy Efficient Engine Preliminary Design and Inte- gration Study," Pratt and Whitney Subcontract Number 20628-3 which included the following: • Definition of airframe design and technology features • Aircraft and mission definition • Aircraft performance and mission sensitivities • Aircraft-engine integration evaluation .During the previous study effort, Lockheed Report LR 28351, two air- craft configurations were developed; one for a domestic mission and one for an intercontinental mission. These domestic and intercontinental aircraft c (using the JT9D-7A engine) were characterized for the following technology features and mission criteria:' e `f a r r: E# I-1 241' ^ LR 28664 • Technology Features • Supercritical wing • Active controls • Advanced composite structure Mission Criteria • Domestic Intercontinental Design Range (n.mi.) 3,000 6,500 No. passengers 400 Cruise speed M 0.8 M 0.8

1,400

Typical range 3,000-

Configuration 3-Engine-Wide Body 4-Engine-Wide Body For this study,reevaluation of aircraft technology features and mission criteria resulted in the retention of previously established criteria, except for the passenger/payload capacity. A payload capacity of 100,000 pounds (500 passengers) was incorporated in lieu of 80,000 pounds (400 passengers) previously used. This change was made based on a review by Lockheed's Market- ing Development Division relative to potential market demand in the 1990's Reference aircraft design and performance characteristics time frame. con- sistent with the increased payload capacity are included in Table 1. These configurations were established as baseline aircraft to be used for comparison; with aircraft incorporating the Energy Efficient Engine.

The Energy Efficient Engine cycle selected by Pratt and Whitney for in- stallation on the domestic and intercontinental aircraft is the STF 505M-7C with the following characteristics, as compared to the current JT9D-7A engine: JT9D-7A STF505M-7C Technology Level Current 1990's Fan Drive Direct Direct Exhaust Separate Mixed 6.55 Bypass- Ratio 5.0 25.4 38.6 Overall Ratio Turbine Inlet Temp 2290 2450 1-2 LR 28664 TABLE 1. REFERENCE AIRCRAFT DESIGN AND PERFORMANCE CHARACTERISTICS Domestic Intercontinental Mission Characteristics Design Range (n.mi.)

3000 6500 Typical Range (n.mi.) 1400 3000 Cruise Speed (Mach) 0.8 0.8 No. Passengers 500 500 Init. Cruise Altitude (ft) 35,000 33,000 Field Length (ft) 6970 9398 Approach Speed (kt) 135 128 Design Characteristics Configuration 3 Engine-Trijet 4 Engine-Quadjet Power Plant JT9D-7A JT9D-7A Sweep (.25c) 300 (lb/ft 2 118 134 W/S ) 0.260 0.220 T/W AR 10 10 (%) 13 t/c 13 TOGW (lb) 481,357 707,924 OEW (lb) 261,934 303,985 Wing Span (ft) 202.0 229.8 Body Length (ft) 228,3 229.5 Body Diameter (ft) 19.6 19.6 Performance Characteristics Thrust/Engine (SLS, lb). 41,718 38,936 99,99`9 263,686 Block Fuel-Design (lb) -Block Fuel-TYP. (lb) 43;352 101,781 DOC-Design (c/ASM) 1.227 1.414- DOC-TYP. ((, , /ASM) 1.336 1.407 r LR 28664 Table 2 is a tabulation of the aircraft design and performance character- istics of the domestic and intercontinental aircraft with the STF505M-7C engine. Comparison of this data with the performance of the reference air- craft (JT9D-7A engine) indicates mission fuel and direct operating cost (DOC) savings with the STF505M-7C engine as follows; Fuel DOC Design Typical Design Typical 18.1% 18% 5.9% 5.4% Domestic 20.1% 19.1% 9.2% 8.1% Intercontinental General Arrangement Drawings, depicting the domestic and intercontinental aircraft with the STF505M-7C engine, are included as Figures 1 and 2. The size of the STF505-M-7C engine, as supplied by Pratt and Whitney, is compati- ble (thrust class, reverse thrust level, and power extraction) with the Lockheed specified mission/payload characteristics for 1990's aircraft.

Installation layout drawings using the STF505M-7C engine on the domestic aircraft are included as Figures 3 through 5, and depict location of aircraft accessories in the engine pylon and placement of the nacelle with respect to the wing consistent with minimization of interference drag penalties.

The results of this phase of the Energy Efficient Engine Component Devel- opment and Integration study are as follows: • The NASA defined goals for minimum fuel and DOC savings of 12% and 5% respectively are attained with the STF505M-7C engine.

• Installation of the STF505M-7C engine (with mixed exhaust), without a penalty for interference drag, appears feasible.

• Pylon mounting of the aircraft accessories is an acceptable configura- tion and enhances the aerodynamic characteristics of the STF505M-7C nacelle.

• Incorporation of the STF505M-7C engine results in aircraft configura- tions, sized for long range and large payload capacity, which are compatible with existing airport facilities ,(field length, wing span, body length, etc.)

- 1-4 n.

TABLE 2. E 3 A Domestic Intercontinental Mission Characteristics Design Range (n.mi.) 3000 Typical Range (n.mi.) 1400 Cruise Speed (Mach) 0.8 0.8 No. Passengers 500 Init. Cruise Altitude (ft) 37,000 34,000 Field Length (ft) 6976 Approach Speed (kt) -535 Design Characteristics Configuration 3 Engine-Trijet 4 Engine-Quadjet Power Plant STF505M-7C STF505M-7C Sweep (.25c) 30°' 30° W/S (lb/ft 2 ) 114.8 132 T/W 0.255 0.220 AR 10 10 - t/c% 13 13 TOGW (lb) 454,013 630,491 OEW (lb) 255,937 286,974 Wing Span (ft) 198.9 218.6 Body Length (ft) 228.3 229.5 Body Diameter (ft) 19.6 19.6 Performance Characteristics Thrust/Engine (SLS, lb) 38,591 34,677 Block Fuel-Design (lb) 81,862 210,888 Block Fuel-Typ. (lb) 33,513 .82,387 DOC-Design (C/ASM) 1.155 1.`285 DOC-TYP. ((/ASM) 1.263 1.293

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t s 01t 1GINAL PAGE 1g oli POOR QUAITII^ The study effort by Lockheed in support of Pratt and Whitney's Energy Efficient. Engine Component Development and Integration Program consisted of an evaluation of integrating the E 3 propulsion system with the domestic and intercontinental aircraft, as envisioned for the 1990's time frame. The evaluation included first establishing aircraft mission and design definitions and then incorporating the advanced technology engine into the aircraft con- figurations for comparison with the reference aircraft (JT9D-7A engine),_ 2.1 MISSION AND DESIGN DEFINITION Mission and design definitions, along with applicable advanced technology features, were established for both the domestic and intercontinental aircraft a during the previous study effort (Lockheed Report LR 28351). On initiation of this effort, those definitions were reviewed, and updated where applicable, m for the purpose of establishing reference (baseline) configurations and per- formance, characteristics for comparison of those aircraft with the E 3 engine.

Definition of the domestic and intercontinental aircraft is included in Table 3.

2.2 PROPULSION SYSTEM AIRCRAFT INTEGRATION } 2.2.1 STF 505M-7C Engine Evaluation Performance, weight, and pertinent installation data for an advanced technology energy efficient engine (identified as STF505M-7C) was by supplied Pratt and Whitney for incorporation into the reference aircraft. Both the { domestic and intercontinental aircraft were previously optimized (for minimum fuel usage and DOC) using the Pratt and Whitney STF505M-7 Engine. Since- the STF505M-7C represented only a slight variation,_ previously established r?

F` 2-1 LR 28664 TABLE 3. DESIGN AND TECHNOLOGY FEATURES-1990'S TRANSPORT AIRCRAFT Domestic Intercont.

Wide body trijet Wide body quadjet Aircraft Type 235 in. fuse. dia. 235 in. fuse dia.

9 abreast seating 9 abreast seating No. Engines and Location -2-wing mounted 4-wing mounted 1-center mounted Payload Capacity (lb) 100,000 (500 pax) 1001000 (500 pax) TOGW Class (lb) 500,000 750,000 Engine Thrust (lb) 45,000 46,000 ^. _,, Mission Characteristics " f1f f Design Range (n.mi.)_ 3,000 69500 Typical Range (n.mi.) 1,,400 3,000 _ Range L.F. 0.55 0;55 -Typ.

M0.8 MO.8 Cruise Speed 35,000 35,000 Cruise Alt. (ft) 7,000 10,000 TOFL (ft) 15 135 App. Speed (kt) Advanced Technology --3% reduction of -•3% reduction of Supercrit. Wing wing wt - increased wing wt - increased thickness of airfoil thickness of airfoil • AR = 10 • AR, = 10 /c = 13%

• t/c = 13% • t

• Sweep = 30' • Sweep = 30° -5.5% wing wt. -5.5% wing wt.

Active Controls -1% body wt. -1% body wt.

• Load Relief -28% tail size • Relaxed Stability -28% tail size Advanced Composites -8.7% M.E.W. -9.2% M.E.W.

i^ Primary Struct.

• Secondary Struct.

LR 28664 design parameters were retained and aircraft performance evaluated using the revised data for the STF505M-7C Engine. Design and performance characteristics are shown in Table 2 and detailed tabulations of aircraft design and perfor- mance characteristics are included as Appendix A to tris report.

Performance evaluation of the domestic and intercontinental aircraft with the STF505M-7C engine was accomplished using the Lockheed Parametric Analysis (ASSET) Program, Figure 6. The ASSET Analysis Program is a Lockheed proprietary synthesis model to parametrically size and determine the weight, performance, and cost of aircraft sized to meet given mission profiles, payload capacity, and structural criteria using a preselected optimization criteria. For this study, minimum mission fuel and direct operating cost were the optimization criteria utilized for sizing both' the domestic and intercontinental aircraft. The pru cedure for calculating DOC, and the associated cost factors, for this study effort are included in Appendix A.

2.2.2 Sensitivity Analysis Sensitivity factors were calculated for each aircraft, with the STF50514-7C engine, to assess the effects of changes in cruise TSFC, engine weight, and isolated nacelle drag on aircraft performance. As specified by the subcon- tract, the following sensitivity factors were calculated: ±1000 lb ±2% s Eng. Wt.

±5% TSFC Nac. Drag

X

TOGW X OEW X X , K Engine Thrust X X K i Mission Fuel Design X X K X }{ Typical X The resultant sensitivity factors are depicted in Figures 7 through 12. - 2-3 ,;J r_ '7' LR 28664 CANDIDATE CONCEPTS PRELIMINARY DESIGN MASS AERO-

I

I

MATERIALS COST PROPULSION DESIGN I I PROPERTIES DYNAMICS MISSION PROFILE VARY.

THE ASSET PROGRAM ( T /W), IW/SlETC.

TAKEOFF 6 NOISE REQUIREMENTS I PROGRAM OUTPUT I PAYLOAD REQUIREMENTS I MASS PERFORMANCE COST NOISE SIZE L.

• ROT&E • SIDELINE • FLIGHT HISTORY . • BODY • GROSS r BLOCK FUEL •" 'INVESTMENT • FLYOVER • • EMPTY WING PRODUCTION • FOOTPRINTS • STRUCTURAL • BLOCK TIME • TAIL a TAKEOFF • RESERVES TOOLING • MATERIALS • ENGINES • CLIMB d TRANSONIC SPARES & SSE • LANDING DISTRIBUTION • GEOMETRY 1 DATA, ETC. • SONIC BOOM PERFORMANCE • PROPULSION • FUEL CAPACITY • FAA SAL. TAKEOFF • OPERATIONAL • SUBSYSTEMS AND LANDING DOC; IOC: Rol • TOTAL SYSTEM COST Figure 6. ASSET Synthesis Cycle Airframe Noise Estimates 2.2.3 Estimates of airframe noise levels at the 1969 FAR 36 measuring points, along with the aircraft conditions, were made for the domestic and intercon- These estimates are included tinental aircraft with the-STF505M-7C engine.

as Table 4.

io Engine Bleed Requirements and Power Extract n 2.2.4 For this study effort, engine bleed and power extraction requirements were included in the engine performance decks supplied by Pratt and Whitney.

-Estimates of the bleed and power extraction requirements for a 500 passenger in the early 1990's are: aircraft for introduction into service ECS and anti-icing Bleed Air - 9 lb/sec for • Power Extraction - 370 hp for hydraulic pumps and generators, • 2 -4 LR 28664 +10 +80 +60 +40 +20 r C7 W —20 —40 —60 —8a LR 28664 +100 GW +80 +60 +40 -+20 F- C7 W d —2a —4C —6( -s( —10( LR 28664 TOGW +20 DES. FUEL +10 I C7 W d -1C -2( LR 28664.• +100 TOGW +80 )EW +60 +40 +20 o.

r- H W d -20 -40 -60 -80 -100 LR 28664 Tn f!IAI +500 +400 +300 +200 +100 F S C7 0 W '

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-100 -200 -300 -400 -500 LR 28664 TABLE 4. AIRFRAME NOISE ESTIMATES (STF505M-7C ENGINE) Condition Domestic Intercontinental Approach (42° Flap, Geardown, 3 0 Glide) 372,000 419,600 Landing Weight (lb) Approach Speed (knots) 136 131.5 Altitude (ft) 394 Airframe ' Noise (EPNdB) 96.0 95.8 Takeoff (25 * Flap, Gear Up) "limb Angle 6.10 4.530 TOGW (lb) 454,013 - 630,491 Altitude (ft) 1,710 1,101 Distance (n.mi.) 3.5 3.5 Speed (knots) 152.5 159.6 Airframe Noise (EPNdB) 84.3 89.6 Sideline Point Airframe Noise (EPNdB) 81.0 82..9' hese estimates are based on use of current state-of-the-art accessori for the 1990's time frame. Significant savings in mission fuel arepossible by using advanced secondary power systems such as large capacity generators and an all electric aircraft to minimize or eliminate engine bleed requirements.

Lockheed believes such a system is feasible for aircraft introduced into se vice in the 1990's time frame.'

2.3 ENGINE INSTALLATION N

2.3.1 Nacelle Configuration

The nacelle attributes (dimensions and weight) for the STF505M-7C engine were supplied by Pratt and Whitney. The STF505M-7C engine uses a mixed flow i exhaust which requires a full length cowl. As previously detailed 2-11 P: pql LR 28664 (Report LR 20351), use of the full length nacelle requires consideration of the following installation items: • Potential of interference drag - particularly for wing mounted engine • Increase in wetted area drag of nacelle • Potential of increased nacelle weight due to full length cowl a Access to engine hot section and to engine and airframe accessories.

2.3.2 Nacelle-Wing Interference Figure 5 depicts installation of the STF505M-7C engine to the wing of the domestic aircraft. Placement of the engine with respect to the wing is con- sistent with previous Lockheed experience for elimination or minimization of interference drag. Aerodynamic assessments of this installation indicate no drag penalty imposed by wing/nacelle interference. Development testing (wind tunnel tests) and/or tailoring will be required prior to actual installation on the E3 of the STF505M-7C mixed flow engine aircraft. For the aircraft_per- formance analysis, zero interference drag was used, which is compatible with experience on the _L-101i commercial aircraft.

2.3.3 Accessory Location Figures 3 and 4 depict location of aircraft accessories for both the wing and center mounted engines. Aircraft accessories are located in the pylon to provide an improved aerodynamic contour nacelle. All aircraft accessories state of the art with no consideration given for decreasing the are current advanced technologies which may be available for the 1990's size by use of is an assessment of the pylon time frame. Included in the design layouts sized for strength and stiffness requirements. Shape and size of structure, consistent with the incorporation of aircraft accessories, pylon the pylon is - structure, bleed lines, bleed air heat exchanger, fuel lines, hydraulic lines, This pylon layout was used for harnesses, and throttle controls.

electrical of drag interference effect.

assessment on the preliminary design layouts, it appears that 'pylon mounting Based accessories, along with the required aircraft plumbing and of the aircraft x electrical harnesses is a feasible configuration.

2-12 a LR 28664 During this study effort, various aircraft accessory locations were considered, as shown in Table 5, which indicates an assessment of the advan- tages and disadvantages of each location. Locating the aircraft accessories in the engine pylon with the engine accessories core mounted seems to be desirable particularly for minimization of nacelle drag.

Attempts to pylon mount all accessories, for best nacelle aerodynamic shape, requires an increase in pylon size and probable adverse effect on interference drag.

Assessment of maintainability and reliability were also made for pylon _ mounted aircraft accessories. Reliability of components will be enhanced due to the improved environment (as compared to the engine core).

Maintainability> aspects should be similar to those with accessories mounted external to the fan case except that an additional work stand (similar to that required for the center engine on the L-1011) will be required for pylon mounted accessories.

Aircraft accessories, plumbing, and shafting will incorporate the required disconnects to allow all aircraft accessories to remain in place during engine removal.

2.3.4 Access Provisions Access to the engine core and the core mounted engine accessories will be provided by using large cowl doors (similar to those of the JT9D-7A). For the pylon mounted aircraft accessories, maintainability requirements dictate removal of the top of the pylon to provide ready access to components. Since the pylon of panels for access can skin is only subjected to aerodynamic loads, removal be accomplished with nonstructural, quick turn type fasteners: Thrust Reverser 2.3.5 Reverse thrust is provided by a set of cascades, located in the engine translating cowl during the reverse thrust fan stream, which are uncovered by ,a The required levels of reverse thrust are approximately i operating mode.

35 percent of the forward thrust requirement, which is consistent with the sizing criteria incorporated into the STF505M-7C engine by Pratt and Whitney.

Flow directivity is required to minimize impingement on the aircraft control A schematic of the surfaces and to minimize reingestion into the engine.

expected flow directivity requirements is shown in Figure 13.

2-13- A Aircraft Accessories Engine Accessories _ Advantage Disadvantage Pylon Mount Pylon Mount • Best aero shape nacelle • Large pylon • Improved component environment • High speed shaft from engine to pylon • Access to engine not req. for component maint. • Possible effect on interference :^ drag si • Utilize integral gearbox • Requires additional work stands Pylon Mount Cowl Mount • Good aero shape nacelle Large pylon • j' Improved component environment e • Requires added gearbox, high speed shaft, etc.

• Engine access not req. for aircraft accessories • Requires additional work stands • Aircraft and engine components in separate locations p Pylon Mount Core Mount • Good aero shape nacelle • Large pylon

N

• Improved component environ- • Requires added gearbox, high N ment - aircraft accessories" speed shaft, etc.

• Engine access not req. for Requires access to engine hot • aircraft accessories section for maint, of engine components • Hot environment for engine components • Aircraft and engine components in separate locations Cowl Mount Cowl Mount a Improved, component environment • Large nacelle Utilize integral gearbox Revision to nacelle structure • • Y, and thrust reverser • Enhances accessibility to r components F _ • Small pylon Core Mount Core Mount • Utilize integral gearbox s Large nacelle 7d • Small pylon • Hot environment for all

N

components OD • Rigid mount for all components 0% • Requires access to engine hot 0% section for component maint.

• Revision to nacelle structure and thrust reverser r -

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Figure 13. Thrust Reverser Flow Directivity a` Ln LR 28664 2.3.6 Center Engine Installation Primary concern for installation of the mixed flow nacelle in the center engine location is the nacelle overall length and the potential effect on interference and possible scrape of the nacelle during takeoff rotation. For the domestic aircraft design, the STF505M-7C center engine was located such that ground clearance at the nacelle aft end during takeoff rotation was consistent with the current L-1011 installation. Also, the "S" duct inlet configuration of the L-1011 was retained to maintain existing L-1011 flow characteristics to the center engine. As is the case with the wing engine installation, future aerodynamic development testing (wind tunnel tests) and possible tailoring will be required to minimize interference effects. For this st d effort zero interference dra (eonsistent with L-1011 ex erie e) u Y nc g P utilized for the center engine installation.

was PERFORMANCE AND ECONOMIC COMPARISONS 2.4 The previously stated objectives for the Energy Efficient Engine Program regards to fuel and operating cost savings are: with Reduction in specific fuel consumption of 12 percent minimum.

• Reduction in direct operating costs of 5 percent minimum.

• Figures 14 and 15 show the savings in block fuel and DOC, of the domestic aircraft with the STF505M-7C engine when compared to the and intercontinental reference aircraft (JT9D -7A engine). The results show significant savings for the STF505M-7C engine as,follows: Domestic Intercontinental Typ. Range Des. Range Typ. Range Des. Range Block Fuel -18.1% -18% -20.1% -19.1% DOC - 5.9% - 5.4% - 9.2% -`8.1% DOMESTIC INTERCONTINENTAL DESIGN TYPICAL DESIGN TYPICAL 120 ,.00 -20% 80 200' ` ' p V ^ O IJ J W UA !

LL LL j -18 % -19.1 v L) n 0 50 r7. `° 20 0 0' CD LL LL LL y LL y

r

U Figure 14. Block Fuel Advantage with STF505Pi-7C Engine Lei 28664 Figure 16 depicts the advantages in aircraft size when the STF505M-7C engine is used. Incorporation of the energy efficient engine provides an air- WING SPAN INTERCONTINENTAL —5% —2% 5.6°I° C) r.

i n ch U V (.)

I Q I ^ I

Ln o 0 700 o Q^, Qf LL U- CD U.

_ CD ~ to cn N G 0 Figure 16. Size Advantage with STE505M-7C Engine I

LR 28664 SECTION 3 CONCLUSIONS AND RECOMMENDATIONS The results of this study, accomplished with the STF505M-7C engine design and performance characteristics provided by Pratt and Whitney, show that: • The NASA specified goals for minimum fuel and DOC savings are exceeded with the STF505M-7C direct drive, mixed exhaust high bypass turbofan engine.

• Installatiop of the mixed exhaust, high bypass turbofan on both the domestic and intercontinental aircraft appears to be feas i ble with no penalty for interference drag.

• Pylon mounting of the aircraft accessories is an acceptable configura- tion and enhances the aerodynamic characteristics of the STr505M-7C nacelle.

• Incorporation of the $TF505M-7C engine results in aircraft configura- tions, sized for long range and large payload capacity, which are compatible with existing airport facilities (field length, wing span, body length, and gross weight).

• Size of the STF505M-7C engine, as supplied by Pratt and Whitney, is compatible with the Lockheed specified mission and payload charac- teristics for the 1990's aircraft.

3-1 'k ....... .. ..

—

LK 28664 APPENDIX A DIRECT OPERATING COST (DOC) CALCUh,ATIONS — E 3 AIRCRAFT The following factors and formulas were used in calculating Direct Operat- All costs are in January 1976 dollars: ( DOC) for the E 3 aircraft.

ing Cost 3-Engine Domestic 4- Engine Intercont.

$397/blk-hr $476/blk-hr Crew Cost Fuel Cost $0.308/gal $0.387/oal Cost of Fuel $1.00/.lb $1.00/lb Cost of Oil 1.;0123 1.0123 Non Revenue Flying Factor

4% 4%

Salvage Value (SV) 16 YRS 16 YRS Life 0.304% 0.304% Insurance Rate (IR) $9.00/hr Labor Rate (LR) $9.00/hr 2.23` Maint. Burden Factor (MBF) 2.23 0.52 0.52 Airframe Labor/Cycle (AFLC) 0.52 0.52 Airframe Labor/Flt-Hr (AFLH) 0.68 0.68 Airframe Matl/Cycle (AFMC) 0.68 Airframe Matl/Flt-Hr (AFMH) 0.68 0.62 0.62 Engine Labor/Cycle (ELC) 0.62 0.62 Engine Labor/Flt-Hr (ELH) 1.31 Engine Matl/Cycle (ELC) 1.31 1.31 1.31 Engine Matl/Flt-Hr (EMH) A-1

FORMULAS DOC CALCULATIONS

Fuel Cost (FC)

(Cost Fuel x Blk Fuel/Blk Time) + (No. Engines x 0.135 Cost of Oil)

x (Non Revenue Flying Factor)

Unit Air Vehicle Cost (UAVC)', = Airframe + Engine + Avionics + RDT&E/No. of Aircraft

Depreciation Cost (DC) _ (UAVC + Spares - SV)/Life

Insurance Cost (IC) = (UAVC x 1R)

Airframe Weight (AFW) = (MEW -

Engine and Thrust Reverser/103)

Airframe Cost (AFC) = (UAVC

- Engine and Thrust Reverser/106)

Thrust (T) Total

Max. SLS, Uni.nstalled - Std Day (Sum of A:

Engine Price (EP) _

Total Constant Price Including Thrust Reverser

Y

(Sum of all Engines)/105

N :

No. of Engines (NENG)

Flight Time (FT)

AF Labor/Cycle =

[(0.05 x AFW) +

6 - 630 /(120 + AFW)] x LR _x AFL(

AF Labor/Flt Hr _

x(0.05 x AFW) + 6 - 630/(120 + AFW)]

x 0.59 x F7

=

AF Matl/Cycle 6.24 x AFC x AFMC

AF-Matl/Flt-Hr = 3.08 x

AFC x FT x AFMH

Eng. Labor/Cycle =

(0.3 x NENG + 0.03 x T) x LR x ELC

Eng. Labor/Flt-Hr = (0.6 x

NENG + 0.027 x T)-x LR x FT x ELH

N J

w

i t FORMULAS — DOC CALCULATIONS (Continued)

N

Eng. Matl/Cycle 2 x NENG x EP x EMC Eng. Matl/Flt-Hr = 2.5 x NENG x EP x FT ;k E:,fI Maintenance Burden _ (Total AF Labor + Total Eng LAbor) x MBF r

Total Maintenance = Sum of all Airframe and Engine Kaintenance

`i e r F MAY 31 1978 u A R A M F 1 k A N A L r S 1 S- A :. S F 1 I C

SUM'-4AQY IU NG 1

WING QUARTER CHORD SWEEP - 30.00 DE CO CN G IN- I.D. -- 416000 AIRCRAFT - t ULL ---C L X13 ; ^.-17.

0.300 WING TAPER RATIO - --I-hS SLS S L- AL! I.L = 46850 I.C.C. OAT, 4UMoch L'F ENGINES = 3.

-SUBSUNIL UC-S1':Y SPELU b

0.0 0.0

1..%: i1.3 0.0 0.0 0.0 0.0 0.0

6.6 ".V 1.1.0

1 W /S I1-._1 8

r^ 0.0 C.0 0.0 0.6 0.0 0.0 1..G 3.0 J.0 ,.) L.3 0.0 0.0 2 1/W u.2ti•. u...0 r 0.0 0.0 0.0 0.0 0.0 0.0 O.0 L..0 O.(• 0.0 C.0 0.0 0.0 0.0 3 AR 10.1J y u.L 0.0 0.0 0.0

O.G L. C.0 0.0 0.0 0.0 0.0 p

4 T/C 13-, U.0 u.. L.L C.L :.L: 0.0 0.0 0.0 4.0 0.0 0.0 0.0 u. • C,.O L.G G.G 0.0 5 SWEEP 30..,., 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 6 FPQ L.0 u.0 ...0 u..• V.0 0.0 0._ 0.0 0.0 0.0 L.( G.0 G.G 0.0 C.0 0.0 0.0 0.0 U.,^.0 C.0 C.0 7 OP'+ C. 0. O. O.

^^± G. J. (. 0. U. 0. 0.

u. •J. ^. J.

8 T11 a VA 0.0 0.0 p^' y 0.0 0.0 0.0 0.0 0.0 0.0 9 14P u.0 U.0 L.I. (.0 c.L U.0 0_0 0. 0.

P. a. 0. 0. 0. 0. 0.

T U. U. L. U. 0. 0.

10'AUG 0 0 O 0 0 0 0 0 0 kA,)IUS '!1 300,. .• G 0 3 0 11 N.

0 O O 0 0 O C G 0 0 1Z GF:'S`i WLIGKT 4,5.01 1 (, (. ( O 0 0 U 0 0 0 O 0 93 ,1 J U k1 U 13 FJEL WL16HI 0 0 0 t 0 0 0 0 0 0 25'b 4 7I u U u 0 14 _OP. WI. EMPTY 0 0 0 L 0 0 0 0 0 0 ZLRCt VUEL WT. 35597 L G C (' O 0 0 0 U 0 .. u. 0 0 G 0 0 38541. .; s'. 16 THR USI/E NGINF r.p 0.0 0.0 0.0 0.0 0.0 0.0

C,(. G.( 0.( Q.0 G. C• 6.0

y 17 ENGIYC `.,LALL 0.9.1, J.J ^..'

C. 0. 0. 0. O.

U. C. :,. 0. 0. 0.

18 WING ARLA 311,55. J. 3. r ► ^, 0.0 T^ 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 G . 0 L.0 u.0 U.0 U.. U.0 14 W1 4 G SPA T 1118.4 0.0 0.0 0.0 0.0 0.0 L.^ .l 0.0 0.0 0.0 0.0 1. TAIL AREA von.'/ u.0 u.v 0.0 O.L 0.0 0.0 0.0 0.0 0.0 i' 0.0 0.^ C. U.0 U.0 C.G .G 0.0 741L AREA 5n• ► ,} ,, L.:• 21 V.

9 6.0 0.10 0.0 0.0 0.0 0.0 0.0

G.C. t-U ..0 0.0 .0 t;.0 0.0 0.0 Ocr

22 ENG. LLN.T'^ lu.14 0.0 0.0 G.0 0.0 0.0 0.0 0.0 0.0 ¢.G U. 0 ...0 0.0 0.0 LNG. DIANFILR 7.u1 L •,J i.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.. 0. 0 O.0 G ._ U.0 o.0 0.0 f{ 40JY LL-NUTH 22n.5 0.0 0.0 0.0 (J. f; 0.0 0.0 0.0 0.0 .( U., .1.0 .0 u.0 L..0 0.0 25-WING FULL LIMIT 3.uJLI COST DATA 0.0 0.0 0.0 0.0 0.0 0.0 C . ,,.0 J. . 0.0 G.0 0.0 2C.. 2.4aZ 3. :+ J.0 3..) Q .

' - RIL. ROE 0.0 J.0 0.0 0.0 O.0 0.0 0.0 0 0 c.c (.0 (;.0 C..0 0.0 2 1 FLYAAAV - M IL. 50.15 u.v (•.0 - ^. 0.0 0.0 0.0 0.0 11.0 0.0 0.0 0.0 0.0 I%VFSIMNI-'i1L. 1.1- U.0 J. :.. :.0 ..0 .^^ 2a 0.0 0.0 0.0 0.0 0.0 u.0 G.,, 0.0 u.0 0.0 0.0 0.0 u.L 0.`c L.0 2v OOC - L/SM 1.11,5 0.0 0.0 0.0 0.0 0.0 .0 u.G U.0 0.0 0.0 0.0 IOC - C/SM 1.026 u.. v.-. u.L 3j 0.0 0.0 0.0 0.0 0.0 0.0 0.0 O.0 L.0 0.0'_ 0.0 0.0 G.0 ROI A.T. - u/3 35.05 J.0 b.

MISSICN PARAMETcR S O O 0 O O 0 0 G r 0 0 G 1, 1) 373 ti.: C• 32 41St/ V(( 0 O 0 0 0 0 O .. 0 0 0 0 0 33 41St/ Vtl( ► 11 b1062 u 0 0 0 0 0 C 0 0 0 u 0 G C _ L 34 MISN Vl(2 ► 1) 410u:: 0 0 0 0 6 -Z. 0 0 0 0 0 U L. C 35 M1S1 V[(2,11 35`.13 CONSTRAINT OUTPUT 0 O O J 11 0 0 0 0 0 0 0970 u 3o TAKEUFF JSTI1) 0.0 0.0 &.0 0.0 0.0 4.G 0.0 G.0 0.0 f-.0 .1.11.1 ,.0 G.0 .• 0..! 1.0 37 CLIM4 GRAU(Il 0 0 0 0- O 0 0 3 0 0 0 TAKEnFt USTIG ► 6>145 zi u 3e.

0.0 0.0 0.0 0.0 0.0 0.0 U.0 C.0 J.0 0.0 0.0 4 3 GkAU(2) u.U41'+ 0,L 3 4 'ELI' 0 O 0 0 0 0 0 0 L 0 G U 40 -C7t'L LNU> ')111 t'4 0.0 0.0 .J 0.0 0.0 0.0 0.0 0.0 0.0 T (• 0.0 O.0 0.1 O.0 41 AP SPE,U-K JI) 136.2 O 0 0 O 0 J 0 0 0 0 : O U C.

42 SEPI if - r -PS 1l u G 0 6 0 O 0 0 J 0 v 0 ., 0 ., } 43 SEP( 2) - FPS s a% CA 1^ P&WA E 3 DOMESTIC ISTF5MM-7C) U .. ., .... - .. _..... -. ........., .. ...,...-. .... .., ..

w.:. r ^.,. a ,..^:-:.. -..za .. ... .... .. ,x . .. ..nun. <.n c. G' ::r.^Y-._.....-: ..... ..... . '::s.s . .._.,.a MI 7 M .00 MISS E-z^ AIRLKAF7 / ` . JO PASS / 3,000 N N T/C AF W/S T/W J .13.00 1G`.:;J 114.b C. 2:5 ♦ E G h 1 S I A T E M E N T IPERCENTI WEIGHT FRACTION WFIGH7IPJUNDSI -54013.)

GROSS WEIGH7 1 c c FUEL 21.59 4pGh1.

I-J_L AVAILAML.

U.

1 = XTFkNAL " ...

9df36.

INTERNAL.

355973.

7E u 7 FUEL WEIG ,il 22.03 PAYLOAD lOSOut.

PAYLOAD 05000.

PASSE4GLRS 0.

6AG&ASE 1tiDJ0.

CARGO 0.

a1LiRFS 2:597:. - JPE947I944L L4PIY WLlb"7l OPERATIONAL, ITEMS 4.76 OPERATIONAL IlrM. It•1P6.

5407.

STANDARD 17LMS 234374.

EMPTY WEIGHT 29.63 STRUCTURE 135447.

STRUCTURE 41•1b6..

W14G kOTDR 0.

5755.

TAIL $UQY. 56544.

IPH2O.

ALI . SHTI ` JG GEAR.

Ln a1 (>22.

FNGINE %KT10N AND I.ACLLI.0 6.45 PROPULSION 2Q2 03 .

PRL'PUL S I ON CRUISE ENGINE-` 22654.

0.

LIFT EV . INkS 4400.

THRUST REVLKSER EXHAUST SYSTEM 197.

ENGINE CONTVOL S32.

- SIARTING SYSTtM PROPELLERS 0• U.

L114RI AlING SYSTEM 1467.

i FUEL SYSTEM IPC1wEk IhAN') 0..

OkIVE ';YSTEM 69a7U1.

SYSTFMS 4416.

FLIGHT i VNTROLS . Y POWER PLAN[ Ill0.

AUXIL1A v2s• INSTkUMLNTS 2674.

HYDRAULIC AND PNLUMA1)C 5453.

ELECTRICAL 15.35 SYSTEMS 2200.

AVIONIL_, i 0.

ARMAMENT Nl 44293.

t FU:41SHIhGS AND EUUIPME C' 7602.

AIR CUNUITIONINb 41,6. N ANT 1-1C141, U.

PH9TU:RA PHI G ON LOAD AND t+A4DLIN6 ^.

TOTAL 1 100.1 4-- P&WA E 3 DOMESTIC (STF505M-7C) LR 28664 AI Mi. k4 16 1.stj T s SJU PASS / 3000 N MI / M .80 MISS T/C AM W/S T/W 13.0U 10.00 114.• B U. 155 c E F A N 1 G U k 1 I C v U M E T R V C U AI,Lal:.4 6-71 SPANIfII )APE-1 RATIO C/4 bASIC W1'4(.-- SWEEP L.E. SWEEP MAC(Fl) J14S4.8 114tl.tJ7 V.3UJ 30.000 32.260 21.81 WINS PAVLLS-- AnLA1S1; FTI EXP. AkEA AVG 7/C L.k. SWFFk SFLFISO Fl) REF L(FT) 17Hh.7 152,'.2 13.00 32.260 0.0 29.31 1814.h 1''la.b 13.00 32.260 0.0 16.35 AM TOTAL WINO-- An(,A(`U FI) LFF AVG T/C Lk(F71 CTIFT) MACIFTI LIFT) r147.k 4.11 13.UC 40.06 4.17 25.36 71.94 I WsTISL; F1) BWWIFTI LQUIV OIF11 SPIISQ F7) F US=L A k l"- L.N',THlFI 2.'1. . i3 1.'ou4.; lY.S9 19.5q 300.95 3HI1-11 FT) rWIF11 SHw1:,0 t.,SIt 11S14Y.0(!

t y . 4 It TAIL 1-- SHII(SQ Fl StIX1ISJ F71 VLF LI(FT) L HT1(Fl) H71 VOL COEF HUkZ.

',Hn.29 Yul' 74 13`.24 M7. 19 0.918h FT) SHX2(SU F1) kEf H12 VOL HURT. TAIL L-- _Hl.]C2U L21FT) L HT21FT) CCIEF U.v 0.0 C,X 22U. 3e 0.0 PERT. TAIL 1-- SV1117.0 FT) SV Y 11S0 Fl kEF L11FT) L V11IF1) V11 VOL COEF 4h5.9./ 91.7h 5t,5.97 20.63 0.0660 Z-- Fl) SVX2ISO Fl) REF L21FT) L V72IFTI V12 VOL COEF VER1. 141L `)V121:0 U.0 O.0 0.0 226.33 0.0 A PRUVULSIUN- tvG L(Fll INr 01F11 PUU LIFlI P00 Q(FT) PUO S WE7- MI.. PODS INLET L(FTI 1U.2 4 7.03 71.1'+ b.23 10415.03 2. 29.23 TANK:,-- WINW CU FT) PDX(CU 1-1) FUSILU FT) FUEL 1320.93 49414.00 4561. 1U WkT1L() VOLUMES-- LLUUV WI.lr, (AILS P!AS PYLONS PONTOONS TOTAL 'fsi ih.4 ! t 14 7251.84 5331. 8 d1+`_,J. 114 0.0 0.0 61070.31 P&WA E 3 DOMESTIC (STF505M•7C) F - ^ r A-6 F LR 28664 4 A 6 F F r1 U C O c 1 C 1 t N 1 S-- c x

T E R N A

L S T

O R E

S

I kLT ITUDc = FtLT

0.

LIFT L"Er.= v.,J 0.30 0.(-0 0.50 0.60 MACH N0.

v.2..^ O.C.ld3b u..l"Jb u.U1d73 L.02C4Z C.CZ?2G L.02684 G.C31g3 0.41)')

0.01hbU U. 11031

J.vIhyh u.J)865

0.02143

J2508

L. 03016

0.750

u.c1530 C.11hL'1

U.Gl:ou

.01731 9 31003

C2 362

:;.UZJ48

^.4 0. 11hil :.,1''03 v. ,lc.l l •c) 76t r..')2C 0 4 (1.02329 (-.02t'10

.0. U.J161)

u.v161 e 0.016-01:

J.(J1tu1 0. 02 Ube

Co. )2 51 8

0.032bb

=.450

L.+1o43 x..+113, o-.017t2 ...(11`+111

U. 0,'3t)5

::.03234

C.04846

:.4:J u.-iis5-,

)..<<1•• u.OZ117

u.J242u

0.03365 0.05275

0. 08t74

ALT ITU7F =

IUG.n.. FFFT

LIFT CI't:F.= 7,.J

^.iu :: J.?0 C. 50 0.eC M AC H N13.

:.t Ji .LI1 19 1,4 L L .02112 J. 01384 G.02754 L.03?63 0.+. .v U.U1 72v v.. 1141 L•J1 !`t :..01142t L.uZ2U3 f,...2568 G.6306

U. 1 .I1 "J.)I5d) 0.u1-11+2

u...ltit I J.01 7N4

J.JCL55

0.02•+14

C.028QQ

C1. iib93 Ahl-,

0.01ou: 0.n1w1/

0.04056

L..U2380

6.02861

u. 174•

U.v10b1 ,,.J1O1,y

.Clb`1t t•.JZ133

0.02569

('.(%3337

1.-050

U..11G43 J.11 7t1

0.01%'/2 v.0146") U.JZ415

u.u32E5

U. U4b9b

U.4uJ

0...17.. .:• L,'63 0.J[lc<, (•(1247 ".0341" L'.,:`.324 0.08724 ALTITUUE _ 2000). F;^T (.. J ;..1< LIFT Cu 9-F.= U. 30 0.40 J.SO 0.60 M AC r+ NG.

1).1..

0.1)11-`42 1.11v^. .

J.va'...4 1.0214 t 0.02470

u.:,2b4J

0.03349

_'.^^• . J.C1 !^-. .^1 It..

,.01V.J. (.C2OUL

U.0227b U.U2641

0.031`+1

C.7`J -*

0...1 6 b16

v.vlt81 L.01R4/ G.G211d (;.;,2477 C.G21463

11.,11*116

:.u1.^17 J.:)1 +31 L.01840

O.Cel1b

G.G2443

G.C2424

1 r .01,x(."/

).` " 0

G or, I 1.5 G.('1 lot 1;.0191•. ('.C2191 u.?2b31

0.033S9

v.',SJ G.Jtv75

U.

1.02027 ;1.J24-17

.03346

3.G4960

G. 4 .''.. u. U1763 0. J.Joc, "1

(1.02530 »7;

0.03 u.C'5385

0.087b4

ALT I TU.;(- = "-iJUUU.

F t . T

LI LT CLt^-.=

O.0

U .30

0.40 0.50

0.hG

MAC`1 V(,.

1.2

0• .20"

L.L)c1!1

.01101 )2' "/Q (

• 1. 2443 (.C,345c

0.40L

u.O18oz U.J1.)''4

U.vl`01b 620HY u.J"3b6

C.vt730

0.03239

C-.71u

U..,I7[(t h..)uvI .,.vl lit.

C1 Q 2? 0.0+1143 L'.J1552

0.03C38

3.t4GL U-11141 L.,17,1 G.vIt.u5 k U1 4 ' 1 4 O.J2143 C.C.2518 0.0214'+0

0. U.01 d81 '1

145 U.J 1 ::3n ,

.01 4 81' 7. 072h4

J.:,7-105 0.03473 ').r•50 G.u212N 6.c.l'41 1 0.01Vt. 1 G.021GO C.Ot15C 0..•342) 0.05033 J. 4 vv 0.:),83`, v, JVO 1.02633 0.0354-1

O.J5457

C.Otb5b

ALT l TuD) = ti!.UO!.. F F t T

LIF1 Cl":F.= 3.0

C.30 0.4U

L.50 0.60

MACH NU.

).2i,(i 0.^111`, C.1)11ta

(i.c .0241U

).Ozbyp 0.63v63

0.03°71

v. "J3 u., l y C,, .J1'J5b ,., . v.. 0..0214;;

0. 0 6e

:.02P32 0.03341 0.7 .J u. J1dU7 v.L07b u.Uld•1^5 J.02004 v.022t'v 0 03 J.03125

C.^.J A l l It ...1s3h .om.;Ib

L.L26(.j4 C.03C85

).

J. F c) ;1'•e"/ -A • 1 1

4 h u.Jl`+11 u.02G7^

0.J235t L.

27140 u.03^o5v

0.8`0 .;2113 ,.:.t JJ[ ti.vI I'?Z ?.0214'+ O.G2b35

J.;,3SO4 ;, . 05118

,i.t . u.0.

0 3t 3 C .OthFo ^.;)3631

G.U5`41 L.Oh94u

P&WA E 3 DOMESTIC (STF505M-7C) A-7

IS

LR 28664

gZG^Ai' YuALl'1'Y

G^Y^RQ

0 k G C 0 :: F F 1 C I E N T S -- N 0 E X T E R N A L S T 0 R E _S_

ALT 1TUDE _ FF s

`CGOUu. T

LI S

T CUtI-.= 'J.0 :^.1:^ O.cr

0.30 0.40 f•.50 O.bO MACH Nr).

.:.2)U ...,z376 i,.t,:^3"7

U.J241 U2582 U. 02 b5 9

L. 03`24 0, C3732

0 L!2 Id. , U.t.,Zo,43

L.U..I`I(

0.02327 0.026C5

0.:,2Q69

0.03478

0.750

). X 1923 0.;.,1.94 t,.,IQti'+

L.J212r% 0.0439e

0.02755

0.03241

G.Et.4

4.0904 0.1 19:3

C. v1:.G7 i,.0215t ,

0.0239: 1x.02719

C.03200

J.•:2u iyQy

^.JCULi .i.jct :,6

.:00218h

O.'OZ469

0.02905

C.03674

0.."l-7 U.UC3^7 0.(2115

0.6,LIUt.

G2295 0.0274b

0.43615

0.05231

0.4-10 ...303v 0 ..,2' ) y 1 (,.02797 ; . ;73742 0065652 U.09t,51 = T 6000... F E ALT I TUUr LIFT CU(-f .^ v.l:, t,.^^.

0.30 0.40 C .`.0 0.60 •_ MACti 140.

0.L , t1.1 U..'E L.Ucbvb C• 0277b 0.03013 O.U3418 U.03LJ27 v.4.. .."<<.^$^.

1..^2.: ^t ::.C^^^' L.32442 (:.U17'7,) (r.%j3134 0.03643

0.71j

2:.11 L.JZQ-b ;:.02265

0.02`35 U.J28Q5

0.033bu

0.3 0,;131 0 13

; .OZ29- O.U2ti33

U. J2e5P

0.0333Q

1^18 E.C2I31 v.L2113

L.0232r

0.02606 C.(3)41

0.03E11

.uZ,:'). U. L . ,Y 4 1 •.02434

i.02t84 -1.03'Iti 3 .^ 0.05_67 J.

lhy a.U274. Li. ve6&8

Cs02'-s1 1 7 0.038 76 h

G.u5786 0.0916c.

_

FM Cit.S

L.bQ35 CL L , L- G. :566 ' P&WA E 3 DOMESTIC (STF505M•7C) F A-8 N I S `• 1 U S U M M A R Y M = .P.: 9^0103 / tu0 PASS / ?o0c v MI / MISS IJ AIRCRAFT

O AVG AVG MAX

SEGMT iD T At E7(TFRN ENGINE- EXTERN 11:7 1'.1Ttl Sc6M7 701J•L SE:M^'+T IVI7 ;'111 t•E'.Ml OVER THRUST F TANK L/O SFC 01!.7 L•1:I IIMt IIMt STORE ALT PIUU • . htLH atl:.I.T FUf.L FULL I FF/T) PRIES lAII3 ID TAB 10 TAB 1D RAT 10 IN (N "l) ( Ml NI (MINI (Ftl i' ILhI ILPI ILNI Ml 0.317 0.0 0. 416401. 0. C.0 u. 5.0 0.3 - .54013. u. 0.

Pow ER 1 3 0.; 0. 0.0 0.317 0.0 J. 1.3 0. 416401.

. .54u1-1.. t•e>. elt•. ^. 1.3 FO ► :Eic J ^.

0.507 0.0 0. 416201. 0. 19.87 16.1 17.5 CLIME 0.559 0.0 416201. 0. 19.15 14. IC4. 2.0 14.5 0.

446054. 6:4. 6613.

ACC! t 30v,,.•. G.6": 0.569 0.0 U. 416201. O. 19.24 6.7 2h.2 :t(. ,UJ. +GI B r4a:.6. 2G>3. lfbbb. `2. 15b.

CLI1B f.; 0.569 0.0 J. -416101. 0. 19.65 nc 2594. 2750. 339.1 365.3 CItUISF 37 1" 0.t . _ ♦433 x•'+. • 5. 77551.

0.0 364„1 0. 416301. 0. 18.55 -1.499 24. 277 4 . 3.8 O.bJ_ 17E:4t_. 123. 77619.

0: SCENT .1 17.99 -0.656 0.0 370.0 0. 416301. 0.

117':+. 6. 278?. 0.4 30'i.,u. 0.x,3 _37,,33-.. ::9.

DiiCEL 0.0 0. 19.31 -0.613 2@'/S. 16.6 386.6 0. 416301.

3 h. 0-1. 7P 344. P9.

UESCE%l 3,0 "-J. J.6-• +7 c.

0.570 0.0 0. -416101. 0. 19:53 17'. 3UUU. 16.4 4C3.2 ..bu r; .)Sbe4. _9tr0. 613?'+ .

CRUISE 4L;i0.

0. 19.51 0.560 0.0 406.2 0. -416101.

5^:3. F11-62. u. 3LUL. 3.0 15_J. L_335 3716t.4.

LOITER 0.0 0.0 0. 0. 0. 0.0 F. 0.0 406.2 't3.(: 37,,191. C. P1662. -3600.

R&SE1 0, 0.0 0.0 0.0 0. 0. 0.

0. a. 0.0 406.2 )721'•1. C N16C2.

RESEI U. U..0 0. 19.21 0.479 0.0 40 4 .0 0.. 416201.

634h2. 13. 13. 2.0 B J. 7 13 37151. I6^u.

CLIM 0.3 0.487 0.0 0. 19.24 0.0 40 4 .0 U. 416 201.

tt= '2. 0. 13.

ACCEL IU.IU. ► . G.4`h 370`.31. 0. 4 19.15 0.51S 0.0 417.9 0. 416201. 0.

6t. 6.9 U.4'6 37..531. s0. 14 . 67 ) ,11. ti.'.

CLIME IGLJU.

0. 18.23 0.581 0.0 3.5 421.4 0. -416101.

6FP. 6777 0 . t5. 9u.

AO :,a0. u.7,5 33x 9 ,.

CRUISt -0.582 0.0 U. 416301. O. 19.12 152. iC.'+ 432.4 st.t.c34. 3:5. 9P1'4. 62.

UESCE4T 31 U.".

-0.542 0.0 0. 416301. 0. 19.16 0. 1_2. U.G 432.4 U.4_•5 2a5b5'• . 1. PE154.

D`CE'L I::.,,;, 0. 19. L6 -0.674 0.0 437.7 0. 416301.

beat, - . 24. 177. 5.4 UESCEVI 10u0U. L.4` t, 0.0 0. 18.21 O.SBL 441.0 0. -416101.

64U26. 23. 2G:-. 3.3 3UJ_;. 0.7,5 -6.61 3 . 0'2.

CvUIS_ 0. 19.42 0.601 0.4 0. -41610I.

P9404. 0. 2U0. 2 .0 +43.0 3049.1 0. 3 7:!.

C+U1St 1 tiCO. U.3 7 ' 0.0 C 0. I8.I9 0.581 45.0 408.0 0. -416101.

96416. U. 200.

u. 72`• :6 OJ.. L6 2.

GUISE 30G4J •L^ F :4 8 : 1-UEL 41-336.2 c 4'.ti. ^. l l . 'f FU*L A= eL.7 MT O TIC P&W DO MES F3 T 1. U" M A P V P. S I' O N N U. 1 N A T E l t A L 1 t r = 1^..N -tl. / . .!•:.

. / 1 . 2 ( 41S5 f:c!.. Alkc- Ar 1 / '. y llh - AVC MAX ENGINE El17ERN AVG IOIAL EXTERN S`t,PT I:.TrL St:G ',t '..-41 1011L 1 -N'-71 I41T 1.111 1(111 JE SFC OVER THPUST F TANK L/D lIM'. IIME S1UkE DIS1 '_11J1 FU L L FULL ALI ITUJtr MG C'7 41L1^'11T RATIO IFF/11 PRES TAB 10 TAD 1D 441N) TAB 1D (41%) r Nj ILF) ILt- I IV `+11 IN 1411 I -T1 ILR1 IAKF'JF+ 0.0 0. 0.0 0.317 0. 416401.

0.6 F..0 o.

U. U.

IC45L.

PUNFR 1 0. 0.0 0.317 0.0 O. 416401.

1.3 t1.:. Hz S. L.. w 1.3 31".51..

PJ.+IX U. ...L 0.0 19.12 0.506 0. 416201. 0.

b•.. 11.6 13.0 > 110 6005. 64.

369b31.

CLIMB 1,. ' 0.3 tb 0.0 0. 17.63 0.559 0. 416201.

1.3 14.3 44/. 6452. 11-. 7..

l.h^_' ^6Y451.

ACC^L w;. t1, 0.0 0. 18.45 0.568 0. 416X1.

141. 8.7 23.0 0654. C7.

I'L.ts-. 30Y...3. _ _ CLIN4 3..J S.

0.572 0.0 C) 0. 19.43 154.4 0. -416101.

1u 09. 115.:. 131.4 +f,It - I. .:Lnr1. 31521.

J CRUISE 0.0 -3.037 416301. 0. '18.04 159.4 U.

34. IIbY. 4.4 v1`'. 1^0. 3167t.

Fv ► 4_^u.. .y •C ^3

DES

17.08 -0.657 0.0 8

0. 416301. 0.

l .b 160.2 .U. 3170.. 6. 119u.

33+.71 9 , OF C!. 31).,:•. 0. b.,.- 0.0 -0.617 r 0. 18.56 0. 416301.

1b.L 170.2 b.1''. 32313.

i... U.h4: 33.1731.

L•FSfcNI 19.39 0.573 0.0 db -416Y01. 0.

10.4 192.6 U.

125. IYfG, 350-4.

C1tUlSt 19.46 0.562 0.0 -416101. 0. T--' 140J. 3.0 195.0 0.

6551_). L 3354[(•. Y.'.

1-)LU. 0. 32V f` LOIIbt C) 0.0 0.0 0.0 0. O.

U.0..145.6 0.

1 0. L.

L. 35513. -14.

kLS2T _J. U.0 A,YrYL.

0.0 0. 0.0 0.0 0. 0.

... 0.0 195.6 u. 3 1 ,51 U.

3ii•.9ft-. 3.: U. O.J K.ESri 18.41 0.479 0.0 0. 416201. 0.

19t.f, 1460• 464.

a. J.'!c- ±.v.

CLIME 0.0 0. 18.48 0.487 0. 416201.

0.0 14tl.0 3b4•54. C. 11.

. ' 33 A >1^.

1 WJU. L.45 ACC5L 0.515 0.0 0. 18.37 21-5.7 0. 416201.

5 7.7 400"3. 45.

0.4'16 f3f514• 31:..

1J...(..

Cll y ^ 0.581 0.0 0. 17.99 3 0. -416101.

9.:. S.C. 210.7 4O974. +..

;31 b:l .

4JU.ij. (:.'I (Quist -0.582 0.0 416309. 0. 18.33 0.

59. .14'6. 10.4 221.0 3-,h. 41 k tit , .

• 5"t.

3JUUL. U. ti` t . 3C ` UE.SLE ^iT -0.542 0.0 0. 18.37 0. 416301.

L. 14 t. G.0 121.6: 412bb. u.

;_.le y .

1J-1 ,G. t..4:, DECEL 0.0 0. 18.37 -0.674 416301.

5.1 .10.1 0.

23. 177.

4 . --.. r1514.

0. , .`.. 32'010 U^SC O VI 1.• ^1U.

0.581 0.0 -416101. 0. 17.95 0. tt 2U3. 4.0 2311.1 42227. 28.

U.IJU --4.941. 713.

-1JUuJ.

LtiUISt 0.616 0.0 0. 18.71 0. -416101.

2.0 .32.1 3:.1. 42 4 N7. u. :UL.

...3 /H 3!3 -c3.

CRl11SF N 0.0 00 18.01 0.580 -416101. 0.

65 .1, 277.1 0. rn 0. -UL.

( b0' 71 4 tiL`4Y0 i tU.1JJ.. L.t.r1 3^/ L.

CkUlcc, j! P&WA E 3 DOMESTIC (STFSOSM-7C) l ^Ul k"tf^ F'UDUCIIO11 PROCUREMENT PJ TUTAL PER t T07a1 • MATERIAL LAb0R PkUD A/C** PER PROD A/[„ l 6 DEVELOP41"J1 - AUNktLUKk1 ' 41, 11RUCTURk .15'A . 66 13344.8"[ 1750 . 4F TOTAL PRODUCTION 41765.64 tilYO 1 67.11 2730. b6 4597.96 ENf..1NEER1yG Div./v PU7Lk l'.L L.L' G.0 INIFGR LOGISTICS SUPPORT 7L'OLI%' X36.76 1LIL lbS.Av 423.77 6Gb.Q7 PLANNING 30.39 Tt!b T A^TlCLrS 7c.9-s tUUY Iz49.4r. 417V .36 10426.82 D A TA l.. b, ALISPIING G L AR 613.3- 31.tr t 644.98 TRAINING 10.33 SYSTEMS FYGPWG y I tNta 1c CT • ,J.. NACLLL, [42.61 1#74.19 1221.76 CRJISf EvGIVE U... EW,!^ECTION -L G.0 O.G INAINERS 360.68 LIFT E4G1 1 4L U.e.

NACELLE Iva. 1 b80'.61 1073.02 FAN ..0 AIR INDULTIIN 5-.3c 94.38 140.74 HANDBOOKS 39.95 AVIONICS .0 OTHFR SY:1L - M1 r.0 vk(IPULSIO N I16.Sr1 112.03 236 . 94 FACILITIES 0.0 FALILIIIt'. U..: L1.1,IN1 INSTALL 13.0 29 . 57 2'+.57 TUTAL Ali( VEt1ICLt lS..w...B 7Hh,111 REVEFSE' U.0 5.75 5.75 .SSE - CFE 20.88 LXHAUST SYSTEM -C.6 0..0 0.Q 1NTE'.k , ,IS T ICS 3•JPPUKT LO 1%61% CCNTR('LS 3.3u 5.43 8.82 SSE - GFE 751.07 PLANNING SYSTEM 53-42 16.05 63.36 TOTAL 1LS 1213.31 TRA1NIVG 3._+7 PRUVELLER INSTALL U.0 0.0 0.0 HA4DE93KS ln.42 LU66ICATIN6 SY!IEM O.0 0.0 0.0 SSE 4.U6 FUEL SYSTEM Ws!u 71.24 131.44 INITIAL SPARES COST 5646.66 TOTAL 1LS 36.56 SYS1PMk It?N) LUkIV! 4:.1. d.0 0.0 PRODUCTION DEVELOPMENT TUTAL UVLPMNT-VUNR v C 15r6..>4 SYSIEMS 1733.71 7563.46 10297.17 ENGINEERING 314.89 9 FL10HT CCNTRULS 4'19.15 361.13 e40.2tt AUX PONEk PLANT 175.23 22.78 1..8.01 TOOLING -131.60 CD F4 F' DEVELOPMEN7 - RtCUtitPkUIJTYvL11 66.61 1N1Tk.1MENTS 64.95 131.56 HYDRAULIC • PNEIIM 133.21 361.65 434.26 ENGINES 0.0 v^ AIR VEHICLE 69o.c6 ELECTRICAL ti:?y.14 1165.16 1574.29 TOTAL PROD DEW 183.28 SPARES IIle AVION:C INSTALL 29 . x4 329.63 359.08 ARMAMENT 0.0 0.0 0.0 FOTAL UVLPMNI - ktCtfk b7,...0 FUk% AND EQJ1P 1_67.15 4694.40 5761.55 AIR CUNUITIUNIVG 402.50 ?36.03 436.53 TOTAL PROCUREMENT 48809.24 GOVMNT DVLFMNI LOS'. u.v ANl1-ICING 40.60 21.27 28.33 ^ b a P14UTO.RAPHIC G.0 0.0 0.0 LUAU AND HAwDLING 0.0 0.0 0.0 r IOTAL OVLP'*11 '.ISI iv-%I.71 :YSIIMS INTFGk 505.95 SRISr.b9 1095.84 } 1 7AL COST 7514.22 2162 6. 14 29134.37 TVIAL HkS •v 773.'.9 773.99 I 14 LHANGF ORDtk1 958.3S 1USIAINING ENG COST 1782.17 PkOb IOOLINf, CO`_1 1569.48 WALITY ASSURANCE 2412.44 • - MILLIONS Of DOLLARS MIS(..LLANEDUS • A • 1006.55 TOTAL AIRFRAME (.UlT 37363.32 •• -1000 Of DOLLARS OR HOURS PER PROD A/C - RNGINE CCST 3910.11 N AVII1N1CS LOSI 267.19 s•s - INCLUDES PkOD DATA, ON TOTAL NANUFACTU F 1NG COST 41'540.62 SYSTEMS ENGR AND OTHER SYSTEMS .F , CPERATIt_NAL COSTS - JiR =LT L.VI CVERPT ' JPER^ 11U p 4L 1UUCI INUIRFCI ICNAL LOST 1IC'C) MISC. OAIA 7s C/SM* PERCENT C/SM**• PERCENT FLIGHT W 4hI -Y$IkM 2999.95 I CN'w U. to9-1c 1b..j 0.02094 2.04297 RANGE (N. 111.)

1!IEL A'1D OIL : i.248b [ 1 L. CAL c.0'+41.1 ".75i42 BLOCK SPEED IMP") 419.39 INS:-RANCE 0.uzut;- 1.7-)171 AIkCRAf-i CVN1F 4 jL U.3015- L'. 1546C FLOCK TIME ( MRS) 7.15 ' E O^PkEC1AT:U3 ..4.Y1C 3t.i+toSc LAHIN A1IFHDAYT 0.10980 2U.r1o18 FLIGHT TIME IHRS1 7.15 MAIVTSVAW E 0.157~3 . ^3o11 ^ AND dEVERA p E 13.11962 AVG ;STAGE LENGTH IN. MI.) 1144.00 cz F A1;: 4.13462 PASSENGER HANDLING L.12884 12.53341 AVG CARGO PER FLIGHT 17413.00

O

TOTAL DOC 7.1Sr /7 1'.:. JuJ C LARGO HANULIW, 0.05700 5.54071 UTILIZATION IHRS PER YR) 3636..00 Ul1ER PAS,LN!,ER LXPtNSE ..32984 32.09718 FLIGHTS PER A/C PER YEAR 508.31 (1) OTH' K I;ARGD t XPEHSE 0.00583 U. 56699 FARE 268.79 GENERAL + ADMINISTKATIUN 7.76982 ^-j ro L.37484

a

ss• - CENTS PER SEAT N. MILE 7 TCT&l IL'C 1.D2763 106.000 -- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - ^ N RATE OF RETURN i:Y INVtSI.MENl YEAR AVb No AVI-KA6f CUMULATIVE AVF1.A 6 E REVENUE INTEREST OPERATING CASH RO1 "J AIRCRAFT 41RLKAFI ADJE, IvVCSIMLNI DEPREC JAI 10N nuUK EXPENSE EXPENSF FLOW VbLUE OF DUKTN(- OUK1V6 OUR INt.

YEAR YEAR FLEET YEA(.

sM t sM sM tM !M sM fM PERCENT .9

334.90 34.21 1(•4.00 -76.15 28.76

1 6.3 1.+. , 36b.3% 21.31 262.44

I

Vtb.4-P 76.97 x•49.5: 662.34 83.47 270.40 14.57 29.16 s^

2 16.3 10.., 24.0 -64.41 63o.81 93.05 332.80 164.17 30.16 3 3 0.,1 11'....3L 145.34

4 ku.0 0.. 1140..34 ;13.61 Q16.tiy 834.91 82.10 332.80 189.65 31.79 r

5 20.0 4.0 11ru.30 266.2.' 158.4:/ 839:61 71.15 332.80 195.12 33.69 6 2J.J 0.0 11"L.3(. 354.b4 /Q9.bb 639.81 CC;. 21 332.50 200.59 35.92 49.26 7 1u.0 0.-j 114..34 41u.Ub 721.24 639.61 332.80 206.07 38.56 3b.31 41.77 8 :u.0 G.J 1140.3u 401.4;! 652. b.' 834.91 332.50 211.54 y !U.0 J. 11ry-3c 555.04 364.r., 83Q.81 27.37 332.80 217.01 45.72 10 2U.0 ilw 114u.3: 624-31 `•15.44 834.111 16.42 332.80 222.49 50.72 YEAR 35.8, PERCENT AVb 60l UVE4 THE '10 P-R1?U = f N OD `

r

P&WA E3 DOMESTIC (STFS05M-7C)

N

W ^!vEnAll'_M'jRL CUSIS .P :II,iCl ► ' I t :TIU'JAL Ci_1 IL-LL1 1%titRtC1 0PERRl1U 1 4AL COSI IIOC1 MISC. OPTA C/Spss.

PItc1:41 C/SM**+ PERCENT 1-1.14ril LkI H J.2. b" . I6.3'•63a YyI L 0264,1 I. 44u5.. RANGE IN. MI.)

1399.95 .4 FULL ANU 011. U.231 1.. L-Lt L 10.3 i w5 G.I2h67 y.2b102 BLOCK SPEED IMPH) 384.31 IVSUS ANCk .^21pc I. - I- I" All:(-RAFT CLINIkUL (L.u634, {..24,143 ELOLK TIME (MRS) 3.64 UfPRFC1A1 lUN J.4792O 37.9"311. CA'tlN AITENOANT 16.7746?

0.22tIV5 FLIGHT TIME IHRS) 3.64 M41411 NANLL J.;fl. 'I -,t-.66.2c.

► FLCU AND eFVt-RAVE 0.14713 10.77-#72 AVGSTAGE LENGTH IN. MI.)

1144.00 PASSLNG1k ttANDLING u.27iOL.

20.22153 AVG CARGO PER FLIGHT 17413.00 TV] AL UOC l .163 /e I t, U. -,vu c L"h•>U HA%L)LIN', J.li214 6.9491( UTIL17ATION IHRS PER YRI 3636.00 PASSFN(.Ek 1-)Pt U111t h 111 :.E :.329b4 24.1663. FL:GMTS PER A/C PER YEAR 998.16 O1h1'R CARGO FXPE9S1, L.ot-SR3 U.426B9 FARE is) 125.54.

?1

GrNPkeL * AU^t1N151kA11CN .Uvt42

7.21x79

10141 IOC 1.364E7 1GG.60Q CENTS PER SEAT 11. MILE - - - - - - - - - - - - - - - - -- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -

W

RAU OF RLTUPN .'N INVESIMLNI _ Y= AU AV(, Vu A1kCrAFT AV 'kALF CLIMUTATIVE WFkAtE REVENUE INTEREST -- -OPERATING CASH ROl AtkLKArT AUUFu DEPRLL1Al]ON 111VtSIwLNT .IUUK EXPINSE EXPENSE FLOW UL/kI% G DURING UURIN(, VALUI 'JF V LAR VEAk Vt A1. FLE11 iM sM tM sM sm sM sM PERCENT 6-1 k4 1 lu.a ti!-t.. 21. fl 33r.-#6 240.49 3..21 114.76 -92.50 23.0.7 2 16.3 IU..J' VER..4 76.9.1 :,4N.5[ 6.5.27 63.47 298 .37 --27.95 24.15

3 :0.(s L.,1 li4t .3u 14').31+ •+,1r.y1

769.56 93.05 367.23 131.84 24.90

3'O.0 il. ll%,;.3L 713.111 •:t.44 769.56 b2.10 367.23 137.31 26.14 r 5 cv.0 J.0 II,u.3u 2P .2. "'.B.OI 769.50 71.15 367.23 142.78 27.59 6 2. ) 9>f.64 It 4.66 3.0 1141.3(. 769.56 60.21 367.23 148.26 29.29

7 'u.0 Ue.I 114 U.- C 49.26

rly.0p 721.[4 769.56 '567.23 153.73 31.3'1 a P iU.0 O.0 487.40• 769.5h 11.U.3P 3b.31 367.23 159.20 33.75 -# 44.0 Sbr.4. 36.76 4 v.0 Ii.u.3L 5,,5.F- 769.56 27.37 367.23 164.68 10 26.i, c,.0 L1.6.3U 6c..31 515.44 76V-5c j 16.42 367.23 170.15 40.56 AVG KOI OVER THE 1C YE4F PcRIJD = 29.2r PERCENT 7d IJ 00 P&WA E 3 DOMESTIC (5T'F505W7C1 } a a !

7 t I t n F f P L F O k M A N C t POL1k i%:-tP GRAD GROCFC GROCAE XFIELD 4F1S . 1T 1 ' Uw k -LAP CDSTl Y LL L_ FNLO CLRU`: CU41? 4 LL:. TOP ♦ T -. XG1 FT XROTF7 -X08 FT. X$O FT V1.'..FPJ VLfJFo _ • V2 -FPS x V-TALL - -,Vk]FRS- 1.0000 6975.9 _ U.11Y1 1.0000 4S-Ull. ti02. tS...i. 1.

U.0444 0. 1 1. l.ti`:at 1.20LG 3.74e2 6066.0 814.3 zZa. a ^Za..1 244.113 1')...69 ^j P&WA E 3 DOMESTIC (STF505M-7C)

8y

^^

r

OD N Ln 9N a t_J C1` IJ r , r 1 • • f Y IL t •( L 1 4 ,1 P - - 1 P: l H r0wt1, 1 ,1 111:. .. _ • IN' . IYk) VPrk. _ j.

1ttl _r1^1..vCf aLIIIJIt V1 L v t.AMaA L•tLIA •t5 All :,P1+SNU MACH PPS ALT CL t..r+ 1 l .t u... U.0 U../ ..J00 U.J 1144.09 0.0 0.0 0.196 • 0.'?1 »»+ ./ .^ i:F-I C. 2.occ L.0 11.4.09 0.200 C.0 u. 196 !°.rr 1.7 .., r4, to C.0 1'.41b u.0 i1-4.OY 0.213 0.0 1.959 4 r .7'. bUO6.

3^.., ^1.." ti. r :2 11. 7.v 35.L 1143.97 0.720 33.2 I.tl52 5 S r5.it• l:.l 1 ...

1. -' .t b, 1C.e63 ICU.0 1143.73 0.170 95.0 1.535 e 4N.,.» /'•^tl.! (^.. [111.ht• n,4 t3 .7v5 2UU.0 1143. 3t 0.210 I90.4 1.535 t5..lr 6.27Y lo.t'U 1000 11+0.52 0.223 952.1 1.536 v :!v•..,! „11.,,,, ll^.+r, ;tiP,av ;,,,,ati Iu.e.0 :Uf%.P i13^.94 0.227 1903.9 1.531 I%C.n^ !3••ar.v luC+., i,,L.li 11.7 3 b.7 x,OL.0 113!.35 0.231 28 ISIS .4 1.15 IJ Id".v6 -.3t it., I1.P•4 rL00.0 111 v. 75 0.131, 3806.6 1.5$8 11 iv,w S uu.. rnti,b0 '.,I4 !U.uSY 50Uu:1, 11:1, 14 0.240 4757.5 1.539 le t6'r.J3 6, c3.v aJO.'t 173.77 '.^/r. 60UL

I1.t65 U 11ei.1,1 0.244

5708.0 1.539 It (l ,. • rtt 7bill., 2 ► t.tl r.t 1 I11.t10 1000.0 1118.8E 0.248 6658.2 1.540 1 » !7 N.0 I,. LJ.,....

:41.,,, 4.otU lu.r74 b00u.0 II1S.24 0.253 7608.1 1.11.4 .13 lul :be.17 ..+7' !0.677 IUOO.0 1111.59 0.257 8557.7 1.540 It. 4.1.01 II»W'b.3 .'^U.4v 4.1,7 1,'.PPL, I+(•0O.0 1101.0) 0.261 9506.9 1.1,41 I k-n P&WA E 3 DOMESTIC (STF505M 7C)

x

ri co a+ rn 1` L ) 11 I :I ► • , 1 . of 42.0 U F G I L A P LL I,IIfII- LuU LIIl1N UIUtLP JLCSLI CDINO OCCGND CDOG CDI6 0.0•,.11 0. ( • 1"'eV ..( 6321 u.u1U72 -0. 0( 516 0.15400 U.148b4 u.U3llI t.Lr.itit u.UNOU t q U.0C226 - U. 0( ) 113 0.13115 0.13162 . 3t•uu 1 U. o5t.:1 J...1 7! t U. J: 1.1 .utitat- U.0 s9.7 U.UO310 -O.OLISS 0.1223P 0.12083 •nnm :I.w.:,51 ,. .r ...l tl o.(.11,A ..C,,2P" 4,.629t.4 4.01140 -0.00513 0.12111 U. 11536 -IV UV • f l.>1 v.f. U.ul /I G.JI U.0-.2bc U.U2I*lb u.0257(f -0.012119 0.17737 O. 114,.8 v.h" , 6',1 .,.01 11 u.U. • ► 7. 0.(' •i h y L.L151u U.0..70t -0.02351 0.14213 0.11862 ..c.;.. • • .1 .....Ill s U.0.44 U. 'Gin " u.Glu1. 0.07614 -0. 03b40 0.16604 0.12854 1 . ' 4(111(. J.nit,51 u..rt /1 u.J7:n- :,. nkr k J.uUS 11 0.11706 -U. 05853 0.20278 0.14425 0... >0> 1 to 1 11 4.11!1 — t•.0heu u.001 t3 0.16"76 --0.0P438 0.25000 0.16562 l.4nU1 1.o^t.5I 1.1.1 11 ' 0.I-Sit. ..•Jn; r: • 0.6 6.224FI -6.11141 0.30482 0.192,.2 e.POVUU 0.n • ,n^l vl /13 of ;bb u.21.1u2 u. u.0 J.Ib401 -U. 14451 0.36902 0.22452 N 1 t J ( L t --- Ll to l 1 F G 4( l' N 0 -- G L A P U P tr.J 31.4:•0 u.60r1(J Lr.v((.c0 I.Zf.CJ6 1.5(d-UO 1.86060 2.10000 2.40000 2.70000 4(.U.- 540u 0.13. it u. 1., JL 0.12111 0.127.7 u.141II U.It.644 0.20276 O.IS000 0.30482 0.36902 I P U 1 A l 1 • h It ( -- 1 1 u" U U N u-- ( F A It U P fJ IG11.Uf`l -1..,'0..1 J.0 J. it.:uu 0.60J4) 0.4f(•L( I.iGOnU I.SOf)00 1.110000 2.10000 2.40000 2.70000 4l .JL.. J., 0. 14df.. ift.

,.1 l J. /.vl 3 u. IIS IL ,.11.. U.I Ib62 O./lr%4 0.14425 0.16562 0.19242 0.22452 4 4 1 1 '+ u . l I t T C J • i t I L I c N 1--- LL I FIX I Nt • LL1 = 2.66091 PUVA E 3 DOMESTIC (STF606M 7C1

r

x

N

O, I LA •J1'1!.

A!.I L+4.14:• .1 i 4, F1 XgntL,11 %^I A-t.II 1'11 OlSl.f 1 ► 1 fLO L VAPYR,KlS Yu1"a.., I1+ ^ 6115.97 136.15 P&VVA E 3 DOMESTIC (ST F505M M

y

v

r

:J Q, O^ S;IMMA4Y P A R f M 1 1 F I C lJ 'At'. I • f 1 A 1. A l Y I S MAY 31 1978 A I CRA^ 1 M AW 1 k --LL 11 t 1 14 1.t,. -- 4Ic(OL MINE OUAR1fP CHORD SWEEP - 30.00 3fG I uAI --I 'it '. Sl_ 'I.AtI 1.:. = a(t`U kINC. IAPfP RATIO - 0.300 k utI .

N11M1,1 N !,1V1 0.0 l I/' I3..0 ..0 k U.L. J.0 W.. U. 0.0 U.0 0.0 U.0 0.0 0.0 0.0 0.0 _ l T/W .[_. .t .0 .,.0 1 .L L.0 U.0 0.0 0.0 O.c 0.0 0.0 0.0 0.0 3 AP Iu. 33 c.0 U.0 U.0 u 0.0 0.0 0.0 G.0 0.0 0.0 0.0 1/C r 11.7.7 U.0 u.J .0 U.•1 'j. t, U.L 'l.( L.0 U.0 0.0 0.0 0.0 0.0 0.0 0.0 S SWtrP IU.:.1 ( .0 .( 1 .r (.'' (.L G.0 O.L. L.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 1 I Pk (...: 0.1 0.0 0.0 0.0 0.0 0.0 0.0 U.., l .0 .0 C.v C.L 0.6 L.G / if, k (I.0 0.0 0.0 0.0 0.0 L.0 L.v U.G G.0 L c.t 0.0 0.0 d I11 .. 1.. 0.

L. L. G. G. 0. 0. U. 0. 0.

9 VVN G.J U.0 L.+. 7.0 U.0 O.L L.L L..L U.C` U. G.0 O.c 0.0 0.0 0.0 0.0 0. L, o. U. J. 0. 0. 0. (•. 0. 0. 0. 0.

N.

I l •(AOIUS -1 Woo . ( 3 ^ 1. ( 0 c O 0 O 0 0 O "LISS Mt I G tiT AIU4.1 U l l' L, u u 0 0 0 0 0 0 0 la 11 41nL WI I( 3 (1 :r3,11 t l u l 0 0 0 0 0 0 0 l' c 0 14 JP. WI. FMPIY ♦ b1,r 1, C l 0 u 0 0 0 0 0 G L 0 C 0 1 1, Ifk - I flit t dT. IbU4/. 0 0 0 0 v 0 0 L 0 0 0 0 In I-ik , 1`1/1r461Vf 3.6 II I ( U L 0 0 0 0 0 0 0 u G G II ENGI V/ S1 ALI 0. 3 J.v v.0 L.0 v. 0.0 O.0 C.O 0.0 0.0 0.0 0.0 0.0 .J ...0 U.0 114 W14G Akt A 4/ /r.. 1. L. U. 0. 0. 0. 0. 0. 0.

U. ,. L. U. 0. 0.

MINL 2l r.r• O,G 14 "PAN L..l 0.0 U. u.L U.L, U.G 6.0 0.0 0.0 0.0 0.0 0.0 9410./ 0.0 0.0 U.0 0.0 !L• ". TAIL AN FA U.0 'I.0 t.0 O.J Q. U.G O.0 U.0 C.0 0.0 0.0

G

.^ II V. 1 4 11 AREA bs..`• 1.1. 1 .1 L'.6 (I O.c. (,c U.0 13.0 G.0 0.0 0.0 0.0 0.0 0.0 0.0 O.0 21 F4G. LIN'.11f r,Iv v.0 U.0 .. 0.0 _.V U.c L,. C.0 C.0 0.0 0.0 0.0 0.0 0.0 l F 23 FW.. UTAMf It t.I 1 U.0 ..0 J.0 G.0 U.'1 0.0 O.0 U.0 O.G 0.0 0.0 0.0 O.J U.0 0.0 00 74 4(1tlY Lt W1_Tr 2tV.*1 U.. ,L L..( G.I1 U. : ..1. 0.0 C.0 0.0 0.0 0.0 0.0 L.0 0.0 0.0 IS W14 F 1117 1141T I.60J .•.^ L.L. ...0 U.0 L.L v.G L.0 O.G G.0 0.0 0.0 0.0 0.0 0.0 ii.v

xa

CO%T DATA

r

21. N01. - 2.-/hl .,. 3 u /:,L C.O L J. i•. L. G.0 C.C, G.G 0.0 0.0 0.0 0.0 0.0 0.0 NIL.

It YAWAY - I.0 C.O L.,' u.0 0.: 0.0 G.G 0.0 0.0 O.0 0.0 0.0 0.0 2 / MIL. 1'•.1 ` , v.0 IN V SIMN I - .IL. L'.. u.0 U.L U..• L'.0 V. (I C.0 U.0 0.0 0.0 0.0 0.0 0.0 28 I 1.262 v.I L , L'_,, r9 ^^ yI /)M 0.0 L.0 u.0 O.G 0.0 0.0 0.0 0.0 0.0 2-0 ACC - C 1.[d: LI.0 v.l U v.c t..l L...• ► ^^ l J .,.0 0.0 0.0 0.J 0.0 0.0 3u IOC - C/1M 1.2-_ .v 3 .. ,.l ..O U.0 J. G. 41 G.0 0.0 A.T. - v 1'.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 31 001 /G c4, 1` • v.0 (•.0 1 .c O.tl C,.0 U.0 0.0 c 1'l(_ MISSION PAk.`1F G 32 `+ISY VII I.11 34t, 1. U J O G G L O 0 0 0 0 0 it '' M1 )` VLI 1. I I IUPPP , C i. U J t U G U 0 0 L O 0 0 34 °1:,N v11a.11 41 •.. . u l U 0 ti L 0 0 0 0 0 0 0 0 35 MISn v21..1 ► l% 1 u U J U 0 0 0 0 0 0 0 0 e., C 6 4 4 1 4 7 a (1PUT CU411 1 v.nU 0 0 0 0 0 0 0 3o f AKf Lit- f ' 1 ',1111 U V 0.(1 0.0 0.0 0.0 0.0 3/ CL 14' • tkA 3 111 J.v..3. .L .0 O.0 J. L, U.0 O.0 0.0 la 1 0 0 IAKtUFf J11(:) b411 L 1. 0 U 0 U 0 0 0 0 3` • . 18 C.k AuI. 1 J... IS- U.I. L.L J.( O.0 U.L 0.0 0.0 0.0 0.0 0.0 0.0 CL 1 4(, CT A LNO(. 0111 %1:J 3 U C h v U C 0 U 0 0 0 0 L 0

r

AP ;PfIO k.0 .l (•.0 t.n C (.l O.0 4.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 41 x"111 131..

4, , IPA L U n 0 0 0 0 0 0 SEPI II - U u 0 r3 St PI 1 C U L U U p 0 0 0 0 0 0

21 - F 3 N

O^ t P&V11A E3INTERCONTINENTAL ^O (STF505M 7C) c* *j AlRl1.AFl / SJO PASS / 6500 N Ml / M c .80 MISS tJ 'I I/C AF w/S 1 ^ I/w 13.00 10.)0 13[.L U. 220 H 1 l A T M N 7 wF 1 U S E I v WFIGHII UUVU^I WEIGHT FRACTION IPkRCkk7I '.k • ISS wt It. 6314.41.1 HT 1 i^ F 3 UFL AVIV ANl 24 511. FUEL 38.62 FK7fRVAL U.

IYTFIINAL 241SI7.

l(kO FUEL wEIGHI V A YL')A 0 IULULI. PAYLOAD 15.86 VASSkYGLRS BtiC•Ufi.

C Ak!," 1 `•000.

STURL', C• .

'1v1 R Al PINAL k " P I Y wt l"nl 26t4ll.

UPERATIONAt IIIM!^ 162L' , . OPERATIONAL 17EMS 3.96 S TANI)ARU I I f M^, 87814.

'MPTY wf IGHT 26171.

S714LMURF 1')41'!7. STRUCTURE 24.45 w1NU 60F5 r.

k71'7k I L.

T A t 6374.

I 3LOY Si176.

r

HTIVt. G^Ak 244 14 .

ALI r .

I N:lV ` '1 ^4.

I CY 10 At-li 1,ALLLLL 73 3440?2. PROPULSION 5.54 1`40PLILSIUN LF • 11 SL I NUI'AE 26F32.

LIF1 F9'.1'1CS U.

T ,I%USl RI VLkStk 5414, FXt/RUST SYSTI4 0.

F1 46INC C1,114"tIL [36.

STARTING j.Y,1FM 7J4, PR'IPFI CI VS J.

LU9 k lt All`.'. SYSILM U.

FULL SY,IL14 1711.

I1kIVF SYSTEM (POWL 11.Ay'.)

R 0.

Sr:TfMS 7,( to FL1Gt1T CUNTRCIL$ ^bti0, AUAILIAkY v!IWtR VlA':1 Ille.

• I'1STKLIM V1, I11/.

t1Y9RAULIC AVIt VvtUMAIIL 3t 12.

_t ECTw ICAL r.b50.

AVIUNILS [ 0 00. SYSTEMS 11.56 r, Aky AMk NT 0.

S AND I001PIllNI 4%, FLA.VISHIVt , 14 3.

N :.C"10ITIu4ItV.

Alk It-C, 2.

OD ANTI-1CI%C ^0.

ON A Vr1I L 01 Pr1UT U1.k U.

t- LJAU ANU 44,'1(,LI41.

0.

TOTAL 100.)

P&WA E 3 INTERCONTINENTAL (STF505M 7C) k*-I AINC I+AF1

/ SLU PAS! / 6 5

CO N M1 / M c .8( MISS T/L cN W/!^ 1/N 15.0(1 IL.uo 13t.0 C.:20 L U N F I` (- K e I 1 C N G f I: M t 7 k V QASIL

1.1'.1.-- Ak1.i'^j FT) SPIN(1-1) IAPl4 P

4110 C/4 SWEEP L.E.

SWEEP MAC If T)

4 1 76.4

218.55 (•.30) 3L,. U00 32.260 73.97 WIN(. 3 A41LS-- AKIAItu F1) 11(P.

AKA AVG 1/C 1,.t. Limit EP S1-LEISO F1) REF LIFT) 3. , 61.2 2239.0 13.00 3.260 0.0 32.S3 2191.6 a191.c 13.01 32.260 0.0 17.97 IOIAL WING-- F I I EFF Akl41*,U Ak AVG 1/C Ck1F11 L11F71 MACIFTI L I F T I 5:52, 10.04 II.UO 44.2L 10.09 77.94 79.06

F us! lAt" - LtN("IH(1-1) S w111S0 1-11 D I F I )

9MWIFII EOUIV SPIISO F11

2.9. 4 119tiu.0

19.59 19.58 300.95 OW( tll hH(FII SowI^Q fll 14.4p

1 4. SO 1194(1.00

y

i

HIWi. H F I I

I A I L 1-- .', 111:U F11 S-IX 11 S0

kEF LIIF)1 L H1IIF11 "1,l VOL COEf N 9E^'.TO) 752.34 O 13.IIN 1(16.36 0.9195 C ?

I

HOkl. 1

1 AIL -- !hl,1:0 1-71 !.r1Y24SO F11 ktt L 2 1 F I

L H124F11 M12 VOL COFF b _^ L.G O.0 224.49 O.0 ('.0 VERT. 11111, 1-- 1V1I1`.: 14 SVx11S0 1-11 kL1- 11IFII L VIIIFII V11 VOL COEf

^a

0'...9` 652.p'- :2.15 105.63 U. 0661

r

VER1. IAIL 2-- V1:ISU 1-71 SVY2(SO Flt kEF 1-21FT1 L V121FT) VI2 VOL C(IfF u.0 0.(. U.0 724.44 L.0

• r.. r7

PlJL51t-N--

PR O ONO, 11F11 INC Off . 11 PUU LIF11

POU DIFII PUU S wt w1,. PODS IMLET LIF71

9.7 1 0 6.61

2C.2I 7.6L 1087.61 4.

0.0 FULL wINGILU FT) how(CLI Fl) F11 1-U'_ICU 23 t '9109.U^1 1041 .^ wtT1EU Ed,DY 4(-1,("4IS-- WI VG TAILS PODS PYLONS PONTOONS TOTAL 5187'+. 04 7291,.7: 1:60.54 3N77.8i 0.0 0.0 04:57.10

r

N

co ON P&VVA E INTERCONTINENTAL

N

ON ^O (STF505Ai 1C)

A

LR 28664 A G c f 0 Q L p r I ', ENT I S-- N O F x T E k N A L S T O R E S ALT 1 T I )OL T LIFT C IJEr.= J.0 i).1.

0.,u 0.3C 0.40 U.50 0.60 MAIN '49.

0.1 j) v.vl 74 c,.(,1714 u.0 l 118 1,.014414 C.U2225 L.u25QU 0.C30914 7- O.u1j--h ^.Ulhll 0.A7PU 0.02"5: C.02422 0.01931 0.750 (.'.C•1^5^ C.014i5 1^90 .,.01h'.h (1.01VZ7 0.02296 U. 02772 3. 0.,1407 G. I I>- L .1.3169;1 0.31929 (•.(1254 C. 02735 0.d. • J O.: lhl • u.u1:37 3..,1:7, (J.O172b 0.02JU7 0.02443 0.03211 0 .:18014 (..k 1 b ,.011.-7 (..01841 0.U2191 (..U31h0 0.04774 ^.' • .J U.Ut.5$1u C, .1.,2 U.J&C44 x.02'347 U.U32102 O.U5202 o. 08601 _ ALT 1 T(JUF I (,01 0. FFC^T LIFT COEk.= 3.3 1. 1v J..0 3.30 0.40 0.50 C.bG M AC M ")'7.

U._'.•0 i,. il2iuh

J U.UlUrll•

L.C2U1- U.02^1441 C.C261)6 0.03165 0.4JJ j.)1632 ,.,lr.vs U.JItb1, C.Cl637 u.U2115 O.C2479 U. 02488 0.75f: U. 115.13 t4 O.uI-. t,.J1534 ).01705 0.014It, J.J2335 0.C2821 6. '160 0 ri157S i•.',1^)36 C.01'sy( i .0173 1- 0.0197P J.023C3 U.02783 D. h2v '.; 16x7 J..•15u5 L;.J1:,-1 .ul771, 0.')2055 3.02491 C.O3259 0.u Is) U.v1911 11L5 U.uIh`+: U.ol8R1- 0.L7",38 J.v3'08 0.04 b,' -4',J C .,,l6Cl Li.

II , L: .J[Ul.1 .02394 L,.

334 0.05249 0.0864b ALT I T UUE = IUUUU. Ft LT LIFT COFF.= r.

U«30 J. 40 C.`.J 0.60 M AC t , vU.

0.e 1 'j.J16`4 :'.U1 .0209t, O.C2^74 c:.u2738 (J.03247 O..Uu .)17ul v.: 1613 11314 0."1000 U.U[1bS c.0?550 C.0305Q 3.750 )..15h3 0..•1534 u. .1 .017tt U.02U3b 0.02396 C.02881 0.aJO J.^l635 0..11596 O.J _.01794- 0.02037 0.02362 6.02843 O.k:O 11.01716 C,.:lo.5 0.Glc•,-1 r,.n1P34 n.0'11-* (•.0155) :.03319 U.o l i j _A971 3.u17o4 v..11`.

C.319ti7 C-2 3 4# :,.33266 3.C488(j 0.9 )t .

J.UL<+'34 U. j&: ,c 1ti0 i'.02452 U-.03^4I 6.35300 3.0167U6 ALTITUUl- = 3Uu0v. FtLT LIFT C'1t^.= U.0 1.1'1 C.'u v.30 0.4C 0. 5;' 0.60 MACH NO.

J.2 I J. )19W :'. 1 0 5 L .;JL /L i .021Q- 0.02 16 72 CoC2P 35 0.03345 0.4- O..i17l t6 L.ul'^57 ..u1L.l v.01,492 0.0,2t)9 O.J2634 0.03142 D. ISO ,.Jlb3,

u.Llw5 It, 1,i

3.u1P37

C.U21Cb U.U2467

U.02 9%2

1 1.J17',,h

).A33 l•.vlhc,7 v.u1 /[I

J.0187L

J.321Cb v.J2433

".02414

1).92 ti U.'`+1791 0. )1 /l5 ).)I / ti l v.019G4 0.021U5 O.

Ih20 O.Q33b9 3.c53 ').JCG45 r.^ IS3-i c,.Ci<<y n .C2(17 O.C(467 C.C-3336 0.0 4 95u 0.9,1.3 v.ot"7'>?

u.v23l v.u1211 r.:1l'+2C 0.03465 .635375 :..Cb774 ALT I TUA

= ,o., r . FFET

L1F7 C(,''F.- u.v C.1U J._.

0.30 U. tic, L- . SO 0.60 M AC H Nr).

0. 2L' (R.(Il C C .023Ct C,.Oie C,J295J 0.03459 0.4Ju 0. ilhb .. Ii, • ,, ,..A +1'+ .J2JL C.32 ^bc ;•.,1131 (.0324( 0.7'+0 L. .'171 7

C'. I u.JII;3

C•.01`42C 0.0219C O.J2`.53 J.03035 0.d_, u.(.114' L.A 1"o v.. lkv3 U.01Q52 (I.U21v0 u.02s1) U.U29Q6 1I .A2'J J.Jl'i7'' ,.31797 u1QP5 L.u216t U.u2701 U.U3471 0. FSJ :1. ,^ 12'> x.71914 ; ."

1'14..4 L .0209'/ O.U254 1 J.-j3-17 C.05030 J. v_ U J.; 2b 32 Z) ?s43 0'f b ' .0159 q 0.03 5....

C.(' 5-+54 O.C8 E53 P&WA E3INTERCONTINENTAL (ST F505M-7C) A-21 ?.R 28664 ORIGINAL PAGE' OF POOR QUAIXI-f ! Z A G N O '2 E F f I U l L T!) - - N 0 F x T F k N A L S 1 0 R E S AL T ITUC) 97 : 5G('^L( . cccT LIFT 1,LtEF.= 0.0 0.30 0.40 G.`11G 0.60 V'l.

MACH 0.2:.0 to. J2255 I. 7 0 i i i 6.02461 U. 0273 L, G.U3103 O.C3612 0.4;)U U.U7U1? U.vlyb5 v.JLUiv u.UZ214 U.U24107 44.62861 C. 03370 0.750 U.: 1 :2Y '). ) 1 ,79'1 v.U1263 '1.32931 0.Ji330 C.U2660 0.0314'+ U.tb('0 C.vljcob (.(,lr7^b V.(i1,+1, 0.O&lG61 (7.0[30(1 U.U2625 0.03106 0.9iJ •, J. ig87 U..)ly•>5 v.019-11 .0:094 .,.0237~ (.02N11 0.03579 0.9So 'J.G2t33 U.U:U 2 U.02205 0.0245 3°Z,* 0.05138 3.9,IZ? ui %# 3N .027U`, 0.U3650 6.01566 08'+59 (1.

ALT I T F'-ET UUE = '+C')Ol.

r T Ll CUFF-.= (..0 .t..

0..( 0.30 0.43 x.50 3.60 kiA Lr . i 14(j.

0.,e J, 0.0/44L U.U,41, U.Ul4'16 v.Cl'54tj 0.02424 0.u32b8 :,.03747 G.4'^i G.•).'1 !j U.-it 1`.x C. 'J., Jc (,.02-%76 0.026!+•0 0.U3(,18 6.03527 J.750 U. A46 + d U.,1`+-I L'.u1 tiYt, i.-.02162 0.02433 U. 027 o 2 (j.0?279 0.°':.r 1). 3292 c) 0.')1')9:• U.U, ).02193 ,^,.02431 U.'.,2756 0.03237 0.heJ 0. 321 it, C..•1(.3 .02:c5 0.02`.Gt+ C.C2941 0.03710 0.+^`,v 0.v.'3nL J 1 L1: . u...c'14 .02314 n.027t44 j.C3654 U. 05 26 7 O. Q Ju U.03UhF G.J.I^,fb U:83i 0.03777 O.C5687 O. G9 (A!

FM 'JLS = U.b,J:!.

CL CtS = 0.1566 P&WA E 3 INTERCONTINENTAL (STF505NF7C) A-22 U M M A R v P&VVA E 3 INTERCONTINENTAL M l S S l 0 N S (STF505fN 7C) _ M ,P0 F 1 / -,,)L, -,,)L, -,,)L,PASS / t5() ( 4 MI / MISS ^p E* 0 3 AIRC'.A

A

fRIfRN AVG AVG MAIL TUTAI T TOTAL St G M T TOTAL PIER, INGINF SE',MFN7 ! 1411 1r.11 INII SE(.Ml SE (. M SFC 71441 I('RDST F IAH9 LiG OVER ALT I I UUL MACH 4ytI"til EUII 1, U1 DIST Ji,1 TIME SIORI IFF/T) PRIES (MINI TAb TAP ID TAP ID RA11U tart IL n ► IN 041 IN MI) IMINI 10 1*II IL 41 (L! ► TAKEOFF 0.0 0.110 0.0 0.0 0. 416401. U.

6.0 hio4v1. U. C. u. G. 0.0 POWER 1 0.

0. #16401. 0. 0.0 0.318 0.0 POWER 911. 111. O. u. 1.3 1.3 2 U. U.1- o3CKY1.

0.510 0.0 22.4, 0. 416201. 0. 20.19 CLIM9 U. 0.37( • b2 Y5uu. 11Z%I. 121~.. 1/U. 12u. 21.4 0. 416201. 0. 20.37 0.560 0.0 13131. [L. I4... 2.7 25.5 ACCEI 3UJ:.U. ,.. 6 9 , n1:Z43. 1t: r9.

0. 20.41 O.S73 0.0 179.. 3b. 17h. 4.6 30.1 J. 416201.

CL IMP 90000. G.t60 01715 . 1 1, 124.

0. 20.34 O.S71 0.0 216012. 607~. 625'-. 199.0 t:2 3.2 0. -416101.

CRUISE 34060. J. wJ 6153o:. 1•.UbLJ.

41t3O1. 0. 18.bb -1.938 0.0 166IF.. 33. 6262. 4.2 817.4 0.

DISCE47 4Zuuu. U.dL- ..441#. 111.

416301. U. 18.15 -0.6.19 0.0 .5. 206219. u. 6: P 4 . 0.4 b2b.3 0.

3:4000. O.bUJ 42--317.

UECEL Iv.eo -0.619 0.0 6374,. 16.9 P45.2 0. 416301. 0.

3JJJU. 1 2 424171. 171. 2064 , 00. vu.

UESCENT 0.6 0. 20.07 0.57S 0.0 65UU. 14.,. 4,61.1 0. -416101.

;y CRUISE 42JO0. 0. Est: 47351,.. :2ci. 21C1 112.

-416101. 0. 19.20 0.635 0.0 4944 (I OF t-8. C. 69CL. 3.0 864.1 0.

15.'.L. 0.4„0 4 [Ji • l.

N LOITER W 0. 0. 0. 0.0 0.0 O.0 -650L. U. 3.0 b64.1 0. O.0 I. f100eP.

Kt SET 0.0 0.0 0.0 4,64.1 0. 0. 0.

C.c 4I"buj. u. 21Ut4Pb. U. U. 0.0 RESET u.

70.02 O.S76 0.0 4:0.5 0. -416101. 0.

110+u. 227116. ( r. P6.4 CRUISE 420LU. 0.1tL1 41,6J3.

IAKFOF^ 0.318 0.0 YSo. ,. 0. 416401. 6. 0.0 4('.514. I.. 21791t. C. L. U.O POWFR 1 3. O.J 416401. 0. 0.4.) 0.318 0.0 J. v. 1.3 151.E 0.

P']MtR 2 4u[57±. 1. 22P90-4.

U. u.c 0. 19.12 0.480 0.0 11. 11. 2.4 v!)4.2 0. 411201.

U. U.31r %4j 15 161,4. 2366:.3.

CLIM9 416201. 0. 19.19 0.46,4 0.0 23L601. (•. O.c. 954.2 0.

46.454. 3vvju1. 4•. I1.

ACCFL I(•411j0.

0. 0.516 0.0 5. 7.5 961.7 0. 416201. 19.08 rvvLJ1. ±b .1. 2j426.. , .

CLIMP 1JCJu. 6.450 0.0 35. 9 466.4, 0. -416101. 0. 1d.56 0.589 IC .b. 23531P. S.1 CRUISt 9UJJJ. G.7(1 •. 491231.

r

19.04 -0.584 U.0 4 7?.: G. 416301. 0. ,T1 i3576L. 61. 151. IC.7 irScLNT 300.:u. 6.692 :.511.. 4-2.

N

- 0. 416301. 0. 19.08 0.544 0.0 C. 215 164 . t 191. 6.0 477.5 L 1::; 4. 0.-^1 lv#/3l. Oo DEC t (T 0. 19.09 -0.676 O.0 175. j.3 vE:. t 0. 4161101.

U.• - •4 , 3 .4.1 )1. . 41. 1361`4..

DE SCI VT IJ( u4 . r

0. 18.53 0.5614 0.0 3.6 vet.4 0. -416101.

•u,..... C.l^, .944•.1.. 7 23414 /. 2vU.

CYUISf 0.0 30.0 1016.4 0. -416101 . O. 141.37 0.628 ...3 1 ;ti +6 ,03. 243131. C . 2".

L 1 ► UI SF 1560. I.3 M A R Y 0- 1 N N 1 S U M A l I N A 1[ 1 1. S 1 1 I .

INTERCONTINENTAL P&WA E 3 M - (STF505M -7C) AIR LK AF / •-: LF / 'GO( i 14 M I / •h. HISS E 00 3 MAX St IU1AL EX1fRN f NCI Nf fXIIRN AVG AVG 11 I('IAI . ."T 7.11 At SEGMI SE:4E141 I-111 IVI1 1'111 J[ (• t F TANK SFC OVfk (111.1 710-1 110-1 S7wE IHPOST L/O MACtt Mt 1:r+fl tu,t I- U'L v1'.1 At 111110, IAP 1At RATIO IFF/11 vkES TAP ID ID 11) "11 IN 411 (0-1 1 1"1 141 1 ► 11 NU IL41 ILQI IL"1 IN I I TAKEOFF 0.318 0.0 O.L 0. 416401. 0. 0.0 .554th. L. U. 0.0 POwfk 1 G. J.0 0. 0.0 0.318 0.0 1.3 O. ..16401.

V.I. -0"l. L. 1.. 1.3 POwfd i U. U.J 4544bb.

19.98 0.508 0.0 O. 416201. 0.

7311. 6!. 6'. 11 .t 13.1 18 r544-11. 637f..

CL14e 0. 0.3 0. IS.S3 0.560 0.0 14.5 0. 416201.

5>'. 7945. 10. 75. 1.3 ACCEL 3U.IUV. 0.6-02 44LI11.

19.28 0.510 0.0 h.7 U. 416201. 0. C C :4^1. 161 1-e. t 7. 141. t3.1 ELIMh 3UOJL. U.bL.. 4476.3.

r.

1-.

0.576 0.0 364.. 0. -416101. 0. 20.01 77b40. !AC`,. 21` _. 341 .l +t(Juu. 0.tlU.) 4.479.., III t•l.

CRUISE ► z, n 1.630 0.0 0. 416301. 0. 18.16 7h037. '/. 21t,1. 4.6 36-0.1 4+3JG. U.8"%: :7764u.

DESCFN7 416301. 0. 17.06 -0.659 0.0 27-13. O.b 3614." U.

U. 0.PG1 3774 4 1. ±,. 7PCFL. 6.

OFCFL 416301. G. It1.61 -0.619 0.0 1`. .- 1 -IV 5.n 0.

U.6V. 371411. 7. e. 7b7 V r.. e4. ?tI UESCI- 14T 3UJuU.

0. 19.56 O.S79 0.0 1.3. 16.0 401.e 0. --416101.

.YtS. 1 1 721. 3('0( .

CRUISE 44( _. O.b( G 3766'•..

18.20 0.649 0.0 C. -416101. O.

6,.6. t. 3b 7 . .. JUO.. 3.L 4U4.1 1SUG. U.4JG 3I 37:.1.

I L91TER ".^ C13 N U. 0. 0. 0.0 0.0 0.0 b23"7. -3t a. G. 0.0 44- 14 RtStl U. U.0 37311:1.

O.0 0. 0. 0.0 0.0 l•. 0.J 404.1' 0.

0. 0.4 ' 3731:.1. G. t• 73 , 7.

kFSFT 4..`. 1 0. -416101. C. 1'+.bb 0.560 0.0 40 .S +I. 1L,I. -1.7t. t96!.

L,.UISt 45Uu0. G.( JU IAKFOF& 445.E 0. 416401. U. 0.0 0.318 0.0 U.c 3f 5n. J.

POwtk I J. U.j 0. 0.0 0.318 0.0 1.3 446.6 0. 416401.

. u.. 591. 9061... ( ,.

P7w F 4 0. 3W 0.:3• 0.480 0.0 "f . .8 0. 416201. 0. 18.28 -171 11 • Ito. I'.. 2.2 J. J. - I. 36-03.. 1514.

CLI4P 0. 416201. 0. 18.38 0.459 0.0 U. . 1 U. 14-. 0.0 44E.b .•+1•c. 34. 331 L. !I I(.

A,CEL I.rOUJ.

0.0 411:01. 0. 18.25 U.S16 4S3"4. s•,. 41-. 6.6 45!.4 0.

k I:J.J. .45. 36^)31h. J,,4.

Ct14 U. -416101. 0. to _46 0.581 0.0 t'. ' 965' a. 4. . 46. 6.3 461.1 36. J"-. 1 l 1 4.

c P U1St $1'• • T . 0.'

C, 416301. G. Ib.22 -0.5b4 0.0 140. 10.1 471." 0.

0.n •. 7S..JYi. •.11. •t 7Ull. tl.

1 41 AU.LJ.

Utsc` 7v 0. 18.26 -U. S44 U.0 14r. 1,.0 471." 0. 416 1 01. N 3!.+•.7:. .. '.7G1l. +.

UfCEL 1.... 1. G."

Q^ 0. 18.(6 -0.676 0.0 5.0 474,.b 0. 416301.

l•.-'r 1S.•/(•. I. 97279. 171• ' ASL` V I '( ... ,. 41 G. 15.41 0.581 0.0 4.4 461•! O. -416111.

U l f . . __.. 0'i• vhl1/. 26i.

♦ JJJU• CrUlIt 0.641 1011. 11.: 0. -416101. U. 18.50 0.0 (111. 104221. 3u .O X IS.•J. ..311' '513/1.

CkUISI ____.

-rtwwwt. t ( U M M A _^ S T ! k V I WI A'^U PF(ilU(If(lY PRUCUREMiNi tJ IUTAt Pik TUTAI M A I f k I A t PAC() A/C 0• PER PROD A/C**

• LAPIIN

'1tVELOPMtNI - -1491L,lk..ING 1;77.67 TOTAL S1kLILTURf I.It5.vS IV064.61 PPUD(7LT10N 46114.18 3427.49 wIN • . /)55.3. 5183.33 t4GIVtfRING 1101.S1 11 101 G.t 0.0 0.0 INIFGR LOGISTICS SUPPORT TOOLIN'. Sr-t..t,6 loll (n.kl 461.34 t71.115 PLANNING 33.09 TEST A+710 c S /1, Y3 PUUY I21T. L4 8hvb.64 1..115.8E DATA L'.. ALII•HI1'46 GIA LI lvo.31 4G.t:r 1137.20 1RAINING 11.25 • •,AC! Lt t SYS IE AIS I VI,MIdGMT r N:, SI L T ,G3.vS 1 3S2. 11 1656.06 CkUISt ENGINE EN', StLlj"" u.0 L.0 TRAINERS G.L 360.btl LIFT FVoINF L,.0 NALILLt 16n.41 1294.54 15,68 .15C F AN J.J All, INDUCTION 411.011 P7.56 HANDROM S 48.11 51.52 AVILINIC1 (.L OTHER SY11t'4S t,... PkUV Lit SIUN 1«.11 1.4.11 2°4.24 FACI1,11ItS 0.0 FACILIII f S 34.75 0.4. 1NGI'# INST A LL G.L 34.75 TOTAL AIR VLHI CLI 17Ab. 1"k, - S1 1 - II kfV M S! O U.0 7.10 1.10 SSE - CFE 23.09 I XHAU',I SYSIfM U.G G.L G.O INTFGk LOGISTICS SJPPO47 4.06 LAG 111'. CONik Lit S 6.46 1G. 51 SSE - C , FE -151 r17 Y!ItM 84 PLANNING II.uI -TA47IW, ' _ 10.91 13.31 . 2P TOTAL Ill 1227.29 1RAIN1W. x.76 PkJPtLtIF INSTALL J.0 0.0 0.0 HANUNUJKS •.38 LL13' IC A 11N', SYt1LM 0.0 0.0 U.0 SSE I L i t L SYSIfM tt..Ok I04T7IAL SPAkE1 b2.',2 152.60 COST 6332.06 TOTAL It!, 45.'.j UPI VI SY1 IPw 1. T O NS O.J L.0 G.0 PRODUCTION OEVELOPMtNT O C TOTAL DVLPMNT - VtN - ikL11./3 Y11 L ` 15 .4Cti.r'. 1461-.../ 10f11 . FNGINffRING Sc 346.11 FLI:Ail LOAT t .^l' '25.J' 194. U1 v19.39 y 1 AUX P.,wfk PI ANT 12" .4j, 147.66 1'1.63 TOOLING -174.83 N Lit vi- tf'v9fNI - ,+ C114IPkt] *.IYvt:1 INSINI,M'N1S 1-1.73 79.30 161.03 HY..kA'1l If PNtOM 1 1".24 .P3.54 662.83 ENGINE S • 0.0 AIR Vt4ILLI tl...l ll't1•ICAL TOTAL 41U .G1 1331.43 1801.95 PROD DEV 171.28 SPARE I AVICN1L INS T ALL +8.71. 41I.6b 410.43 AkMA-4I NI U.G U.0 0.0 IOIAL JVLP`WI- "Lk v.+. 13 pup : A'', 0k tO.IIP I L%.4,' 4t,63. 13 4724.16 LI•,;']II('V1v:, AIv ra1.it 531.11? 434.38 TOTAL PROCUREMENT 53404.78 GOVMNT ,.

L)VL , 'MNT LOST ALIT-Ill'.:. .I.4r 24.1. 51.10 v-Ji Of. 61APHII O.0 0.0 0.0 L L A J A'/(' MANDL ING O.v G. U G.0 JVLP-.1 1C1AL . t-Sl 1t J.u7 SYSTI"S 1N1 1 1F , r.bl 654. C 4 1114.65 It IAi Cl %1 frv6.P4 :1453.:2 31450.35 I" J A L AL . c •• r, 1.72 P21.72 f N6 CNA NGt L k V- Q !- 1033.03 1` 146 t.USIAINIV'- COSY 1496.31 PkV , I1101,1Nf. CW-I 16b6.28 • LUAl I1Y ASSI kA VCt :042.07 • - MILLIONS OF DOLLARS M1.','.tl t A,,.O(1. --- IOt.P. 63

r

71'11 AI Alk/k A41 LWI 40306,64 •• -1000 OF DOLLARS UR

x

HOIIRS PER PROD A/C N 01SI 4671.74 t'lt.l^/t co AVl."11CS 1111-1 1111.00 000 - INCLUDES PRUU UA1AA tT 7l 1 AL MAtiUF At 101-1 41 EVS1 45429.41 SYSIEMS iNGR AND O+

r

3 UTHER SYSTEMS INTERCONTINENTAL P&WA E a tr 2r 4. 7r w4^ (STF"5M-7C) Pk'1 r1 1L11C': 1' , Itl C"'.I 40114.1E 1 VF F Al Il'NAL C( S11 ulkf%t INDIN"(1 O Ptr. 1 N .n' 1 4.•111 •• +rL CU'.1 Iu_LI A11( A L CUS1 IIUCI MISC. DATA L/%M o o . Pt kU N1 ► it ^rSl FM FLU, 1 i w '_N .1 1•..t'.1^ 0.02211 0 I.L164b'^ RANGf IN.

M1.) 65U0.60 i v.l' ► r, 394 F U_L & %3 tfI L ,.3 /4An L' CAL 6.l', 1.68796 bLOCK SPftO IMP") 435.66 r A11 LRAF Y U OI0kJL U. 17G4A 111NS1 U.JG2Ur BLOCK II"S 14. 111 L : . .11 N A! 11 -J( ,A-47 DEPkf CI tl l{1"I ft - 6'.

(•.: S: r 1 20.65',01 FLIGHT TIMF IHRS) II.0 . t.

14.78 MAI y l L14A'.1,f .lot',.

f J'1 ANV IfV/kA 'a L.U78b1 44 'U.h..• 6. 140 AV(6 SIAGF tfNGTH IN. M1.) 2547.OU + P1:' ' NGF ► 14 AN[`I.INI• AVG ;1.1 3 4 03 10.'+648: CARGO PfR F L'iGHI 18386.00 7117 AL WC I.:b,." . at•L C-bc, L AP•.I l "AN[ L 14(- 0.04.62 3.4b 7C 3 UTIt12AT1ON (MRS PIR YR) 4133.00 CIt" k VASSfY(.fk tAPFYSt U.4b4UU 39.59520 FLIGHTS PfR A/( PIR YEAk 279.Sb U7w k CARGL . 1PtNSt (•.OU`9S U.4bt.83 FARE IS) 504.40 G I N' kAt • AU'4I4I!,1kA110N ► .I(•5e3 8.64141 if 1' r 1.:2231100.000 •1• - CENIS PIR SEAT N. MILE I N LT\ kATE L' F RE TUkN ;N+ INVE STMENI AVG. N" Yt A4 AIRCI . lil AVLI- A , t LOWUl a71 V1 AVf kAUF kt VE NOE IN1tk: ST CrPf RAT INC. ( AStt ROI A I',4AII An.,' , 1 ..t^.l y l%1 Ut Pk{' IA I1(1N t•O'l1t fxPf NS1 1XPFNSf FI.Ow OUR 1 146 DU lr 1'y .1 JR 1W, V ! t U ► (IF A4 Yt V 4, Yt An 1 t I F7 S.

1M fM 111 fti •M fM PERLENV ; 3,61 1 6.? 1 .^ tvti..! 4IU.7n e%3.bv 37.86 142.38 -116.21 20.15 In.3 / 1 Iv1S.:•. 14 ••40.3 , 6t.O.I 1 92. 3v 37C.1@ -66.88 20.33 f c6.C. (,., I.6..15 16: .o. - I1LI f3 812.45 1G2.o4 .4 S 1, 101.61+ 10.88 IN. .I 4 .., I.o..l' 23• .6' 1 ,.S ! 7 P12.4, 90.81 4SS.60 101.74 21.83 4 31.•, 4a 26.. 4 .. l.oc.l5 v49.7+ Tr. 76 612.45 45%. be 113.80 22.93 b iU.0 ., i.o..l' 38F.11 74.04 b12.45 66.64 4'•S.6^ 11y.b6 24.23 7 ^C.( •. 1.'r,..l' 46'.P-, 7`rb.31 b1i.45 54.12 4'.'.60 125.92 25.76 53`•. • •7 b12.45 4.•.41 451.60 131.08 21.63 '+ . .0 ..4•, .15 +11.3, n46.h5 N12.4% 3u. 24 41%.60 138.03 210.42 13 !( .t . , +.6(.15 6101 . ,.'4 11.1 f 12.45 IC. 17 4IS.60 144.09 32.83

r

UVI1- YIA11 AV., t01 11+1 IV P "IW 4.1•• 11ERCtN1 N CU PBVYA E 3 INTLHCONTINENTAL rn tT ISTF505M 70 tJ J no Lk Allcf.AI COSTS . v l II- "Ct k.)IJNAI C'11 tU:'LI ItUIFE(I UVfkATIUNAt COST I IUCI DATA N MISC.

t ► a L/SMs I•tklt'.1 VE ic[ NI C/5,M ►► • FLI' ,I II i.M'w U..ZL.I• ► /.>Jte7 SvSIE•' (,.02 5901 1.6473 t• RANGE IN. M1.1 3000.00 FO I L A*0 ('IL '1.31/1'+ 14.:33;4 LILAL C.Z'.361 12. v ZOPI . fIOCK SPFFO (MPH) 420.76 1f0bt'k4vCE P +.ulti.: l.5,+.:+ AI.1.1.01 LL+NT JL ...JU4S1 0.78641, BLOCK E IHRSf TI M 7.13 of Pk( (IATJUN -421.. CFNIV AIT'NUAN1 (1.263b1 16.74011 TIME INNS) FLIGHT 7.13 MAINIEVANL( V. -4.1 4 t , F;'^3 A'IU I, tY • kaSE u.081!> 7.22'+81 AVG STAGE LENGI.1 IN. M1.! 2547.00 V ► ^,*-INCEk M IIANULI9: (1.Zv040 1r.42819 AVG CARGO PER Fl, IGHI 16136.00 TOTAL J IC 1.1V.(!. ......

1 u LoRtIt HANI • tivL G.ovi3S S.86US3 LIT ILIIATION IHP5 PER YRI ♦133.00 UlHt M VAsstNGtk t IVtN`_E 0.4840L 30.71364 FLIGHTS PER A/C PER YEAR 579.67 CAR(.(. 9%9E-4A UIHEk C.0G545 U. 37763 FARE Isl 231.40 6f4tPAL • AUMI4IS1kAl1O% T. 79872 0.lc[90 1( • lAL ]UC 1.S7S8N IGL.000 CENTS PER SEAT N. MILE

a

- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -

N

v kATE OF R ETURN ON INVESTMENT YLAR AV(. N1 AIRCkAt?

AVEkA(,t CU'tUtAIIVI AVIkAGE REVENUE INTEREST OPERATING CASH 001 AI Q LAA . I AU.+t , I +vtSIMi 1 41 UEPkt(IAIIUN FO 'II EKPENSt fx"NSf FLOW nt1RIN , 11URIIV UIJR IN( VALUI Of Ytik Y,Ak YZ Ak tiE!7 sM sM sH sM sM f ► J sM PFIMENI I 6.3 1(.v 3v II- 21.011 3 It- It. 242.VB 37.86 155.90 -128.43 16.85 2 16.3 11,., 1:115.51, b_.lv vo0.3u 631.74 91. 31, 405.31 -96.65, 16.95 3 ZO.J U., 1:6..1` Ib,.v7 1101.13 717.'3 102.Qv 498.87 62.59 17.33 4 ZU.0 C'.•.r 1cr...IS 236.65 777.53 1015.5'1 90.87 498.87 66.65 16.02 S 2().0 3.. 1.0..1:' 1,49. 11 '1:.IS 711.51 78.76 49P.P7 74.71 16.62 (U.0 1^6..1ti 3NF.11 C7%. 777.5? 66.64 498.87 80.17 19.75 7 :G.1, 1,.. ::c..l + 46'.Pr 79P.31 777.51 %4.5,2 41,8.81, 66.81 20.67 A J: `.1 771 S3 aL.0 .57 /:2.:b 4: .41 4vP.F 92. b8 72.77 0146.As . ....IS 61`.3., 777.53 30.20 49b.87 9;.94 23.88 IU Z(•.L (. 1 .(i I. • 71.11 777.53 it! .17 408.P7 25.99 10s.CA)

r

AV- F1,1 LVE G - Trll IJ Y C A.' Vtkl(lU = 1 1, .7± PE R LENT $U N PJLVVA E 3 INTERCONTINENTAL Q` ISTFS05M 7C1 r u tkA11r1.A1 C:1S1S IPrk4I IJ`tAL !D 71 CI C'.'ll IUl • LI ► ICY of671UNAI 1KUI OP CO,,I I IUCI MISC. DATA L/S I•t kLI III C/sMoA• Pt U LF NI FLI(srll LRPW 0.1ZL.c.

II 5•`11 (.0a"••vn 1.6413H RANGE IN. M1.1 31000.00 FUEL A40 UIL 'x.3111'+ (4.533;9 L i L L L C.:u361 1!. 9 !0811 FLOCK SPFfO IMPHI 410.16 INSl1K ANCE 1.-., ,:r a R J. ulb.: AI 3 r.1 41 1 LL • NT 7L ... JU • .S1 U.:8649 eLVCK LIME IHRS) 7.13 0tPRtC1 A1IGN .42% ► ».u:o ±! LAI1I'4 All I NUAN1 O.!6361 lt^.1~012 ► LIGHT 1IMF iHRS1 7.13 MAINTFNANLE U.3ul9tn 5,r)',4: F:'Pfl A •4u ••1v•{.A;,E U.UbzjlS 5.22'01 AVG STAGE LENGTH IN. M1.1 2541.00 VAS tNIFf r P ANU1.14' u.2v0%0 16.42819 AVG CAR(.n PER FLIGhl 1,8386.00 TOTAL U I.tvlr• IU .,PU, LtRt I. "A 14(A 1 1 t;1 1..04235 5. N6U51 UT1l I1Al 100 IHRS PER YR) ..I 33.170 tilt* R • VASSF IGFP It XPI y SE ,'.4 B 4C 30.71364 FLIGHTS PER A/C PER YEAR ?79.67 CAR(.( , VP & NSt 111411'k L.UG595 G.377C3 FARE 111 232.80 Y Uf4IRAL • ADMINI' . 14oIi(IN V I , ev0 1.714b71

C

II IF ► IUC I.`7%k 10(.00( •^• - CENIS PLR SEAT N.

P41LE ;- a ..

-.

- - - - - - - Iv - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 4ATE GF R ;1N ETUMN INVF^TMENI Yt AR AVt. N 1 AIMLKA ► 1 A.VE1,ALA CU•'t.1I AIIVt +V14ACE ME Vt.Not INTEREST OPERATING ► 1 CAS Rol ► 0 , AI W L4A 4 AO - 1 vV(jIMf NI 1. U. 'k UEPI.Ef IA11J r1 fXPt NSE IF XPF NSF FLOW OUR IN, DURING UUR1Vl VttUt OF Yt AK Y. A4 YEA4 {l ► °_1 S . 1'" fM sm 1M TM JIM PERLENT I L. t 1' ... tvr...

7 A.6/ f 14'.76 242.48 37, bn 155.90 -128.43 16.85 2 IL.+ lu.. l.:15.. r .?., 611./4 ,41. 3 4 40'5.33 -98.65 16.95 .U.0 U.. 1:6.,1 16 . 9 1 I toil .11 71/.13 IU2.4v 4,48.81 61.59 11.33 26.( .. I.r..l'• i3n.t•`• l.)LS.'•l 171.5! 90.87 49111,87 68.65 18_02 5 1U.0 J.. 1.

..1' • •1: . ?h 949.77 111.51 1121,76 4vP.P1 14.71 IB.b2 6 'r U-) .I1 14.U•. 711.5+ 64". 64 49b.h1 00.77 19.71+ .L.t t :O..1'• 46'.1'•. I'll .11 711.53 14.12 498.011' 66.81 20.57 M :L.^ U., 110:,1. •-3••.'+1 1.2 t 711.5; 4..41 4vP•P/ ,4l. b0 22.21 ^.' .^ ...1`• r.l•. • • I F1.%3 30.io 491'.87 98.94 23.88 .

:t•.L / 10 1. L: .1', r.41.411 /1.1 1 771.53 IF.17 490.07 105.00 25.99 r AV.. V • lr,F Y • a • PtI lHU {• .,1 1J - 19.1 • PEOCrNI ^u l.)

(S 4- PILVVA E3INTERCONTINENTAL (STF505M 7C1 I I A M 1 n F P E R F O R M A F N C E W O I O I 0E16 1 4T PL,vo FLAV VLLALI tN(.:^P GR AD GRDCFD GRDCAE XFIEL0 I CLRU'. CUkU', CLSII'P CDS1,-I ILLU FNLO VM ALL Vk11PS VI FMS VLDFPS V2 FPS XG Fl XL, I F1 XRO7F7 XUP FT X10 FT 630-l.

A. J2 e5.00 I. 4. 0.OP30 1.0000 1.0000 9+60.5 I 0.7 4" O.1CO5 P.It103 1.v!2,. I.20tiO 6~67.4 21—"I Z4b.1? ab.5 .161.v4 Z6,).ZI 0.0 1993.11 1965.3 6226.5 !

I r P&VVA E 3 INTERCONTINENTAL (STF505M 7C)

a

I N ^D

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