Fuel conservation possibilities for terminal area compatible aircraft

Section 6.2.2 presents a detailed technical comparison of the four airplane concepts that were central to the study, in terms of principal airplane design parameters. Section 6.2.3 presents an economic comparison of the four concepts in a manner that accounts for assumed differences in delay that would be experienced during operational use. From the technical standpoint, comparisons are provided for fuel usage, takeoff gross weight, and auxiliary airplane performance parameters such as field length capability, altitude capability, range factors, and others. The following paragraphs, therefore, provide a comparison of both the technical and economic characteristics of the concepts, in terms of the principal terminal-compatibility characteristics: namely, congestion, noise, and emissions, as well as fuel usage.

6.2.2 TECHNICAL ASSESSMENT AND COMPARISON OF STUDY AIRPLANES Table 18 contains a comparison of the principal size, geometry, performance, and noise characteristics of the study concepts. As shown, all aircraft were designed for the same 196-passenger, 5556-km (3000-nmi) range. The principal design parameters that affect the concept size are range and payload specification, takeoff field length and approach speed constraints, and initial cruise altitude capability.

The CWB-E wing and engine sizes were established by takeoff field length and minimum fuel usage. Similarly, the ATT-E was thrust sized by takeoff field length and minimum fuel usage.

The TAC-E wing and engine sizes were selected primarily from takeoff noise considerations.

The engine size of the TAC-E was sized to achieve sufficient climb gradient to reduce takeoff noise to approximately -15 EPNdB relative to FAR Part 36.

In contrast, the TAC/Energy wing and engine sizes were established by approach speed and minimum fuel usage. Approach speed is not an independent consideration from approach noise measured at the FAR Part 36 noise station.

This situation arises because additional wing area can be traded against decreased flap setting in order to achieve a desired noise level and still satisfy the approach speed constraint.

The TAC/Energy concept wing was sized using maximum flap deflection of 50". It could have been sized at a lower flap deflection (30*) with a slightly larger wing. This would reduce approach noise by approximately 3 EPNdB at a fuel and DOC penalty of 0.5% and 1.5%, respectively.

However, the configuration would have a maximum glide slope capability of -3.5".

The engines were sized for minimum fuel usage and not on the ability to achieve sufficient climb gradient to reduce takeoff noise to -15 EPNdB below FAR Part 36. Had the airplane concept been constrained for takeoff noise, a 3.5% penalty to fuel usage would have resulted. This noise level could have been achieved by an engine size increase or use of acoustical engine treatment. A typical engine and wing sizing chart is shown in figure 101 for the TAC/Energy concept showing the effect of performance constraints.

The airplane concept as sized assumes that the wake turbulence problem can be solved by some "zero penalty fix" or at worst by a very small penalty to fuel usage. At this time, the current definition of "programmed" flaps (retracting the outboard half of the flaps) to achieve a lift distribution with acceptable levels of wake turbulence probably allows Table 18.-Summary Aircraft Characteristics Characteristics CWB-E ATT-E TAC-E TAC/Energy Cruise speed, Mach 0.82 0.885 0.885 0.80 Takeoff gross weight, kg (Ib) 147 200 (324 600) 132 400 (291 800) 140 800 (310 500) 115 300 (254 200) Operating weight empty, kg (Ib) 85 890 (189 360) 74 080 (163 320) 82 120 (181 030) 67 351 (148 480) 2 2 (111.4) 5390 (112.6) 5690(118.9) W/S, N/m (lb/ft ) 5290 (110.5) 5330 ( 0.282) 2.83 (0.288) 2.34 ( 0.239) T/W, N/kg 2.94 ( 0.30) 2.76 ( 27 400) 99 600 (22 400) 67 600 ( 15 200) Thrust/engine, sea level static, N (Ib) 144 100 ( 32 400) 121 900 4/ATSA 4-2800-24 4/BPR 6.0 Number of engines/type 3/CF6-6D 3/ATSA 4-2800-24 2 rings/1 splitter Peripheral lining Acoustics Peripheral lining 2 rings/1 splitter 2 2 Wing area, m (ft ) 272.9 (2937) 243.2 (2 618) 256.1 (2 757) 198.6 (2 138) Sweep, deg 35.0 36.5 36.5 25.0 7.6 9.0 12.0 Aspect ratio 6.8 Performance 2 530 (8 300) 2 530 ( 8 300) Takeoff field length 2 530 ( 8 300) 2 530 ( 8 300) at 302 m, 32.20 C (1000 ft, 900 F), m (ft) speed, m/s (kn) 69.4 (135) 69.4 (135) 61.7 (120) 61.7 (120) Approach 500) 11 550 (37 900) 12 130 (39 800) 11 050 (36 300) Initial cruise altitude 11 130 (36 capability, m (ft) Range factor, km (nmi) 21 020 (11 350) 21 520 (11 620) 23 020 (12 430) 24 800 (13 390) *Fuel usage, kg lb ) 0.0328 (0.134) 0.0301 (0.123) 0.0306 (0.125) 0.0225 (0.092) seats x km seats x nmi Community noise (Est/est-FAR Part 36) using FAR abatement procedures: Takeoff noise with cutback power 96.6/-7.0 95.4/-7.4 88.1/-15.0 92.1/-9.6 Sideline noise 95.5/-10.7 93.3/-12.6 90.3/-15.8 86.2/-19.3 Approach noise 97.7/-8.5 96.9/-9.1 96.7/-9.4 102.1/-3.4 Traded noise -8.7 -9.4 -11.4 -5.4 5556 km (3000 nmi) range; 18 140 kg (40 000 Ib) payload *5556 km (3000 nmi) stage length Range = 5 556 km (3 000 nmi) Payload = 18 140 kg (40 000 Ib) N/kg 4.0 - (Ib/lb) .40 - Block fuel, kg (Ib) 26 310 (58 000) 25 860 (57 000) weight, Gross 3.5 1000 kg (-1000 Ib) .35 127(280) 122.5 (270) CL ratio T.O. noise-15.0 - 3.5% block fuel penalty 1.00 T/W 3.0 / 950 (55 000) .30 / / 117.9 (260) TAC/Energy -1* 0.90 S24 490 (54 000) 1\\ 113.4 (250) 2.5 .25 0.80 2.0 .20 - VAPP, m/s (kn) 61.7 (120) I I I I I I 90 100 110 120 130 140 150 (lb/ft ) I I I I I I I I 7000 7500 4500 5000 5500 6000 6500 N/m W/S Figure 101.-TAC/Energy Sizing Thumbprint on the order of 1.85 to 5.56 km (1 to 3 nmi) for those airplanes separation distances currently required to fly 5.56 to 9.26 km (3 to 5 nmi) apart. However, the loss of low-speed definition "programmed" flaps required decreasing wing performance from the current speed constraint and increasing thrust loading by 19% loading by 25% to meet the approach fuel burn. This would cause a large fuel burn (7%) and economics (9%) penalty to optimize defined.

for the "programmed" flaps as currently 6.2.2.1 Fuel Usage comparison of the fuel usage and the Figures 102, 103, and 104 present a summary range factor, and OEW fraction. Relative to the CWB-E principal performance parameters, usage is approximately' 8% less as a result of improved cruise L/D concept, the ATT-E fuel technology structures), and decreased (advanced airfoils), improved OEW fraction (advanced propulsive efficiency (S-duct and acoustical lining). The range factor, on the other hand, and shows only slight improvement. This is due to a combination of increased cruise speed L/D (AR and advanced airfoil technology) offset by propulsion installation inefficiency TAC-E fuel incurred because of the extensive use of acoustical treatment. Similarly, the 10% better, except a 4% fuel burn penalty was economy would have been approximately TAC-E concept incurs traded for a -15-EPNdB takeoff noise relative to FAR Part 36. The penalty in OEW due to the many features added to provide terminal compatibility. Of these to features, the changes to improve low-speed aerodynamics also provided, improvement both cruise range factor, primarily from the increased aspect ratio and increased wing area, of which tend to improve cruise L/D.

TAC/Energy concept yields a Assuming the full fuel reduction potential can be realized, the fuel usage approximately 30% better than the resized CWB-E (32% better than the CWB), which was approximately 2%. The including the weight penalty of the TAC features, preceding 30% does not include effects due to delay reduction. The TAC/Energy concept TAC-E, approximately 26%, with 3% of this differs dramatically in fuel usage from the with the revised list of terminal compatibility features. The substantial saving associated of other changes (cruise speed, wing geometry, engine cycle, and propulsion number efficiency) accounts for the remaining fuel savings.

Figure 105 presents a slightly more detailed comparison showing approach speed, cruise Mach number, and a breakdown of range factor into its aerodynamic and propulsive efficiency factors. A comparative weight breakdown of the study concepts is shown in table 19, while comparative cruise efficiency is shown in figure 106. One additional fact indicated by figure 105 is a slight improvement in propulsive efficiency of the TAC-E compared to the ATT-E. This improvement in efficiency is a result of (1) changing the four-engine engine configuration from one that includes S-duct internal losses to the on-the-wing configurations used for the TAC airplanes, and (2) an improvement realized in the cruise SFC for the primary engines as a result of using an in-flight dedicated auxiliary power unit for the TAC-E concept. The SFC savings in the engines result from reduced penalties associated with bleed and horsepower extraction which were offset by fuel expended by the APU units (i.e., the fuel expended by the APU unit is about equivalent to that saved by the engines).

(seat x nmin kg Seat x km = Range 5 556 km (3000 nmi) Payload = 18 140 kg (40 000 Ib) .040 .16 .035 .14 - .14 -- Fuel usage .030 .

.025 .10 .020 - .08 ATT-E TAC-E TAC/Energy CWB - CWB-E at (resized) LRC Figure 102.-Fuel Burn Characteristics of Study Airplanes 1 000 km km (3 000 nmi) Range = 5 556 - Payload = 18 140 kg (40 000 Ib) nmi) (1 000 26 - 14 - 25 - 13 - 24 - - factor 23 Range 22 - 21 - 11 - 20 - 19 - 0 CWB CWB-E ATT-E TAC-E TAGC/Energy at (resized) LRC Figure 103.-Summary Performance Characteristics of Study Airplanes Range = 5 556 km (3 000 nmi) = (40 000 Ib) Payload 18 140 kg .59 .58 OEW OEW .57 - TOGW .56 -

Frf l i

o

TAC/Energy ATT-E TAC-E CWB-E CWB at LRC (resized) LRC Figure 104. -Summary Performance Characteristics of Study Airplanes m/s (kn) Range = 5 556 km (3 000 nmi) Ib) 75 Payload = 18 140 kg (40 000 .9

APP'

MCRUISE

.8 -

60 U

- 3 VAPP kmN x 10 kg (nmi) 15 - 800- 14- V/SFC L/DCRUISE CWB-E ATT-E TAC-E TAC/Energy CWB-E ATT-E TAC-E TAC/Energy Figure 105.-Additional Performance Characteristics Comparison Table 19.-Comparative Weights Breakdown ATT-E, CWB-E, TAC-E, TAC/Energy, kg kg kg kg (Ib) (Ib) (Ib) (Ib) Item.

15 663 19 069 13 749 group 15 654 Wing (42 040) (30 310) (34 510) (34 530) 1 483 1 275 807 Vertical tail 885 ( 1 780) 950) ( 3270) ( 2810) 447 1 383 1 497 3 298 1 Horizontal tail ( 3050) (3 300) ( 7270) ( 3190) 18 765 16 670 group/nacelle & strut 17 123 22 145 Body (41 370) (36 750) (37 750) (48 820) 6 396 5 779 5 493 7 847 Landing gear (12110) (17300) (14 100) (12740) 10 809 10 496 6 378 group 10 301 Propulsion (23 140) (14 060) (23 830) (22 710) 18171 18 720 16 670 Fixed equipment 17 210 (36 750) (37 940) (40 060) (41 270) 157 4 940 items 5 094 5 484 5 Standard & operational (10890) (11 230) (12090) (11 370) 871 975 871 1 021 Miscellaneous 1920) (2150) ( 1920) (2250) ( OEW 74128 85921 82196 67351 (181 210) (148 480) (163 420) (189 420) (considering the lower of the TAC/Energy engine cycle The overall efficiency (V/SFC) treated engines) is equivalent ECS system, and BPR 6.0 peripherally cruise speed, redesigned (- 3.5%) than the TAC-E concept.

ATT-E concept but slightly worse to the for the four loading and wing loading characteristics 107 summarizes the basic thrust Figure previously.

for the data trends have been discussed study airplanes. The reasons 6.2.2.2 Congestion (wing trailing to the TAC airplane for congestion relief All the features that were added deceleration capability, reduced approach speed, increased ground vortex modification, included in the final improved avionics) are powered wheels, and high-speed turnoff, affect the congestion.

the other changes to the airplane low-energy version. None of the programmed flap vortex effects can be minimized by use of Therefore, assuming wake the will also be realized for for the TAC airplane congestion relief projected approach, the TAC/Energy airplane.

-- -

T77

20- TAC/Energy L/D 14.01 b e = 0.75, Cf = 0.0030) TAC-E ATT-E CWB-E O L/DLRC co 0 _ I I I I I I I 1.7 1.4 1.5 1.6 1.0 1.1 1.2 1.3 Span/A-e Figure 106.-Effect of Wetted Area and Span on Cruise Efficiency Range = 5 556 km (3 000 nmi) = 18 140 kg (40 000 Ib) Payload N/kg (Ib/Ib) .30

T/W

2 - .20 - N/m ) (lb/ft W/S 5500 5000 - CWB-E ATT-E TAC-E TAC/Energy Figure 107.-Engine and Wing Size on Noise 6.2.2.3 Low-Energy Impact on Figure 108 shows that the TAC/Energy airplane gave up some noise reduction potential the TAC-E the FAR 36 30 approach and cutback but is quieter on the sideline. In particular, achieved a quieter 30 approach and cutback noise than the TAC/Energy. The TAC-E engine ft) over the community and to permit thrust cutback was oversized to achieve 520 m (1700 to 56%. In addition, the TAC-E used a bypass-4 engine (optimum for Mach 0.9 cruise) and engine, rings, and inlet rings and a fall duct splitter to quiet the engine noise. The oversized splitter all caused fuel penalties. To conserve fuel, the TAC/Energy uses a bypass-6 engine However, the peripherally lined (optimum for Mach 0.8 cruise) with peripheral lining.

bypass-6 engine is about 2 EPNdB quieter than a similarly treated bypass-4 engine at both sideline.

cutback and approach conditions and 5 EPNdB quieter on the When noise contours and two-segment approach are considered for the different airplanes, the TAC/Energy. Figure 109 shows the 90-EPNdB the noise comparison changes to favor contour for both takeoff and approach. The takeoff area encompassed by the TAC/Energy TAC-E but the TAC/Energy 60/30 two-segment approach area is slightly is less than the than the TAC-E 90/30 approach. The net effect is that the TAC/Energy total contour larger area is 28% less than the TAC-E contour area. (See fig. 110.)

A further reduction of the TAC/Energy approach area could be achieved by lengthening the lined inlet to an L/D of 1.0. About 1- and 2-1/2-EPNdB reduction at cutback and approach,.

inlet or with respectively, could be achieved. Additional reduction is possible with a sonic inlet rings and splitter. The addition of inlet rings and splitter, at an airplane OEW penalty of about 1% and less attractive economics, would result in the 90-EPNdB approach contour contour approximates shown in figure 111. The figure shows that the TAC/Energy approach the TAC-E approach contour. Noise reduction concepts described in this paragraph would require substantial research dollars and testing to verify they are both feasible and practical.

6.2.2.4 Emissions A summary of the emissions impact is contained in figure 112, which shows the total pollutant emissions of CO, HC, and NO for the EPA landing takeoff cycle for the x TAC/Energy airplane. The data are compared to the previous Mach 0.9 TAC (ref. 2) concept and the TAC goals for the Mach 0.8 and 0.9 airplanes.

Various differences exist between the two airplanes shown. The TAC/Energy airplane has engines of smaller airflow size because of the decrease in cruise Mach number from 0.9 to 0.8 and improved aerodynamic data base. The TAC/Energy concept uses an APU sized for ground operations, which supplies power to the powered wheel system. The TAC airplane used two much larger in-flight operable APU's sized to provide anti-icing at the low engine thrust settings for steep (9*/3 )descents and, consequently, ran at a less efficient part throttle setting for ground operations. This difference in the APU's is responsible for the large difference in the CO emission.

Traded Approach Takeoff FAR 36 Sideline (Cutback) (3/2 trade) FAR 36 0 Noise relative -10 to FAR part 36 *Approach noise (AEPNdB) dominated by CWB-E airframe noise ATT-E -20 T A C /Energy TACre L Figure 108.-Airplane Noise Comparison TAKEOFF APPROACH (CUTBACK) CWB-E TAC-E 800- ATT-E TAC/Energy TAC-E E W " TAC/Energy CWB-E -2000 d and ATT-E 6o 40 000 000 20 000 30 000 -10 000 0 10 -40 000 -30000 -20000 (ft) (ft) 00 1 10 000 0 5 000 -10 000 -5 000 m m Distance from threshold Distance from brake release TAC-E 1600- 5000 ATT-E TAC/Energy CWB-E TAG/Energy 800 0 -- :--TAC-E

and ATT-E

-1000 30000 40000 10000 20 000 -40 000 000 00 .-20 000 E 800 Note: Includes airframe noise ~ ._ -V 1600 Figure 109.-Noise Footprint Comparison (90 EPNdB) TAC-E -3000 CWB-E ATT-E TAC/Energy CWB-E and ATT-E E TAC/Energy - 2000 APPROACH TAKEOFF(CUTBACK) 9 0 -40000 -30000 -20000 -10000 0 10000 20000 30000 40000 (ft) (ft) -10000 -5000 0 5000 10000 0 m m from brake release Distance from threshold Distance CWB-E 1.0 - T.O.

ATT-E (C/B) Relative 90 EPNdB TAC-E .0- footprint acreas I O TAC/Energy (C/B) T.O.

- approach - 60/30 approach S- approach acres = 9.05 st mi 1. Base area = 5780 2. Includes airframe noise Compatibility/Low-Energy Benefits Figure 110.-Terminal 90 EPNdB Noise Contour Area Comparison Approach profiles TAC/Energy ° 60/3 Drag brakes 0o 600 (e 3 TAC-E e Engine idle 90/30 * APU anti-icing 90 EPNdB APPROACH CONTOURS scale) (approximate TAC/Energy TAC/Energy 6e/3 (periph) 60/30 (2R/1S) TAC/Energy TAC-E 90/30(2R/1S) 30(periph) _ 1_ 4 3 2 (nmi) 4 2 8 6 km Distance from threshold Figure 1 11.-Steep Descent TAC kg/cycle Mach 0.9 TAC/Energy Mach 0.8 15 - (Ib/cycle) 7r/rnrm/r 30 - " .. .

30 777777 TAC goal 10- 20 - APU CO CO II 5- 10 - I -- I I Main S I engine 3- nmn7777nr HC 6 - HC 2- 4- r- - -i 0 - 0 I I 15 - mnn7 .-- 7 6C I NO NO 10 xI 4I I II 0 - 0I Figure 112. -Emissions Comparisons AND COMPARISONS 6.2.3 ECONOMIC ASSESSMENT 6.2.3.1 Method of Analysis costs and the value of investment was derived after The method of calculating operating airlines. Every effort was made to establish considerable discussion with the subcontractor and the real world sensitivity of operating costs methods that would correctly reflect characteristics of the aircraft concepts. Table investment potential to technological the comparison airplane the basic concept used to evaluate the TAC/Energy and shows concepts. (See sec. 6.2.3.3.)

Method Table 20.-Economic Evaluation assembly area $/Ib calculated on basis of study of complexity of each major 1. Price: Pay: constant $/FH 2. Crew and 30d/gal) 3. Fuel Cost: fuel burned times prices 5.8/liter and 7.9/liter (22d/gal 4. Maintenance a. Airframe to 727 and 747. Special analysis of subject aircraft compared System-by-system unlike systems studies of b. Engine American Airlines (NASA) formula 5. Trips per Year formula American Airlines 6. Depreciation (includes spares) 14 years to 10% residual value .7. Insurance 1% of aircraft price 8. Spares price 6% of airframe 30% of engine price 9. Cabin Attendants of seats: $20.39/F H/attendant. Food Number is function number load factor, and block time varies with number of seats, 10. Landing Fees Function of maximum landing weight: $0.42/454 kg (1000 Ib) and Administrative Costs 11. Passenger Related km (1000 nmi) flight Constant cost on all airplanes evaluated: $1426/1852 Value 12. Net Present 15% discount rate is year of Payment made at time of delivery; computational base delivery 7% tax investment credit tax purposes 8 yr sum of digits method for calculating depreciation for price of the TAC/Energy aircraft by major assembly areas.

Table 21 shows the estimated these costs was based on the contractor's experience. They have been Development of more complex manufacturing procedures, advanced materials, or other escalated to reflect deviation from standard current model manufacturing procedures. A 300-airplane run such was assumed. All calculations are in mid-1974 dollars. This table also indicates the recurring of weight.

and nonrecurring costs as dollars per kilogram Table 21.-Airplane Price Calculation; TAC/Energy Airplane Costs (mid-1974 dollars) Average cost per ($/Ib) Ncosts, $M Cost factor airplane, $M $/kg (300 airplanes) Wing 2.217 130.74 (59.32) 125.621 Body 3.368 189.74 (86.09) 211.201 Empennage 0.638 228.62 (103.73) 41.205 Landing gear 0.711 106.92 (48.5,1) 27.811 Nacelle 0.322 101.54 (46.07) 15.812 Power rack 1.116 456.85 (207.28) 44.453 Electrical 0.484 237.44 (107.73) 23.957 Electronics 0.955 497.86 (235.89) 19.514 Controls 0.581 234.11 (106.22) 46.125 Hydraulics/pneumatics 0.452 366.06 (166.09) 21.023 Air conditioning 0:311 201.38 ( 91.37) 14.480 Interiors 1.660 154.74 ( 70.21) 34.300 625.500 Total airframe 12.815 184.41 (83.67) Engines 2.344 Total 15.159 calculated on a constant dollar per flight-hour basis for all airplanes being Crew pay was compared. A value of $275 per flight-hour was used as the constant. This is above current airline crew pay, chosen to reflect the likely level for this capacity aircraft and with average that it would be the prime route choice aircraft.

the assumption simply fuel burn times price per unit. To evaluate the effect of potential future Fuel price is are fuel prices, two values, 5.80/liter (220/gal) and 7.90/liter (300/gal) were used. Results at both fuel prices.

shown calculated on an annual basis and allocated to the average Depreciation and insurance were The estimated number of similar trips per year obviously has significant trip being evaluated.

studied this problem and derived a formula for effect on this allocation. American Airlines was used except that by ground rule calculating trips per year. (See sec. 4.6.2.) This formula of change in cruise velocity of less than no change in trips/year was indicated as the result spares. Spares were estimated Mach 0.03. Depreciation was calculated on the airplane plus was calculated to be 6% of the airframe price and 30% of the engine price. Depreciation was based on the assuming a 14-year life of the airplane to 10% residual value. Insurance at 1% per year (estimated to be an average over the life of the airplane price and estimated as the previous TAC study, a special study was made to aircraft). Using the same procedures costs. The airframe maintenance costs calculated for the calculate airframe maintenance updated to reflect higher (1974) labor rates and CWB-E, ATT-E, and TAC-E concepts were costs above those estimated in reference 2.

inflated material in a manner similar to that The TAC/Energy transport airframe maintenance was calculated basis were 2. Hourly and cyclic maintenance costs on a system-by-system used in reference maintenance. Airline data devised from CAB-reported data relative to current transport and cyclic basis available to the contractor were used to allocate costs not only on an hourly on each system. As a result of the but also by the division between labor and materials study, all four concepts were evaluated updated information from reference 2 and this each system with the like system in current transports.

comparing used herein was derived by the method propulsion system maintenance cost estimate The (ref. 6). This method relates the propulsion developed by American Airlines for NASA costs to a repair rate and a cost of systems (ATA systems 71 through 80) maintenance a function of the flight length (flight-hours per flight cycle), the repair. The rate of repair is engine type), operational maturity of the engine (years of operational fleet service of the and replace the powerplant assembly and the turbine inlet temperature. The labor to remove build up the assembly in the shop are in the airplane and the labor to tear down and by engine weight). The shop labor to repair the base functions of the engine size (defined turbofan, engine was assumed to be 1400 manhours per shop visit for a conventional estimated to be proportional to independent of engine size. The material cost per repair was an estimate for the thrust reverser and engine price. In addition to the engine maintenance, starter systems, both proportional to flight-hours, was included.

formula were divided The accounts listed as "indirect operating costs" (IOC) in the ATA airplane-related. Flight those which were passenger-related and those which were into number of attendants were calculated as constant dollars per flight-hour, the attendant costs factor.

and tourist seats in the aircraft but not varying with load varying with the first-class as a function of aircraft seats and the assumed load factor, these Food costs were calculated trip. Aircraft servicing was calculated as a function of costs varying with block time for the same general size but with differing weights gross weight. It is recognized that aircraft of the the amount is small and it was decided to will probably require the same servicing. However, criterion. Landing fees this deficiency rather than make a study to determine a better accept differences between as a function of maximum landing weight to reflect were calculated aircraft of differing size and landing weights.

a special case. All concepts except TAC-E and Ground facilities were handled as ground facilities. A small deduction were charged a constant cost per trip per TAC/Energy of powered wheels in the TAC-E and from this account was made to reflect the advantage TAC/Energy type concepts, assuming elimination of the requirement for tugs to push the aircraft back from the gate. Estimate of the value of this saving was made with allowance for a backup tug being available.

All other accounts were held constant for all concepts. A representative number, estimated from previous work with the Boeing-Lockheed IOC formula, was used.

Net present value (NPV) was calculated for each concept, using the following ground rules: * Discount ratio = 15%.

* Payment for aircraft and spares.is made at the time of delivery (no prepayments).

* Net present value is computed to the year of delivery.

* A 7% tax investment credit was allowed.

* An 8-year sum-of-the-digits method was used for calculating depreciation for tax purposes.

6.2.3.2 Economic Evaluation of TAC/Energy Airplane Evaluation of the TAC/Energy airplane was made using the methods described in section 6.2.3.1.

Based on the analysis conducted during the TAC study (ref. 2), it was assumed that all airplanes in the fleet were equipped with TAC features, with delay being held to the 6 min of air maneuver objective established by DOT. This assumption, based on the estimated capability of TAC features to reduce congestion, makes the block time of the TAC/Energy aircraft less than the higher cruise velocity comparative airplanes when all are evaluated at the expected delay level. The combination of lower fuel consumption and minimum delay gives the TAC/Energy concept a tremendous advantage from the standpoint of fuel conservation as well as airline economics. The lower fuel consumption airplane obviously pays a smaller penalty as fuel prices increase. Basic evaluation was made at a fuel price of 5.84/liter (220/gal). Therefore, the evaluation of the economics at 7.90/liter (304/gal) fuel price shows an even greater advantage when compared with the other airplanes. (See sec. 6.2.3.3.)

Figure 113 shows the distribution of operating costs for the TAC/Energy concept, indicating costs at both fuel prices. Note should be made that maintenance costs on the complex features of the aircraft are greater than fuel costs, even when fuel is assumed to be 7.9/liter (30/gal). Part of the reason for this high level of maintenance cost is a conservatism relative to maintenance of composite materials.

All evidence to date shows good maintainability; however, since the sample from which the evidence is taken is small, maintenance of composites was held at a high level.

Fuel (19.2%) $601.66 (24.4%) $601.66 $817.80 Crew (22.1%) Crew $692.38 (20.7%) $692.38 Direct cost Maintenance accounts (26.8%) Maintenance S () $840.85 (25.1%) 40 $840.85 Insurance (4%) $124.25 Insurance (3.7%) $124.25 Depreciation Depreciation (27.9%) (26.1%) $876.35 $876.35 /gal) 7.9d/Q (30d/gal) 5.8/Q (22d Fuel Fuel TAC/Energy Aircraft Figure 113.-Cost Distribution, [1852-km (1000-nmi) Stage Length] 6.2.3.3 Economic Comparison Between TAC/Energy and Comparative Concepts (CWB-E, ATT-E, TAC-E) An economic evaluation was made of the TAC/Energy concept with three comparative concepts established in the previous portion of the study. Economic evaluation was made on the updated versions of these comparative airplanes (CWB-E, ATT-E, TAC-E), as described in section 6.1.2. Evaluation was made in the areas of operating costs (DOC accounts) and in the relative value of the investment in each type of aircraft.

Results are shown at two fuel prices, 5.84/liter (224/gal) and 7.94/liter (304/gal). The results are also shbwn under two delay conditions: (1) standard delay which provides for 6 min of air maneuver in the terminal area and (2) expected delay, which is the standard delay for TAC-E and TAC/Energy since the congestion-relieving features are estimated to increase runway acceptance rates to the point that delay will be minimal. An average expected delay of 36 min was applied for aircraft not so equipped (CWB-E and ATT-E).

DOC, block fuel, and block time comparisons are tabulated in the pull-out chart on page 229.

It is recognized that a few TAC-type aircraft in a mix of aircraft operating at an airport will make little or no difference in congestion. The "expected delay" case indicates conditions when all or most aircraft are TAC-equipped, thereby minimizing congestion. The economic potential of a 30-min delay reduction is greater than the economic potential of fuel reduction, assuming fuel to be priced in the range studied.

Figure 114 shows operating costs relative to costs for the current wide-body aircraft Two scales are indicated-one a comparison with the resized CWB-E and the second a comparison with the CWBLRC. When all aircraft are evaluated assuming each to have the standard 6 min of airborne maneuver, the TAC/Energy has a slight advantage in operating costs with the delta being greater when fuel price is higher, as would be expected. The basic TAC-E airplane concept has the highest operating costs when the costs of the noise reduction, emission reduction, and congestion reduction modifications are assessed with no offsetting credit being made for congestion reduction and consequent reduction in delay.

The advantage of slightly higher cruise velocity, and consequently lower block time, cannot offset the higher acquisition price and higher maintenance costs. When compared to the TAC-E concept, the primary economic advantage of the TAC/Energy concept is attributable to its low fuel usage. The CWBLRC is 1% to 3% more expensive to operate, varying with conditions evaluated, and therefore the percentage improvement is slightly greater in each case.

The lower part of figure 114 shows the situation for the TAC-E and TAC/Energy concepts if the TAC-type aircraft become a large portion of the fleet.

These aircraft are evaluated and compared with the CWB-E and ATT-E, assuming the predicted delay for standard aircraft in future traffic situations (36-min average delay for all flights). The TAC-E now shows lower operating costs than ATT-E or CWB-E. However, the TAC/Energy concept, with benefits in congestion reduction as well as lower fuel consumption, is estimated to operate at 84% of the cost of current aircraft with a fuel cost of 5.84/liter (224/gal).

Further advantage, due to lower fuel consumption, will be realized if fuel price is increased.

Inasmuch as aircraft prices are not greatly different, net present value calculations follow the same patterns. See figure 115.

fuel - 7.9/Q (30d/gal) 1.10 (22/gal) fuel 1.10 - 5.8d/Q -1.05 1.05 1.05 1.05 - TAC-E TAC-E 1.0 = CWBLRC 1.0 =CWBLRC CWB-E = 1.0 CWB-E = 1.0 -- - .95 - .95 .95 -- .95 TAC/Energy - 90 -. 90 ATT-E TAC/Energy S - .90 - ATT-E .90 .85 .85 .85 - .85 Noise a direct Relative suppression costs (%) operating penalty 1.10 1.10 - = 1.05 ------ --- 1.0 CWBLRC.

CWB-E= 1.0 - 1.05

- 1.0 = CWBLRC j .95

= .95 - CWB-E 1.0 ATT-E .90 TAC-E .95 .90 .95 ATT-E TAC-E .90 .- 85 .85 .90 TAC/Energy .80 .85 .80 .85 - TAC-Energy .80 Versus Comparison Aircraft Figure 114.-Relative Operating Costs: TAC/Energy [1852-km (1000-nmi) Stage Length] DELAY STANDARD 7.9d/Q (30d/gal) fuel 5.84/Q (22d/gal) fuel 2.5 TAC/ TAC/ ATT-E Energy Energy ATT-E - 2.0 1.5 - TAC-E TAC-E 1 I CWB-E net 1.0 , present value Relative value of investment EXPECTED DELAY (NPV) (%) 3.0 TAC/ TAC/ energy Energy 2:5 2.0 TAC-E TAC-E TAC-E ATT-E TAC-E ATT-E 1.5 - 1.0 1CWB-E net present value .5 Value ofnvestment (NPV [185-km mi Stage ength] Figure 115.-Relative Value of Investment (NPV) [1852-km (lO0-nmi) Stage Length] by measures to advantage offered significant potential economic reflects the The preceding reduction is shown to be, delay reduce delay and congestion. Important as fuel consumption fuel price considered. The economic payoff at the reduction has a greater potential the important in both areas, indicates with significant advantages airplane, TAC/Energy factors (price, maintenance, crew, etc.) are economic payoff even though the other in showing the difference 116 illustrates this point, higher. Figure comparable or slightly operating costs.

well as differences in total TAC/Energy and CWB-E as fuel costs between are evaluated at standard delay, the fuel benefits of the TAC/Energy When both airplanes fuel that is not is a delay difference, the However, when there are somewhat offset.

airplane is significant for the because of congestion reduction in holding in the terminal area wasted holding fuel was importance in this study, fuel was of prime Because TAC/Energy airplane.

and at an assumed speed for fuel conservation, for all concepts at optimum calculated fuel reduction ft). For all models compared, significant altitude of 6100 m (20 000 holding in reference 2 when of a time increment as reported by the simple addition was achieved study of optimum holding fuel consumption was simply an economic parameter. Future delay, the TAC/Energy might prove to be fruitful. Without procedures for fuel conservation relative to time-oriented than CWB-E and the advantages at a lower block time operates etc.) add to the total savings.

variables (crew, hourly maintenance, other than noise abatement equipment calculation assumed that no The preceding added for the If rings and splitters are had been added to the engines.

peripheral lining Fuel bum is DOC would be expected.

levels, a change in purpose of reducing noise The hatched result of these noise abatement procedures.

estimated to increase 3.5% as a indicate the expected variations.

areas in figures 114 and 117 Comparative Fuel Burn 6.2.3.4 provide study is to minimize the fuel necessary to Inasmuch as the prime objective of this airline system, analysis was made passenger service on the United States domestic adequate on a partial system on a per passenger kilometer basis and also of comparative fuel burn indicate substantial fuel savings aircraft and routes. The results wide basis involving several aircraft (CWB-E), even when concept relative to current wide-body for the TAC/Energy occur if the TAC-E and to be equal. Significantly greater savings delay is considered saved due to minimal delay. Figure 117 concepts are given credit for fuel TAC/Energy delay.

shows relative fuel usage for both standard and expected accomplished by the TAC/Energy Figure 117 indicates the savings in fuel consumption aircraft. Fuel burn by a TAC/Energy airplane when compared with current wide-body by the CWB-E, even airplane on an 1852-km (1000-nmi) trip is only 70% of that consumed If compared to the CWBLR the when delay is assumed to be the same for both aircraft.

C , airplane. When the TAC/Energy airplane TAC/Energy airplane uses 68% as much fuel as this for fleets of TAC-equipped airplanes, the fuel is credited with the delay reduction estimated with its expected 36-min average delay. If burned drops to 64% of that burned by CWB-E , the fuel burn of the TAC/Energy airplane is made with the un-resized CWBLR comparison C and lower fuel consumption are considered.

is 62% of CWBLR when both delay reduction C CWB-E costs

EXPECTED DELAY Jc

STANDARD DELAY 0- Fuel TOC Fuel TOC -100 , savings savings - - 300 L Relative cost of TAC/Energy - 400 variable (dollars) -500 - Fuel price = 5.8 /V (22d/gal) Fuel price = 7.9d/ (30d/gal) - 600 - 700 -800 Figure 116.-Economic Advantage of Fuel Reduction and Delay Reduction- TA C/Energy Airplane STANDARD DELAY EXPECTED DELAY 1.10

.....-- -

------------- CWBLR C 1.00 I I I CWB-E TAC-E TAC-E ATT-E ATT-E .90 - 32% 38% Relative fuel consumption .80 - per passenger km, .70 - T Energy TAC/ .60 Energy .50 - Noise abatement penalty 117.-Relative Fuel Consumption Per Passenger Kilometer Figure 55% Load Factor] (1000-nmi) Stage Length, [1852-km 6.2.3.5 Comparative Fuel Burn on a System Basis Comparisons of fuel consumption and block times were made using the current United Air Lines DC-10 route structure as a basis for comparison. Actual July 1974 schedules were used to determine the block times. Allowance for delay varies as a function of place and time of day. These block times were used for the system labeled as CWB-E.

Block times for the comparative aircraft (ATT-E, TAC-E, and TAC/Energy) were derived from computer computations. Expected delay, 36 min, was calculated for ATT-E since this aircraft has no delay-reducing features. TAC-E and TAC/Energy airplanes are given credit for reducing delay down to the standard 6 min of air maneuver. Therefore, block time for these airplanes was calculated with the expected lower level of delay. Calculations of block time for all of the comparative airplanes used the system average for delay and applied it equally at all airports and at all times of day.

Block times were not significantly different between the slower cruise velocity TAC/Energy airplane and current schedule times. An exception was on nonstop transcontinental flights where time differences in excess of 30 min are observed in spite of allowance for reduced delay.

In summary, reduction in delay allows the TAC/Energy airplane to operate at lower block times on ranges up to approximately 1852 km (1000 nmi). Block times are approximately equal on ranges between 1852 and 3700 km (1000 and 2000 nmi) and longer on ranges in excess of 3700 km (2000 nmi).

Fuel burn savings for the TAC/Energy airplane are impressive. (See fig. 118.) Fuel savings of 272 000 kg (600 000 lb) per day may be realized for the 25-airplane portion of a major airlines total fleet, even if no credit is given for delay reduction. When calculated on an expected delay basis, savings mount to 386 000 kg (850 000 lb) per day on this system. The TAC/Energy aircraft concept serving this system, with minimal delay, would use only 70% of the fuel current wide-body aircraft use in flying the same routes. This system total varies from the per passenger total shown in figure 117, because the routes covered vary from a short 254-km (137-nmi) route from Portland, Oregon, to Seattle, Washington, to the longest from Washington, D.C., to Los Angeles, California (4890 km, 2639 nmi). Savings vary from essentially zero to 8000 to 9000 kg (17 000 to 18 000 lb) per trip.

STANDARD DELAY EXPECTED DELAY 1.10 CWB-E fleet fuel burn CWB-E fleet 1.00 ATT-E TAC-E ATT-E fuel burn TAC/ TAC- E TAC/ Energy Energy .90 - Relative fuel burn; fleets of aircraft (%) .80 .70 ure 118.-Comparative Fuel Burn on a Representative Route System.60 Figure 1 .- Comparative Fuel Burn on a Representative Route System

7.0 CONCLUDING REMARKS

The TAC study (ref. 2) forecast expected unconstrained air traffic growth and the potential impact on airplane and airport compatibility to the year 2000. The emphasis of that study was on reducing congestion, noise, and emissions. Technical and economic assessments were made of the design modifications that were proposed in that part of the study. This present TAC/Energy study was limited to airplanes with four engines on the wing. With reducing fuel consumption as the major objective, but with attention to the concerns of the TAC study, the following conclusions can be made relative to a 5556-km (3000-nmi) range, 18 140-kg (40 000-lb) payload transport: I. If all of the calculated potential can be realized in an acceptable design, an advanced technology aircraft designed for fuel conservation and terminal area compatibility may achieve a 30% reduction in fuel use compared to a current technology wide-body sized and operated at long-range cruise. It is estimated that this could erode to concept as low as 23%, as the design proceeds to production status acceptable to operating airlines. To achieve this reduction, it will be necessary (1) for the aircraft to contain a composites in the structure; (2) for new substantial percentage of lightweight advanced wings to be developed with aerodynamics based upon the most 'advanced airfoil data; (3) for the airplane to be designed with full-time stability augmentation; (4) for a new advanced, quiet and efficient turbofan engine to be developed; and (5) for the airplane to be designed to a lower cruise speed.

2. If flight delays could be reduced, the fuel reduction would increase to 36%. This would require an efficient, practical, and safe solution to the problem of trailing vorticity, assuming that the solution would not have any adverse effect on fuel consumption.

3. Of items 1 and 2, the largest single improvement (approximately two-thirds of the total) hinges upon advanced airfoils. The best way to exploit advanced airfoils for of moderate sweep and lower fuel consumption is to apply them to thin (8%) wings rather high aspect ratio. Such wings would tend to operate at higher altitudes and significantly higher lift coefficients than current wings. This aspect needs special verification with wind tunnel data and perhaps also full-scale flight data.

4. The best engine for fuel consumption is a moderate bypass ratio conventional turbofan with BPR of about 6. Noise treatment should be limited to only peripheral lining and all inlet obstructions such as rings or splitters should be avoided as extremely costly in fuel use. Advanced turbofan engines with such features as very high OPR or variable well enough understood at this time to be pitch or camber fan blades are not recommended. The application of regenerative engine cycles to aircraft propulsion still requires the invention or development of a very lightweight efficient heat exchanger.

5. Today's turboprop engines will not yield an airplane with low fuel consumption. To achieve a fuel consumption reduction with a turboprop engine, a new lightweight and efficient design will have to be developed and the aircraft speed reduced to Mach 0.6.

Even though fuel reduction then occurs, the economic penalties (DOC) may be significant.

that the can be calculated if it is assumed 6. A very large reduction in fuel consumption at near seats or more) operated in very large airplanes (500 public will accept flying capacity levels.

8.0 RESEARCH AND TECHNOLOGY RECOMMENDATIONS

A continuous evaluation was made of the technologies involved throughout the study, particularly during the definition of the candidate TAC/Energy concept. This allowed identification of the advancements that would be required to make similar fuel conservative concepts available in the future. The TAC/Energy concept was defined on the basis of an assumed 1985 operational introduction date.

This was done by applying technologies that, with adequate R&T, were believed to be achievable by that time, while recognizing that even further identified advancements would also be beneficial during a later operational period.

The candidate TAC/Energy concept definition relied on two categories of potential technology advancement: (1) those defined during the previous ATT and TAC studies and (2) the additional advancements required to support unique design approaches to fuel conservation, the focus of this study. The contributions of the first category (ATT and TAC) were significant. Assuming the fuel conservation objectives of this study are important to the future of United States aviation, then the results of the study provide a basis that is even more significant than previously established for pursuit of recommendations made during the ATT and TAC system studies. Table 22 lists the individual R&T program recommendations that resulted from this study.

Table 22.- TA C/Energy Recommended R& T Program Subjects * Aerodynamics Technology-Wing design and drag prediction * Propulsion/Noise Technology * Systems Technology-Reduced inflow dabin air-conditioning system -Energy conversion techniques * Operational Techniques and Requirements-Airplane/ATC interface -Fuel reserves -Short range aircraft The following paragraphs provide summaries of applicable technologies and the recom- mended advancement R&T programs significant to fuel conservation, which are related to the ATT and TAC system studies, as well as those unique to this study.

8.1 ATT STUDY RELATED RECOMMENDATIONS The basic ATT system studies were directed primarily toward the cruise portion of the flight profile of long-range transports and to noise reduction in the terminal area. The high subsonic/transonic speed regime was emphasized. It was found that weight reduction and aerodynamic/propulsion efficiency, together with noise reduction techniques, were impor- tant to meet study objectives. The same fundamental needs are important to the TAC/Energy concept, although at a lower cruise speed.

The R&T recommendations that resulted from the ATT studies are listed in table 23 and are detailed in reference 8. Table 23 is coded to indicate the individual ATT technology advancement recommendation items that, if pursued, would also benefit fuel conservation.

.

Table 23.--Principal.R&T Recommended Work Packages Resulting Fromtb e ATTStudV Ie1-9) - Total

4 16 I 18 10 (Dollars in

1 2 1

1millions) Design integration Configurations 1 .. 75 configuration 1.75 * -Low noise/congestion 2.0 2.0 2.0 2.0 2.0 2.0 1.0 1.0 10.0 R&D configuration integration 1.6 * Maintenance and delay reduction 0.1 1.5 17.35 3.85 3.5 2.0 2.0 2.0 2.0 1.0 1.0 Subtotal Aerodynamic configuration Exploratory programs 1.2 1.2 1.2 1.2 1.2 1.2 1.2 1.2 9.6 * Advanced airfoils 1.5 High critical Mach cowls 0.8 0.7 area ruling 0.5 0.7 1.2 Subsonic/transonic * Analytical methods improved design and 0.3 0.3 2.4 analysis tools 0.6 0.6 0.6 3.6 18.0 Novel design concepts 2.4 3.6 2.4 3.6 2.4 programs Verification 1.0 2.6 "Interference free" transonic test section 1.2 0.4 1.8 * Flight/wind tunnel correlation 1.2 0.6 0.6 * Roughness and excrescence drag 0.2 0.4 0.5 Sonic boom during cruise 0.4 0.1 Slightly below Mach 1.0 8.5 8.3 5.2 5.1 3.9 4.8 1.2 1.2 38.2 Subtotal Structures and materials processes Materials * Advanced filamentary composite 10.12 5.02 0.92 0.33 168.49 structure 4.15 13.92 30.56 43.94 37.24 22.29 6.9 3.0 18.9 * Bonded aluminum structure' 0.8 8.2 0.09 * Improved corrosion protection 0.08 0.01 * Titanium structure steel structure * Improved * Load alleviation 6.9 control system 1.90 2.0 2.73 1.25 WIntegrated active capability 2.0 2.0 1.0 0.50 5.5 *Improved analysis and design 7.93 26.13 41.21 48.69 37.24 22.29 10.12 5.02 0.92 0.33 199.88 Subtotal * These items also offer fuel conservation potential.

Table 23.-(Continued) Total 7 8 9 10 (dollarsin 2 3 5 6

I I I I 6 1 millions)

Power systems * Propulsion system demonstration 3.0 2.0 3.0 32.0 45.0 50.0 135.0 * Cycle selection Engine cycle impact' 0.25 0.25 0.12 0.12 0.12 0.12 0.12 0.12 0.12 Turbomachinery noise 3.0 3.0 3.0 3.0 3.0 3.0 3.0 1.0 1.0 50.0 Jet noise 3.0 3.0 3.0 3.0 3.0 3.0 3.0 1.0 1.0 0.28 0.28 0.28 0.28 0.32 0.32 Improved propulsion controls 0.12 0.12 0.20 * Optimized nacelle 1.25 1.25 1.25 Inlet development 0.68 0.68 0.80 0.80 1.25 1.25 Nacelle configuration development 0.36 0.36 0.80 0.80 1.30 1.25 1.35 1.25 1.25 Nozzle development 0.06 0.06 0.40 0.40 0.48 0.48 0.88 0.88 Thrust reverser development 0.32 0.32 0.40 0.40 0.44 0.44 0.80 0.80 2.0 2.0 2.0 2.0 2.0 2.0 2.0 2.0 2.0 Acoustic lining development 0.50 0.50 0.50 0.50 0.50 0.50 0.50 .Optimized nacelle development 0.10 * Nacelle/airplane integration Nacelle/airplane development 0.28 0.56 0.56 5.0 Optimized nacelle development 0.22 0.56 0.56 0.66 0.60 0.50 0.50 1 Community ramp and interior noise 3.0 3.0 3.0 3.0 3.0 3.0 3.0 1.0 1.0 5.0 23.0 Auxiliary system 1.72 2.60 3.35 2.66 1.85 0.58 Hydraulic systems protective * Pneumatic, conditioning, and 0.39 0.27 0.20 0.21 0.12 systems and distribution 0.60 1.60 1.40 0.20 Electrical power * Powered wheel 0.20 1.20 0.60 35.77 0.60 0.76 0.80 1.69 0.54 * Landing gear

1.50 0.25 I

Integrated engine generator 0.38 0.90 1.50 1.75 2.00 0.42 0.17 Fuel systems 0.73 0.47 0.25 0.52 0.58 0.15 * Carbon brake Subtotal 21.35 28.73 27.28 53.79 65.48 67.95 16.78 10.12 7.94 0.16 295.77 Table 23.-(Concluded) Total in (dollars 1 2 3 1 4 5 6 7 8 9 10 millions) Control of flight

Flight controls I I I

* Application of load-alleviating controls See structures and materials integrated active control system.

to configuration optimization * Application to CCV philosophy 0.25 0.25 Transonic trailing-edge controls development 0.20 0.05 0.25 Development of improved atmospheric models for mesoscale turbulence discrete gusts 0.20 0.20 * Flightpath control system development 0.10 0.10 0.20 0 Improved aeroelastic methods for flight controls analysis 0.10 0.10 * Avionics Advanced navigation systems development 1.60 4.00 5.60 Advanced digital systems application 1.00 1.10 2.10 Advanced flight deck development 0.75 1.15 1.70 3.10 0.10 0.10 6.90 Low-cost inertial sensors/systems 0.20 1.50 0.75 2.45 ATC/operations Airplane requirements 0.93 0.93 Subtotal 5.33 7.90 2.45 3.10 0.10 0.10 18.98 Total 46.96 171.56 178.141 111.891108.65 97.14 1 29.15 117.341 8.86 0.49 570.18 I Key items that were relied on during definition of the TAC/Energy concept are described in the following paragraphs under headings that reflect the thrust of their impact on fuel conservative aircraft.

8.1.1 AERODYNAMIC EFFICIENCY Under "aerodynamic configuration," table 23 lists several exploratory and verification-type programs that are directed toward improved aerodynamic performance. When modified to the speed regime of this study, a majority of those programs would benefit future fuel conservative aircraft designs. This was borne out by the results of the wing planform sensitivity study when it was found that aerodynamic data on improved airfoil/wing designs were badly needed. Specific recommendations on this subject are covered insection 8.3.

8.1.2 PROPULSION EFFICIENCY AND NOISE REDUCTION The ATT study identified several propulsion design areas (listed under the "power systems" heading in table 23) that were postured toward reduced SFC, noise, and nacelle drag.

Similar improvements are needed for fuel conservation for engines at BPR's of 6 to 7, rather than at 4 for the ATT. The objective of noise reduction is somewhat in conflict with fuel conservation objectives because of inherent weight and complexity increase associated with noise reduction techniques. However, if low noise objectives are also important for fuel conservative designs, then much R&T is needed to minimize the impact.

8.1.3 WEIGHT REDUCTION Significant weight reduction potential was identified in two major areas during the ATT system studies.

The first study, advanced structures (including composite materials, bonded aluminum, and improved steel alloy), offers from 10% to 25% reduction in the structural weight of future aircraft. The TAC/Energy concept is based on achieving the 10% reduction by a 1985 operation date. In addition, composite structure is particularly adaptable to the high aspect ratio wing type of design used on the TAC/Energy concept. Its use reduced the wing flutter penalty that was found to exist on the aluminum structure version. An additional 3% reduction in fuel use could be attained for a later airplane when technology is available to support the 25% structural weight reduction value. It is recommended that the NASA structures program that is now underway be continued and expanded to cover all aspects of the program during the ATT program (ref. 8).

The second study, advancement of active controls technology, was directed toward proving the avionics, hydraulics, and overall controls systems associated with stability augmentation systems as well as load and flutter alleviation. The stability augmentation system will allow flight under aft stability loading conditions and will provide the reduced drag that was assumed on the TAC/Energy concept. The high AR wing included on concepts similar to the TAC/Energy will be subjected to high wing root bending moments and will be more susceptible to flutter. Such designs may benefit from load and flutter alleviation type, active controls.

NASA is encouraged to continue and expand the current active controls program, now underway, in a manner that will yield the technology confidence required by the airline adoption.

and airframe industry to permit its 8.2 TAC STUDY RELATED RECOMMENDATIONS The TAC system study was directed toward the reduction of congestion, noise, and emissions in the terminal area. The TAC/Energy concept incorporates most of the features defined by that study. This was accomplished with some negative impact on weight.

However, this study confirmed the conclusions revealed in the TAC study that, on an overall basis, many of the features actually contributed to fuel conservation when they were included in the basic design. This was particularly true for the high AR wing and the powered wheel which were included to meet noise and emission reduction objectives, respectively. All facets of the TAC concept that were related to the reduction of congestion will be vital to fuel conservation as traffic increases in the future. The objective, of course, is to hold operational delay to a minimum which in turn will minimize fuel consumption, by decreased holding and increased runway acceptance rates. Needed key features include a high-energy brake system, high-speed turnoff landing gear, and the advanced avionics required for improving the airplane/ATC control loop to permit low-tolerance position accuracy, touchdown dispersion, and, of major significance, the reduction of separation distance between parallel runways.

The subject of wake vortex dissipation is of particular importance to congestion reduction.

Based on broad theoretical models, vortex dissipation methods have been proposed, and promising qualitative test results have been obtained. Although the system impact of some of these vortex dissipation methods may impose large penalties to economics and energy usage, the first step is nevertheless to find workable solutions to the problem. To date, some of the most promising results were obtained from the following: * Differential flap for specific lift distributions * Differential engine thrust * Splines * Wing fences As was the case in the ATT studies in regard to noise reduction, the TAC study provided recommendations for R&T needed to reduce emissions through improved combustion and burners and to reduce noise through use of steep approach glidepaths. Again the objectives were at cross purposes with fuel conservation objectives because of weight impact. In a similar manner, however, if those objectives are realistic, they must be attained at minimum technical and economic cost. The R&T programs recommended from the TAC study were directed toward that end.

In summary, essentially all of the R&T recommendations resulting from the TAC study are of great importance to future transport designs that would be similar to the candidate TAC/Energy concept. A summary list of those recommendations is shown in table 24. Key items that are of particular importance to fuel reduction for the candidate concept, either Table 24.-Principal R& T Recommended Work Packages Resulting From the TAC Study (Ref. 3) Item Approach Ground operations Takeoff Cruise * 1. Wing trailing vortex research. 16. Engine emission control 038. Reduced penalty high-lift and low- 41. Low-cost inertial systems * 2. Airframe noise research and control 017. Auxiliary power unit development noise concepts 42. VLF navigation 3. Aerodynamic drag device development 18. Tire technology 039. Advanced low-noise engine cycles 43. Impact of cruise Mach constraint 40. Water injection systems 044. Structural implications of outboard 4. In-flight thrust reversers 19. Landing gear/brake dynamics engine location 20. Automated stopping system 5. Landing guidance 45. Weight implications of high aspect 21. Brake material and configuration 0 6. Improved anti-icing techniques * 7. Impact of turbulence and wind shear on 22. Ground steering systems 046 Guidance and controlintegration airplane design 23. Reduced gate space design techniques 47. Secondary power system redundancy 047 Secondary power system redundancy and landing study 24. Airframe and engine maintainablility Airplane 0 8. Noise-abatement approach 048. Airplane design/energy relationship N) 9. Stopping system trades 025. Large payload airplane scheduling/marketing 049. Iirplane concepts for TAC n ativ noise abatement assessment4 0 10. Noise treatment versus procedures 26. Aircraft maintenance monitor 50. Advanced fuel airplane concepts liq H 011. Engine component noise research 027. Powered-wheel systemfuel airplane concepts liq H 012. Engine noise reduction with lining/treatment 029. Powered-wheel operational assessment Ground 030. Surface traffic control compatibility 13. Steep approach operational facilitiers tolerances 31. Passenger Passenger 14. Passenger compartment and flight deck noise 32. Schedule spreading incentives 51. Cabin and interior noise effect of MCR Aircraft interior improvements 33.

0 34. Ambient air quality standards review and assessment 35. Terminal area air polution model Community 15. Post-1985 noise criteria 36. Terminal area meteorology model 36. Terminal area meteorology model Pollution containment attitude 37.

also offer fuel conservation potential.

*These items li-4 GO directly or indirectly by minimizing the impact of noise or emission reduction, are noted by symbol in the figure. The programs defined during the TAC study for each item are essentially adequate from the standpoint of emphasizing fuel reduction. The current NASA Terminal Configured Vehicle program is headed in the needed direction and should be continued and expanded. NASA is encouraged to pursue the previously defined recommen- dations for each of the items noted in table 24.

8.3 TAC/ENERGY STUDY RECOMMENDATIONS With energy conservation as the driving force, this study provided the framework for evaluating additional technical and operational features, as well as extensions of previously considered features. The attractive additions and extensions were found to focus primarily on the aerodynamics, propulsion, and systems portion of aircraft design as well as potential revised operational procedures that could affect the aircraft/ATC interface. Although the quantity of new technology advancement items was not large (as compared to the ATT and TAC studies), they will be vitally important to future aircraft that are designed to objectives similar to those of the candidate concept. The following paragraphs identify those items under headings that reflect the technology involved.

8.3.1 AERODYNAMICS TECHNOLOGY-WING DESIGN AND DRAG PREDICTION 8.3.1.1 Potential Payoff Airplanes developed for low fuel consumption will tend to have lower cruise speeds and wings of higher aspect ratio and less sweep than those directed toward best economics.

These characteristics lead to cruise at higher lift coefficients than current subsonic jets.

Research is needed to develop the most efficient aerodynamic shapes for such wings and to establish a data base for use in further preliminary design studies.

8.3.1.2 State of Readiness The necessary design methods and test facilities are available.

8.3.1.3 Recommended Action A three-phase program is recommended: * Phase I. Two-Dimensional Airfoils.-A family of airfoils should be developed by two-dimensional transonic analysis and wind tunnel testing. A set of nine airfoils is suggested, covering the range of thickness and design lift coefficients shown in table 25.

* Phase II. Three-Dimensional Wings.-A family of wings should be designed, using the airfoils developed in Phase I. The characteristics of these wings should be measured by wind tunnel testing at both on-design and off-design conditions. A recommended set of 16 wings is shown in figure 119.

* Phase Ill. Parametric Design Chart.-A set of design charts should be developed from the wind tunnel data. These charts would be used during preliminary design to select airfoil sections and to predict wing drag characterisitics.

Table 25.- Two-Dimensional Airfoil Matrix t/c, M CQ n desig percent 0.45 0.55 0.6 7 X X X 9 X X X 11 X X X 8.3.1.4 Cost and Schedule The 18-month program shown in figure 120 is estimated to cost $0.5 million.

8.3.2 PROPULSION/NOISE TECHNOLOGY 8.3.2.1 Potential Payoff A conventional turbofan with 1980 inservice technology does not have any particular advanced technology requirements. This fuel conservation study indicated that a BPR turbofan provides close to optimum performance from the standpoint of fuel utilization and operating costs. Although no new engine technology requirements for 1980 inservice operation are needed, it is important during engine and airplane installation development to preserve high engine reliability and maintainability.

In the brief study of advanced engine cycles in this program, a single engine was not identified that was significantly superior to the conventional turbofan in fuel utilization with acceptable economics. However, high potential for energy conservation was indicated in some of the advanced cycles studied. Many additional exploratory design studies and analyses, coupled with application of advanced technology, are required to fully explore, evaluate, and validate advanced low-energy engine concepts.

8.3.2.2 State of Readiness Since a leading candidate low-energy cycle for the post-1980 time period has not been identified, specific recommendations for technology cannot be provided. However, the following advanced technology items would improve all engines including the turbofan: * Higher pressure ratio per stage capability at good efficiency levels, for fans, compressors, and turbines * Improved turbine airfoil, endwall, and secondary cooling capability root treatment * Wing twist and methods 3-D subsonic designed using at one or two stations Surface pressures * flow visualization * Fluorescent oil 2-D studies Airfoil section from o model Sting-mounted force * TEST PLAN A = 300 A = 25 o ' . 2-D design CV (camber) Design C and t/c 1 2 3 1 2 3 from 2-D studies 0.25 9 Priority 0.5 1.0 0.25 Wing Design and Test Figure 1 19.- Three-Dimensional Complete 2-D Complete wing Complete report design and test design and test and design charts Months Figure 120.- Three-Dimensional Wing Design and Test Schedule * Improved compressor, burner, and turbine durability to minimize performance degradation and maintenance costs * Improved clearance controls throughout the engine, to improve component perform- ance and reduce leakages and secondary flows * Improved materials such as composite fan blades, high-temperature titanium for compressor disks, high-strength nickel base alloys for turbine disks, and high- temperature turbine airfoil materials and coatings * Shorter burners with improved performance and low emissions technology * Fabrication technology, including reduced-cost turbine airfoils, small cooling hole technology, near-net and net-shape disks, large cast cases, lightweight structures, and fabricated sheet metal cases * Improved bearings; improved seals including abradable/abrasive seal technology on airfoil/cases for tighter clearance control e Improved accessories, including an electronic control to improve performance, hot section life, control maintenance, and reliability 8.3.2.2 Recommended Action It is recommended that engine/airframe studies be conducted that are similar to those conducted in this study but expanded to permit indepth evaluation of potential alternate approaches to given objectives, considering alternate engine cycles. Engine and airframe manufacturers should participate in the study, and both participants should be involved in establishing objectives and evaluation criteria. The airframe organization should participate to provide advisory and evaluation assistance during cycle and engine evaluation by the engine organization.

Overall integrated evaluation should be conducted by the airframe organization.

One or more contracts funded to a minimum of $500 000 and involving engine and airframe manufacturers is recommended.

8.3.3 SYSTEMS TECHNOLOGY 8.3.3.1 Reduced Inflow Cabin Air Conditioning System Potential Payoff -The airplane secondary power systems (SPS) extract up to nearly 5% of the usable energy from the airplane engine, resulting in significant fuel consumption. Of the SPS, the airplane pneumatic system is the greatest user of this energy which, during cruise, uses engine bleed air extraction to air-condition and pressurize the airplane cabin. The potential exists to reduce by at least 50% the quantity of bleed air required, with an associated reduction in block fuel of approximately 3%.

State of Readiness.-A -3 3 fresh air ventilation standard of 9.44 x 10 m /sec (20 cfm) has been established for commercial jet airplanes. Under high load density conditions, this rate can be reduced somewhat, provided that proper air movement is attained. To achieve required air movement on some airplanes, unfiltered (not revitalized) recirculated cabin air is used.

As the amount of fresh air is reduced and cabin recirculated air is increased, odor carryover, stuffiness, and general discomfort result.

If a high degree of recirculated cabin air is to be used, it is necessary to establish guidelines as to the allowable quantities that can be recirculated without cabin air pressurization problems and the degree of filtration required to make the air physiologically acceptable.

Recommended Action.-A program to establish suitable cabin air recirculation levels is recommended.

The program would consist of: I. The analysis and testing necessary to establish recirculation air rate and filtration levels to be used on a suitable inservice airplane 2. The design of a filtration kit 3. Inservice testing of the kit to verify results of item 1 A 707 vapor cycle equipped airplane would be typical of a suitable airplane as it has a built-in recirculation system.

The recirculated air is not now revitalized. Filtration would have to be added and necessary ducting changes made to allow variations in both filtration capability and recirculated airflow quantity.

Cost and Schedule.-The 18- to 24-month program (fig. 121) would cost approximately $0.15 million, not including airplane modification.

WORK STATEMENT * Determine criteria.

* Design filtration kit and install on airplane.

* Inservice flight test.

PAYOFF power extraction * Fuel savings from reduced engine SCHEDULE Complete Flight criteria test Test analyses filtration complete Document Design filtration kit 0 3 6 9 12 15 18 24 Months Figure 121.-Reduced Energy Cabin Air-Conditioning System Research and Technology Summary 8.3.3.2 Energy Conversion Techniques Payoff-To accomplish the required fuel conservation required on future Potential to review all energy-consuming systems and determine the most airplanes, it is necessary to convert fuel energy to the required airplane function. The SPS directly efficient means of the total airplane fuel energy to power hydraulic, pneumatic, and electric uses about 5% and about twice that again to carry the systems weights. Considerable fuel savings systems, reduce the could result if improved energy conversion techniques could be found that would energy use through higher efficiency of conversion.

total been conducted under NASA State of Readiness.-Secondary power systems studies have studies established several contract (ref. 9) and in conjunction with the TAC program. These potential fuel savings. They include use of dedicated APU's, engine shaft power in areas of bleed, integrated engine generators, and combination generators/starters.

preference to of these studies is recommended with emphasis on power conversion techniques Extension a minimal fuel and integration of the various power sources and functions to produce consumption airplane. In particular, an independent power generation system should be explored to determine potential fuel savings benefits.

involving the integration of the airplane SPS energy Recommended Action.-A program the methods of conversion techniques is recommended. The program would explore and electric) through converting fuel energy to power for the SPS (hydraulic, pneumatic, other engine bleed and shaft power extraction, separate power generation units, and normal methods determined from the study.

recommended.

Cost and Schedule.-A 12-month study funded at $250 000 is AND REQUIREMENTS 8.3.4 OPERATIONAL TECHNIQUES A limited analysis of the flight profile conducted during the sensitivity part of the study techniques that reduce fuel consump- resulted in the identification of several operational concept; in fact, portions are being tion. Some were applied directly to the TAC/Energy reduced climb speed, flight at long-range applied to current flight operation. These include and step climb during cruise. Other cruise (LRC) rather than minimum DOC cruise speed, possibilities were identified but could not be applied because of regulatory constraints. The if following paragraphs provide recommendations relative to these possibilities which, NASA/DOT/FAA and/or the followed, may involve coordinated or joint action by airline/airframe industries.

Interface 8.3.4.1 Airplane/ATC by airframe manufacturers has Potential Payoff-Action by the airlines with assistance demonstrated substantial fuel savings by application of revised operational procedures, within current regulating constraints and with no changes to the aircraft or the ATC system.

altitude change as a function of gross weight and Two areas were also identified-optimized to accommodate aircraft with different optimum fuel-use dispersed lateral spacing constraints would not allow the 2% to 3% block fuel altitude/speed characteristics. Current of these techniques to be included for the candidate TAC/Energy reduction potential concept. Other possibilities may exist.

State of Readiness. -The necessary definition of the current and planned future ATC system the necessary evaluation, with fuel and potential aircraft capabilities are available to perform conservation as a prime driving force.

is recommended that an initial study be conducted considering Recommended Action.-It of aircraft/ATC interaction over the complete trip profile. This should include the all facets of fleet operations, considering different aircraft designed for specific stage implication lengths and planned changes to the ATC system. The study could be conducted by an airframe manufacturer in cooperation with airlines and Government agencies, under DOT or least the following activities: NASA sponsorship. The study should include at airlines, and DOT and NASA fuel conservation studies for potential 1. Survey of industry, fuel-saving operational techniques 2. Use, as appropriate, of models to determine fuel-saving payoffs of candidate techniques as functions of ATC and airplane system parameters 3. Identification of promising operational techniques 4.

Identification and recommendation of aircraft and ATC system technology advances necessary for the best use of fuel-saving operational techniques Cost and Schedule. -A 12-month program estimated to cost $300 000 is recommended.

8.3.4.2 Fuel Reserves Potential Payoff.-During the sensitivity studies (sec. 4.2), it was determined that a significant amount of fuel is consumed because of the amount of reserve fuel required by current regulations. At the 5556-km (3000-nmi) design range, reserve fuel equivalent to approximately 25% of the total block fuel is required. To correlate the effect on fuel usage (although undoubtedly not practical), if the reserves could be reduced to 50% of current rule requirements, a saving of 6% total block fuel could be realized for the design range mission. For a 1296-km (700-nmi) stage length, where reserve fuel is equivalent to block fuel, a 7% fuel saving could be realized.

State of Readiness.-Necessary data on current and future ATC provisions, weather, airports and airways, and aircraft capabilities are available to determine the potential for possible revision to rules that control the amount of required fuel reserve.

Such rules have been in effect for several years and deserve reevaluation in light of the current and predicted technical and operational environments that govern their content.

Recommended Action.-It is recommended that a study be conducted involving an airframe manufacturer, airlines, and governmental agencies. The study should consider the following as a minimum: 1. Isolation of the factors that supported definition of the current rules 2. Correlation of technical and operational conditions that affect factors identified in item 1 initially, at the current time, and as projected in the future 3.

Determination if conditions that would affect reserve have or are projected to change 4. Determination of technical advancements that, if made available, would allow desirable rule changes Cost and Schedule.-A 6-month study funded at $450 000 is recommended.

8.3.4.3 Short Range Aircraft Potential Payoff. -Evaluation was made of transport aircraft fuel consumption as a function of payload capacity and stage length. This revealed that nearly 60% of the fuel is consumed by all types of aircraft operating over stage lengths of less than 2880 km (1500 nmi). Most 2 0 0 of those aircraft are of the less than 100- to -passenger-capacity type. The cruise speed of aircraft operating over 370- to 1110-km (200- to 600-nmi) stage lengths is from M = 0.1 to 0.15 less than for the design range considered by this study. The fuel used during the climb/descent portions of those shorter missions is, however, considerably greater.

Because of the altitude, speed, effect of reserves, climb, and other characteristics of the short-range aircraft, desirable fuel conservation features and the need for R&T will be different in several design areas, as compared to the longer design range concept considered in this study. In particular, at least the propulsion cycle and wing geometry of such aircraft are expected to be different. Weight is expected to be more important than aerodynamic efficiency. Because such short-range aircraft consume a large portion of the aviation fuel, their optimum design characteristics and need for R&T should be evaluated.

State of Readiness.-The techniques and background data needed to evaluate the special characteristics of short-range aircraft are available. Although a short-range study was sponsored by NASA/Ames (contract NAS2-6995), that study concentrated on a quite short field length aircraft.

Although possibly adaptable to future conditions, the short field lengths are not consistent with the probable majority of short-range fleet operations.

Recommended Action.-It is recommended that a 9-month study funded at $250 000 be issued with the objective of defining the unique technical and economic characteristics of short-range transports when operating in a fuel shortage and/or high-cost environment.

APPENDIX A

ENGINE SUBCONTRACTOR

REPORT

Portions of the complete final report submitted by Pratt & Whitney Aircraft in support of this study are presented. Those parts of the report considered most pertinent to the airplane studies using unconventional engines are included.

INTRODUCTION P&WA support of the Boeing effort included the following tasks: 1. Review the cycle and installation assumptions and the performance and weight trends of the Boeing baseline turbofan engines.

2. Suggest advanced engine concepts that could provide fuel conservation opportunities, and coordinate with Boeing the selection of three of these for further study.

3. Provide technical support to Boeing in simulating the performance characteristics of the selected advanced engines.

4. Estimate the performance, dimensions, weight, and price of each of the selected advanced engines.

ENGINE PERFORMANCE Baseline Cycles The baseline cycles used by Boeing in the study consisted of a family of conventional turbofans covering a range of bypass ratios, at one level of overall pressure ratio and turbine temperature. These cycles were used as a reference point in evaluating the merits of the advanced cycles.

Boeing provided P&WA with the performance of this family of engines along with the component assumptions that were used to calculate the performance. Based on the cycle definitions and component assumptions used by Boeing, the performance of the baseline cycles could be very closely approximated by P&WA.

The small difference in level (less than 1%), which was later resolved to be due to a difference in the method of handling turbine cooling air, was considered not important since the trends were consistent.

This comparison ensured that Boeing and P&WA were calculating cycle performance consistently.

Because of the possibility that the procedure used to define and rate engine cycles may be different than the method used by Boeing, P&WA established its own base turbofan cycle as a starting point from which to define the advanced cycles.

The performance increments for the advanced cycles relative to base cycles were established by P&WA, and Boeing applied these increments to their own base cycles to achieve the same relative performance change for advanced cycles.

The P&WA base cycle is defined as follows: FPR 1.6 BPR 6.1 OPR 25 Maximum TIT 1315 C (24000 F) is also shown: For comparison, the Boeing bypass 6 cycle FPR 1.66 BPR 6.0 OPR Maximum TIT 12820 C (23400 F) Turboprop The turboprop cycle selected for the study was based on the same component performance and technology level as the base turbofan. Since the base turbofan had a benefit in OPR from the fan root supercharging which was not available for the turboprop, the overall pressure ratio was reduced to 20 to 1 for the turboprop. This requires an increase in high compressor pressure ratio from 17 for the base turbofan to 20 for the turboprop, which is achieved by adding two stages to the rear of the compressor. The cycle definition of the turboprop is given in the following list: OPR 20 Maximum TIT 13150 C (24000 F) Prop characteristic Hamilton Standard PDB 6101, used for thrust Generalized Method of Propeller sizing Performance Estimation 140 AF/0. 15 CL. blades / 4 Diameter 5.9 m (19.4 ft) i Tip velocity 213 m/s (700 ft/sec) Performance for the turboprop cycle is provided in table A-1.

High Overall Pressure Ratio Turbofan Before selecting a level of pressure ratio for a high overall pressure turbofan, a trend study was completed which showed the benefits to be gained for increases in overall pressure ratio.

The following levels of component technology were considered in making the study: 1. Component Technology Consistent With the Base Turbofan: For this case the turbine cooling air was increased as overall pressure ratio increased in order to maintain a constant turbine airfoil metal temperature. High turbine efficiency was penalized for the increased cooling air.

2. Improved Turbine Cooling Technology: For this case the level of turbine cooling air was held constant at the base turbofan level as OPR increased. This case would be Table A-1. -Summary of Engine Cycles Variable Base High OPR pitch fan Cycle definition turbofan Turbopropa turbofan (variable AJD) FPR 1.6 - 1.6 1.4 BPR 6.1 - 6.4 8.4 OPR 25 20 40 25 Max TIT, K (0 F) 1 590 (2400) 1 590 (2 400) 1 590 (2400) 1 590 (2400) Performance SLS T/O FN, N (Ib) 63 000 (14 160) 77 300 (17 380) 63 200 (14 200) 66 100 (14 850) FN lapse to 0.3 M, % -29% -34% -29% -35% Max Cr., 30 000 ft 0.8 M FN, N (Ib) 16 500 (3 700) 11 800 (2 650) 16 500 (3 700) 16 500 (3 700) TSFC Base +11% -5.1% -2.1% 0.7 M FN 16 600 (3 730) 15 600 (3 500) 16 700 (3 760) 16 600 (3 730) TSFC Base -21% -5.2% -2.4% Design Point Airflows WAT , kg/sec (Ib/sec) 279 (615) 25.4 (56) 281 (619) 362 (798) Technology Components 1980 1980 1980 1980 Turbine cooling 1980 1980 1985-1990 1980 - 140AF/0.15C - - Propeller (0.7 M design) i a AF/0.30Ck propeller was used for 0.6 M cruise with performance 1 4 0 i as shown in section 4.4.3.6.

representative of either more effective turbine cooling, increased metal temperatures, or cooled turbine cooling air.

For this trend study the fan pressure ratio, high compressor polytropic efficiency, and nozzle jet velocity were held constant as the overall pressure ratio was increased.

This resulted in a decreasing bypass ratio as overall pressure ratio increased.

At the base turbofan level of technology it was found that the optimum overall pressure ratio is about 33 to 1 and the TSFC improvement is only 1.3%. At higher pressure ratios, the increase in turbine cooling and high turbine efficiency penalty due to more cooling offset the pressure ratio effect and TSFC gets poorer. However, the case where turbine cooling air flow was held constant at the base turbofan level shows TSFC improving by about 3% up to pressure ratios of 40 to 1 and above.

Based on the results of this study it was decided that increasing OPR alone, without a turbine cooling technology improvement, did not offer a significant enough improvement in fuel consumption. For this reason, advanced turbine cooling technology was assumed to provide a 50% reduction in cooling flow at 40 to 1 pressure ratio relative to the base 25 to 1 engine. This would require extensive advances in materials, cooling configurations, and/or cooled cooling air. The reduction in turbine cooling air results in an increase in BPR relative to the base cycle.

Variable-Pitch Fan Turbofan A fan pressure ratio of 1.4 at the aero design point was selected for the VPF turbofan as a reasonable compromise between: * Higher FPR, which the baseline studies indicate would provide better. airplane performance * Lower FPR, which previous studies have indicated would show more operational advantage for the variable-pitch feature Engines with low fan pressure ratios have a fan stability problem at sea level static since the fan operating line moves up relative to altitude cruise, because of the unchoking effect of the fan nozzle. In fixed-pitch fans, this effect can be controlled by opening up the fan duct nozzle to lower the fan operating line. This requires either a two-position nozzle, which results in an installation penalty, or a permanent lowering of the cruise operating line, which results in a performance penalty.

The variable-pitch fan offers the possibility of providing increased sea level stability through a fan pitch change. However, evaluation of the 1.4 pressure ratio fan indicates that the variable-pitch capability does not provide enough surge margin control at sea level to completely eliminate the need for one of the other methods of operating line control. For this study a two-position nozzle was selected for takeoff and cruise operation.

A third position of the nozzle would be required for reverse thrust.

Performance of the variable-pitch turbofan is shown in table A-1.

ENGINE INSTALLATION This section summarizes the weight information provided for the base turbofan and the advanced engines.

WEIGHT Baseline Turbofan.-The recommended specific weight (weight/airflow) band versus bypass ratio, shown in figure A-1, assumes that the engines will be in service by 1980.

6.0 5.5 Max. cruise thrust 34 700 N (7800 Ib) 9144 m (30 000 ft), 0.80 M 5.0 4.5 0J, 4.0 2 4 6 8 10 12 Bypass ratio Figure A- 1.-Specific Weight of Turbofan Engines Turbofan Weight Scaling With Airflow.-The turbofan bare engine weight scaling curve of figure A-2 is independent of bypass ratio and overall pressure ratio.

Ref FN T/O = 133 000 N (30 000 Ib) 1.6 Bare engine weight 1.4 1.2 M.

1.0 C" Thrust-to-weight .8 -' .6 .4 .6 .8 1.0 1.2 1.4 Relative thrust size Figure A-2. - Weight/Thrust Scaling Turbofan Engines Turbofan Weight Variations With Overall Pressure Ratio.-For constant technology, the specific weight is essentially constant for a given bypass ratio engine for overall pressure ratios of 20 to I to 40 to 1. This conclusion is based on results of previous preliminary concept studies.

Turboprop Turboprop Versus Turbofan Weight. -The base turbofan's weight includes the fan, cases, gas generator, controls, and accessories, but no installation (i.e., cowl, nozzles, etc.). The turboprop's weight includes only the gas generator, propeller drive turbine, controls, and accessories. Propeller, gearbox, and installation are not included.

The turboprop has an overall pressure ratio of 20 to 1, as opposed to 25 to 1 OPR for the base turbofan. Since the technology levels are assumed constant for these two cycles, the turboprop must add two stages to the rear of the base turbofan's high compressor to partially offset the loss of fan supercharging.

The turboprop weight is estimated to be 1.25 times the. base turbofan's weight, assuming equal inlet "core" airflow sizes at 30 000 feet, Mach 0.7, maximum cruise.

Propeller and Gearbox Weight.-Propeller preliminary weight estimates were summarized as based on Hamilton Standard data: follows, 0.12 1.0)0.5 SHP/D 00.6 N .5 (M+ 0Weight = 80 I0 2000) ]Weight where D = diameter, feet B = number of blades AF = activity factor per blade N = propeller maximum speed, rpm M = flight Mach number at design condition SHP = design SLTO standard day power, horsepower The weight assumes fixed camber, lightweight fiberglass blades with an integral gearbox type hub and includes the spinner, control, de-icing and oil, but not the gearbox.

The gearbox weight estimates were also generalized on Hamilton Standard data, as Weight = 0.075 (Q)0.84 Q is the maximum continuous output torque in foot-pounds. The following where assumptions apply to this gearbox weight formula: * Two-stage gearbox with one planetary set and offset parallel drive * Reduction ratios from 10 to 1 to 14 to I cross-shaft drive, declutching, or special accessory drives 0 No provision included for High Overall Pressure Ratio Turbofan It was noted that, for constant technology, there is essentially no specific weight (weight/airflow) variation with overall pressure ratios ranging from 20 to 1 to 40 to 1.

However, Boeing requested information on a high OPR engine with a more advanced technology in the areas of either high-pressure turbine materials ahd coatings or cooling scheme effectiveness or both, relative to the base turbofan. This resulted, in the 40 to 1 overall pressure ratio turbofan with 1985-90 inservice cooling technology for which P&WA provided information to Boeing.

The advanced technology permitted a turbine cooling airflow reduction for the 40 to 1 OPR turbofan, relative to the base turbofan. This cooling airflow reduction resulted in an increase in turbine efficiency and, at constant combustor exit temperature, a decrease in the cooling air dilution effect on high-pressure turbine exit temperature. Bypass ratio had to be increased to offset the effects of the efficiency improvement and decrease in dilution effect.

The increase in BPR caused an increase in engine total airflow, resulting in a thrust increase at cruise. Therefore, the entire high OPR turbofan with advanced cooling technology was scaled down to maintain constant cruise thrust. This scaling, in effect a scaling down of the gas generator, resulted in a 1% weight reduction for the high overall pressure ratio turbofan relative to the base turbofan.

Variable Pitch The variable-pitch fan was assumed to use shroudless advanced composite fan blades rather than titanium blades, as in the shrouded fixed-pitch fan. The weight saving of the composite fan blades, plus the resulting saving in fan disk, shaft, and containment weight approximately offsets the weight increase due to the pitch change mechanism, lower aspect ratio, and higher hub-tip ratio of the variable-pitch fan. This results in the estimate of no weight difference between the fixed- and variable-pitch engines.

Summary The bare engine weight results are summarized in the following list. Each engine is sized to provide 16 500 N (3700 lb) of thrust at 9144 m (30 000 ft), Mach 0.8, maximum cruise (turboprop sized for 15 600 N (3500 lb) of thrust at 9144 m (30 000 ft), Mach 0.7, maximum cruise): Engine Relative Weight Base turbofan 1.0 0.81 Turboprop gas generator High OPR turbofan 0.99 VPF 1.27 ADVANCED TECHNOLOGY REQUIREMENTS Baseline The definition of the base turbofan (1980 inservice technology) excludes the need for any advanced technology requirement.

particular Turboprop Gas Generator.-The gas generator and propeller drive turbine are based on technology levels consistent with the base turbofan; therefore, no specific advanced technology requirements exist for the gas generator or propeller drive turbine.

Gear. -Advanced high-powered gearing concepts should be investigated, and specific tests on new technologies such as high contact ratio gears should be conducted.

System health diagnostics (e.g., multigraduated mesh metallic particle sensors, vibration/acoustic analyzers) need to be developed. The potential of advanced lube system components such as integral centrifugal air-oil separators and centrifugal self-cleaning filters should also be investigated.

Heat rejection and efficiency must be considered in the design of any advanced high-powered gears.

Propeller.-Advanced composite blade technology needs to be developed, and design investigations into advanced maintenance concepts should be conducted. In addition, the aircraft-engine-propeller integration has to be optimized.

High OPR Turbofan Further improvements in advanced technology are required for the high OPR engine relative to the lower OPR base engine because of the higher exit temperature, smaller airfoil sizes, and advanced cooling technology (i.e., 1985-90 inservice engines) of the 40 to 1 OPR turbofan.

These further improvements are: * Airfoil materials with higher metal temperature capability * Improved temperature airfoil coatings to go along with the higher airfoil metal temperature * Improved cooling system effectiveness * Further advances in clearance control, especially at the rear of the high compressor and the first stage of the high turbine where the airfoil spans are smaller than usual * Improved disk materials such as: (1) high-temperature titanium to minimize the use of more expensive nickel base alloy high-compressor disks and (2) high-strength, improved low-cycle fatigue (LCF) nickel base alloys to provide reduced weight and increased life of high turbine disk configurations * Reduced NOx emissions technology, since NO increases as pressure ratio x and compressor exit temperature are increased with today's technology * Improved sealing technology, due to the higher pressure differential across seals in the burner-high turbine area, to reduce leakages * Improved burner liner materials, to accommodate the higher burner inlet temperature without degrading burner durability or performance VPF Gas Generator.-Since the gas generator and fan drive turbine are based on technologies consistent with the base turbofan, there are no advanced technology requirements for the VPF's gas generator and fan drive turbine.

Variable-Pitch Fan.-Composite fan blade technology must be developed in areas such as improved foreign object damage (FOD) resistance and low-cost fabrication methods.

The reverse thrust/reverse flow air inlet configuration should be optimized, and concepts for variable fan duct exit areas as applicable to high-bypass-ratio designs need to be investigated.

Gear.-The advanced technology gear requirements noted for the turboprop are also applicable to the VPF.

General Further advances in the following items, although not essential, would improve all engines including the base turbofan: * Higher pressure ratio per stage capability at good efficiency levels, for fans, compressors, and turbines * Improved turbine airfoil, endwall, and secondary cooling capability * Improved compressor, burner, and turbine durability to minimize performance degradation and maintenance costs * Improved clearance controls throughout the engine, to improve component perform- ance and reduce leakages and secondary flows * Improved materials such as composite fan blades, high-temperature titanium for compressor disks, high-strength nickel base alloys for turbine disks, and high- temperature turbine airfoil materials and coatings * Shorter burners with improved performance and low emissions technology * Fabrication technology including reduced cost turbine airfoils, small cooling hole technology, near-net and net-shape disks, large cast cases, lightweight structures, and fabricated sheet metal cases * Improved bearings; improved seals including abradable/abrasive seal technology on airfoil/cases for tighter clearance control * Improved accessories including an electronic control to improve performance, hot section life, control maintenance, and reliability

REFERENCES

1. The Boeing Company, Final Report, Study of the Application of Advanced Technologies to Long-Range Transport Aircraft, Volume 1-Advanced Transport Technology Final Results, NASA CR-1 12092, May, 1972.

2.

The Boeing Company, Advanced Subsonic Long-Haul Transport Terminal Area Compatibility Study, Volume I-Compatibility Assessment, NASA CR-132367, Febru- ary, 1974.

3. The Boeing Company, Advanced Subsonic Long-Haul Transport Terminal Area Compatibility Study, Volume II-Research and Technology Recommendations, NASA CR-132368, February 1974.

4.

The National Aviation System Plan 1973 to 1982, DOT/FAA, document no. 1000.27 appendix 2, 1972; and Study and Concept Formulation of a Fourth-Generation'Air Traffic Control System, The Boeing Company, DOT report no. DOT-TSC-306-1, April, 1972.

5. Lockheed Georgia Company, A Proposed Standard Method for Estimating Airline Indirect Operating Expense, Robert F.

Stoessel, LW70-500R, May 1970.

6.

Economic Effects of Propulsion System Technology on Existing and Future Transport Aircraft, NASA-Lewis, NASA CR-134645, 1974.

7. Hamilton Standard, Generalized Method of Propeller Performance Estimation, PDB 6101, revision A, June 1963.

8.

The Boeing Company, Final Report, Study of the Application of Advanced Technologies to Long-Range Transport Aircraft, Volume II-Advanced Technology Program Recommendations, NASA CR-112093, May 1972.

9. The Boeing Company, Final Report, An Advanced Concept Secondary Power System Study for An Advanced Transport Technology Aircraft, NASA CR-1 12103.

TAC/Energy Study 1974 TAC study 1973 Standard Delay CWBLRC.

CWB-E ATT-E TAC.E TAC/Energy DOG, S/trip 1852 km (1000 nmi) Resized to meet mission = 5.8d/ (2 d/gal) fuel requirements at long-range cruise LR 0.82 C Crew 0 700.78 623.44 623.33 692.38 Fuel , 837.06 839.34 869.28 601.66 Maintenance ; 704.11 638.67 874.07 840.85 Insurance 136.25 116.76 131.20 124.25 085 MLRC.82 Depreciation 960.49 822.77 926.66 876.35 TOGW= 152 400 kg (336000 ib) TOGW = 147 200 kg (324600 Ib) = Total 3402.13* 3338.69 3040.98 3424.54 3135.49 OEW 88 o g 3194 O (3.) eW E= 8 90 g (189 3605 Ibl ATT tudy 197172 Aspect ratio = 6.8 Aspect ratio 6.8 DOC not calculated by ATA formula (see sec. 6.2.3.1) Block fuel, kg Il) 1852 km (1000 nmi) 11975' (26400) 11 340 (25 000) 11 340 (25 000) 11 793 (26000) 8165 (18000) 55% LF reuiremets at ongGrmnge 749 (72 200) 33 158 (73 100) 24 313 (53 600) 5556 km (3000 nmi) 38 465' (84 800) 35 562 (78 400) 32 100% LF cruise MLR =0885 C plus latestaero and Block time, hr propulsion data base = 2.523 MCR o.90 MLRC 0.885 1852 km (1000 nmi) 2.473 2.473 2.267 2.267 = 6.207 6.207 6.873 NAS document TOGW 137 10O kg (302 200 Ib TOGW 132 40 kg (291 80 Ib 5556 km (3000 nmi) 6.796 6.796 = *Estimated at long-range cruise conditions CR OEW 742 200 g (159 200 dgI SEW Ep74 00 lc10 63 3206 dm Aspect ratio = 7.6 Aspect ratio = 7.6 Expected Delay CWBLRC* CWB E ATT-E TAC-E TAC/Energy DOC, y/trip Resized to meet mission 1852 km (1000 nmi) - requirements at long-range I = .d/ (22d/gall fuel 0885 C 6e plus latest aero and propulsion data base Crew 838.28 760.94 623.33 692.38 Fuel 947.47 985.78 869.28 601.66 TAE Maintenance 78337 738.59 874.07 840.85 Desgn criteria (al configuratiTAC = 0.885 Insurance Z 136.25 118.76 131.20 124.25 Design pavoad 18 140 k (40 lb Mission 0.90 TOG = 140800 kg (31500 Ib Depreciation 960.49 835.77 926.66 876.35 T.O. field length - 2530m 00 tt TG 0 900 kg (Se c i ( 8 16 320 Ibd at 305 (1000 ftI at TOGW= 141 100 kg (311 100 5b Sweep c 4) =0(637 rad 136.5 d Total 3740.00* 3665.86 3439.84 3424.54 3135.49 an 0 5 K 9 l dSn /4 0.7 Aec ra3io. 9.0 = DOC not calculated by ATA formula (see sec. 6.2.3.1) Amect rati 9.0 Sized to meet mission requirements at long-range cruBlock fuise MLR =0.80 C Block fuel. kg (Ib) Cruise speed sizing criteria code plus latest aero and MCR Critcal Mach number (minimum cost propulsion data baSe 1852 km (1000 nmi) 13585- (29 950) 12 882 (28 400) 13 381 (29 500) 11 793 (26 000) 8 165 (18 000) 55% LF MLR = Longangecruis Machnumber (minimum fuel) C 5556 km (3000 nmi) 40 642 (89 600) 37 988 (83 750) 34 994 (77 150) 33 158 (73 100) 24 313 153 600) TAC Energy = 100% LF MLRC 0.0 TOGW = 115 300 kg (254 200 Ib) = OEW 67 310 kg(148 390 Ib) Block time, hr Swee (c4 = 0.436 rad 125.0del 2.767 2.267 2.523 1852 km (1000 nmi) 2.973 2.973 5556 km (3000 nmi) 7.296 7.296 6.707 6.207 6.873 *Estimated at long-range cruise conditions TAG/Energy Concept and Comparison Concepts Evolution PAGE IS ORIGINAL OF POOR QUALITY

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