Evaluation of active control technology for short haul aircraft

Page Section 5.0 Turboprop Airplane Characteristics with Active 66 Controls 5.1 Characteristics of Airplanes Incorporating Ride 66 Quality Systems 5.2 Characteristics of Airplanes Incorporating Gust 66 Load Alleviation Systems Characteristics of Airplanes Incorporating Artificial 74 5.3 Stability Systems 6.0 Comparison of Airplane Characteristics With and 79 Section Without Active Controls 6.1 Comparison of Turboprop Airplane Characteristics 79 6.2 Turbofan-Powered MF Airplane Characteristics 115 and Comparison of Concepts Section 7.0 Conclusions and Recommendations 121 List of References 123 vi LIST OF FIGURES Page Figure Title Active controls) xviii (278 km; 150 n.mi.) vs. Field Length and Mach No. (No S-1 DOC-2 xx 610 m (2000 ft) Field Length Design Point Aircraft S-2 xxi S-3 914 m (3000 ft) Field Length Design Point Aircraft xxii S-4 Ride Quality Analysis xxiii Change due to GLA S-5 Wing Box Weight xxiii Capability and Redundancy S-6 System Designation, xxvi of Aircraft Characteristics S-7 (3 sheets) Comparisons xxv Effect of Active Controls on Horizontal Stabilizer Weight S-8 xxx S-9 Effect of Active Controls on Fuel Usage Per Passenger xxxii S-10 MF Vehicle xxxiii S-11 OTW/IBF Vehicle vii LIST OF FIGURES (CONTINUED) Figure Title Page 1 Engine and Propeller Weight and Cost 2 Wing Weight Correlation 19 3 Example Thrust/Weight Ratio vs. Wing Loading (4-Engines) .24 4 Example Computer Sizing Print-out 26 5 Example DOC-2 (150 n.mi.) vs. Wing Loading 27 6 DOC-2 vs. Aspect Ratio (278 km; 150 n.mi) 28 DOC-2 vs. Aspect Ratio (926 km; 500 n.mi) 28 8 Example Thrust/Weight Ratio vs. Wing Loading (2-Engines) 30 9 DOC-2 (278 km; 150 n.mi) vs. Field Length and Mach No. (No Active 31 Controls) 10 610 m (2000 ft) Field Length Design Point Aircraft 32 11 914 m (3000 ft) Field Length Design Point Aircraft 33 12 Ride Quality Evaluation (W/S = 215 kg/sq.m; 44 Ib/sq.ft) 37 13 Ride Quality Evaluation (W/S = 287 kg/sq.m; 58.71 lb/sq. ft) 38 14 Typical Load Factor Envelopes 40 Ride Quality Control System 45 Ride Quality System Effects (610 m; 2000 ft. F.L.) 48 17 Ride Quality System Effects (914 m; 3000 ft. F.L.) 49 18 Ride Quality and Gust Alleviation Control Flap 19 Horizontal Tail Sizing 20 RQ System Weights vs. Wing Loading 68 viii LIST OF FIGURES (CONTINUED) Figure Title Page 21 RQ System Weights vs. Number of Passengers 68 22 RQ System Costs (Scaled) 69 23 GLA System Weights - Scaled 73 24 GLA System Costs - Scaled 73 25 VB Speed vs. Field Length 75 26 Gust Load Factor vs. Field Length 27 Wing Box Weight Change Due to GLA 76 Controls) 80 28 Direct Operating Cost vs. Field Length (44 Pax - Without Active 29 Direct Operating Cost vs. Field Length (100 Pax - Without Active Controls) 81 (148 Pax - Without Active Controls) 82 30 Direct Operating Cost vs. Field Length DOC-2 (278 km, 150 n.m) vs. Field Length (44 Pax) 83 32 DOC-2 (278 km, 150 n.m) vs. Field Length (100 Pax) 84 33 DOC-2 (278 km, 150 n.m) vs. Field Length (148 Pax) Controls 86 34 Percent Change in DOC-2 (278 km, 150 n.m) Due to Active to Active Controls 87 35 Percent Change DOC-4 (278 km, 150 n.m) Due 36 Percent Change in DOC-2 (926 km, 500 n.m) Due to Active Controls 88 Active Controls 88 37 Percent Change in DOC-4 (926 km, 500 n.m) Due to Percent Change in DOC (278 km, 150 n.m) vs. Passenger Size - 610 m (2000 ft) 89 F.L.

39 Percent Change in DOC (926 km, 500 n.m) vs. Passenger Size - 610 m (2000 ft) 90 F.L.

ix LIST OF FIGURES (CONTINUED) Page Title Figure Size - 914 m (3000 ft) 91 Change in DOC (278 km, 150 n.m) vs. Passenger 40 Percent F.L.

km, 500 n.m) vs. Passenger Size - 914 m (3000 ft) 41 Percent Change in DOC (926 F.L on Mission Fuel (278 km, 150 n.m) - 44 Pax 42 Effect of Active Controls Mission Fuel (278 km, 150 n.mi) - 100 Pax 43 Effect of Active Controls on Mission Fuel (926 km, 500 n.m) - 148 Pax 44 Effect of Active Controls on Fuel (926 km, 500 n.m)- 44 Pax 97 45 Effect of Active Controls on Mission km, 500 n.m) - 100 Pax 98 Effect of Active Controls on Mission Fuel (926 500 n.m) - 148 Pax 99 of Active Controls on Mission Fuel (926 km, 47 Effect Per Passenger 100 48 Effect of Active Controls on Fuel Usage of Active Controls on RGW - 44 Pax 49 Effect 50 Effect of Active Controls on RGW - 100 Pax 51 Effect of Active Controls on RGW - 148 Pax OWE vs. Passenger Capacity - 4 Engines, 610 m (2000 ft) F.L 104 52 RGW and 53 RGW and OWE vs. Passenger Capacity - 2 Engines, 914 m (3000 ft) F.L 105 54 Percent Change in OWE due to Active Controls 106 55 Effect of Active Controls on Horizontal Stabilizer Weight 56 Effect of Active Controls on Airframe and Aircraft Price - 44 Pax 108 57 Effect of Active Controls on Airframe and Aircraft Price - 100 Pax 109 58 Effect of Active Controls on Airframe and Aircraft Price - 148 Pax Capacity - 4 Engines, 610 m (2000 ft) 111 59 Airframe and Aircraft Price vs. Passenger F.L x LIST OF FIGURES (CONTINUED) Figure Title Page - 2 Engines, 914 m (3000 ft) 112 60 Airframe and Aircraft Price vs. Passenger Capacity F.L 61 Percent Change in Airframe Price due to Active Controls - 44 Pax - 100 Pax 114 62 Percent Change in Airframe Price due to Active Controls Percent Change in Airframe Price due to Active Controls - 148 Pax 114 64 Percent Change in Airframe Price due to Active Controls vs. Passenger Capacity 116 65 MF Vehicle 117 66 OTW/IBF Vehicle 119 xi OF TABLES LIST Table Title Page S-1 Effect of Active Controls on Wing Loading xxv Ratio and Aspect Concepts xxxiv S-I! Comparison of I Design Requirements and Selection Criteria 6 II (Sheet 1 of 2) Deflected Slipstream Aero Data 11 -4 Engines (Takeoff) Deflected Slipstream Aero Data 12 (Sheet 2 of 2) Engines (Landing) -4 III (Sheet 1 of 2) Deflected Slipstream Aero Data -3 Engines (Takeoff) (Sheet 2 of 2) Deflected Slipstream Aero Data -3 Engines (Landing) IV T-56 Engine/Quiet Propeller Characteristics V Comparison of Analytical and Statistical Wing Weights 20 Points 31 VI Selected Baseline Design Design Point Aircraft Characteristics (No Active Controls) VII VIII Baseline Aircraft Derivatives 35 IX Flight Conditions 36 X Airplane Characteristics (No Active Controls) 43 System Designation, Capability and Redundancy XI XII Weight and Cost - RQ System 52 xii LIST OF TABLES (CONTINUED) Title Page Table XIII Weight and Cost - GLA System 58 XIV Wing Weight Change due to Gust Loading 58 Characteristics of Airplanes with RQ System 67 XV XVI Optimization of Aspect Ratio with GLA System 71 XVII Characteristics of Airplanes with GLA System.

Characteristics of Airplanes with AS System 78 XVIII XIX Comparison of Concepts xiii AND ABBREVIATIONS SYMBOLS airplane aspect ratio AR AW augmenter wing b span coefficient drag CD lift coefficient CL coefficient thrust C T force coefficient CX axial c chord seat statute mile c/ASSM cents/available conventional takeoff and landing CTOL decibel dB direct operating cost DOC DOC at 11.5c/gallon of fuel DOC-1 23c/gallon of fuel DOC-2 DOC at DOC at 46c/gallon of fuel DOC-4 of fuel DOC-10 DOC at $1.15/gallon decibel EPNdB equivalent perceived noise FAR Federal Aviation Requirements FPR fan pressure ratio xiv AND ABBREVIATIONS (CONTINUED) SYMBOLS constant gravitational g IBF internally blown flap Mach number or Meter M m meter MF mechanical flap OTW over-the-wing hybrid over-the-wing/internally blown flap OTW/IBF OWE operating weight empty q dynamic pressure RGW ramp gross weight square rms root mean SFC specific fuel consumption sea level static SLS STOL short takeoff and landing wing thickness t setting T/O takeoff power takeoff field length TOFL ratio T/W airplane thrust/weight velocity V weight or airplane weight W angle of attack a of A increment of wing span fraction power setting or xv SUMMARY determined that turboprop and studies conducted for NASA Ames (Ref. 1) it was From with powered were economically competitive turbofan powered mechanical flap airplanes flap field length. The turbofan powered, mechanical lift concepts for 914 m (3000 ft.)

lift concept, but the than the most promising powered airplane has higher fuel consumption therefore be considered a fuel consumption and must turboprop design provides lower the turbofan and turboprop concepts contender for short-haul operation. Both major required for this short-field because of the low wing loadings would have poor ride quality performance.

to evaluate the economics in this report was The primary objective of the program described and low wing-loading. The active control technology of short-haul aircraft incorporating to the program was: overall approach with active controls.

o Determine the airplane characteristics the airplanes.

o Define active controls systems for incorporating active control systems.

the characteristics of airplanes o Determine - The turboprop-powered and Evaluation of Active Controls Turboprop Characteristics power plant selected was a study utilized 2- and 4- engines. The concept chosen for the a rubberized, low tip-speed Detroit Diesel Allison T-56 engine combined with rubberized for previous Lockheed work (Ref. 2).

"quiet" propeller defined by Hamilton Standard Figure S-1 which indicates that resulted in the data shown in A preliminary analysis as DOC-2) for operating cost at twice pre-energy crisis fuel price (identified minimum direct at - km (150 nm) stage length is provided a 278 (3000 ft.) by a 4-engined configuration Field lengths shorter than 914 m o effect and meeting FAR Part-XX regulations utilizing the deflected slipstream cruising at 0.5M or less.

(Ref. 3), and by either a 2- or 4- engine design o Field lengths of 914 m (3000 ft.) or greater M and meeting to 0.6 m. The 2-engine design cruising at 0.6 cruising at 0.5 since it provided almost Part 25 (Ref. 4) was selected for further study FAR 2-engine configuration as a variable.

miminum DOC and retained the PRlCEDING PAGE BLANK NOT FILMED xvii 100 PAX T-56 BEST ALTITUDE, SPEED AND AR FOR 278 KM (150 N.MI.)

4.8 4.4 DOC-2 I2-ENGINES DO -0.6M 4.0 0.5 ¢/ASSM% 3.6 0.55 DESIGN POINT 0.5 4-ENGINES POINT DESIGN 3.2 0 1.5 A 12 2.5 FT 5 6 7 8 9 10 100M FIELD LENGTH Figure S-1 DOC-2 (278 km; 150 n.mi.) vs. Field Length and Mach No. (No Active controls) xviii were one with 2-engines and one with 4-engines, Two 44 passenger baseline configurations, ride quality analyses were in Figures S-2 and S-3. Longitudinal configured and are shown computer program.

these airplanes utilizing a 3 degree-of-freedom digital conducted for criteria did not meet the r.m.s. vertical acceleration Figure S-4 indicates that these aircraft of Ref. 5 & 6 for the descent case.

trailing edge flap segments and (RQ) systems, utilizing aileron, Ride quality control quality of the airplanes to equal or better were synthesized to improve the ride elevators were in Figure S-4. Duplicated electronics of estimated Boeing 737 data as shown that but against run-away of the complete system in these RQ systems to safeguard provided less comfortable since it would only result in a possible failure of the system was accepted, ride for the passengers.

supplies, dual piping channel, larger hydraulic power By the addition of an extra electronics surface sizes and each surface actuator, and modified and electro-hydraulic valves to for the two baseline system were synthesized electronic gains, gust load alleviation (GLA) cases were no more reduced the gust load factors such that the gust aircraft. These systems Gust load factors cases in designing the wing box structure.

critical than the maneuver a 20 m/sec (66 fps) gust at m/sec (50 fps) gust at cruise speed and resulting from a 15 gust load alleviation, were gust speeds (VB) for a series of airplanes, with maximum critical This reduction in must lower these load factors to below 2.5g.

calculated. The GLA system aspect ratio wings for which the aircraft to be reoptimized with higher gust effects permit due to the GLA system are shown in Figure equivalent wing box weight savings example S-5.

and additional channel in the pitch control electronics The addition of a fourth system to be modified to provides the necessary redundancy for the electro-hydraulic valves size of the horizontal "relaxed static stability" capability which permits the also provide for briefness as the "artificial stability" stabilizer to be reduced. This system is identified alleviation. Figure provides ride quality control and gust load (AS) system although it also control systems just described.

the capabilities of the three alternate active S-6 summarizes were developed of the two point designs, parametric computer methods Based on analyses and speeds, wing box weight change, gust estimate such parameters as gust load factors to system weights surface dimensions, and actuator, hydraulic system and electronic alleviation of relaxed static stability a horizontal stabilizer and costs. In order to determine the effects stability margins was developed and routine, accounting for the appropriate sizing incorporated into the airplane sizing program.

xix

n

n

PAYLOAD: 44 PASSENGERS RGW: 22,251 KG (49,055 LB) OWE: 16,005 KG (35,285 LB) WING AREA: 103.3 SQ.M. (1, 112 SQ.

FT.)

WING LOADING: 215 KG/SQ.M.

(44.0 LB/SQ. FT.)

ENG/PROP S.L.S.T.: 19.12 KN (4,300 LB) 926 KM (500 N.MI.) CRUISE: 0.5M ' 7620 M (25,000 FT.)

278 KM (150 N.MI.) CRUISE: 463 KM/HR (250 KEAS) C' 4570 M (15,000 FT.)

Figure S-2 610 m (2000 ft) Field Length Design Point Aircraft xx

Or

44 PASSENGERS PAYLOAD: (49,080 LB) RGW: 22,262 KG KG (35, 140 LB) OWE: 15,940 (832 SQ. FT.)

AREA: 77.3 SQ.M.

WING 287 KG/SQ.M. (58.8 LB/SQ.FT.)

WING LOADING: LB) 45.64 KN (10,261 S.L.S.T.: ENG//PROP (25,000 FT.)

0.6 M @ 7620 M CRUISE: 500 N.MI.

C 4570 M (15,000 FT.)

556 KM/HR (300 KEAS) 150 N.MI. CRUISE: S-3 914 m (3000 ft) Field Length Design Point Aircraft Figure xxi 215 KG/SQ.M.

M' 2000 FT. FIELD LENGTH, W/S = 44 LB./SQ. FT.

.16 a AIRPLANE WITHOUT RQ -j CRITERIA u .12 U •D.IiJ JIJI J I IIAJh.J j -j

a a a 737

U

- 08

> a 0 AIRPLANE WITH RQ % I S04 1.74 M/SEC 2.5 M/SEC 3.0 M/SEC Ou = 5.7 FT./SEC.

ou = 8.2 FT./SEC.

au = 9.8 FT./SEC.

I I I 1 I I I FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION 914 M 287 KG/SQ.M.

3000 FT. FIELD LENGTH, W/S 58.8 LB./SQ. FT.

.16 AIRPLANE WITHOUT RQ S.12 , CRITERIA ,, , ,I u /L/ -// U S737 .08- -

u0 0 a

- - .. - AIRPLANE WITH RQ : .04 1.74 M/SEC 2.5 M/SEC 3.0 M/SEC a = 5.7 FT./SEC. o = 8.2 FT./SEC. o = 9.8 FT./SEC.

u U U 0 1 II I I I I FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION Figure S-4 Ride Quality Analysis xxii o o o o o 0 0 4-ENGINE 2-ENGINE 0.5 M 0.6 M Z -3 I- 3 -2

z -3

/

-4Z.

-2 0 2 3 4 1000 FT 12 100 M FIELD LENGTH Figure S-5 Wing Box Weight Change due to GLA Designation Capability Redundancy Baseline No Active Controls Not Applicable RQ Ride Quality Control only Multiple surfaces and hydraulic (Ride Quality) systems with individual actuators.

Two electronic channels - FAIL SAFE GLA Ride Quality Control plus As for RQ plus third electronic (Gust Load Alleviation) Gust Load Alleviation channel and duplicated hydraulic supplies to each surface - FAIL OPERATIVE AS Ride Quality Control plus As for GLA plus fourth electronic (Artificial Stability) Gust Load Alleviation plus channel and third hydraulic Artificial Stability supply to pitch control - FAIL OPERATIVE after two identical failures.

Figure S-6 System Designation, Capability and Redundancy xxiii Airplanes were initially sized without active controls and then sized with each of the 3 levels of active control for field lengths of 457 m through 1067 m (1500 through 3500 ft) and for 44,000 and 148 passenger capacities. The resulting characteristics of these aircraft are summarized as a function of design field length in Table S-1 and Figure S-7. As mentioned earlier the 4-engine configurations utilize deflected slipstream effects and meet FAR Part XX requirements while the 2-engine designs meet FAR-25 which accounts for the differences in the wing loadings. The small changes in wing loading are due to the reoptimization to the higher aspect ratios with gust load alleviation. It was found that all the aircraft without gust load alleviation required aspect ratio 8, the minimum investigated, when optimized on the basis of minimum DOC-2 for 278 km (150 n.m.) stage length and 926 km (500 n.m.) range.

The weight data show the expected increase in gross weight for the 2-engined configuration relative to the 4-engine configuration, partly due to the higher cruise speed and partly due to the additional installed thrust and lower wing loading required to meet the field performance.

The provision of ride quality control incurs a weight penalty at all field lengths and for all passenger sizes, while the adoption of gust load alleviation, with or without relaxed static stability, offers weight saving benefits which are highest at the lowest wing loadings and increase with increase in passenger size. At the highest wing loading, weight penalties can actually be incurred due to the wing weight change at the higher aspect ratio and the subsystem weight exceeding the saving in fuel weight. The effects of the various systems, particularly the AS system, on horizontal stabilizer weight are shown separately in Figure S-8. These data are for identical types of stabilizer systems with and without active controls and therefore do not contain any reductions possible through the use of higher values of stabilizer lift coefficient.

As indicated in Figure S-7 large fuel savings are possible while still retaining minimum direct operating cost for short stage lengths (278 km, 150 nm) by the adoption of higher aspect ratio wings combined with a gust load alleviation system.

The figure shows the savings for the longest stage length (926 km, 500 nm). Smaller, but still worth while, savings are achievable for shorter stage-lengths but only at the longer-field lengths and with the 4-engine configuration. Note the relatively poor fuel consumption of the 2-engined configurations which is caused by a combination of the following reasons.

The selected 2-engine designs cruise at a higher speed than the 4-engine configurations which results in increased fuel consumption. However, even if the speeds were identical, the 2-engine design would still have poorer fuel consumption because it requires a lower wing loading (larger wing area) and higher thrust to weight ratio (larger engines) to meet the required field performance. The xxiv FIELD LENGTH - M 457 610 914 1067 (FT) (1500) (2000) (3000) (3500) WING LOADING - KG/SQ.M (PSF) 4-ENGINE BASELINE, RQ 156 (32) 215 (44) 347 (71) 395 (81) (78) GLA, AS 156 (32) 215 (44) 322 (66) 381 2-ENGINE SYSTEMS - 287 (59) 347 (71) ALL ASPECT RATIO 4-ENGINE 8 8 8 12 BASELINE, RQ 8 12 14 AS 8 GLA, 2-ENGINE - 8 8 BASELINE, RQ - AS - - 8 GLA, on Wing Loading and Aspect Ratio Table S-I Effect of Active Controls 1000 LB RQ PAX - - -GLA 1000 KG 2.0- 100 4-ENGINES 2-ENGINES 1.5 -0.5M 0.6M HORIZONTAL STABILIZER WEIGHT 1.0- 44 % PAX 0.5 S44 0 1 2 3 3 3.5 1000 FT 4 6 8 10 9 11 100M FIELD LENGTH Figure S-8 Effect of Active Controls on Horizontal Stabilizer Weight xxv z z U -- : , ..

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Characteristics S-7 (Sheet 2 of 3) Comparisons of Aircraft

O poop. QUATLM Figure

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YI, 'Iz (I"O)-O NI"NV~ IJ (1 z ~"4 ('WN0 )V0 I 30 V: z.W. OS -0 N01( I C a Mee1 I 1Ia w an ~ .. x 'A - z z Figure S-7 (Sheet 3 of 3) Comparisons of Aircraft Characteristics ii2 C- v ,i I 1 C ; zi+ p oWp, O O J o 9'~~6) ~IfN(W 96) WN 22 Figure 5-7 (Sheet 3 of 3) Comparisons of Aircraft Characteristics xxviii xx vijj result larger engines actually more fuel, while the thrust and hence larger wing requires more than the 4-engined designs, despite lower percent power setting in the airplane cruising at a fuel makes the specific percent power setting in speed. The lower the difference even further. Figure increases the fuel consumption poorer which of course consumption for the aircraft miles per gallon seat statute measure of fuel efficiency, S-9 presents, as a Note again the poor a function of field length.

RQ and AS systems, as optimized with the field length to 457 m drastic effects of reducing 2-engine design and the performance of the the the AS system has aircraft incorporating largest 4-engined ft). As expected the (1500 while also providing excellent ride quality.

best fuel efficiency increase due to the in Figure S-7 show a price price estimates shown The aircraft In at the shorter field lengths.

for all but the larger aircraft of active controls introduction basic the same aspect ratio as the GLA and AS retains the reoptimization with these cases, cost the weight and sufficient to offset weight and cost savings and results in airplane a cost the RQ system incurs system. The introduction of introduced by the active control and all passenger sizes.

increase for all field lengths are shown prices (DOC-2 and DOC-4) 2 and 4 times 1972 fuel The direct operating cost at and of 278 and 926 km (150 airplanes for stage lengths Figure S-7 for the baseline in the rapid to increase in passenger size and expected large reductions due 500 n.m). Note the length is approached.

as the shortest field increase m (3000 ft) field configurations at 914 to compare the 2- and 4-engine It is interesting but in some cases the more and use more fuel, designs are heavier, cost length. The 2-engine passengers, 4-engine designs. For 44 lower DOC than for the higher speed results in a the 100 both stage lengths. For for both fuel prices and minimum DOC 2-engines provides are equal two configurations are better at 926 km (500 n.m); the passenger case, 2-engines 278 km design is better for DOC-4 and km (150 n.m), while the 4-engine at DOC-2 and 278 for 926 km (500 n.m).

4-engine design is best except For 148 passengers the (150 n.m).

and DOC-2.

S-7 as percent on DOC are presented in Figure the active control systems The effects of results in a cases. The RQ system 926 km (500 n.m) DOC-2 and DOC-4, change for the passenger size. The GLA on field length and percent increase in DOC dependent 0.75 to 2.2 field lengths while providing provide reductions in DOC at the shortest and AS systems can field length, these two 914 m (2000 and 3000 ft) ride comfort. Between 610 and excellent passenger sizes while for the 100 and 148 the improved ride quality systems will provide can be as much as 1.7 the basic airplane value; DOC-4 DOC-2 within 0.5 percent of holding weight and cost 44 passenger size the increased baseline airplane. For the percent below the a further increase in DOC above offset any savings and result in of the GLA and AS systems the RQ system.

xxix 926 KM (500 N.MI.)

SEAT ST. MI./GAL.

90r

SEAT I

KM/ KG 4-ENGINE

r IM/

0.5

M

45 -

40 -

"~ P9-

LU W, PAX

S 60

,L '- - V

pn I 4I 25 0 - .0- 50 - /0 2-ENGINE 40/ 1 2 ' 3 4 1000 FT 6 8 10 12 100 M FIELD LENGTH Figure S-9 Effect of Active Controls on Fuel Usage Per Passenger xxx Turbofan Powered MF Characteristics and Comparison of Concepts In an earlier phase of the study (Ref. 1) it was determined that the mechanical flap (MF) concept which is illustrated in Figure S-10 and which is powered by two 1.35 FPR engines, provided optimum DOC-2 for 926 km (500 n.m.) at 0.70 M. This compares to the OTW/IBF concept which is illustrated in Figure S-11 and which optimized with four engines at 0.75 M. The results of the present study cannot be directly compared to the data from the previous phase for these airplanes because of different economic assumptions, updating of the computer program and some differences in equipment standards. Table S-II presents the characteristics of OTW/IBF, turbofan MF and turboprop aircraft which are all sized on a consistent basis and can be directly compared.

The wing loading of the OTW/IBF is high enough to obviate the need for a RQ system but both the MF and turboprop concepts are shown with and without active control systems.

The turboprop configuration is shown to have the lowest fuel consumption, DOC-2 and DOC-4 of the three concepts. It should be noted however that the turboprop engine performance and cost are based on a rubberized Detroit Diesel Allison T-56. It is obviously not possible to achieve these costs except at the actual T-56 size which would result in a 200 passenger aircraft rather than 148. Alternate sizes are possible for other speeds, field lengths and configurations. For example, a 2-engined, 48 passenger vehicle can be sized for 1067 m (3500 ft) field length and 0.6 M. Further alternates can be sized using other available engines while approximating the rubberized T-56 data. If a new advanced turboprop engine is used the DOC-2 and DOC-4 are increased by 11 and 6.5 percent respectively and the turboprop is then only competitive with the other concepts at and above DOC-4 fuel price.

It is concluded that any new turboprop aircraft must be sized to use an existing engine in order to keep the engine price down, and the aircraft economically competitive.

The MF with a ride quality system and the OTW/IBF are almost identical based on DOC-2 but the OTW/IBF has the advantage of a lower fuel consumption and a lower DOC-4.

In order to be competitive at DOC-4 the MF must use the AS system.

Conclusions and Recommendations The ride quality of short-haul airplanes with low wing loading can be improved to the standard of contemporary high wing loading airplanes by the use of active control systems.

The direct operating cost penalty for improved ride quality is 2 percent or less for all cases; incorporation of gust load alleviation and augmented stability overcomes this penalty and gives better DOC than aircraft without active controls in all the very low wing loading xxxi 148 PASSENGERS 0.70 MACH OPTIMIZED FOR DOC-2 ; . .. ... j - -- -- --- Figure S-10 MF Vehicle xxxii 148 PASSENGERS 0.75 MACH OPTIMIZED FOR DOC-2 Figure S-11 OTW/IBF Vehicle xxxiii 914M (3000 FT.) F.L., 926 KM (500 N.MI.)

A/C OPTIMIZED FOR DOC-2, TURBOPROP D.S.

MF 1.35 FPR CONCEPT OTW/IBF 1.35 FPR NONE RQ GLA AS NONE RQ GLA AS ACTIVE CONTROL NONE 4 4 NO. OF ENG. 4 2 2 2 2 4 4 0.70 0.5 0.5 0.5 0.5 MACH NO. 0.75 0.70 0.70 0.70 OWE - KG 36,510 39,687 39,850 39,189 38,970 34,179 34,303 35,135 34,981 S- (LB) 80,490 87,494 87,853 86,396 85,913 75,351 75,623 77,458 77,119 RGW - KG 56,446 59,848 60,026 59,387 59,110 52,694 52,828 53,278 53,084 0(LB) 124,440 131,940 132,332 130,924 130,314 116,168 116,464 117,457 117,028 "a RATED THRUST - KN 55.33 119.6 119.9 112.8 107.8 41.5 41.59 37.19 37.05 0 (LB) 12,440 26,890 26,948 25,365 24,231 9,330 9,350 8,355 8,332 MISSION FUEL - KG 4,400 4,790 4,802 4,749 4,708 3,601 3,609 3,335 3,304 o (LB) 9,700 10,560 10,586 10,470 10,380 7,938 -7,956 7,352 7,285 W/S - KG/SQ. M 554 287 287 287 287 347 347 322 322 T.O (LB/SQ. FT) 113.5 58.8 58.8 58.8 58.8 71.0 71.0 66.0 66.0 12 8 8 10 10 8 8 12 AR 1.911 1.897 1.909 1.884 1.876 1.7866 1.799 1.793 1.788 DOC-2 ¢/ASSM 2.347 2.336 2.304 2.117 2.129 2.097 2.090 DOC-4 ¢/ASSM 2.326 2.333 5.7143 A/C PRICE $M 9.103 8.2736 8.3984 8.4015 8.3778 5.5163 5.6248 5.7253 airplanes (457 m, 1500 ft field length) and in the 100 and 148 passenger airplanes at 610 m (2000 ft.) and 914 m (3000 ft.) field length. For small aircraft (44 passengers) a GLA or AS field lengths but for field lengths of 610 m (2000 ft.)

system is recommended for very short and longer the simpler RQ system results in a smaller DOC penalty. For larger airplanes (100-148 passengers) the GLA and AS systems generally provide smaller DOC penalties than the RQ system for field lengths of less than 914 m (3000 ft.). Above this field length the RQ system appears to minimize DOC effects except at the longer ranges and higher fuel prices where the increased aspect ratio of the GLA and AS systems results in improved fuel consumption and an advantage in DOC. Fuel savings of 11% were obtained by use of active controls in a 148 passenger airplane at 914 m (3000 ft.) field length and 347 kg/sq.m (71 Ib/sq. ft.) wing loading.

Weight savings were obtained with the GLA and AS system at the lower wing loadings where reoptimization did not increase wing aspect ratio. At longer field lengths and higher wing loadings the best economics of aircraft with active controls were obtained at increased aspect ratios. Fuel consumption was improved but small weight and cost penalties were incurred compared to baseline aircraft. Generally, the active control systems increased the initial cost of the airplane; the only exception being the largest aircraft at the shorter field lengths.

Due to the favorable fuel consumption and competitive direct operating costs, the turboprop-powered configuration with active controls must be considered a major contender for the short-haul low/medium density market, particularly for the shorter route segments where the time increase due to low speed is negligible. It must be stressed however that these turboprop aircraft, to be competitive, must be designed to match existing turboprop engines. The increased cost of a new turboprop engine will nullify most of the advantage of this configuration.

It may be that the low s.f.c. of a diesel engine might make consideration of this engine cycle worthwhile. Similarly since the development of a new turboprop engine is questionable it may be that a new, very high bypass ratio fan would be advantageous at these speeds and field lengths. The incorporation of active controls in the turbofan MF airplane results in it being equal to the OTW/IBF hybrid in terms of DOC and ride quality. However, the OTW/IBF, because of its higher wing loading, retains its advantage of lower fuel consumption.

This study has been limited to short-haul; it is likely that larger fuel savings are available by the use of active control systems on long haul aircraft which stand to gain so much more from higher aspect ratio wings, providing the wing weight increases can be minimized.

Active control systems combining features such as ride quality improvement, gust load alleviation, flutter control and relaxed static stability could result in very efficient high aspect ratio wings. It is recommended that such a program be considered with the final step being the flight demonstration of the wing design.

xxxv 1.0 INTRODUCTION NAS2-6995, Lockheed addressed the medium to high Background - Under Contract density short haul areas with 926 km (500 n.m.) as a key design range; airplanes were sized for 148 passengers and the studies were concentrated on low-noise fan-powered aircraft (Ref. 1, 5). It was shown that the "simple" mechanical flap (MF) configuration has better direct operating costs than both the externally blown flap (EBF) and the augmentor wing (AW) concepts at all field lengths longer than 914 m (3000 ft.) and is competitive with the less-well-developed hybrid over-the-wing/internally blown flap (OTW/IBF) concept at field lengths longer than 1067 m (3500 ft.)

Additionally, it was shown that turboprop powered aircraft operating at cruise speeds of .6 M and below, and field lengths of 914 m (3000 ft) and less, have the best fuel consumption and direct operating costs of all the concepts studies. This indicated superiority of the slower turboprop airplane is expected to be accentuated as shorter stage lengths are considered. At design ranges longer than 926 km (500 n.m.), the economic value of higher speed becomes overriding (as well as passenger preference).

Thus, it can be seen that turboprop powered deflected-slipstream airplanes must be considered as major contenders for the short-haul market operating at field lengths of 914 m (3000 ft) or less.

The turboprop must overcome the problem of passenger appeal, and since these concepts have low wing loadings, they must overcome problems such as poor passenger ride comfort and perhaps gust criticality of the structure.

With the recent development of active control technology, systems can be designed to improve ride quality, alleviate gust loading and provide artificial stability in addition to other possible applications.

A system providing ride quality improvement only will not be required to achieve "failoperative" operation and would accept poor ride quality if failure of the system occurs. While such a system would provide some gust load alleviation, the structure cannot be designed to the reduced loading since failure of the system could result in failure of the structure. However, the structural relief provided by the system will provide an improved fatigue life or can be used to provide a weight saving due to the reduced fatigue loading. The degree to which these benefits offset the system weight and cost has not been established previously.

alleviation, which permits the structure to be designed to reduced To provide gust load loads, requires a similar system but with redundancies incorporated to ensure continued system operation after a failure. Such a system will provide ride quality control, structural design load reduction and reduced fatigue loading. It must be established whether these weight to offset the system weight and cost.

benefits will save sufficient structure (relatively) wing area and generally a large horizontal Low wing-loading means a large the size of the horizontal stabilizer area. Active control technology can be used to reduce failure of such a system would mean an stabilizer by providing artificial stability. Since ensure continued operation unstable airplane, sufficient redundancy must be provided to after one or more failures. Again the economic tradeoff of the benefits versus the system and cost must be evaluated for short-haul aircraft.

weight of short-haul aircraft Objectives - The primary objective is to evaluate the economics designed with active controls technology and low wing-loading to achieve short field To fulfill this objective the following secondary performance with good ride quality.

objectives are necessary: ride qualities of two typical short-haul aircraft o Determine the unaugmented with low wing-loadings, suitable for 610 m (2000 ft) and 914 m (3000 ft) field length performance.

o Determine the gust criticality of these aircraft.

o Define active control systems for these aircraft to provide: o Ride quality equivalent to present short-haul aircraft such as the B737 and DC9.

o The above plus gust load alleviation and reduced structure weight.

o The above plus artificial stability and reduced horizontal stabilizer area.

o Determine the weight and cost of these systems.

o Determine the effect of these systems on the weight, first cost and operating costs of short-haul aircraft covering a range of field lengths and passenger sizes.

Approach - The general approach to the program was to divide it into three primary tasks, namely: o The determination of airplane characteristics without active controls.

o The definition of active control systems for these airplanes.

o The determination of airplane characteristics with active controls.

To achieve the primary objective of this study, the choice of airplane concept is not critical.

Data are required to compare the turbofan MF concept with the powered lift concepts at 914 m (3000 ft) field length but the major portion of the study has been conducted with the turboprop concept since: o It provides lowest operating cost.

o It provides lowest fuel consumption.

o It is expected to rate even better at shorter ranges.

o It is already operating from short runways at major hubs.

o Quiet turboprop engine/propeller combinations exist.

The first task required the parametric sizing of turboprop aircraft for short-haul operation.

Two- and four-engined configurations have been sized for Mach numbers of 0.5 to 0.6 and field lengths of 457 m through 1070 m (1500 - 3500 ft) which results in wing loading ranging from 161 kg/sq m through 415 kg/sq m (33-85 lb/sq ft.). From this family of aircraft two baselines were selected and their ride qualities determined and compared to those of a present-day turbofan short-haul aircraft known to provide a satisfactory ride.

This work is described in Sections 2 and 3.

Active control systems were then defined for the baseline airplanes to provide: o Ride quality improvement.

o Gust load alleviation and ride quality improvement.

o Artificial stability, gust load alleviation and ride quality improvement.

The weight and cost of these systems were determined and then used to develop parametric weight and cost relationships for different wing loadings and aircraft sizes.

These data are described in Section 4.

The third task involved the incorporation of these active control system data into the sizing program and the resizing and reoptimization of the baseline and parametric family of aircraft to determine the characteristics of the aircraft including active control systems as described in Section 5.

Section 6 then compares the characteristics of the aircraft before and after the introduction of active controls and determines the benefits and penalties associated with each system as a function of field length (wing loading) and passenger size. Finally the aircraft with active controls are compared to equivalent data for the powered lift aircraft developed under the earlier NASA contracts (Ref. 1, 5) and recommendations for further research and development listed.

DESIGN REQUIREMENTS AND EVALUATION CRITERIA 2.0 2.1 Design Requirements n.mi.) range are as stated in Ref. 1.

The general performance requirements for 926 km (500 used in the present study are summarized in The requirements and evaluation criteria Table I.

both 2 - The primary concept used throughout the study is the.turboprop with Power Plant of the Detroit Diesel and 4 engines. The primary power plant considered is a combination both Allison T-56 engine and the Hamilton Standard 4.9 m (16 ft) diameter quiet propeller, pricing for the engine and propeller is based rubberized to provide the required thrust. The price for the propeller. The airplane cost data on the current price of the T-56 and a study only at the design points using the actual generated using this engine are therefore accurate since a new turboprop engine is engine size. This is considered the most desirable approach new turboprop aircraft is most likely to use unlikely to be initiated and therefore any by rubberizing these engine data existing or modified engines. The relative trends provided are of course unaffected.

initial study to determine optimum parameters such as aspect Passenger Capacity - The Mach number and altitude used a 100 passenger capacity. Later studies, to ratio, cruise illustrate the effect of passenger capacity, used values of 44, 100 and 148. The baseline airplanes used for the structural and ride quality studies accommodate 44 passengers since system operating from this size is considered reasonable for introducing a new short-haul is representative of the aircraft suitable short runways (914 m; 3000 ft) at major hubs, and stage lengths in the low density market.

for short Range - As in the previous studies a range of 926 km (500 n.mi.) is required. Extended is not range up to 2778 km (1500 n.mi.) is not included since the turboprop concept such ranges.

considered suitable for To provide an adequate range of wing loadings for the evaluation, field Field Length - lengths of 457,610, 914 and 1067 m (1500, 2000, 3000 and 3500 ft) were included.

on the turboprop data generated in Ref. 1, 926 km (500 n.mi.) cruise Cruise Speed - Based Mach numbers of 0.5, 0.55 and 0.6 were considered and the optimums determined.

to results of previous studies involves comparison of aircraft with design speeds Comparison up to M 0.8.

REQUIREMENTS o POWER PLANT: RUBBERIZED T-56 AND "QUIET" PROPELLER o NUMBER OF ENGINES: 2 AND 4 o PASSENGER CAPACITY: 44, 100, AND 148 o RANGE: 500 N.MI.(926 KM) o FIELD LENGTH: 1500, 2000, 3000 AND 3500 FT.

(457, 610, 914 AND 1067 M) AND 250 KEAS (463 KM/HR) MINIMUM o SPEED: 0.5 - 0.6M o ALTITUDE: UP TO 25,000 FT (7620 M) o ASPECT RATIO: 8, 10 AND 12 o FAR 25 FOR TURBOFAN AND 2-ENGINED TURBOPROP DESIGNS o FAR XX FOR 4-ENGINED TURBOPROP (DS) DESIGNS SELECTION CRITERIA (PRIMARY) o MINIMUM DOC AT FUEL PRICE OF 23c/GALLON OF FUEL (DOC-2) FOR 150 N.M.

(278 KM) STAGE LENGTH AT BEST ALTITUDE AND SPEED Table I Design Requirements and Selection Criteria

For a stage length of 278 km (150 n.mi.) alternate speeds were considered as described in Section 2.2.

Cruise Altitude - The aircraft were sized for 916 km (500 n.mi.) cruise at 7620 m (25,000 ft) altitude. For shorter stage lengths the altitude was optimized as described in Section 2.2 Aspect Ratio - Aspect ratios of 8, 10 and 12 were considered and the optimum selected for each case.

Federal Aviation Requirements -The requirements of FAR Part 25 (Ref. 4) were applied to the turbofan mechanical flap and 2-engined turboprop designs while the requirements of FAR Part XX (Ref. 3) were applied to the 4-engined turboprop deflected slipstream and the hybrid powered lift designs.

Costing Methods - Airframe and engine costing were on the same basis as used in Ref. 1 except for inflation of 12% (6%/year) from 1972 to 1974.

The 1967 DOC estimating methodology was used except for the following changes, by agreement with NASA.

1. Block time minus flight = 10 min.

2. Block fuel as determined from the flight profile, using 6 minutes ground time and 4 minutes air maneuver time.

3. Reserve fuel for 370 km (200 n.mi.) at cruise altitude plus 15 minutes at 3050 m (10,000 ft) altitude, maximum endurance speed.

4. Crew cost = 2 x (0.05 x WG + 63) 5. Hull insurance: 1% rather than ATA 2%.

6. Utilization: 2500 hr. per year.

7. Labor rate: $7.45/hour.

8. Maintenance cost: 75% of 1967 ATA value.

9. Maintenance burden: Retain ATA factor of 1.8 25% engine spares.

15 years, 15% residual, 10. Depreciation: Fuel costs: 11.5c/gallon 11.

as DOC-1.

to 11.5c/gallon is identified DOC applicable also the following: Note DOC-2 - 23c/gallon - 46c/gallon DOC-4 - $1.15/gallon DOC-10 Selection Criteria 2.2 cruising at 7620 m were sized for 926 km (500 n.mi.) range Although the aircraft were selected and design cruise Mach number the optimum aspect ratio (25,000 ft) altitude, cost (identified as DOC-2) for a Operating Cost at 23c/gallon fuel on the basis of Direct this realistic for short-haul. At 278 km (150 n.mi.) which is considered more stage length of will reach 7620 m (25,000 ft); the length it is unlikely that the cruise altitude shorter stage and 6100 m by the computer at altitudes of 3050, 4572 aircraft have therefore been flown for this shorter stage ft). Similarly the optimum cruise speed (10,000, 15,000 and 20,000 926 km (500 n.mi.) range.

the design cruise Mach number for length may be different then determined for have been flown at each altitude and the value Accordingly 3 cruise speeds consumption provide a definite choice, DOC-4 and fuel minimum DOC-2. If DOC-2 did not were considered.

3.0 AIRPLANE CHARACTERISTICS WITHOUT ACTIVE CONTROLS controls into low wing loading In order to determine the effects of incorporating active aircraft, it is first necessary to size low wing loading aircraft without active controls and determine the characteristics of the parameters which may be affected by the incorporation of active control systems. To identify the magnitude of the effects at various wing loadings a family of aircraft with field performance varying from 457 m through 1070 m (1500 - sized. To identify the effects of size, passenger capacities of 44, 100 and 148 3500 ft) were have been configured.

Before sizing this complete matrix of airplanes a parametric study of the 100 passenger capacity aircraft was conducted to identify the optimum number of engines (2 or 4), and the optimum aspect ratio, cruise altitude and speed to provide minimum DOC-2 for a 278 km (150 n.mi.) stage length, and for field lengths of 457, 610, 914 and 1066 m (1500, 2000, 3000 and 3500 ft). This work is described in Section 3.1.

The optimum parameters obtained from this study were then used to size and configure the two baseline airplanes defined in Section 3.2. The ride qualities of these aircraft were then determined and compared to a contemporary short/medium haul aircraft known to have satisfactory ride qualities. These analyses are described in Section 3.3 and were used to design the ride quality control systems described in Section 4.0.

The two baseline aircraft were also used to conduct a wing structural analysis with the object of confirming or updating the sizing program weight routine. These analyses identified the degree of gust criticality of the baseline wings for use in designing the gust load alleviation systems of Section 4.0, and are described in Section 3.4.

Finally in Section 3.5 the matrix of airplanes without active controls are defined, reflecting the optimization of the configuration parameters and the updating of the weight routines.

3.1 Initial Parametric Sizing As explained in Section 1.0 the turboprop was chosen as the primary concept for this study. Both 2- and 4-engined configurations have been sized. The 2-engined designs were treated as conventional propeller driven airplanes with lift margins meeting the requirements of FAR Part 25, while the 4-engined designs make use of the slipstream-generated lift which qualifies as a powered lift concept under FAR Part XX performance ground rules. The aircraft could be either low-wing or high-wing arrangements, each having some advantages and disadvantages. The choice between them is too detailed to be determined in this study and will have little, if any effect on the conclusions. It was therefore assumed that all the airplanes would have a conventional low-wing configuration similar to the Convair 580 and Lockheed Electra.

for the deflected performance The aerodynamic Data.

Basic Aerodynamic 3.1.1 data from which operating and terminal C-130 cruise is based on detailed concept slipstream for requirements and wing area thrust In determining been derived.

effects have slipstream used as inoperative, were one-engine stall speeds, including landing the power-on takeoff and operating and one-engine data for the all-engines XX. CL, CX, CT by FAR Part permitted angle.

ranges of angle of attack and flap in Tables II and III for failed cases are provided the slipstream for effects of deflected do not consider the configurations The 2-engine speeds the power-off stall compatible with therefore made case and are engine-failed these airplanes III were used in sizing of Table II and Part 25. The data dictated by FAR Table III were used.

the CT = 0 data of cases only but for the engine-out by the Detroit Diesel manufactured The T-56-A-15 engine, Data.

3.1.2 Basic Propulsion While for this study.

baseline turboprop engine was selected as the Division of GMC, Allison available were immediately technology, data of the latest engine is not representative this advanced 1 showed that a new Additionally, Ref.

program (Ref. 1).

from the previous DOC due result in an increase in to the T-56 would identical shaft horse power turboprop of new the T-56. It,is considered that any of the new engine compared to to the higher price the T-56 data are modified engines and will use existing or powered by turboprops aircraft typical.

therefore on the existing T-56-A-15 are based the turboprop concepts data used for The propulsion generated by Hamilton which data had been propeller, for with a quiet engine combined 95 EPNdB is designed for (Ref. 2). This propeller Lockheed study Standard for a previous to 4.9 m propeller diameter by increasing the which was achieved at 152 m (500 ft) sideline takes The propeller design the disk loading.

and lowering the tip speed (16 ft), lowering results in only construction and spar and shell composite advantage of advanced technology gearbox change to associated with a T-56 includes the penalty small weight penalty, which a the this propeller, including Cost increases for shaft speeds required.

provide the lower a were found in and the gearbox changes costs of the propeller development distributed at takeoff and the an increase in thrust than offset by (Ref. 2) to be more previous study increased slightly.

at cruise only cost/thrust ratio identical to for this program are essentially propeller data generated The T-56 engine and improve the accuracy incorporated to 1; slight alterations have been those described in Ref.

effects, bleed airflow of the installation altitudes. A discussion at the lower cruise also estimation is used in the performance and bookkeeping procedures corrections of Ref. 1.

in Section 7.5.3 included AR= 10 8f C I .3264 .653 T 1.224 a -4. 0.

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05. -. 087 -.

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2.1 3.

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-. 738 . 74;-2 .. 935 .ii. -. 3 -. 57 IS ORIGINAL PAGE Table II (Sheet 1 of 2) Defelcted Slipstream Aero Data -4 Engines (Takeoff) OF POOR QUALITY AR = 10 36.

.653 1.224 .3264 C n.

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a 2.78 2.63 2.74 2,.27 1.78 1.31 /.80 3.38 3.35 3,.23 2.85 2.29 1.74 1.16 3.91 3.93 3.80 3.32 2.67 1.4 2.04 L 4.63 4e43 4*59 3.83 3.13 2.41 [,1.68 -. 83 -. 632 -,475 222 -. 312 -. 177 -.

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3.24 3.23 2.55 3.02 2.04 1.58 1.08 3.855 3.82 3.70 3.39 2.22 2.85 1.59 ".41 4.40 3.98 4*33 3.40 2.67 1.985 5.10 5.03 5.10 4.66 4.13 2.50 3.30 -. 86 559 -. 699 -. 465 -.

319 -. 382 -. 272 -.

676 -. 825 -. 540 -.

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3.30 3.30 2.59 3.05 2.05 1.57 1.10 3.91 3.86 3.45 3.75 2.90 1.65 2.29 4.46 435 4.44 4.03 2.79 3.48 2.09 5.15 5.15 4.78 5.15 4.25 2.65 3.44 -. 86 -,705 -. 484 -o573 -. 393 -. 280 -. 328 -. 825 -. 552 -.69 -. 447 -. 33 -. 116 -. 216 -. 805 -. 535 -*670 -. 406 -. 273 .. 02 -. 12 -. 77 -. 52 -*641 -. 298 .106 -. 13 .31 81.

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3.26 3.30 3.06 2.06 2.60 1.13 1.60 3.92 3.90 3.53 3.90 2.99 1.81 2.39 4.50 4.43 4.49 4.13 2.96 3.62 2.30 5.18 5.14 5.18 4.46 498 3.71 2.93 717 -. 835 -. 60 -.

-.485 295 -. 377 -. 238 -.

725 -. 825 -. 625 -.

-.495 229 -. 363 -. 123 -.

-. 82 580 -. 700 -. 448 -.

-. 308 .002 -. 14L -. 64 -. 76 -. 515 -,375 .045 -. 2 .27 Table II (Sheet 2 of 2) Deflected Slipstream Aero Dataj-4 Engines (Landing) AR = 10 CT .l.

.24±80 .48975 .9180,) a - . ... 01 Hi; 4.00ji)) 8.0 i)l(I 'n 12. 00((00 16. OO00)(00 20. U lif)( -. >, t . 0 .6700 1.09500 1.53000 1.6200i 1.560 1i; C -. 18590 .37875 .83626 1.32470 1.83251 2.07376 2.08939 L -. 170 i;"29r . .990 1 3 1.54505 2.11006 2.41508 2.50i),19 -. 1212 6 .5365 1.173011 1 .79721 2.42351 2.80922 2.9 894 S w.i, -. 045, -. 050, -. 08700 -. 13500 -. 1860) -. 2 i40 : .12701 .05270 -. 02593 -. 10812 C .1813 .19670 .17345 X 4 .- i- .'. h5 .4059 .33 854 .23854 .13303 .0 17953 . il .F5; .- .81191 .72198 .58958 .42959 .24i; .i . .24 8 .48975 .91 R80 -W. 4. Ol ; 8.0 )i 12.0;, 16.0 ti, 20.0 .

-. .A . .8400, 1 .2600: 1 . 68 1 . 83. I .81 n .l; .Uq49877 1.0201 1 . 5251 P2.fn 96AQ 2.29376 2.3, 18R9 .07( :1 .62, 2 1 .20 4 1 .77505F 2.31756 2.62 : 17 2.7350cQ .1- ;16 .78292 1 .41994 2.0390 2.63457 3.02720 3.254;7 -. i..; -. l ,i'0.. -. 6701 -. H 105 IWl: -. 16201) -. 2500 - 39(1 .7482 .18214 .15139 .09632 .01345 -. 09237 -. 219"3 .15b .4 1805 .37605 .3 02n4 .18954 .06403 -. 7 /97 .83745 .83595 .78077 .68A49 .53973 .3670( .16791 15.f, . . .

A2Lu,80 .48975 .9180i -'4 ., I .. . A; 4. 01)n i;, ;.00 1 2.0 11 1 6.0 ' i 20.0i; .160 ;( .95850' 1 .020ti, I .450i0;: 1.870)1h, 2.5090 2. 8 , .24250 .7059 1.23064 1.77251 2.23876 2.51314 2.57 ',L S.32 0 .85:13 1 .435 2.02( 6 2 . .5P2016 2.84508 2.9 0(i C .43 110 1.06,59 1.70357 2.31947 2.86753 3.281.30 3. 52 19" -. 0 78. -. 07201. -. 0920, -. 128O -. 201 0 -. 32Lnii -. 52500 .16474 .159P2 .12132 *06'01 -. 04137 -. 17512 -. 34793 .4035 .3910i .3 904 .25804 .12703 -. 02'47 -. 20147 S821 27 .8-70 .736 48 . 63.)74 .46581 .27697 .0 n5736 25.0.

.24 80 .48975 .918 -4 . .. . ,;:,. 4 i.0. 8.0'.. 12.0 ' 16.00;' 20.u ' .4u ., .8750; 1.31 1 I .740'i 2.16011i 2.350(0i 2.420 ; . 7626 1.06(782 1.60876 2.13313 2.57314 2.83314 P2.881 .7( .3 1 .24504 1 . 4 0- 2.420 6 2.905i7 3.2100 3..320, .8..) 51 1. 50625 2.17.F 2.783'. 3.34284 3.74580 3.,359 -. 1 ; -. 10-.1 01 -10.1.840; -. 2880 -. 40ii -. 6951i .13264 .1 i9 . .63 9 -. 1 318A -. 1 465 -. 3, 605 -. 57 81 .0133 -. 18397 -. 45098 4 .3- .3. .2654 .16704 q.72;:'9 .63 ., , . 50278 .305 0 . 0 214 -. 1932 . 7h 73 .; .24 .((i .4 897 .91R 0 0 .l 1 .3 1 .78; 2.27W!' 2.630; 2.740'i 2.780 1.0i77 51 ,64 a 3 2.170 -4 2.71189 3.08190 3.20;'65 3.23- 14 1.2805 1 . F1,; 2. 480 ,7 3.0 (85 i18 3 . 51:0,9 ,3.6630 3.6' i 1 .528> 2.,0q75 2.87973 3.54853 4.0(i6 4.21623 4.2Pa,, -. 15 1' -. 17,'(): -. 2: i -. 3120' -. 4750 -. 63201 -. 830I, . ('382 -. 1,31 - .0794' - .2 181 -. 483 -. 8A, ,2 -. 7Q462 . 1 6S .ii7803 -. 0777 -. 351 019 -. 32 -. 76)(49 .23,04 .59137 .50 1 i .37629 .1 2 1 -. 1 2 -. 41 r'1 -. 70i927 (Sheet 1 of 2) Deflected Slipstream Aero Data -3 Engines (Takeoff) Table III AR = i'l I S 36.0 .91800 .24480 .48975 CT .0II0n0 8.00000000.00000 16.00000 20.000{0 a .I' -. O()l0 .00000 4.00000 2,78000 2.63000 2.74000 1. 78000 2.27000 .801Ju19 1.31000 3.23314 3.08190 3.20065 2.71189 1.64063 2.17064 C 1.07751 3.65509 3.51509 3.63009 2.48007 3.08508 CL 1.28005 1.89006 4.24686 4.21623 3.54853 4.08006 2.20975 2.87973 3.52824 -. 83000 -1.7500 -. 63200 -. 22200 -. 31200 -. 1500 -. 17700 --. 58662 -. 79462 -. 20818 -. 40831 01030 -. 07949 C .03582 -.

-. 76049 31098 -. 52999 '-.07747 -.

.16954 .07803 X .23104 -. 41551 -. 70927 .19221 -. 11526 .50105 .37629 .59137 51 I i*0 II I0 .48975 .91800 .niiil .2LL80 00i0 12.00000 16.00000 20.O 4.00000 8.00000 -4. 000 (, .000n 3.09000 3.110121 2.50000 2.94000 1.99000 1.0 1.5 1.520:)1 3.56252 3 .43752 3.55002 3.10627 1 . 9481.4 2.56751 .3639 3.98509 3,9009 3.57510 3.91010 2.99510 1.630116 2.29508 4.58033 4.58096 4.12527 4.52219 2.75159 3.53959 1.99794 -.86501 -. 53700 -. 69200 -. 33000 -. 42000 -. 2490(i -. 27900 -. 65662 -. 83294 -. 37250 -. 50100 -. 16324 -. 26062 -. 09631 62249 -. 80649 -. 45749 -.

-. 17998 -. 31249 .05203 -. 04998 -. 56877 -. 77182 -. 17838 -. 36766 .16520 -. 02302 .31789 .n) (i i n E,0 .48975 .91800 .0 1(0 .24480 16.00000 20. l 0 0 8.00000 12.00000 0tI()(I ( .0000 4.00000 -4.

3.2401)[ 3.02000 3.23000 2.04000 2.55000 1.08000 1.58000 3.67315 3.7987 3.19564 3.53315 1.46970 2.07189 2.66377 4.132513 4.,01510 4.11009 3.12510 3.68511 1.78757 2.4508 4.70819 4.71388 4.28984 4.64184 2.96198 3.73763 2.22087 -.860H,, -. 55900 -. 69900 -. 38200 -. 46500 -. 27200 -. 31900 -. 68194 -. 83312 -. 42987 -. 54625 -. 21718 -. 32593 -. 12700 -. 51850 -. 66300 -. 81250 -. 26049 -. 38599 -. 00697 -. 12848 44936 -. 62071 -. 79029 -. 14362 -. 29978 -.

.20969 .03849 66. r00 ni .48975 .91800 .Ou 11 j( i 24480 20.0 00( 0 8.00(10 12.00000 16.00000 )0(0 .000(o 0 4. 00(0 ( - 4.0 3.30000 3.300')1 2.59000 3.05000 1 .. 10C)(I 1.570(0) 2.050(l0 3.76127 3.58127 3.71877 2.70439 3.25252 1.51.939 2.12376 4.18509 3.74011 4.05010 4.15009 1.870(j8 2.540119 3.19011 4.77213 4.71510 4.75805 3.83602 4.37630 2.34541 3.08967 70500 -. 86000 -. 48400 -. 57300 -.

-. 32800 -. 39300 -. 280011 -. 83281 45650 -. 55700 -. 69406 -. 24301) -. 34537 -.

-. 15525 -. 81500 -. 54350 -. 68000 -. 16798 -. 30149 -.42649 -. 04798 -. 65689 -. 78834 -. 35468 -. 52846 -. 01258 -. 20491 •.15723 81 * 000 ii0 .48975 .91800 .0HO)n .24480 20. 00000() 8.00 a 1i) 12.00000 16.0000 -4 . n ,: .0')0 4.000 (1(1( t) 3.26000 3.300 2.06000 2.60000 3.06000 1 .3 309 ! 1.60,,t; 3.750;02 3.70939 3.76064 2.77627 3.31814 1.651,9 2.20627 4.20509 4.200n(9 3.83012 4.16510 2.05509 2.67510 3.30512 4.80170 4.80235 .4.00871 4.52641 4.76026 2.58947 3.30528 -. 83500 -. 60000 -. 71700 29500 -. 37700 -. 48500 -. 23800 -.

-. 82739 -. 49606 -. 62312 -. 72506 -. 24668 -. 36906 -. 35237 -. 82250 -. 60250 -. 71250 -. 33549 -. 47150 -. 06048 -. 18649 -. 79018 -. 55225 -. 67248 05 36 -. 25858 -. 41602 .1278h -.

Table III ,(Sheet 2 of 2) Deflected Slipstream Aero Data\-3 Engines (Landing) IV. All with the quiet propeller are shown in Table Data for the T-56-A-15 baseline engine airplane performance scaling of the propulsion system is carried out in the appropriate to match the airplane thrust computer programs. The installed engines are scaled cost based on the factors shown in requirements with the scaling of engine weight and Figure 1.

logic within the sizing program is 3.1.3 Basic Weights Data - The weight estimation Short-Haul Aircraft Studies (Ref. 1).

essentially the same as that used for earlier air conditioning, furnishings, Modifications were made to the wing, electrical, instruments, set of weight estimating and operating equipment weight items to obtain a more general relationships.

logic, the wing box weight is determined by subtracting base weights for In the wing weight structure from a base weight for the total wing as the control surfaces and secondary was from a statistically correlated wing weight equation; this procedure determined assessed at 6.58 in Ref. 1. In this earlier procedure, the base aileron weight was described assessment was suitable for the psf for an aileron area equal to 5.3% of the wing area. This for the present study, previous studies which involved relatively high wing loadings; low wing this assessment resulted in high base aileron weights because of the however, to loadings under consideration. Therefore, the base aileron weight assessment was changed use the following logic: chord per unit wing chord = 0.27 o (E AlL) B = Aileron Aileron span per unit wing span = 0.3 o ( A q AIL)B = L)B = Aileron spanwise centroid per unit wing span = 0.85 o ( ( AI = 2 (1-(1-TR) ( Al L) B)(E AIL) B ( A i AlL)B) + (1 +TR) o (SAIL/SW)B 9 5 4 ( AIL). ) o (W/SAIL=0.1445(WG) - (SW(T/C)'2A = o (WAIL)B (W/S)AIL (SAIL/SW)B (SW) Where, = Wing taper ratio TR SW = Wing area (sq. ft.)

ratio at the aileron centroid (%) (T/C)A = Wing thickness-to-chord WG = Airplane gross weight Uninstalled Installed Overall Pressure Ratio 9.5 Airflow (power Section - Kg/Sec.

(Lb./Sec.)

14.67 (32.35) ESHP Thrust (Total, Prop.

+ Power Section) KN (Lb.)

52.48 (11,798) 48.55 (10,915) Weight (Power Section - Kg (lb.)

871. (1,920) (Propeller) - Kg (lb.)

502 (1,107) Diameter (Max., Power Section + Gearbox - m (In.)

0.99 (39) (Propeller) - m (Ft.)

4.88 (16.0) Length (Power Section) - m (In.)

3.71 (146) Stages - Compressor Turbine TIT 1F TW - T/O 3.89 2.58 Price/Lb. Thrust - T/O* 20.70 31.90 Speed Lapse (M - 0.2)** 0.788 At 9140 m (30,000 Ft., M - 0.6 Thrust (KN) Lb.

9.1 (2046) Lapse .1736 SFCP .5744 * Uninstalled T/O Thrust/Engine + Prop.

and Controls Weights ** Engine + Prop. and Controls Price T/O Thrust Including Estimated Prop. Development Cost.

Table IV T-56 Engine/Quiet Propeller Characteristics BASED ON T-56 AND QUIET PROPELLER 1000 LB 1000 Kg 6 2.5 2.0 WT 4 1.5 0 1.0 .5 I I I I I LB 12 16 20 1000 0 4 8 I I I I KN 20 40 60 80 THRUST (S.L.S.)

COST - $1000 50 I 16 20 1000 LB 0 4 8 12 KN 0 20 40 60 80 THRUST (S.L.S.)

Engine and Propeller Weight and Cost Figure 1 The correlation of the total wing weight estimating relationship was presented in Figure 127 of Ref. 1, page 203, for a broad range of contemporary transport aircraft.

In the initial phase of the present study, further correlation of this estimating relationship was derived for aircraft with low-to-medium wing loading and medium-to-high aspect ratio. This correlation is presented by Figure 2 which illustrates that this relationship also adequately predicts the wing weight for these aircraft. It should be recognized that the high aspect ratio wings on these contemporary aircraft are relatively straight wings (i.e., little or no sweep) and that this correlation does not imply that high aspect ratio swept wings would be accurately predicted by this technique. For the present study, however, this proof is sufficient since the low-speed configurations being studied have essentially straight wings.

The effects of aspect ratio and gust loads on wing weight was also a primary concern in the present study. To correlate the statistical weight relati =,hip used in the sizing program, a weight study was conducted using Lockheed-Georgia Company's analytical "Wing Weight Analysis Program".

The two initial baseline aircraft were analyzed using this analytical method and the wing statistical weight estimating relationship.

The aspect ratios were then varied over the range of 7 to 10 (the baseline aspect ratio was 8) to view the effects of aspect ratio and the comparative weight quantities from the two methods. To isolate these effects, only the baseline configurations are sized to perform the design mission with the remaining configurations defined by changing the Aspect Ratio to 7, 9, and 10 from the baseline of 8. The results of this analysis is shown by Table V which indicate close correlation between the two methods. The analytical method uses estimated maneuver, gust, and ground load conditions in conjunction with stiffness, strength, and geometric constraints.

For the above described study, all of the wing configurations were gust critical indicating that the statistical method does produce satisfactory results for a slightly swept wing, low wing loadings, medium to high aspect ratios which is the spectrum of interest for this study.

The electrical, instruments, air conditioning, furnishings, and operating equipment weight estimating relationships were modified to reflect the wide range of fuselage sizes caused by passenger capacity variations from 44 to 148. These relationships were developed from statistical correlation of passenger transport weight and design data. These relationships are as follows: (1) Electrical .5 8.6 (NPASS) + 114 (WG/1000) W ELEC = where, NPASS = Number of Passengers WG = Gross Weight (Ibs) LOADING HIGH ASPECT RATIO - LOW WING C-133A OC-124C 1649A 15 B377 0 A CL-440-4 C-130H ACTUAL WING WT ADC7-C UNITS

a P3B

188A P2V-71 AR 0 12 CV 440 CV240-16AA &C-8A a 10 4 F-270 ,. 7.5 A DC-3 G-159 4 6 8 10 15 20 30 WT - UNITS ESTIMATED WING Wing Weight Correlation Figure 2 44 PAX GUST CRITICAL -I WING WEIGHT - KG (LB) % DIFF.

< AIRCRAFT AR STATISTICAL* ANALYTICAL.** WEIGHT o 4 ENGINES 7 2874 2817 +2.01 0 M = 0.5 o (6335) (6210) 3038 2983 +1.84 SW/S = 215 KG/SQ.M. 8 (6577) (44 PSF) (6698) WG =23,130 KG 9 3192 3136 + 1.78 o (51,000 LB) (7037) (6914) SW = 107.3 SQ. M. 10 3337 3314 + 0.70 -- (1155 SQ. FT.) (7357) (7306) o 2 ENGINES M = 0.6 7 2313 2272 + 1.82 (5099) (5008) W/S = 287 KG/SQ. M. 8 2449 2430 + 0.77 (58.8 PS F) (5398) (5357) WG = 23,100 KG 9 2576 2593 - 0.31 (50,900 LB) (5678) (5716) SW = 80 SQ. M. 10 2695 2761 - 2.38 (862 SQ. FT.) (5942) (6087) * DETERMINED BY EQUATIONS IN AIRPLANE SIZING PROGRAM ** DETERMINED BY ANALYSIS OF LOADS, STIFFNESS, STRENGTH AND MATERIAL DISTRIBUTION.

(2) Instruments 4 5 1 (FL + BW).

(NCREW). X (CIN)(NENG) W INSTR = 9.21 of Crew Members Where, NCREW = Number for Turbo-Fan Engines CIN = 1.0 for Turbo-Prop Engines CIN = .883 NENG = Number of Engines (ft.)

FL = Fuselage Length BW = Wing Span (ft.)

(3) Air Conditioning 0 3 5 (DELP + 5.) (VP) .

WAC = (6) = 8.0 for the present study.

Design Pressure Differential (psi) Where, DELP = Volume (cu. ft.)

= Pressurized VP Furnishings (4) 5 6 2 (SF). + .68 (DNAC)(LNAC)(NENG) WFUR = 13.6 + 70.5 (KTP)(FW) (KFUR)(LCC)(WCC) +9.36 Wetted Area(sq. ft.)

Where, SF = Fuselage Diameter (ft.)

DNAC = Nacelle LNAC = Nacelle Length (ft.)

Furnishings Factor KFUR = Type of short-haul austere furnishings = 0.7 for for normal domestic furnishings = 1.0 used with long-haul 1.3 for plush furnishings = aircraft (i.e., intercontinental).

compartment length (ft.) including LCC = passenger galleys and toilets.

WCC = passenger compartment width (ft.)

sound insulation factor KTP = turbo-prop = 1.0 for turbo-prop; 0. for turbo-fan.

FW = Fuselage width (ft.)

(5) Operating Equipment WOE = WCREW + WATTND + WFOOD + ENOIL + UNFUEL Where, WCREW = 190 (NCREW) WATTND = 143 (NATTND) NATTND = Number of Attendants WFOOD = 3 (NPASS) + 10 (NCREW) ENOIL .003 (THR) (NENG) THR = thrust per engine (Ibs.)

UNFUEL = .008 (FCAP) FCAP = fuel capacity (lbs.)

The foregoing relationships yield comparable results for the 148 passenger, turbofan powered aircraft with mechanical flaps as was reported in Ref. 1.

3.1.4 Initial Sizing Data. The aerodynamic, propulsion, and weight data discused in the previous sections and the modified cost data of Section 2.1 were incorporated into the computer sizing program with which aircraft were sized for the combinations of parameters defined in the following table: Field Length Cruise Aspect Cruise m (ft) Mach No. Ratio Power Setting 457 (1500) 610 (2000) 0.5 8 0.6 0.55 10 through 914 (3000) 0.6 12 1.0 1067 (3500) (500 n.mi.) and a capacity of 100, a range of 926 km All the airplanes have a passenger Two- and 4-engined configurations were sized, all cruise altitude of 7620 m (25,000 ft).

of power settings was varied with the T-56/quiet propeller data. The range using rubberized aircraft field length and Mach number.

particular size, weight and thrust the program is used to calculate aircraft The cruise sizing portion of loadings for each combination of Mach number, aspect characteristics for a range of wing thrust to weight ratio (in terms Figure 3 shows an example plot of ratio and power setting.

to cruise the airplanes plotted against wing loading of takeoff thrust and weight) required settings.

for the three aspect ratios and three engine power thrust to weight of the program is used to calculate the The takeoff and landing portion loading for each field length and takeoff and landing as a function of wing ratio required for to meet field length and climbout For each case flap angle is optimized each aspect ratio.

data. Note that aspect ratio shows an example for the 4-engined requirements. Figure 3 single line represents all aspect ratios effect on the landing capability and a change has little at each field length.

weight requirements are found just meeting cruise, takeoff and landing thrust to Airplanes landing lines for each aspect ratio; an example intersection of the cruise, takeoff and at the can be selected for 910 m (3000 ft) field length. Airplanes is identified by a square symbol just meet the cruise and landing of the cruise and landing lines which will at the intersection is takeoff performance; an example while providing better than required requirment ratio for such an figure. Although the thrust to weight identified by the circle symbol in the cruise critical airplane, it may than that of the equally landing, takeoff and airplane is higher be lower.

wing loading is higher and its DOC may be the better choice since its (Ref. 1, 5) the optimum designs were previous studies covered by this contract In the km (500 n.mi.) or in some cases on basis of DOC for a stage length of 926 selected on the criteria have been modified to fuel consumption. In this study the the basis of minimum of 278 km (150 n,mi.)

designs on the basis of DOC-2 for a stage length select the optimum range the average short-haul stage length. The 926 km (500 n.mi.)

which is more typical of modify the airplane requirement. It was therefore necessary to is still retained as the design km (150 n.mi.) for selection of size for 926 km (500 n.mi.) but use 278 sizing program to n.mi.)

further complicated by the fact that for 926 km (500 the optimum. The process is 278 km (150 n.mi.)

is close to 7620 m (25000 ft) whereas for the optimum cruise altitude achieved at a lower cruise altitude and possibly at a the optimum DOC will probably be different speed.

4-ENGINES 100 PAX 0.5M T-56 25,000' ALT 926 KM (500 N.MI.)

0.40 * 8 AR -- - 10 0-- - -- 0.36 0 \ QS , 0.28, - 0 /. 0.8 , I O 0 % * -0 0.24

- -1.0

0.20 SI I I I I I I 6096 m at 3048, 4572, and the sized airplanes modified to fly was therefore The program altitude. The speeds at each and at 3 cruise ft) cruise altitudes 15,000 and 20,000 (10,000, Mach number as follows: vary with the design cruise cruise speeds used VD EAS km (150 n.mi.) Cruise Speeds Mach No. 278 KM/HR (KEAS) KM/HR (KEAS) KM/HR (KEAS) KM/HR (KEAS) 556 (300) 426 (230) 463 (250) 389 (210) 0.5 598 (323) 482 (260) 506 (273) 444 (240) 0.55 556 (300) 648 (350) 519 (280) 0.6 482 (260) of shows an example 250 KEAS or less. Figure 4 FAA limits speed to Below 10,000 ft the for the data available the program and illustrates for this portion of the computer print-out is shown aspect ratio at one power setting Only one speed and altitude combination.

each CPM2 in DOC-2 (identified as column loading the minimum figure. For each wing in the In the figure the best altitude and speed.

can be selected which then identifies the figure) c/ASSM at 3048 m minimum DOC-2 is 4.22 m (30.8 Ib/sq ft) the for W/ST.O. = 150 kg/sq.

for all was repeated This process km/hr (250 KEAS).

and 463 (10,000ft) altitude power wing loading, aspect ratio, and number of engines, Mach number, combinations of shown in Figure 5 loading such as those plots of DOC-2 vs. wing From these data, setting.

have been prepared.

are cruise requirements landing and at which takeoff, loading value transferring the wing By for that field 3 to Figure 5 the DOC-2 length from Figure exceeded for each field met or the short lines on as shown by can be determined and power setting aspect ratio, length, minimum DOC-2 field lengths the ratios and all seen that for all aspect 5. It can be Figure At 457 and 610 m (1500 sized for 0.8 cruise power setting.

achieved with the airplane was slight would have produced a power setting ( r = 0.7) 2000 ft) field length a smaller and has minimized.

field lengths the DOC-2 in DOC-2; at the longer further reduction M the 4-engine, 0.5 field length for ratio for each versus aspect 6 presents DOC-2 Figure and lengths of 457 optimum for field ratio 8 is that aspect It can be concluded designs.

aircraft and 3500 ft) the and 1066 m (3000 ft) while at 914 m (1500 and 2000 ratio variation.

to aspect is very insensitive operating cost choice aspect ratio. The n.mi.) versus at 926 km (500 of DOC-2 7 shows an example Figure for the a shorter field length than and occurs at ratio is more apparent of the higher aspect n.mi.) case.

278 km (150 c9H LOW/ RU'' H "0. 3 MATRIX 0/12/74 rlFOTh6/oQ STOL UIST.= 30 .FT M=.bO AIGE= -11 .HiM PAYLO'U 20 .L SEAT-. 10,.

SWE P= h.ODEG CI ':= .I 1 CDMISC= .0 D M=.~ IVER=1 IMACH= .IErIG= 56 11986=0 IVE =0) ITH=I) IGEAH=1 IVHARH=O IRANGE=0 SJNO. ENG USED IIl TIAL CRUISE=. ENG COST FACTOR=1.0' WSP2 CPM1 C""? CI'14 CP'Iln ETA FUEL MACH WSTO RNG2 ALT VT FTM CLID ROC2 .3829 3:0.8 150. 1'. 24 -. ' .72 19.7 2170. 30.6 4.0', 4.-L2 5.27 7.81 .596 4150.

.4191- 30.8 150. 10. 267.6 .67 12.5 2179. 30.6 3.86 4.30 5.17 '.80 .730 4284.

I&K n .458, 30.8 150. 10. 290.0 .63 12.4 218-. 30.6 3.76 4.22 5.15 7.92 .836 4528..

.427' 7 30.8 150. 15. 264.7 .69 22.5 1807. 30.4 3.95 4.*-1 5.32 8.06 .718 4 67.

S.4629 30.8 150. 15. 289." .65 2.'5 1817. 30.4 3.8,L '".31 5.26 8.1' .875 461LS.

.5032 3:0.8 15. . 315.2 .62 2'.4 1825. 30.4 3.7' U.27 5.28 '4.3P 1.069 4964.

.468LL 30.8 150. 20. 287.7 .68 37.0 1438. 30.2 3.95 4.'in 1.';.U 8.43 .H71 4870.

.5130 30.8 150. 20. 315.1 .64 36.9 147. 30.2 3.87 4.3q 5.4 8.57 1.064 5114. :Ol .5 76 30.8 150. 20. 342.5 .62 36.5 14t6. 30.2 3.83 4. 0 5. ,4 8.95 1.290 5')78.

,3829 46., 150. 1 0. 2 4. .72 17.8 1617. 46.1 3.32 3.60 4.15 -. 79 .590 2676.h1 - .419. 16.- 150. 10. 267.6 .67 17.8 1627. 46.1 3.20 3.48 4.0 '" 5.71 .6q9 2714. vie *829 2804.ji5 1635. 46.1 3.11 3.39 3.97 5.69 ."L5t8 46.,, 150. 10. 290.9 .63 17.9 m.

15. 264.7 .70 33.6 1294. 45.8 3.30 3.60 4.21 6.05 .710 2984.

x .42:17 46., 150.

01 .4629 46.-, 150. 15. 289.1 .6- 3.:. 130 . '5.9 3.20 3.51 4 13 6.00 .838 3041' ).

" .9l" 3151.

.5032 46.i, 150. 15. 315.2 .',3 3 .5 131 . 45.9 3.12 3.44 ".09 6.02 .4684 -6.6 150. 20. 287.7 .70 57.0 96. 45.5 3.34 3.68 4.38 6.47 .860 3401..

.5130 46.'.150. 20. 315.1 .67 57.5 976. 45.5 3.26 3.61 4.32 6.41 1.017 3460.

I0 .5:76 46., 150. 20. 342.5 .64 57.6 983. 45.5 3.20 3.57 4.31 6.52 1.205 3604.

.3829 62.6 150. 10. 24".' .72 '1.' 1419. 61.9 3.14 3.:8 3.87 5.32 .587 2361.

5.20 .674 2361.

267.6 .67 41.j 1429. 61.9 3.02 3.26 3.75 .4194 62.6 150. 10.

S.45',8 62.,) 150. 10. 290.P .63 21.3 1437. 6t.9 2.93 '3.l? 3.67 5.15 .780 2405. -.

.42'-7 62.( 150. 15. 264.7 .70 41.4 1093. 61.5 3.12 3.39 3.94 5.60 .705 2688.

3.03 i.31 .3.86 5.53 .806 2707.

.4629 62.t( 150. 15. 289.9 .66 40.6 1103. 61.5 .5032 62.6 150. 15. 315.2 .63 41.1 11' .. 61.5 2.95 3.23 i.80 5.49 .93'" 2752.

-N .4684 62.6 150. 20. 287.7 .70 73.8 765. 60.9 3.17 3.49 4.14 6.08 .851 3152.

.5130 62.6 150. 20. 315.1 .68 73.1 775. 60.9 3.11 3.43 4.n8 6.03 .976 3166.

.5676 62.6 150. 20. 342.5 .65 73.9 784. 61.0 3.05 3.38 4.o0L 6.01 1.'.30 3216.

" .3829 84.3 150. 10. 24a.4 .72 P3.9 131i. 83.3 .5.05 3.29 3.76 5.17 .586 2290.

.4194 84.3 150. 10. 267.6 .67 24.0 1343. 83.4 2.94 3.17 3.63 5.02 .648 2260.

.T7 2273.

.63 23.5 1353. 83.4'2.84 3.08 3.54 L .94 .45b8 84.3 150. 10. 290.9 C .4227 84.3 150. 15. 264.7 .69 47.7 982. 82.7 3.03 3. iO 3.85 5.47 .702 2648.

2631.

*4629 84.3 150. 15. 289.9 .66 47.2 994. 82.7 2.94 3.P1 3.75 5.37 .774 .5032 84.3 150. 15.

316.2 .63 47.0 1014. 82.7 2.87 3.14 3.68 5.31 .76' 2642_ .4684 84.3 150. 20. 287.7 .69 89.7 636. 81.7 3.09 3.42 4.07 6.02 .844 3178.

.5130 84.3 150. 20. 315.1 .67 89.8 648. A1.o 3.03 3..i5 3.99 5.93 .. 934 3148.

.5576 84.3 150. 20. 342.5 .65 8A. 658. 81.,4 2.9Q 3.31 3.96 5.90 1.051 3161.

.3829 30.8 150. 10. 244.4 .72 12.4 2173. 30.5 4.10 4.53 5.40 8.02 .597 4269.

.4194 30.8 150. 10. 267.6 .67 12.5 2183. 30.5 3.96 4.41 5.31 8.02 .736 4u19.

o .4558 30.8 150. 10. 290. .63 12.4 2192. 30.5 3.86 4434 5.29 8.16 *&W 4689.

.4227 30.8 150. 15. 264.7. .69 29.3 1826. 30.4 L.0f5 4.51 5.45 8.26 .720 4584.

.4629 30.8 150. 15. 289.9 .65 2P.4 1835. 30.4. 3.93 ".L2 5.40 8.33 .883 47R4..

.5032 30.8 150. 15. 311.2 .61 27. 1843. 30.4 3.86 4.38 5.43 8.58 1.083 5136.

.68 36.4 1"71. 30.P 1.04 *'.5, ';..,7 8.62 .874 4987.

.4684 30.8 150. 20. P2747 .5130 30.8 150. 20. 315.1 .64 3h.u 1480. 3.2P 3.96 4.50 r,.',7 8.7" 1.074 5260.

.5676 30.8 150. 20. 342.5 .61 3 -. 0 148.-. 3f0. 3:.o0 4.51 6:.6 ".21 '.308 5764. o .3829 46.5 150. 10. 244. .72 17.7 1603. 4,h.0 3. .. l 1i.j5 -. 76 .)92 26' 3.

.4194 "6.5 150. 10. 267., .68 17.6 1 .2. 46.0 3 3.2 j.50 .)k 5.69 .709 2673.

) .45F8 46.5 150'. 10. 290.' .64 17.7 lh20. 4 t.0 3.1, 1.41 o.98 5.~9 .848 277 .

0.5M T-56 BEST ALTITUDE AND SPEED 4-ENGINES 100 PAX 2 78 KM (150 N.MI.)

4.6 ', \ 4.2 DOC-2 \\ ¢/ASSM \\\ 3.8 \ r4.6 3.4 4.2 DOC-2 3.0 ¢/ASSM 3.8 - - - 0.8 71 4.6 1.0 0.9 3.4 457M 1500' F. L.

4.2 DOC-2 ¢/ASSM 8 AR 3.8 610M 2000' F. 914M 106L7M 3500' F. L.

F L.

3000' 3.4 3.4 3.0 90 LB/SQ. FT 60 70 80 30 40 50 KG/SQ.M.

350 400 250 300 150 200 WING LOADING (T.O.)

Figure 5 Example DOC-2 (150 n.mi.) vs. Wing Loading 150 N.MI.

278 KM 5.0 4.6 DO C-2 4.2 ¢/ASSM 100 PAX 0.5M 3.8 610M ( 000FT 4-ENGINE 3.4 914M (3000 FT) 1067M (3500 FT) 8 9 10 11 12 ASPECT RATIO Figure 6 DOC-2 vs. Aspect Ratio (278*km; 150 n.mi) 500 N.MI.

926 KM 2.8 51 0 fl 2.6 DO C-2 100 PAX 0.5 M FT) 2.4 610M (2000 ¢/ASSM 4-ENGINE 2.2 914M (3000 FT) 2.0 067M (3500 FT) 1.8 11 12 8 9 10 ASPECT RATIO Figure 7 DOC-2 vs. Aspect Ratio (926 km; 500 n.mi) Thrust to weight ratio versus wing loading data for 2-engined configurations are presented in Figure 8. Note the lower wing loadings, relative to the 4-engine deflected slipstream configurations of Figure 3 to achieve the same field performance.

Also note the higher thrust to weight ratios required for a given field length which when intersected with the cruise requirement results in lower percent cruise power settings than for the 4-engined configurations.

From similar data for other Mach numbers, Figure 9 was prepared showing DOC-2 for km (150 n.mi.) stage length versus field length for the 4-engine and 2-engine configurations at the different design cruise Mach numbers. The 4-engined deflected slipstream airplanes have equivalent DOC-2 to the 2-engined configurations at approximately 914 m (3000 ft) field length. As field length shortens DOC-2 increases but the deflected slipstream configuration rapidly becomes increasingly superior to the 2-engined configuration. At longer field lengths than 914 m (3000 ft) the 2-engined design should improve and be slightly superior to the 4-engined design due to its lower engine cost per unit of thrust overcoming the effect of the lower wing loading and higher thrust to weight effects on DOC.

Note that at this stage length the extra cost of meeting the higher cruise speeds results in poorer DOC at all field lengths shorter than approximately 914 m (3000 ft); this is due to the large wing areas of these airplanes. From these data the 0.5 Mach, 610 m (2000 ft) field length, 4-engine configuration and the 0.6 Mach, 914 m (3000 ft) field length, 2-engine configuration, indicated as "design point" on Figure 9, were selected as baseline design points for a more detailed analysis.

3.2 Baseline Airplane Configurations The parametric study conducted in Section 3.1 was based on the 100 passenger size.

However, it is considered that the airplane size required in the low/medium density market for operation into auxiliary short fields at major hubs is more likely to be smaller than 100 passengers.

It was therefore decided to size the baseline airplanes at the small end of the passenger sizes being considered.

The selected baseline design points determined in Section 3.1 have the characteristics shown in Table VI.

Baseline airplanes were sized with these characteristics for 44 passengers.

Configuration arrangements for these airplanes are presented in Figures 10 and 11 while weight and economic data are presented in Table VII.

3.3 Ride Quality Analyses This section describes the analyses conducted to determine the ride quality characteristics of the two baseline airplanes, and for comparison purposes, an airplane similar to the B737.

The study was limited to the longitudinal axis for which the airplane derivatives are shown in Table VIII.

PROP 0.5 M 2-ENGINES T-56 QUIET 926 KM (500 N. MI.); 7620 M (25,000') AR 8 --- AR 10 .56 AR .52 914 M M FT 1067 r FT 0.5 FT .48 1500 457 M .44 610 M \\ 2000 FT T/W T.40 .40 0.6 .36 .320.8 .24 .20 20 30 40 50 60 70 80 LB/SQ. FT.

300 350 400 K G/SQ.M.

150 200 250 W/s T.O.

Figure 8 Example Thrust/Weight Ratio vs.

Wing Loading (2-Engines) 100 PAX T-56 BEST ALTITUDE, SPEED AND AR FOR 278 KM (150 N.MI.)

4.8 4.4 2-ENGINES DOC-2 V 0. 6M 4.0 0.5 /ASSM 0.6M

POINT

3.6 0.55DESIGN

4-ENGINES - 3.2 DESIGN POINT 3.2 0 1.5 12 2.5 3 1000 FT 7 8 9 10 100M 5 6 LENGTH FIELD Figure 9 DOC-2 (278 km; 150 n.mi) vs. Field Length and Mach No. (No Active Controls) 914 (3000) 610 (2000) LENGTH - m (FT) FIELD NO. OF ENGINES (44.0) 287 (58.8) (LB/SQ. FT) 215 W/S T.O kg/sq. m, .40 .323 T/W T.O.

8 8 RATIO ASPECT 7620 (25,000) 7620 (25,000) CRUISE ALT. m (FT) 926 km (500 N.MI.)

0.5 0.6 CRUISE SPEED M 926 km (500 N.MI.)

0.78 SETTING 0.8 km (500 N.MI.) CRUISE POWER 463 (250) 556 (300) N.MI.) CRUISE SPEED - km/hr (KEAS) 278 km (150 (15000) 4572 (15000) CRUISE ALT. m (FT) 4572 278 km (150 N.MI.)

Selected Baseline Design Points Table VI PAYLOAD: 44 PASSENGERS RGW: 22,251 KG (49,055 LB) OWE: 16,005 KG (35,285 LB) WING AREA: 103.3 SQ.M. (1,112 SQ. FT.)

WING LOADING: 215 KG/SQ.M.

(44.0 LB/SQ. FT.)

ENG/PROP S.L.S.T.: 19.12 KN (4,300 LB) 926 KM (500 N.MI.) CRUISE: 0.5M C 7620 M (25,000 FT.)

278 KM (150 N.MI.) CRUISE: 463 KM/HR (250 KEAS) @ 4570 M (15,000 FT.)

Figure 10 610 m (2000 ft) Field Length Design Point Aircraft

r

T

1W

PAYLOAD: 44 PASSENGERS RGW: 22,262 KG (49,080 LB) 15,940 KG (35,140 LB) OWE: WING AREA: 77.3 SQ.M. (832 SQ. FT.)

LOADING: 287 KG/SQ.M. (58.8 LB/SQ.FT.)

WING LB) ENG/PROP S.L.S.T.: 45.64 KN (10,261 (25,000 FT.)

500 N.MI. CRUISE: 0.6 M @ 7620 M KEAS) @ 4570 M (15,000 FT.)

150 N.MI. CRUISE: 556 KM/HR (300 Figure 11 914 m (3000 ft) Field Length Design Point Aircraft No. of Engines 4 2 Field Length m (ft) 610 (2000) 914 (3000) 0.6 Cruise Mach No. 0.5 No. of Pax 44 44 (44.0) 287 (58.8) kg/sq.m (lb/sq ft) 215 W/S T.O .403 Installed T/W .325 Percent Cruise Power 78.0 75.5 OWE kg (Ib) 16,005 (35,285) 15,939 (35,139) RGW kg (Ib) 22,251 (49,055) 22,262 (49,078) Rated Thrust kN (Ib) 19.12 (4,299) 45.64' (10,261) 500 n.mi. Fuel kg (Ib) 1,622 (3,575) 1,653 (3,645) 150 n.mi. Fuel kg (Ib) 758 (1,671) 799 (1,762) Airframe $ M 2.5158 2.6713 Total Aircraft Price $M 3.4019 3.2687 926 km DOC-2 500 n.mi. ¢/ASSM 3.732 3.075 926 km DOC-4 500 n.mi. ¢/ASSM 4.235 3.583 278 km DOC-2 150 n.mi. ¢/ASSM 5.617 4.857 278 km DOC-4 150 n.mi. ¢/ASSM 6.398 5.675 Table VII Design Point Aircraft Characteristics (No Active Controls) No. of Engines W/S 215 kg/sq.m. (44 Ib/sq. ft.) 287 (58.8) Flight Condition Cruise Descent Approach Cruise Descent Approach CL /RAD 5.3 4.85 4.6 5.4 4.85 4.6 CM /RAD -. 1 62 .044 CL CD/ CM /RAD/SEC -20 q CM /RAD/SEC -8 CM /RAD -3.44 e CL f/RAD 1.60 1.49 2.1 1.08 1.00 1.66 CM /RAD -. 332 -. 298 -. 400 -. 223 -. 200 -. 322 Table VIII Baseline Aircraft Derivatives digital computer utilizing a linear 3 degrees-of-freedom The analyses were performed flexibility not include any structural model did The airplane mathematical program.

developed power turbulence was modeled with a Lockheed effects. Random atmospheric A vertical gust the Von Karman function.

which is similar to spectral density function -3 Corresponding rms of 10 was selected for this study.

velocity exceedance probability level 1.74, 2.5 approach conditions are descent and landing for the selected cruise, gust velocities based level values are These turbulence 8.2 and 9.8 fps), respectively.

and 3.0 m/sec (5.7, used are identical to the levels ASD-TR-61-235 and in NACA TN 4332 and on data reported for NASA Ames and Langley.

in previous ride quality studies wing are shown in Table IX for the two aircraft The three flight conditions considered airplane.

for the Boeing 737 type loadings evaluated, and APPROACH DESCENT CRUISE M M KM/HR M KM/HR KM/HR V(KTAS) ALT(Ft) (KTAS) ALT (Ft) (KTAS) ALT (Ft). V Airplane V 463 1524 556 7620 215 kg/sq.m S. L.

(5000) (77) (25,000) (250) (44 lb/sq ft) (300) W/S = 463 1524 181 667 7620 287 kg/sq.m S. L.

(5000) (98) (25,000) (250) (588 I b/sqft) (360) W/S = 463 1524 819 9140 S.L.

(250) (5000) (120) (442) (30,000) 737 Type Conditions Table IX Flight pitch rate evaluations vertical acceleration and r.m.s.

and 13 present the r.m.s.

Figures 12 airplanes. The figures also cases for the two baseline the cruise, descent and approach for These criteria are based acceptability of ride qualities.

criteria levels for determining show locations of 0.11 g's.

at three passenger compartment a r.m.s. vertical linear acceleration on in developed by Rustenberg is a similar to the R criterion is an A criterion which This frequency.

human discomfort function of 7, except that A does not include the Ref.

studies conducted for NASA Langley used is based on ride qualities The particular level of A contract (NAS2-6995), Ref 5. The work statement of this and is specified in the original deg/sec for cruise and descent angular pitch rate are 0.5 corresponding limit values on r.m.s.

to however, the criterion used approach. For this specific study and 1.0 deg/sec for landing haul aircraft by contemporary short/medium ride comfort is that provided define acceptable passengers.

be acceptable to a majority of B737, which have proven to such as the M (2000 FT) FIELD LENGTH W/S = 215 KG/SQ.

M. (44 LB/SQ.FT.)

.16 .16 AIRPLANE WITHOUT RQ 0) u .12 CRITERIA U 0 a a 737.

.08 > a o 0 .04 1.74 M/SEC 2.5 M/SEC 3.0 M/SEC u =5.7 FT./SEC.

o = 8.2 FT./SEC. a = 9.8 FT./SEC.

I I I I A I I A FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION I #,' U 1.0 LII Su 0.8 Lu 0.6 CRITERIA I- I II I 0.4 - AIRPLANE WITHOUT RQ U S0.2 0 737 , 0 I I I I I I I I I FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION Figure 12 Ride Quality Evaluation (W/S = 215 kg/sq.m; 44 Ib/sq.ft) M (3000 FT) FIELD LENGTH W/S = 287 KG/SQ.M.

(58.8 LB/SQ.FT.)

.16 AIRPLANE WITHOUT RQ .12 u ' 1 2,; CRITERIA •737

0 a a

- S.04 1.74 M/SEC, 2.5 M/SEC 3.0 M/SEC = a = 5.7 FT./SEC. a = 8.2 FT./SEC.

o 9.8 FT./SEC.

u u u 0 I l l I I I I I I FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION U . 1.0 u0.8 - 0.6 CRITERIA I 0.4 U AIRPLANE WITHOUT RO V 0.2 0 737 FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION Figure 13 Ride Quality Evaluation (W/S = 287 kg/sq.m; 58.71 Ib/sq. ft) < for The B737; it was possible to obtain directly comparable ride quality data It has not been type airplane using the same therefore necessary to calculate ride qualities for a B737 purposes the B737 type methodology as used for the baseline airplanes. For comparison meet the 13 which illustrate that both baseline airplanes data are shown on Figures 12 and the vertical in cruise and approach conditions but do not meet ride quality criteria airplane is better than the criterion in the descent conditions; the B737 type acceleration ride qualities are acceptably close to criteria in all flight conditions., The baseline airplane are unacceptably worse than the B737 the B737 for the cruise and approach conditions but in the descent case.

control systems which would improve the It was therefore necessary to define ride quality the descent condition.

baseline airplanes to the standard of the B737, particularly for 3.4 Structural Analyses wing described in Section 3.2 were used to conduct The two baseline configurations - structural analyses to determine the structure (1) the gust effects on (2) the fatigue effects on the structure modifications to the weight estimation logic.

(3) Weight Analysis Program" These analyses were conducted using Lockheed-Georgia's "Wing adapted for advanced design studies.

which includes structural analysis procedures gust loads were analyzed at gust velocities of ± 15.24 3.4.1 Gust Analyses - The wing 50 fps) at cruise speed (Vc) and ± 20.12 meters per second (± 66 fps) meters per second (± corresponding to CL max (flaps up) in the presence of the gust. The at the airspeed (VB) corresponding to these conditions along with the conditions at gust load factor envelope is illustrated in Figure 14. Each of the baseline configurations was analyzed dive speed (VD) loads for the VB, VC, and VD speeds at their to determine the gust load factor and conditions. The gust load factors were computed in accordance with the corresponding gust contained in FAR Parts 25 and XX (Ref. 3 & 4). The resulting gust load factors formulae configuration are as follows: for each baseline A.

44 Passengers, 610 m (2000 ft) Field-Length, 4 Engine Turbo-Prop, 0.5 Cruise Mach Number Configuration: o NG = 1 + A NG = km/hr (160 KEAS); A NG =± 1.92 o @ VB 296 GUST

VA VB. VC VD

4 II

GUST VELOCITY I!

3 M/SEC (FPS)'

7. 6 k5)

LOAD MANEUVER FACTO R 'N' k.

,

-;.6

-2 100 200 300 KEAS 0 200 400 600 KM/HR VELOCITY W/S-PSF 44 58.8 V -BKEAS 160 -296 KM/HR NVB 2.92 3.0 VC -KEAS 250 KM/HR 463; 556 NVC 3.52 3.52 Figure 14 Typical Load Factor Envelopes KEAS); A NG = + 2.52 @ VC = 463 km/hr (250 o Field-Length, 2 engine turbo-prop.

44 Passengers, 914 m (3000 ft) B.

Number Configuration: 0.6 Cruise Mach = o NG 1+ ANG NG = ±2.00 (194 KEAS); A @ VB = 359 km/hr o = A NG = ± 2.52 VC 556 km/hr (300 KEAS); o @ load case the major designing that, for these two configurations, These analyses indicated encounter at VC.

would be the gust baseline aircraft analysis for the two Analysis - The wing structural 3.4.2 Fatigue A method for advanced design of their fatigue characteristics.

included consideration and cyclic operational usage, design stress, was employed which related fatigue estimation data for the 1972 upon passengers carried versus range to the fatigue life. Based allowables 0 and 926 km (0 and the average stage length between S. Domestic Regional Carriers, U.

passenger load factor.

km (225 n.mi.) with a 40% n.mi.) was calculated to be 417 the foregoing average stage a 30,000 hour fatigue life, average cruise speeds, and Considering fatigue allowable was 241,325 kN/sq.m (35000 psi) and passenger load factor, length estimated.

variations of the present insensitive to the parameter This fatigue allowable is relatively detail study, of course, would to be constant. Further study and, thereby, was assumed Gust Load and assumption since the show slight variations in this assessment undoubtedly load encounter.

lower the cyclic rate of maximum Alleviation System, in particular, should loading aircraft involved in this would still be small for the low wing The benefits, however, study.

were for the two baseline aircraft Effects on Aircraft Weight - The wing weight 3.4.3 in Table V in and the results are given Analysis Program" by the "Wing Weight analyzed of external loading the wing structure for a number section 3.1.3. This program synthesizes (e.g., maneuver, gust, taxi, etc.) along with constraints imposed by geometry, conditions in Section 3.1.3, the "Wing requirements. As indicated fatigue, stiffness and manufacturing that, verified the results of the weight estimation logic, so Weight Analysis Program" The program was the airplane without active controls.

modifications were not required for gust load alleviation degree of damping required from a also manipulated to determine the The application and formulation structural weight saving.

system to achieve the maximum is presented in System" weight estimation procedure of the "Gust Load Alleviation Section 4.2.

Airplanes Without Active Controls 3.5 Final Sizing of the parametric sizing routine were used to confirm or update The baseline airplane studies of airplanes: the sizing of the following array before proceeding with Turboprop Design 2-Engines 4-Engines 0.6 MACH 0.5 MACH Field Length Field Length m (ft) m (ft) No. Of 610 (2000) 914 (3000) 914 (3000) 1067 (3500) Passengers 457 (1500) O O O 44 O O O O O 0 100 0 O O O 148 O Mechanical Flap Turbofan field length.

2-engined, 0.7 Mach, 910 m (300 ft) 148 passengers, optimization of each of the The scope of the program did not permit absolutely precise by varying the cruise power airplanes; airplanes close to the optimum were obtained above setting until the design wing loading and thrust to weight ratios, optimized in Section 3.1.4, were closely matched.

Table X presents the principal characteristics of the turboprop airplanes sized without active controls. The turbofan data are presented in Section 6.2 # ENGINES 4 PAX 100 148 M 0.5 0.5 0.6 0.6 0.6 F.L. - FT.

1,500 2,000 3,000 1,500 '2,000 3,000 2,000 3,000 3,000 3,500 3,000 3,500 3,000 3,500 W/S T.O.

- PSF 32.7 44.0 71.1 32.7 44.0 71.0 44.0 71.0 58.8 71.0 58.8 71.0 58.8 71.0 T/WT.O.

.383 .325 .301 .383 .327 .300 .326 .298 .403 .405 .398 .406 .401 .402 'x 500 .765 .780 .802 .760 .765 .788 .750 .782 .755 .725 .745 .705 .725 .696 RGW - LB 61,004 49,055 42,670 121,369 94,644 80,921 130,163 110,217 49,078 47,026 95,276 90,892 132,329 125,506 OWE - LB 45,406 35,285 29,635 87,794 65,072 52,907 87,514 69,765 35,139 33,364 65,401 61,775 89,091 83,145 SLST - LB 6,285 4,299 3,458 12,527 8,324 6,549 .11,431 8,846 10,643 10,261 20,418 19,841 28,553 27,152 DOC-2 (500) 4.207 3.732 3.487 2.649 2.230 2.030 1.790 1.605 3.075 2.999 1.903 1.841 1.561 1.499 DOC-4 (500) 4.895 4.235 3.916 3.246 2.649 2.380 2.173 1.923 3.583 3.479 2.329 2.242 1.956 1.868 V.

DOC-2 (150) 6.414 5.617 5.189 4.128 3.426 3.066 2.769 2.442 4.857 4.716 3.068 2.955 2.540 2.425 m, CR.

ALT (DOC-2, 150) 15 15 15 15 15 15 15 15 15 15 15 15 15 0 KEAS (DOC-2, 150) 250 250 250 250 250 250 250 250 300 300 300 300 300 "DOC-4 (150) 7.510 6.398 5.823 5.073 4.068 3.575 3.352 2.890 5.675 5.478 3.735 3.587 3.154 2.994 CR. ALT (DOC-4, 150) 15 15 15 10 10 10 10 10 15 15 10 10 10 10 500 N.M.

FUEL - LB 4,913 3,575 3,042 9,751 6,822 5,696 9,264 7,665 3,645 3,443 6,957 6,557 9,571 8,936 CA 150 N.M.

FUEL (DOC-2) 2,352 1.671 1,353 4,678 3,202 2,537 4,346 3,420 1,762 1,641 3,368 3,132 4,639 4,269 WING AREA - FT 1,856 1,112 599 3,694 2,146 1,136 2,949 1.548 832 659 1,614 1,275 2,242 1,760 WING WEIGHT - LB 10,147 6,454 3,957 22,851 14,064 8,479 20,585 12,314 5,211 4,356 11,438 9,519 16,935 14,031 Z SURFACE CONTR. - LB 1,537 1,331 1,214 2,421 2,054 1,853 2,535 2,272 1,332 1,295 2,063 2,000 2,563 2,475 0 HYDRAULICS - LB.

512 444 405 807 685 618 845 757 444 432 688 667 854 825 . HOR. STAB. V .789 .727 .770 .826 .779 .843 .842 .933 .625 .658 .676 .716 .739 .790 HOR. STAB AREA - FT 729 312 131 1,477 616 257 848 357 173 129 349 259 493 367 0 HOR. STAB WT - LB 1,946 1,005 524 3,922 1,974 1,021 2,716 1,410 778 624 1,549 1,241 2,179 1,742 O VERT. STAB V .078 .069 .067 .112 .099 .097 .125 .123 .055 .056 .080 .082 .102 .104 VERT. STAB AREA - FT 533 218 84 1,486 582 219 922 351 114 82 308 219 504 357 AIRFRAME COST $M 2.8926 2.5633 2.3711 4.4232 2.8136 3.4699 4.6846 4.2272 2.7187 2.6517 4.0720 3.9569 5.0305 4.8634 TOTAL COST $I 3.8969 3.4494 3.1959 5.6837 4.9153 4.4879 5.9077 5.3512 3.3162 3.2418 4.8123 4.6903 5.8574 5.6766 AIRPLANES WITH MANUAL CONTROLS (COSTS REDUCED BY $47,460) AIRFRAME COST $M 2.8451 2.5158 2.3236 2.6712 2.6042 DOC-2 (500) 4.190 3.715 3.47 3.060 2.984 DOC-4 (500) 4.879 4.219 3.899 3.569 3.464 DOC-2 (150) 6.391 5.594 5.166 4.837 4.695 DOC-4 (150) 7.487 6.376 5.800 5.655 5.458 CONTROL SYSTEMS 4.0 ACTIVE recent origin but it is worth noting that Orville The term "Active Control Technology" is of A his work in automatic stabilization in 1914.

Wright was awarded the Collier Trophy for application of active controls was the Lockheed-Georgia Company's significant recent the C-5A in order to incorporation of an Active Lift Distribution Control System into wing mode frequency range. Other reduce wing bending loads in the short period and first include flutter control, envelope limiting, load potential uses of active control technology and ride quality control.

control, relaxed static stability to provide the in this study, systems have been developed for the baseline airplanes alternate capabilities: following a B737 aircraft.

(a) Passenger ride comfort equivalent to that of gusts.

(b) The above ride comfort plus structural load alleviation due to of a (c) The above ride quality and gust load alleviation capabilities, plus the use reduced static margin.

are described, their weights and costs defined arid Systems for each of the above capabilities for variations in wing loading, parametric weight and cost relationships determined Table XI summarizes the capability and passenger size and the other related parameters.

of the three systems.

redundancy standard 4.1 Ride Quality Control System control systems, together with the associated This section describes the ride quality parametric system weight and cost data.

ride quality control systems 4.1.1 Ride Quality Control System Synthesis. Longitudinal 3-degree of freedom digital computer program to have been synthesized utilizing a linear the two baseline airplanes to that estimated for a contemporary improve the ride comfort of in such as the B737. The system finally developed is shown schematically jet transport (44 and 15 together with values of the gains determined for the 215 and 287 kg/sq.m Figure feedback loops, ft) wing loading baseline airplanes. The system consists of two 58.8 lb/sq.

aft-segment flap the first being vertical acceleration at the center of gravity driving three at panels plus the aileron on each wing semi-span, and the second being vertical acceleration the flight station driving the elevators. The acceleration feedback to the wing surfaces provides Redundancy Capabi lity Designation Not Applicable No Active Controls Baseline surfaces and hydraulic only Multiple Ride Quality Control RQ actuators.

systems with individual (Ride Quality) - Two electronic channels FAIL SAFE third electronic Quality Control plus As for RQ plus GLA Ride and duplicated hydraulic Load Alleviation channel (Gust Load Alleviation) Gust surface - supplies to each FAIL OPERATIVE fourth electronic Ride Quality Control plus As for GLA plus AS and third hydraulic Gust Load Alleviation plus channel (Artificial Stability) control - Stability supply to pitch Artificial after two FAIL OPERATIVE identical failures.

and Redundancy Table XI System Designation, Capability A NZCOCKPIT O AIRPLANE DYNAMICS ANZC.

G.COMPARATO 6f .167$ +1 ACTUATOR 8e .167$ +1 K K W/S FLAPS (LB./SQ.FT) KG/SQ.M.

-.40 UP .023 215 (44) -1.0 .10 (44) DOWN 215 (44) .10 215 DOWN -1.0 (Ke & Kf UNITS ARE RAD./g) -. 40 UP .023 287 (58.8) -1.0 DOWN .07 (58.8) Figure 15 Ride Quality Control System levels. The addition of the flight ride smoothing but greatly increases the r.m.s. pitch rate vertical accelerations due to pitching station accelerometer permits the sensing of the to the wing. This signal is then used by the ride control surfaces on moments generated to the The acceleration feedback to counter these pitching moments.

drive the elevators synthesis of frequency range. The reduces r.m.s. g's in the short period pitch control also did not include a consideration of the system filtering necessary for the ride quality system would not that this consideration stability requirements. It was assumed system/structural or weight of the ride quality system.

have a significant effect on the cost pilot control during aircraft maneuvering To meet handling qualities requirements for good to the wing surfaces and cancel out the ride control commands it would be necessary to could be determined by a stick position function. The stick positiori function elevator with evaluated Other feedback parameters were an analysis of stick force per g characteristics.

surfaces and pitch angular such as angle of attack feedback to the flap/aileron lift control improvement in r.m.s.

In both cases there was no significant rate feedback to the elevator.

which were not However, for handling qualities requirements, acceleration levels.

likely be necessary to include the scope of this study, it would more than considered within loop in the system.

a pitch damping indicated an optimum gain for the cockpit acceleration feedback to the elevator, results For gain the amount of pitch rate r.m.s. pitch rate. Above or below the optimum attenuating effectiveness is significantly attenuation is reduced. In spite of the fact that the control flap to increase the ride control system gains for with flaps extended, it was necessary increased the Boeing 737 ride quality level.

the landing approach case in order to meet size required to match 737 ride Wing loading had a significant effect on the control flap kg/sq. m (44 psf) airplane required the surfaces to qualities. The wing loading of the 215 m (58.8 psf) wing loading airplane required have a 15% wing chord size while the 387 kg/sq.

wing lift control. Alternatively, 15% chord surfaces full span 10% chord surfaces for can be used for the higher wing loading case; this covering only part of the landing flap span method.

is probably the more efficient of 6 rad/sec with a no load rate The ride quality system synthesis used an actuator bandpass limit under loaded conditions would be around 40 limit of 60 deg/sec. This no load rate surface actuator bandpass and deflection rate limit have not deg/sec. The effects of control but studies performed for NASA Langley (Ref 8)indicate that the bandpass been evaluated by NASA rate limit used are adequate. Furthermore, preliminary data just released and with shows that even for higher speed jet aircraft most of the energy associated Langley aircraft vertical motion is at frequencies below 6 rad/sec.

acceleration and r.m.s. pitch rate for the two Figures 16 and 17 present r.m.s. vertical a ride quality system installed. Data are presented for baseline airplanes, with and without For both baseline airplanes the ride quality systems the cruise, descent and approach cases.

than the the ride in the descent case so that it was slightly better were designed to improve In doing this the cruise and approach ride estimated B737 r.m.s. vertical acceleration.

to equal the B737 estimate.

qualities were generally better than required System Design. The lift changes required for ride control are 4.1.2 Ride Quality Control and the aft automatically controlling the angular position of the aileron provided by design shown of the double slotted trailing edge flap as shown in Figure 18. The segment flap tracks and an aft consists of a forward flap segment which travels aft on conventional point 'A' to provide the landing flap position. For ride flap segment which pivots about flap segment pivots about point 'B' and is actuated by a hydraulic quality control the aft is which is attached to the flap carriage 'D'. The actuator piston-rod end, actuator 'C' this point being coincident, with no connected to the flap ride control operating lever at 'A', about which the aft segment rotates for normal ride control deflection, with the pivot point flap segment can be actuated for ride quality control in the landing flap operation. The aft landing flap position through ± 15 degrees in the design shown. The retracted or extended nose as arranged, provides the required chord (.15C) and allows the overhanging pivot point, of the surface to be used for mass balancing purposes.

this aft segment of the flap, used for ride The lowest wing loadings examined required edge flap. To avoid problems due to control, to extend the full span of the trailing segment on each wing is split into 3 equal span structural deflection under load this flap actuator and a single hydraulic supply. Each pair of flap sections, each operated by a single different hydraulic system so that failure of a single hydraulic segments is supplied by a flap sections on each wing available for ride quality system still leaves the aileron and two of the control flap and the number of control. As the wing loading increases the span the aileron is used. At high wing loadings, approaching sections can be reduced until only is unnecessary to meet the standards set for this that of the B737, ride quality control equipped with powered controls the elevator and aileron study. On airplanes servo-actuators are modified to include an electrical input signal to the valves in addition to the manual pilot input.

In airplanes with manual controls, such as the smaller airplanes in the matrix being studied, it is necessary to change to powered controls on introduction of a ride quality system. This the cost.

does not change the weight of the control systems but does increase 610 M (2000 FT) FIELD LENGTH W/S = 215 KG/SQ. M. (44 LB/SQ.FT.)

% AIRPLANE WITHOUT RQ 0) J u .12 CRITERIA SOa a 737

.08

> AIRPLANE WITH RQ 0 % 0 I S.04 1.74 M/SEC 2.5 M/SEC 3.0 M/SEC ou =5.7 FT./SEC.

ou = 8.2 FT./SEC.

au =9.8 FT./SEC.

I I I I I I I a FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION U 1.0 I, S0.8 I , 0.6 CRITERIA 0.4 - AIRPLANE WITHOUT RQ u AIRPLANE WITH RQ ---

K 0.2 737

Ln 737 FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION Figure 16 Ride Quality System Effects (610 m; 2000ft. F.L.)

914 M (3000 FT) FIELD LENGTH W/S = 287 KG/SQ.M. (58.8 LB/SQ.FT.)

.16 WITHOUT RO a I AIRPLANE .12 U . 1//I2 CRITERIA, i U -. 0 LU 0 I S. .-- . AIRPLANE WITH RQ S.04 3.0 M/SEC M/SEC 2.5 M/SEC 1.74 a := 9.8 FT./SEC.

a = 8.2 FT./SEC.

a = 5.7 FT./SEC.

U U U I A I I I II MID AFT AFT FWD FWD MID FWD MID AFT APPROACH DESCENT CRUISE LOCATION PASSENGER U LU 1.0 0.8 0.6 CRITERIA I 0.4 U AIRPLANE WITHOUT RO -0.2 ---- AIRPLANE WITH RQ 0 V. 0.2 0 737 0 I l II I I I I FWD MID AFT FWD MID AFT FWD MID AFT CRUISE DESCENT APPROACH PASSENGER LOCATION Figure 17 Ride Quality System Effects (914 m; 3000 ft.

F.L.)

TRAILING EDGE FLAP - RETRACTED WING STATION AT 0.5 SEMI-SPAN Q85C IO C 065 C FLAP SCREW JACK O D' FLAP CARRIAGE 'C'GUST ALLEVIATION ACTUATOR A , 25' WITH LVD TO SIGNAL POSITION 1 * FOR 25F LAP ACTION 'A'PIVOT 'B'PIVOT FOR 15'GUST ALLEVIATION TRAILING EDGE FLAP - EXTENDED Figure 18 Ride Quality and Gust Alleviation Control Flap The hydraulic power provided in the airplane is not increased, relative to the airplanes without the ride quality system, since it is not considered an essential system in an emergency condition; priority valves are included in the hydraulic system so that in an emergency or under exceptional demand conditions the ride quality system will momentarily stop functioning until priority services have been satisfied.

While failure of a single actuator or hydraulic system is considered to be acceptable, a false signal from the sensors to all actuators is considered not acceptable. To prevent such an occurrence the sensing and electronics components are completely duplicated and arranged so that their signals to the actuators are compared. If these signals do not agree with each other the ride quality control system is shut down. In the event of a hard-over of a single in the cockpit will alert the pilot to switch off control flap surface, position indicators power to that surface and its twin on the opposite wing.

4.1.3 Ride Quality Control System Weight and Cost - The components and systems affected by the incorporation of a Ride Quality (RQ) Control System include the control surfaces, control servo actuators, control system, hydraulic system and the electronics systems. The weights and costs of the two design point airplanes were estimated using the system design descriptions from Section 4.1.2; Table XII summarizes the data for one of the airplanes. Parametric curves summarizing the weight and cost effects for a range of airplanes are presented in Section 5.1. The following describes the logic utilized for each effects.

major component to formulate and quantify these Control Surfaces - The aileron and the second segment of the trailing edge flap act as the ride quality control surfaces. Since these surfaces are also used for maneuver (the ailerons) and during landing and take-off, the surface weight and cost is not expected to change to any appreciable degree.

Surface Controls - The surface controls effects include the addition of servo-actuators for the RQ segment of the trailing edge flaps and the addition of aileron and elevator servo-actuators with automatic electrical input capability in lieu of actuators without this capability. On the smaller aircraft (set arbitrarily at below 70 passengers) without RQ, the aileron and elevator will probably be operated by manual controls. Addition of RQ would require installation of full power aileron and elevator controls for these smaller aircraft. The weight difference between manual and powered controls was assessed to be negligible for this study; the cost difference, however, was evaluated and included in the cost estimation logic. For the 44 passenger baseline airplane, the cost of powered versus manual primary controls was estimated to be $42,000 before the inclusion of profit. For all sizes of airplanes, addition of electro-hydraulic valves to the aileron and elevator servo actuators cost an additional $18,500 per airplane.

610 M (2000 FT) BASELINE W/S - 215 KG/SQ.M.

(44 PSF) A WEIGHT LB KG A COST -$ R.F.A .

... ,CE ____ _ _) (35) (96,960) * POWERED AILERON AND ELEVATOR • 0 0 47,460 AILERON AND ELEV. E-H VALVES 15 6.8 62 28.2 49,500 CONTROL FLAP ACTUATION A HYDRAULIC SYSTEM (36) (16.3) (2,200) A ELECTRONICS (100) (45.3) (67,800) 2 AILERON/FLAP CHANNELS 30 13.6 2 ELEVATOR CHANNELS 30 13.6 ACCELEROMETERS AND INDIC.

10 4.5 WIRING AND MISC.

30 13.6 A TOTAL (213) (96.6) (166,960) * MANUAL CONTROLS IN BASELINE Table XII Weight and Cost - RQ System full wing loading requires the 215 kg/sq.m (44 psf) baseline airplane with The 44 passenger into 3 the surface is divided control. Assuming for ride quality the trailing edge flap span of and each was estimated to weigh wing side, 6 actuators would be required sections per servo actuator elevator electro-hydraulic The aileron and kg (9 Ib) and cost $4000.

4.1 per actuator; 5.4 kg (12 estimated to be 1.8 kg (4 Ib) for this baseline was weight increment was estimated to weight, therefore, airplane surface controls The baseline Ib) per airplane.

be (28 kg) (6 x 9.x 1.15) = 62 Ibs.

servo-actuators and installation (1) flap = 15 Ibs. (6.8 kg) modification (3 x 4x 1.25) and elevator servo-actuator (2) aileron kg/sq.m (44 psf) kg (77 Ibs.) per airplane for the 215 weight estimates total to 35 The above baseline.

loading, 44 passenger, wing of the degree of was based upon evaluation weight and cost data Parametric scaling of these upon the acceptability was based The basis for ride quality quality control required.

ride were is, rough air damping requirements B737 type airplane. That ride quality level of the for the B737; approximately wing loading relative to that predicated upon the airplane flap segments, The trim requirements for the aft to 576 kg/sq.m (100 to 118 psf).

relative to a 488 kg/sq.m, proportional to the airplane wing loading therefore, are inversely at adequate ride control alone will provide assuming that the aileron (100 psf) wing loading is expressed by the following (100 psf). This trim requirement 488 kg/sq.m proportionality: - .01) C No. of Actuators ACL/CLO (S/W hence the number of the required span, and was used to establish This relationship (100 with six actuators and 488 kg/sq.m using the 215 kg/sq.m (44 psf) baseline actuators, being and cost was evaluated as flap RQ actuator weight for zero actuators. The psf) of to the product which was in turn proportional the surface hinge moment proportional to controls aileron and elevator surface average chord. The gross weight and the wing airplane kg/sq.m (100 psf).

for wing loadings below 488 were assumed constant is composed of surface required for RQ control - The additional electronics Electronics the surfaces and two channels of electronics for two aileron and RQ flap position indicators, These items are considered and aircraft acceleration sensors.

for the elevator RQ operation, kg/sq.m (100 psf) and were below a wing loading of 488 be invariant with airplane size to The weight and cost for employing RQ control.

considered constant for airplanes to be as follows: electronics was estimated $ 2,400 o Surface Position Indicators 1.8 kg (4 Ibs.)

27 kg (60 Ibs.) $60,000 o Four Electronic Channels - o Accelerometers, Wiring, & Misc. 16 kg (36 lbs.)

kg (100 Ibs.) $62,400* TOTAL 45 * Profit not included.

was assumed to consist of - Additional hydraulic system weight for RQ control Hydraulics plumbing lines and the running between the primary control main flex-hoses by added priority valves. The flap actuators with flow being controlled second-segment for the baseline aircraft and related to the weights of these components were estimated wing geometric parameters. The cost of the additional hydraulics number of actuators and the sizing program's hydraulic cost estimating relationship.

was evaluated using System 4.2 Gust Load Alleviation (GLA) System, together with the This section describes the Gust Load Alleviation associated parametric system weight and cost data.

Systems were synthesized for the baseline 4.2.1 Gust Load Alleviation System Synthesis.

to be equivalent in criticality to the airplanes to reduce the gust loading conditions in Section 3.4, the following requirements maneuver cases. From the analyses described were obtained for the system synthesis.

W/S Gust 215 kg/sq.m 287 kg/sq.m 15 mps (50 fps) Vertical (44 PSF) (59 PSF) kg 21,092 20,965 (46,500) (46,220) G. Wt. (Lb.)

kN/sq.m 10.1 14.56 q (PSF) (211.1) (304.1) m 4,572 4.572 Alt. (ft) (15,000) (15,000) Ag Required from System 1.5 ±1.5 .177 .191 CL TRIM W/s kg/sq. m 215 kg/sq.m 287 fps) Vertical Gust 20 mps (66 (59 PSF) (44 PSF) 20,965 21,092 kg (46,500) (46,220) Wt (Lb.)

6.09 4.14 kN/sq.m (86.5) (127.1) q (PSF) (0) (0) Alt. (ft) ±.9 ±11.0 from System Ag Required .422 .465 CL TRIM in Section 4.1.1 quality control system described synthesis show that the ride Results of the system will meet the gust load alleviation the W/S = 215 kg/sq.m (44 psf) airplane for to a is increased 3.0 degrees surface deflection capability if the wing control requirement along with the hinge moment requirement of 18.0 degrees. There is no increased maximum psf) airplane however the the W/S = 287 kg/sq m (58.8 surface deflection. For increased in therefore had to be increased meet the gust load needs and ride quality system will not is used for in span if 15% chord gust flap to 15% chord or extended chordwise length of the those are significantly larger than For this case, the hinge moments the ride quality system.

to 18.0 deflection had to be increased system. The maximum surface for the ride quality in the ride from the values used gains are not changed flap/aileron and elevator degrees. The For the W/S = 287 kg/sq.m kg/sq.m (44 psf) airplane.

system for the W/S = 215 quality would not determine gain the ride quality requirements psf) airplane however (58.8 on the a higher requirement alleviation system places because the gust load magnitude feedback gains.

to the ride The system is basically similar Alleviation System Design.

4.2.2 Gust Load integrity of However since the structural described in Section 4.1.2.

quality control system to reconsider the of this system it is necessary is dependent on the operation the airplane failure cases involved.

splitting of the surfaces and the use of single actuators for the flap The general arrangement, acceptable, since in as for the ride quality system. Single actuators are sections are retained gust conditions, the flaps can be the unlikely event of a surface or actuator failure in severe and the airplane speed reduced. However it is extended to the approach condition to provide duplicated electro-hydraulic valves and duplicated hydraulic considered necessary actuator since these components are more likely to fail than the actuators.

supplies to each the of a hydraulic system, a shuttle valve, located at the actuator, senses On failure off the ports from differential pressure between the two systems and automatically closes failed system. Only one of the two electro-hydraulic control- valves is used during the Awill be the gust lnoad alleviatin system. If, this control valve fails it normal operation of signal to the comparison of the surface position feedback signal and the input detected by will and/or comparison with signals from the other surfaces. A discrepancy valve, automatically signal a solenoid-operated shuttle-valve to redirect the hydraulic supplies through the standby control valve.

would momentarily stop It was accepted in the previous section that the ride quality system exceptionally high hydraulic system demand by the priority systems.

functioning under it is itself a priority situation is not acceptable for the gust load alleviation system since This system.

It is therefore necessary to increase the hydraulic power system and the distribution pipes feeding the flight controls and flaps along the wing rear spar.

and associated wiring have been incorporated, making a An additional electronics channel channel, comparison of the 3 channels automatically total of 3 channels. On failure of one the faulty channel.

identifies and disconnects Arguments can be advanced for over-designing the total system to avoid reducing speed after of an actuator, and for further redundancy to cater for double failures (e.g., failure failure systems); it was decided to take an optimistic approach for the purposes of two electronic of this study.

4.2.3 Gust Load Alleviation System Weight and Cost From the structural analyses in Section 3.4 and the system description given by Section 4.2.2, a computerized estimation technique was developed for determining the GLA system parametric and weight relationships. The cost estimating relationships are similar to those the added actuators, RQ system and are evaluated on a weight basis for employed for the components, increased additional electronic surface controls system modifications, for weight. Weight and cost data and decreased wing structural hydraulic system capacity, a function in Table XIII while variations as point airplanes are summarized one of the design size parameters are given in Section 5.2.

of airplane study consists of a set of technique developed for this The GLA computerized estimation includes estimation of: logic in the sizing program which o Gust load factors at VB and VC increment due to gust loads o Wing box weight GLA system lift coefficients required from the o Trim gust flap for use with the aileron of the required size of a o Determination actuators; for increments for the gust flap, aileron, and elevator o System weight and components; and for the additional the additional hydraulic capacity electronic components.

of for the GLA system is based upon analysis incorporated in the sizing program The logic which follow.

described in the paragraphs design basepoints and parametric correlation the Gust Load Factor Estimation o NG = 1 +ANG + NG = KG UDE VE CL /(498 x W/S) o (5.3 wa CL C + 2 W/S) where KG = 1.76 W/S / a (fps) Velocity = Gust UDE = +50 at Vc; ± 66 at VB (knots) Airspeed = Equivalent VE = V or VB C 610 M (2000 FT) BASELINE W/S - 215 KG/SQ.M.

(44 PSF) AWEIGHT LB KG A COST -$ A SURFACE CONTROLS (137) (62.1) (122,910) * POWERED AILERON AND ELEVATOR 0 0 , 47,460 AILERON AND ELEV. E-H VALVES 20 9.0 CONTROL FLAP ACTUATION 32.7 75,450 SHUTTLE VALVES AND MISC. 45 20.4 A HYDRAULIC SYSTEM (204) (92.5) (12,240) PUMPS 34 15.4 POWER TRANSFER UNITS 13 5.9 RESERVOIRS 12.2 PLUMBING 130 59.0 A ELECTRONICS (155) (70.3) (90,400) 3 AILERON/FLAP CHANNELS 45 20.4 3 ELEVATOR CHANNELS 45 20.4 ACCELEROMETERS AND INDIC.

10 4.5 WIRING AND MISC.

55 25.0 A TOTAL (496) (225.0) (225,550) * MANUAL CONTROLS IN BASELINE and Cost - GLA System Table XIII Weight 23, 134 KG 44 PAX AR = 8 RWG = 51,000 LB W/S - PSF (KG/SQ.M.)

WITH GUST NO GUST WING WING WING WEIGHT WEIGHT WEIGHT CHANGE MAX MIN NG LB (KG) LB (KG) LB (KG) 44 40.25 3.52 6,577 6,080 - 497 (215) (197) (2983) (2758) (-225) 59 53.8 3.04 5,148 4,914 -234 (287) (263) (2335) (2229) (-106) 64.75 2.77 4,415 4,309 -106 (347) (316) (2002) (1955) (-48) 100 91.2 2.35 3,454 3,451 - 3 (488) (445) '(1567) (1565) (-1.36) Table XIV Wing Weight Change due to Gust Loading Cruise Speed VC = VB = Speed with max gust at CL MAX.

W/S = minimum wing loading (psf) altitude (pcf) wa = air density at ft. used in the logic where wa = 0.0481 pcf) (15,000 radian) slope (per lift curve CLa= C = wing average chord (ft) VB definition) for = 1.4 (Assumed CL MAX Change Wing Box Weight was used to analyze various wing parameters to arrive The "Wing Weight Analysis Program" structural weight reduction with GLA. In preparing at gust load factors and potential wing from 215 to estimate, four aircraft were analyzed with wing loading variation this weight condition. Wing loading 488 kg/sq. m (44 to 100 psf) at the maximum gross was used since wing loading is the major influencing variation at constant aspect ratio on the gust loads. The four configurations analyzed were assumed to have the parameter was analyzed same gross weight with wing loading changes obtained by wing area and each due to gust. The results of the with and without gust loads to determine the weight penalty are shown in Table XIV. Examination of these "Wing Weight Analysis" for these cases analyses indicated that a 2g gust load factor ( A NG = 1) produced equivalent loads to those with a 2.5 g maneuver load factor; therefore, a criteria was selected based upon A NG indicate gust criticality. As was indicated in Section 3.1.3, the being greater than 1.0 to parametric wing weight relationship predicts fairly accurate weights for a gust critical wing for the wing at low wing loadings. A parametric relationship, therefore, was developed to the weight from the basic weight change with GLA and applied as a weight reduction wing weight equation.

Gust Flap Size Derivation in the GLA computerized technique by determining the maximum The gust flap was sized of A CL trim required. This required trim was then compared with allowable trim amount available from the aileron to determine the trim required from the aft segment of the trailing edge flaps. With the gust flap required trim identified, the inboard end of the gust flaps was derived for a 15 percent chord gust flap deflected 18 degrees. With this information, the gust flap hinge moment was determined for use in sizing the actuators and other system components.

System Weight Estimates For the baseline aircraft in Table XIV, the gust flap was determined to run the full span of the trailing edge flaps. Systems analysis yielded the following for the baseline configuration: (a) Surface Controls,+ 62 kg (+137 lb.)

o Flap Actuators & Installation (6) = 32.7 kg (72 lb.)

o Aileron & Elevator Actuator Mod. = 6.8 kg (15 lb.)

o Shuttle Valves (9) = 12.2 kg (27 lb.)

o Miscellaneous = 10.4 kg (23 lb.)

(b) Electronics, + 70 kg (+155 lb.)

o Accelerometers and Surface Position Indicators = 4.5 kg (10 lb.)

o Three Electronic Channels = 40.8 kg (90 lb.)

o Wiring and Installation = 25 kg (55 lb.)

(c) Hydraulics, + 92.5 kg (204 lb.)

o Hydraulic Pump Increase = 15.4 kg (34 lb.)

o Power Transfer Units = 5.9 kg (13 lb.)

o Reservoirs = 12.2 kg (27 lb.)

o Main Line Plumbing and Fluid = 23.1 kg (51 lb.)

o Gust Flap Plumbing and Fluid = 10.4 kg (23 lb.)

o Hoses and Valves at Actuators = 25.4 kg (56 lb.)

The above baseline represented a gust flap hinge moment of 45,920 M-N (33,879 ft. - Ib) per aircraft and added hydraulic flow rate amounting to 0.009 cu.m/min. (2.7 gpm) to each of the six gust flap actuators. Since the gust flaps will be primary controls for GLA, added baseline analysis, parametric is required. From this capacity and redundancy hydraulic incorporated into the sizing were derived for each system component and relationships program.

Costs the RQ system, the costing of the actuators is similar to The arrangement of the system and have electro-hydraulic valves of actuators may vary. Additional however size and number which for actuator installation control flap, aileron.and elevator been incorporated at the the baseline amounts to: = $16,500 for control flaps @ $2750 6 E-H valves $4500 = 9,000 2 E-H valves for aileron @ $5500 = 5,500 E-H valve for pitch control @ = 4,000 13% profit $35,000 TOTAL of $80,000 + 13% profit - amounts to $20,000 for a total The additional electronic channel $90,400.

is calculated within the and the installation of all systems The increased hydraulic system individual costs.

cannot be separated readily into sizing program costing routine and System 4.3 Artificial Stability control, structural gust load the system which provides ride quality This section describes margin, which results in smaller the relaxation of the static stability alleviation and permits baseline airplanes.

stabilizers than required for the horizontal static margin it on airplane size and cost of this reduced In order to determine the effects automatic horizontal stabilizer in the parametric sizing program an was necessary to include Section 4.3.1 describes the routine static margin as a variable.

sizing routine which included designs described in airplanes in this report except the initial developed and used for all the Section 3.1.

Section 4.3.2 describes the system design while Section 4.3.3 discusses the weights and costs associated with the system.

4.3.1 Horizontal Stabilizer Sizing - The horizontal stabilizers were sized to perform three major functions: 1. Provide trim capability throughout the operational envelope.

2. Provide sufficient longitudinal control capability during commanded (plus inadvertent) speed and altitude excursions in smooth and turbulent air to meet operational requirements.

3.

Provide sufficient longitudinal stability throughout the operational envelope for pilot control of the aircraft during maneuvers and in steady flight.

The critical conditions for these functions for horizontal stabilizer sizing are: 1. (a) Trim at the landing approach speed with most forward center of gravity.

(b) Flare in Ground effect during the landing maneuver.

2. Sufficient pitch acceleration capability at the most forward center of gravity so that all necessary maneuvers, including go-around, can be performed during landing.

3. Adequate static margin at the most aft center of gravity during landing or high speed, low altitude cruise.

A trimmable incidence horizontal stabilizer was used in conjunction with a 30 percent chord elevator to provide the required trim and control capabilities. Horizontal tail sizing to provide these capabilities requires definition of the following parameters: coefficient = tail off pitching moment CMO R = required center of gravity range = minimum allowable static margin elevator deflection) tail lift, coefficient (a function of CLTMAX = max speed at a given = lift coefficient CL slope lift curve = tail = a CLaT curve slope = wing lift CLa = a de = rate of change of downwash with incidence = radius of gyration in pitch kyy constant = gravitational g C = wing mean aerodynamic chord pitching A6 = change in angular horizontal stabilizer by the primary parameters involved in sizing the Figure 19 illustrates (VH) versus C.G. position. The solid sloping line presenting tail volume coefficient (VH) required for neutral point-tail off) defines the tail size originating at hN (neutral other solid line, originating at CMO/CL (trim stability as the C.G moves aft from "hN". The remain trimmed as the C.G is the tail size for the airplane to position - tail off) defines intersect each other identifies the The point where the two solid lines moved forward.

trimmed and neutrally stable. This C.G location for the airplane to be both tail-size and the C.G travel the tail size must no C.G travel is available. To provide is a point and therefore line" provides an further. The dashed line identified as "stability be increased still neutral stability line, while margin ( 8 CM/ 8 CL) relative to the acceptable positive static control power for maneuver (K 6 ) relative the dashed line identified as "trim line" provides lines that the point of intersection of the two dashed to the basic trim requirement. Note be a range of C.G locations (R) which can increases the tail size required. To provide increased; the larger the required C.G range, the trimmed and stable the tail size must be larger the required tail size.

the line" is a function of lift curve slopes of the wing and tail and The slope of the "stability systems increase the slope which results in a larger tail being wing downwash. Powered lift and The slope of the "trim line" is a function of the wing required for a given C.G. range.

lift coefficient available from the tail (flying-tail, inverted tail lift coefficients; the higher the size leading edge, etc.) the lower the slope and hence the smaller the tail camber, slotted required for a given C.G range.

ITRIM LINE S- STABiLiTY LINE

- ~

8CM/8 CL

I"

ARTIFICIAL STABILITY

v

H

hN K C MO/C L C.G. - % MAC Figure 19 Horizontal Tail Sizing which is "static margin" ( SCM/ o CL) to the present study is the Of primary importance by increasing the tail value of 3% MAC and is provided in the figure to have a positive shown of active controls a negative stability requirement. By the use size relative to the neutral line in the position identified as be accepted, which places the stability stability margin can the tail size for a given C.G range which stability." This has the effect of reducing "artificial due to active controls.

weight saving and possibly a cost saving results in a and incorporated routine which included the above parameters was developed A computer into the main sizing program. Thus by the input of the correct value of 8 CM/ 8 CL, the effect of these static margin could be modified, the horizontal stabilizer resized and the was changes on the complete airplane automatically computed. A similar type of routine also developed for sizing the vertical stabilizer.

required to cover all maneuvers, over and above trim, has been shown in Ref.

The A " level 9 to be a function of the desired landing field length. The real parameter involved is the and set of approach speed which can be shown to be a single value for a given field length from the control will also be a function of the landing criteria. The level of A Y required and flare to touchdown. Thus the actual value technique used for controlling the approach and will depend upon the glide path control of A Y was chosen for each configuration characteristics and requirements for that configuration. The suggested levels recommended the A" level.

in Refs. 10, 11, 12 and 13 were used as guides in selecting is identified as an "Artificial 4.3.2 Artificial Stability System Design - The system It is in fact a combined gust load alleviation system Stability" (AS) system for convenience.

system. The gust load alleviation system is identical to that described and artificial stability same redundancy considerations. This system is modified to in Section 4.2 and meets the plane by the addition of a third electro-hydraulic provide artificial stability in the pitching system, and miscellaneous switching logic valve, and a fourth electronic channel in the pitch these additional components and additional sensors. The artificial stability system utilizes with the actuators, electro-hydraulic valves, electronics, sensors and wiring provided for the gust load alleviation system.

provided permits the failure of any two critical components while The redundancy thus providing a fully functioning artificial stability system. The reduction in hinge moment if two actuators fail is probably acceptable as an emergency condition.

components for this 4.3.3. Artificial Stability System Weights and Costs - The additional do not vary with wing loading or aircraft size. These components weigh 13.6 kg system (30 lb.) and add an additional $39,550 to the aircraft cost. The effects of the gust load weight are similar to those alleviation portion of the system on wing weight and system 4.2.3.

described in Section 5.0 TURBOPROP AIRPLANE CHARACTERISTICS WITH ACTIVE CONTROLS The matrix of airplanes defined and sized without active controls in Section 3.5 were resized with the effects of the three levels of active control systems included. This section describes the characteristics of these airplanes while the following section, 6.0, compares the effects of the various systems.

5.1 Characteristics of Airplanes Incorporating Ride Quality Systems This system increases the weight and cost of the airplanes and slightly improves the fatigue life. Large wing weight savings are not expected and it is unlikely that the optimum aspect ratio would differ because of the introduction of this system. The airplanes sized with ride systems therefore retain the aspect ratios, wing loadings and thrust to weight ratios quality without active controls.

of the airplanes sized Table XV presents the primary size and economic characteristics of the matrix of airplanes.

Included for convenience in the lower portion of this table are weight and cost changes relative to the airplanes without active controls. These data are discussed in Section 6.0.

Figure 20 illustrates the variation of the ride quality system weight with wing loading for the 4-engined, 100 passenger size. Note that the weights for the individual systems, e.g., surface controls, are the additional weights due to the ride quality system. Figure 21 presents similar data to illustrate the variation of system weight with passenger size. The example shown is for the 610 m (2000 ft.) field length; data for the other field lengths are included in Table XV.

Figure 22 presents system cost variations with wing loading and passenger size. The steps in the "surface controls" and "total RQ system" curves at 70 passengers are due to the necessity of incorporating powered aileron & elevator controls into the basic airplane. The weight and cost trends are as expected, increasing with reduction of wing loading and increase of passenger size except for the electronics values which remain constant.

5.2 Characteristics of Airplanes Incorporating Gust Load Alleviation Systems Since the gust load alleviation system will reduce the wing box weight, it can favor the use of higher wing aspect ratios. The degree of change in aspect ratio will vary with field length and it was therefore necessary with this system to reoptimize the aspect ratio for each of the AR= 8 # ENGINES 4 2 100 148 44 100 148 # PAX 44 0.6 0.6 M 0.5 0.5 0.5 0.6 1,500 2,000 3,000 2,000 3,000 3,000 3,500 3,000 3,500 3,000 3,500 F.L. - FT. 1,500 2,000 3,000 58.8 71.0 58.8 71.0 58.8 71.0 W/S T.O. -PSF 32.7 44.0 71.1 32.7 44.0 71.0 44.0 71.0 .398 .405 .401 .402 .300 .383 .326 .300 .326 .298 .403 .405 T/WT. O. .382 .325 .725 .705 .765 .780 .802 .760 .765 .788 .750 .782 .755 .725 .745 .705 '7500 132,736 125,812 61,778 49,459 42,903 122,534 95,163 81,180 130,765 110,493 49,379 47,275 95,637 91,150 RGW - LB 88,067 70,022 35,419 33,597 65,736 62,029 89,468 83,430 OWE - LB 46,105 35,657 29,852 88,841 65.550 53,149 3,470 12,639 8,361 6,563 11,475 8,862 10,696 10,295 20,472 19,877 28,617 27,197 SLST - LB 6,356 4,326 2.688 2.256 2.048 1.808 1.617 3.114 3.034 1.922 1.857 1.573 1.510 DOC-2 (500) 4.280 3.785 3.527 2.259 1.969 1.879 4.977 4.292 3.958 3.291 2.677 2.399 2.193 1.936 3.624 3.516 2.348 DOC-4 (500) 2.796 2.459 4.916 4.769 3.095 2.979 2.559 2.441 DOC-2 (150) 6.522 5.694 5.246 4.186 3.463 3.092 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 15,000 CR. ALT (DOC-2, 150) 250 250 250 250 250. 300 300 300 300 300 300 KEAS (DOC-2, 150) 250 250 250 2.909 5.738 5.534 3.764 3.612 3.175 3.012 DOC-4 (150) . 7.631 6.480 5.883 5.14T 4.109 3.602 3.382 10,000 10,000 10,000 10,000 CR. ALT (DOC-4, 150) 15,000 15,000 15,000 10,000 10,000 10,000 10,000 10,000 15,000 15,000 6,571 9,594 8,952 N.M. FUEL - LB 4,970 3,599 3,055 9,840 6,854 5,710 9,301 7,680 3,661 3,455 5,977 2,543 4,363 3,427 1,770 1,647 3,378 3,138 4,650 4,277 150 N.M. FUEL(DOC-2) 2,379 1,682 1,358 4,720 3,217 1,879 1,120 602 3,728 2,157 1,139 2,962 1,551 836 662 1,620 1.278 2,249 1,764 WING AREA - FT 8,509 20,694 12,347 5,246 4,381 11,487 9,549 16,994 14,069 WING WT. - LB. 10,296 6,514 3,980 23,107 14,152 2,649 2,173 1,896 2,678 2,320 1,383 1,330 2,132 2,045 2,645 2,527 SURFACE CONTR. - LB 1,687 1,414 1,249 778 468 450 716 687 885 847 HYDRAULICS - LB 574 480 423 888 729 638 895 .789 .769 .825 .778 .843 .842 .933 .624 .657 .675 .715 .739 HORIZ. STAB. V .787 .726 358 174 129 350 260 495 367 HORIZ. STAB. AREA - FT 740.8 315 131 620 258 853 1,015 527 3,967 1,986 1,024 2,730 1,414 783 627 1,556 1,245 2,186 1,747 HORIZ. STAB. WT. - LB 1,976 .112 .099 .097 .125 .123 .055 .056 .080 .082 .102 .104 VERT. STAB. V .078 .069 .067 939 352 115 82 310 220 506 358 VERT. STAB. AREA - FT 544 221 328 1,509 587 220 4.9714 4.8287 4.3337 2.8316 2.7563 4.1891 4.0630 5.1509 - $M 3.0540 2.6920 2.4746 4.6070 3.9514 3.5751 A/F COST 4.9301 4.7968 5.9783 5.7850 3.3004 5.8712 5.0546 4.5938 6.0533 5.4583 3.4299 3.3411 TOTAL COST - SM 4.0620 3.5799 143 48 51 35 69 45 82 - LB 150 83 35 228 119 43 WSCR 32 64 41 77 48 76 31 214 111 39 135 45 46 WSCRA - LB 138 36,090 29,395 40,222 31,580 43,140 33,156 CSCRA - S 72,624 49,493 29,166 89,844 57,559 31,188 62,959 32,544 100 100 100 100 100 100 100 100 100 100 100 100 100 100 WAVR - LB 67,800 67,800 67,800 67,800 67,800 67,800 67,800 CAVR - $ 67,800 67,800 67,800 67,800 67,800 67,800 67,800 28 20 30 22 62 36 18 81 44 20 50 21 24 18 WHR - LB 2.52 4.15 3.94 2.88 2.68 2.39 2.22 5.58 5.02 4.51 4.16 3.61 3.63 3.01 A A/F COST - % .1065 .1209 .1084 .1975 .1393 .1059 .1456 .1071 .1137 .1053 .1178 STOTAL COST .1651 .1305 .1045 242 553 257 280 233 335 254 377 285 A OWE - LB 699 372 217 1,047 478 361 258 407 306 A RGW - LB 774 404 233 1,165 519 259 602 276 301 249 1.0 .75 1.27 1.17 1.0 .81 .9 .75 % DOC-2 (500) 1.7 1.4 1.15 1.41 1.17 .89 .86 .76 .72 .6 1.68 1.35 1.07 1.39 1.06 .8 .92 .68 1.14 1.06 % DOC-4 (500) 0.96 0.7 1.2 1.12 0.88 0.81 0.75 0.66 % DOC-2 (150) 1.68 1.37 1.1 1.4 1.08 0.85 0.70 0.67 0.60 1.03 1.34 1.01 0.76 0.89 0.66 1.11 .102 0.78 % DOC-4 (150) 1.61 1.28 0.25 0.40 0.20 0.44 0.35 0.29 0.21 0.24 0.18 % 500 NM FUEL 1.16 0.67 0.43 0.91 0.47 0.39 0.20 0.45 0.37 0.30 0.19 0.24 0.19 % 150 NM FUEL 1.15 0.66 0.37 0.90 0.47 0.24 0.51 0.41 0.42 0.34 0.73 1.19 0.73 0.46 0.63 0.37 0.80 0.70 % OWE CHANGE 1.54 1.05 CONTROLS RELATIVE TO MANUAL 5.9 5.7 7.22 6.87 6.37 % AF COST 1.76 1.68 1.88 1.64 % DOC-2 (500) 2.15 1.5 1.54 % DOC-4 (500) 2.01 1.73 1.51 1.63 1.58 1.79 1.55 % DOC-2 (150) 2.05 1.47 1.39 1.63 1.43 % DOC-4 (150) 1.92 IS PAGE ORIGINAL OF POOR QUALITY Table XV Characteristics of Airplanes with RQ System 100 LB 100 PAX 4-ENGINES 0.5M 100 KG 1.5 1.0 , ) WEIGHT

05 1

HYDRAULICS 0 30 40 50 60 70 LB/SQ.FT 300 .350 KG/SQ.M.

200 250 WING LOADING (T.O.)

Figure 20 RQ System Weights vs. Wing Loading 100 LB 4-ENGINES W/S = 44 LB/SQ.FT (215 KG/SQ.M) 0.5M 100 KG 1.25 1.0 WEIGHT 0.75 SURC.jCNRCL 0.5 ELECTRONICS 0.25 HYDRAULICS I I I a 0 40 NUMBER OF PASSENGERS Figure 21 RQ System Weights vs. Number of Passengers 0.5M 4-ENGINES = 44 PSF 100 PAX 160 WING OADING 215 KG/S. M 120 S ,- CO ST $1000 E80 LECTRONICS ELECTRONICS CO ,CE SURF CO HYDRAULICS HYDRAULICS 70 PSF 40 50 60 120 140 30 60 80 100 200 250 300 NUMBER OF PASSENGERS KG/S.M.

LOADING WING Figure 22 RQ System Costs (Scaled) 2- and 4-engine turboprop configurations and for the turbo-fan design. This reoptimization of aspect ratios was conducted for the 100 passenger size for each field length. From consideration of the initial parametric data discussed in Section 3.1.4, it was determined that the optimum altitude and speed would probably remain as 4570 m (15,000 ft.) and maximum EAS for the 278 km (150 n.mi.) stage length. The reoptimized aspect ratios were therefore selected for minimum DOC-2 at these conditions. Table XVI presents the size and economic characteristics of the 100 passenger airplanes for each field length and the selected ranges of aspect ratios. The best value of each fuel and economic characteristic from the range of aspect ratios is underlined for each field length in the table and from these data the optimum aspect ratios were selected. From the table it can be seen that in some cases it was necessary to make compromises in selecting the overall optimum. The following aspect ratios were selected for sizing the airplanes with the gust load alleviation system installed.

4-engine configurations Field length - m (Ft) 457 (1500) 610 (2000) 914 (300) 1067(3500) Aspect ratio 8 8 12 14 2-engine configurations Field length - m (Ft) 914 (3000 1067 (3500) Aspect ratio 8 10 The scope of the program did not permit the reoptimization of the aspect ratio for each of the three passenger sizes; it was therefore assumed that the values obtained for the 100 passenger size would be close to optimum for the other two sizes.

The characteristics of the matrix of airplanes, incorporating gust load alleviation are presented in Table XVII. For convenience, the lower portion of this table contains comparison type data which are discussed in Section 6.0.

Figure 23 presents an example of the variation of gust load alleviation subsystem weights as a function of wing loading and passenger size. Similar data in terms of initial A airframe cost as a function of wing loading and passenger size are presented in Figure 24. Equivalent weight and cost data for the complete matrix of passenger sizes and field lengths are provided in Table XVII. Note the sharp increase in weight for the hydraulic system relative to that required for the RQ system. This is due to the provision of additional hydraulic - -0 0000 (NO Q 0.0( (N (N . 1 (N 0 '0 N (' ( ~00 N (N (N NN N( (N00 (N '0 M0 '00 a ahoo M ,01 0 0 0 M N ( (- ( 0 . . - 0 O W ,( O(N~'0' vl ( v , 2 TN 7 '0 - - C C41 W '0Cv- c~~~ '#1 - " ONO c! o 6n 8; (N 8f (( -I CN( - I 2 n ev, gj Nq j M 2 8 % - c4 n 0n V; :9 N 1 Q % (4 ,4 cc N k m M c4 c4 m v N 1 c4 7 '00I o ' 'o ol mN v0 .4N (N 6 (c (N(N ( ( ( N M ( ( c " " ; O' '0 (N( 00 0 ( NI ORIGINAL PAGE IS In(N . 0 I4 (N ' 00 0 Z o -J . . . . -J -OP .. J ~~a N .. N In - n - ~I o z 0 0Z 0 Table XVI Optimization of Aspect Ratio with GLA System # ENGINES 4 100 148 PAX 44 100 148 44 # 0.6 0.6 M 0.5 0.5 0.5 0.6 F.L. - FT 1,500 2,000 3,000 1,500 2,000 3,000 1,500 2,000 3,000 3,000 3,500 3,000 3,500 3,000 3,500 W/S T.O. - PSF 32.0 44.0 66.0 32.0 44.0 66.0 32.0 44.0 66.0 58.8 71.0 58.8 71.0 58.8 71.0 .324 .264 .400 .397 .400 .397 .400 .397 T/W T.O. .368 .325 .265 .369 .325 .264 .369 " 500 .806 .782 .818 .80 .770 .804 .790 .756 .780 .761 .713 .742 .693 .729 .678 RGW - LB 58,556 48,832 43,177 113,695 93,042 81,593 157,958 127,297 111,120 48,987 47,604 94,747 91,950 130,608 126,637 OWE - LB 43,204 35,083 30,445 80,937 63,626 54,179 110,701 84,920 71,571 35,067 34,021 64,905 62,773 87,521 84,42 SLST - LB 5,799 4,272 3,082 11,280 8,151 5,808 15,693 11,123 7,907 10,552 10,165 20,411 19,642 28,106 27,044 DOC-2 (500) 4.192 3.783 3.535 2.576 2.235 2.039 2.106 1.783 1.600 3.119 3.047 1.918 1.863 1.561 1.513 DOC-4 (500) 4.854 4.284 3.934 3.138 2.648 2.364 2.629 2.158 1.891 3.625 3.519 2.342 2.257 1.952 1.876 DOC-2 (150) 6.371 5.685 5.288 4.00 3.426 3.097 3.297 2.754 2.456 4.921 4.802 3.088 2.995 2.538 2.455 500 NM FUEL - LB 4,729 3,560 2,824 9,154 6,714 5,264 12,632 9,067 7,009 3,633 3,385 6,929 6,447 9,455 8,789 150 NM FUEL - LB 2,263 1,665 1,304 4,391 3,151 2,431 6,055 4,263 3,257 1,756 1.639 336A1 3,127 4,591 4,274 % A 500 FUEL -3.75 -0.42 -7.17 -6.12 -1.58 -7.58 -2.13 -8.56 -0.33 -1.68 -0.40 -1.68 -1.21 -1.65 -3.78 -0.36 -3.62 -6.14 -1.59 -4.18 -1.91 -4.77 -0.34 -0.12 -0.2 -0.16 -1.03 +0.12 % A 150 FUEL WING AREA - FT 1,823 1,107 653 3;535 2,109 1,232 4,920 2.883 1.679 830 688 1,604 1,289 2,213 1,776 WING WEIGHT - LB 8,545 5,877 4,779 18,534 12,514 10,111 27,413 18,163 14,614 4,855 4,624 10,520 10,026 15,408 14,664 SURFACE CONTR - LB 1,640 1,453 1,329 2,483 2,177 1,985 3,050 2,653 2,415 1,442 1,408 2,184 2,134 2,677 2,616 HYDRAULICS - LB 776 616 510 1,320 1,013 825 1,733 1,306 1,055 556 525 90 854 1,167 1,096 .848 1.031 .625 .713 .678 .782 .742 .864 HOR. STAB. V .779 .728 .837 .822 .782 .927 .870 HOR. STAB. AREA - FT 701 310 132 1,375 603 261 1,886 826 364 173 127.2 346 258 485.4 363.6 HOR. STAB. WT. - LB 1,871 1,000 530 3,657 1,934 1,034 5,037 2,648 1,433 776 621 1,539 1,240 2,147 1,737 VERT. STAB. V .076 .069 .062 .109 .099 .090 .136 .125 .114 .055 .056 .080 .082 .102 .103 VERT. STAB. AREA - FT 506 217 109 1,354 567 283 2,187 902 448 114 93.2 305 250 494 403 AIRFRAME COST $M 2.9986 2.7225 2.5538 4.4211 3.9395 3.6622 5.4328 4.7839 4.4284 2.8740 2.8272 4.2136 4.1343 5.1453 5.0506 3.4156 4.9538 4.8702 5.9679 5.8628 COST $M 3.9766 3.6067 3.3479 5.6388 5.0336 4.6406 6.7905 5.9960 5.5115 3.4696 TOTAL 8 12 8 10 8 10 8 10 8 8 12 8 8 12 8 AR 580 379.7 347.5 506 456 611 931 637 481 1163 782 A WT. SYSTEM - LB 576 454 362.5 -569 -1,193 -881 -4,346 -1,812 -982 -317 -228 -777 -1,266 -514 -265 -2,917 -1,213 -642 A W BOX - LB -113 -582 -333 -3,183 -1,030 -402 . +62.7 +119.5 -271 -690 -60.3 +97.5 -1986 -576 -161 A W GLA - LB 0.00 0.93 -0.31 1.43 1.60- 479 1.20 1.37 1.38 -2.76 0.22 0.44 -0.39 % DOC-2 (500) -0.36 1.15 0.56 0.67 -0.20 0.43 -0.04 -0.67 -0.69 -1.66 1.17 % DOC-4 (500) -0.84 1.16 0.46 -3.33 0.57 1.32 1.82 0.65 1.35 -0.08 1.24 1.21 1.91 -3.10 0.00 1.01 -0.54 % DOC-2 (150) -0.67 0.55 1.07 1.55 1.02 1.25 0.51 1.64 -1.12 1.03 1.32 -3.49 0.07 0.59 -0.30 % DOC-4 (150) 2.59 -0.20 1.97 -0.76 1.62 -1.76 1.60 -0.57 2.73 -7.81 -2.22 2.4 -2.96 % OWE -4.85 3.48 4.48 2.28 3.85 3.30 5.54 2.12 4.75 5.71 6.62 % A/F COST 3.66 6.21 7.70 -0.05 112 103 128 119 136 127 125 105 228 146 122 235 155 131 A WSCR - LB 150 74,084 67,810 66,939 62,664 66,687 63,348 65,936 63,193 $ 81,220 75,447 66,362 90,000 75,224 67,636 95,000 A CSCR- 294 113 90 223 182 320 173 102 547 336 204 773 473 A WHYD - LB 277 71.07 61.65 89.62 65.82 46.53 38.1 63.3 53.77 85 60.3 39.867 150 80.742 56.4 154 WACT - LB 2.3180 2.4492 2.2354 2.3799 2.1877 2.3909 1.9438 2.6547 2.3000 1.9058 2.5745 DNGB (CRUISE) 3.1197 2.5495 2.0099 2.8158 .40966 .40286 .36601 .37064 .33945 .35133 .48038 .46280 .42214 .42185 .43197 .36769 .38706 .41361 DCLG .51382 .42533 .43707 .36776 .34310 .36350 .31636 .36065 .38519 .40725 .15476 .20566 .29408 .2825 .29036 .33570 ETAIGF 1.7060 1.7182 1.6385 1.6683 1.8608 1.8600 1.8898 1.7104 1.7621 1.8362 1.8249 1.8057 DNGB (UB) 2.1556 2.0414 1.9845 193.48 173.15 191.77 155.96 193.15 128.81 153.43 191.46 178.96 190.41 175.35 VEB - KEAS 138.93 160.77 196.07 132.41 300 300 250 250 250 300 300 300 300 250 250 250 250 250 250 VEC - KEAS REL. TO MANUAL CONTROLS 7.59 8.56 9.91 % A/F PRICE 5.4 8.22 1.93 2.11 1.87 WAVR 155 LB % DOC-2 (500) +.05 1.83 CAVR =90,400 1.57 1.59 -0.51 1.54 0.90 % DOC-4 (500) 2.28 1.74 % DOC-2 (150) -0.31 1.63 2.36 1.43 1.92 1.72 % DOC-4 (150) -0.81 1.38

IS

PAGE

ORIGINA

QUALITY POOR OF of Airplanes with GLA System.

Table XVII Characteristics 4-ENGINES 0.5M 100 LB WING LOADING (T.O.)

100 PAX = 44 LB/SQ.FT.

100 KG O WEIGHT ELECTRONICS LECTRONICS SURFACE CONTROLS SURFACE CONTROLS 0 -40 60 80 100 120 140 w30 40 50 60 70 LB/SQ.FT NUMBER OF PASSENGERS 150 250 350KG/SQ.M WING LOADING (T.O.)

Figure 23 GLA System Weights - Scaled 4-ENGINES 0.5M WING LOADING (T.O.)

100 PAX = 44 LB/SQ.FT.

s, 200 OTAL \PT COST - $1000 120ELECTRONICS ELECTRONICS

80 LCE0NT0

HYDRAULICS -AULs 0 40 60 80 100 120 140 30 40 50 60 70 LB/SQ.FT.

NUMBER OF PASSENGERS 150 250 350 KG/SQ.M.

WING LOADING (T.O.)

Figure 24 GLA System Costs - Scaled capacity for the GLA system. The cost/kg (Ib) of hydraulic components is lower than that of surface controls and electronics and this sharp increase in weight does not therefore reflect so strongly in the costs.

The variation of the "rough-airspeed" VB with field length is shown in Figure 25 for airplanes without GLA. This speed corresponds to CL MAX (flaps up) in the presence of a 20.12 m/sec (66 fps) gust. The shorter field length airplanes have a much lower VB than the longer field length airplanesdue to the difference in wing loading. In fact, at the longer field lengths, VB is approaching the cruise speed for 4570 m (15,000 ft) altitude. The effect of this variation of VB can be seen in Figure 26 which presents load factor as a function of field length. The maneuver load factor is shown constant for all field lengths at 2.5. The 15.25 m/sec. (50 fps) gust at cruise speed is shown to produce greater load factors than the 20.12 m/sec (66 fps) gust at speed VB at the shorter field lengths. This is due to VC being much greater than VB. As VB approaches VC, with increase in field length, so the effect of the higher gust velocity increases the gust load factor at VB relative to VC. Although the load factor is higher at VC than at VB for the shorter field lengths, it is the lower speed which actually determines the A CL required from the gust alleviation system and therefore designs the size of the GLA surfaces.

The effect of the gust load alleviation system on wing weight is illustrated in Figure 27 which shows the wing-box weight saving due to the gust load alleviation system plotted against field length. The weight saving is defined as the difference in weight between the weight of the wing sized for airplanes with GLA incorporated and the weight of the wing with the same geometry but without gust load alleviation, This non-alleviated wing does not include the weight increase required by resizing to make the heavier airplane meet the required performance and the weight savings presented are therefore conservative.

5.3 Characteristics of Airplanes Incorporating Artificial Stability Systems The term "artificial stability system" in this report includes the gust load alleviation capabilities of the systems discussed in Section 5.2.

When considering the resizing of the matrix of airplanes with the artificial stability system included it was assumed that the optimum wing aspect ratios would be determined by the gust load alleviation effects rather than the resizing of the horizontal stabilizer. This means that the optimum aspect ratios determined in Section 5.2 are close to optimum for airplanes with artificial stability systems and were therefore used when resizing.

KEAS 4-ENGINES 2-ENGINES KM/HR 0.5M 0. 6M = 463 KM/HR V = 556 KM/HR . 300 KEAS 250 KEAS VB / - 320 / 160 / 280 PAX 100 / 3 3.5 1000 FT 1 2 3 9 11 100M 8 10 4 6 LENGTH FIELD Figure 25 VB Speed vs. Field Length PAX 4-ENGINES 2-ENGINES 4.0 44 0.5M 0. 6M C 3.6 148 LOAD 3S FACTOR 3.2 V 66FPS GUST - 100 - - - - ..

- 2 .8 - 148 - CRITERIA MANEUVER 2.4 1 2 3 3 3.5 1000 FT 11 100M 6 8 10 9 FIELD LENGTH Figure 26 Gust Load Factor vs. Field Length Y-J 0 - 4-ENGINE 2-ENGINE 0.6 M 0.5 M -2 -2 x LU -4 -3 z CONSTANT AR =8 -4" -2L

) 2

4 1000 FT

12 100 M 4 6 8 10 FIELD LENGTH Box Weight Change Due to GLA Figure 27 Wing by modifying relaxed stability on horizontal stabilizer size were achieved The effects of the margin input to the sizing program as described in Section 4.3.1 and by the stability described in Section 4.3.3.

including the subsystem weight and cost changes load alleviation and relaxed of the matrix of airplanes, incorporating gust The characteristics stability, are presented in Table XVIII.

static with the artificial stability system The weight and cost changes of the subsystems associated alleviation system but include the similar trends as those presented for the gust load follow additional increment defined in Section 4.3.3.

SENGINES 4 PAX 100 148 100 148 M 0.5 0.5 0.5 0.6 0.6 0.6 F.L. - FT.

1,500 2,000 3,000 1,500 2,000 3,000 1,500 2,000 3,000 3,000 3,500 3,000 3,500 3,000 3,500 W/S T.O. - PSF 32.0 44.0 66.0 32.0 44.0 66.0 32.0 44.0 46.0 58.8 71.0 1.8 71.0 5S.8 71.0 T T.O.

.368 .325 .264 .369 .325 .264 .369 .325 .264 .400 .397 .400 .397 .39 .397 '500 .783 .768 .812 .780 .758 .798 .772 .744 .774 .751 .705 .734 .686 .721 .672 RGW - LB 57,585 48,402 43,022 111,771 92,175 81,292 155,017 126,280 110,759 48,652 47,414 94,045 91,541 129,710 126,074 OWE - LB 42,447 34,754 30,329 79,406 62,939 53,947 100,362 84,108 71,218 34,806 33,874 64,339 62,447 86,799 84,017 SLST- LB 5,709 4,232 3,064 11,090 8,065 5,781 15,385 11,038 7,879 10,470 10,128 20,222 19,559 27, 890 26,919 DOC-2 (500) 4.154 3.769 3.539 2.544 2.224 2.039 2.074 1.774 1.600 3.115 3.050 1.911 1.861 1.555 1.510 DOC-4 (500) 4.794 4.260 3.934 3.088 2.629 2.360 2.579 2.144 1.688 3.614 3.518 2.329 2.252 1.940 1.870 DOC-2 (150) 6.306 5.669 5.283 3.945 3.405 3.096 3.248 2.740 2.454 4.913 4.804 3.076 2.992 2.527 2.450 DOC-4 (150) 7.325 6.434 5.888 4.812 4.037 3.591 4.055 3.319 2.901 5.717 5.559 3.750 3.625 3.151 3.033 500 N.M. FUEL - LB 4,568 3,483 2,794 8,857 6,581 5,212 12,191 8,913 6,950 3,579 3,354 6,.30 6,387 9,.3 8,713 N.M. FUEL - LB 2,186 1,633 1,289 4,244 3,092 2,412 5,858 4,195 3,236 1,732 1,625 3,312 3,104 4,533 4,239 WING AREA - FT 1,794 1,097 650 3,482 2,088 1,228 4,830 2,862 1,674 824 665 1,59 1,284 2,198 1,769 WING WEIGHT - LB 8,387 5,820 4,760 18,196 12,377 10,070 26,835 18,001 14,562 '4,818 4,602 10,433 9,976 15,289 14,591 SURF. CONT. - LB 1,623 1,445 1,326 2,457 2,164 1,980 3,014 2,639 2,410 1,436 1,405 2,174 2,128 2,665 2,609 HYDRAULICS - LB 767 612 509 1,303 1,005 823 1,707 1,297 1,052 553 *523 903 851 1,161 1,092 HOR. STAB. V .610 .563 .674 .652 .619 .764 .701 .684 .868 .481 .563 .534 .632 .598 .715 HOR. STAB. AREA - FT 536 237 105.5 1,066 470 213.5 1,460 658 305 131.4 100 270 207 387 299 HOR. STAB. WT.

- LB 1,542 826 453 3,045 1,620 898 4,227 2,253 1,265 639 523.4 1,291 1,062 1,829 1,512 VERT. STAB. V .076 .069 .062 .109 .099 .090 .136 .125 .114 .055 .056 .00 .082 .102 .103 VERT. AREA - FT2 493 214 108.7 1,320 559 281 2,124 891 446 113 92.6 302 248 489 " 401 AIRFRAME COST - $M 3.0113 2.7495 2.5887 4.4164 3.9579 3.6941 5.4115 4.8004 4.4594 2.9031 2.8607 4.2351 4.1680 5.1638 5.0769 TOTAL COST - $M 3.9842 3.6311 3.3813 5.6273 5.0481 4.6710 6.7603 6.0095 5.5413 3.4972 3.4483 4.9731 4.8979 5.9842 5.8876 AR 8 8 .12 8 8 12 8 8 12 8 1010 8 10 8 A WT SYSTEM - LB 601.1 & W BOX - LB -1,240 -508.4 -263 -2,868 -1,196 -639 -4,265 -1,795 -978 -314 -226 -770 -566 -1,183 -876 &W GLA - LB -638.9 -26.3 128.2 -1,916 -533 -129 -3,100 -986 -370 94.8 151 -236 -81 -545 -300 % DOC-2 (500) -1.26 0.99 1.49 -3.96 -0.27 0.44 -0.89 -0.31 1.30 1.70 0.42 1.09 -0.38 0.73 % DOC-4 (500) -2.06 0.59 0.46 -4.87 -0.76 -0.84 -1.33 -1.82 0.87 1.12 0.00 0.45 -0.82 0.11 % DOC-2 (150) -1.68 0.93 1.81 -4.43 -0.61 +0.98 -1.05 0.49 1.15 1.87 0.26 1.25 -0.51 1.03 % DOC-4 (150) -2.46 0.56 1.12 -5.14 -0.76 0.45 -0.98 0.38 0.74 1.48 0.40 1.06 -0.10 1.30 % OWE -6.52 -1.50 2.34 -9.56 -3.28 +1.97 -3.89 2.18 -0.95 1.53 -1.62 1.09 -2.57 1.05 % A/F COST 4.10 7.26 9.18 -0.15 +3.78 +6.46 2.47 5.49 6.78 7.88 4.0 5.33 2.65 4.39 A WSCR -LB 150 125 105 228 145 122 235 154 131 111 103 128 119 136 127 A CSCR - $ 81,260 75,438 66,348 90,000 75,203 67,621 95,000 74,103 67,796 66,914 62,632 66,677 63,333 75,938 63,182 - A WHYD - LB 273 172 102 539 332 203 758 469 293 112 89 221 181 318 265 WACT - LB 85 60 39.8 150 80.4 56.7 154 89.4 65.7 46.3 38 63.1 53.6 70.9 61.5 GLF 4.1232 3.5496 3.0095 3.8236 3.3901 2.9436 3.6617 3.3013 2.9057 3.5734 3.3165 3.4496 3.2353 3.3805 3.1877 DCLG .51546 .48081 .4628 .42451 .42242 .43205 .37069 .38783 .41372 .40972 .40271 .. 36639 .37076 .33988 .35151 ETAIGF .15237 .2051 .2941 .27893 .28955 .33560 .36339 .34193 .36337 .31730 .3609 .38468 .40711 .42464 .43681 % FUEL 500 -7.02 -2.57 -8.15 -9.17 -3.53 -8.5 -3.79 -9.33 -1.81 -2.58 -1.83 -2.59 -2.52 -2.5 % FUEL 150 -7.06 -2.27 -4.73 -9.28 -3.44 -4.93 -3.47 -5.38 -1.70 -0.98 -1.66 -0.89 -2.28 -0.70 RELATIVE TO MANUAL CONTROLS % A/F COST 5.84 9.29 11.41 8.68 9.85 % DOC-2 (500) -0.86 1.45 1.99 1.80 2.21 % DOC-4 (500) -1.74 0.97 0.90 1.26 1.56 % DOC-2 (150) -1.33 1.34 2.26 1.57 2.32 % DOC-4 (150) -2.16 0.91 1.52 1.10 1.85 ORIGINAL PAGE I 'OF POOR QUALITY Table XVIII Characteristics of Airplanes with AS System OF AIRPLANE CHARACTERISTICS 6.0 COMPARISON WITH AND WITHOUT ACTIVE CONTROLS 3.5, 5.1, 5.2 the turboprop powered airplanes contained in Sections The characteristics of as direct operating compared to illustrate the effects on such parameters and 5.3 are first the three types of active fuel, weights and initial price, due to incorporating cost, mission turbofan powered MF configuration are The effects of these systems on a control system.

a the competitiveness of the aircraft with and the effects of these systems on then illustrated lift concept are discussed.

powered Characteristics 6.1 Comparison of Turboprop Airplane - Direct operating cost at DOC-2 and DOC-4 6.1.1 Direct Operating Cost Comparisons fuel prices as a function of field length are presented in Figures 28, 29 and 30 for the 44, respectively. The data are presented for the 2- and 100 and 148 passenger baseline airplanes The km (150 and 500 n.mi,) stage lengths.

4-engine configurations and for 278 and 926 apparent by due to increase in passenger capacity is expected large reduction in DOC in DOC as the shortest field length is the figures. Note also the rapid increase comparing approached.

at 914 m (3000 ft) field to compare the 2- and 4-engine configurations It is interesting and use more fuel but in some cases the length. The 2-engine designs are heavier, cost more higher speed results in a lower DOC than for the 4-engine designs. For 44 passengers, for both fuel prices and both stage lengths. For the 2-engines provides minimum DOC at 926 km (500 n.mi.), the two configurations are equal passenger case, 2-engines are better and n.mi.), while the 4-engine design is better for DOC-4 at DOC-2 and 278 km (150 for 926 km For 148 passengers, the 4-engine design is best except 278 km (150 n.mi.).

DOC-2.

(500 n.mi.) and (150 n.mi.), DOC-2 the three active control systems on 278 km The effects of incorporating 44, 100 and 148 passenger capacity are presented in Figures 31, 32, and 33 for the and all It can be seen that the RQ system increases DOC at all field lengths configurations.

at the shorter field lengths but sizes. The GLA and AS systems reduce DOC passenger passenger airplane.

at the longer field lengths except in the case of the 148 increase DOC difficult to use for adequate to show the general trends of the data but are These figures are prepared showing quantitative effects; a series of figures have therefore been determining controls as a function of field percentage change in DOC due to introduction of the active length.

4-ENGINE 2-ENGINE 0.5 M' 0.6 M 7.0 DOC (278 K 6.0 (278 KM) .(150 .)

5.0 N (926 DOC-4

DOC-2

(500

N.

Mi.

(926 KM) 3.0 2 3 4 1000 FT 6 8 10 12 100 M FIELD LENGTH Figure 28 Direct Operating Cost vs. Field Length (44 Pax - Without Active Controls) 6.0 4-ENGINE 2-ENGI NE 0.5 M 0.6M 5.0 ) Oc-4 ( 78 O4.0 S4.0 -(IS N MI.)

S3.0.D

-4

(926

KM)

"C-2(500

N.A41oM.)

-

2.0 (926 KM) 1 2 3 4 1000 FT 10 12 100 M 6 8 FIELD LENGTH Pax - Without Active Controls) Direct Operating Cost vs. Field Length (100 Figure 29

6.0

5.0 2-ENGINE 4-ENGINE 0.5 M 0.6M S4.0 DOC - (278 KM) DOC2 (278 KM)

DO-2

~(50

N.Ml,.)

(926 KM) DO C-4 (500 N.M) DOC-2 (500 N.MI.)

(926 KM) 1.0 4 1000 FT M 10 12 100 4 6 FIELD LENGTH - Without Active Controls) Operating Cost vs. Field Length (148 Pax Figure 30 Direct BASELINE ----- RQ RQ 6.4 6.4- - - GLA - -- - AS 6.0 2 ENG 4 ENG M 0.6 0.5 M S5.6 5.2 c C!

4.8 4.4 I = 1 2 3 4 1000 FT I I I I I 8 10 12 100 KM 4 6 FIELD LENGTH Pax) Figure 31 DOC-2 (278 km, 150 n.m) vs. Field Length (44 4.4 100 PAX BASELINE - 1 .....

RQ ( 4.o- -- GLA - - - - AS Z 4-ENG 2-ENG L 3.6 0.5 M 0.6 M c 3.2 2.81 1 2 3 4 1000 FT I I I I I 8 10 12 100 M 4 6 FIELD LENGTH DOC-2 (278 km, 150 n.m) vs. Field Length (100 Pax) Figure 32 BASELINE 148 PAX ....

RQ GLA 5 3.2 --- ---- AS 4-ENG 2-ENG 0.5 M 0.6 M 2.8 co , 2.4 CN I 22.0 FT 3 4 1000 8 10 12 100 M 4 6 FIELD LENGTH Pax) Figure 33 DOC-2 (278 km, 150 n.m) vs. Field Length (148 Figure 34, for example, illustrates percentage change in DOC-2 for 278 km (150 n.mi.) stage on length. The RQ system can be seen to increase DOC by 0.65 to 2.0 percent dependent increase being for the shortest field length.

field length and passenger size, with the largest The effects of the GLA and AS systems in improving aircraft efficiency, particularly at the shorter field lengths, can be seen for all but the 44 passenger aircraft at the longer field below that of the baseline at lengths. For the 44 passenger airplane DOC can be reduced ft) but increases above that of the RQ design at field field lengths below 579 m (1900 lengths longer than 686 m (2250 ft). In other words, GLA or AS systems should not be incorporated into the 44 passenger size vehicle at field lengths longer than 686 m (2250 ft).

quality The 100 passenger size with a GLA or AS system can be provided with excellent ride DOC of 3 to 4 percent. At at the shortest field length while achieving a reduction in system is more desirable than the GLA or AS approximately 610m (2900ft.), the RQ configuration the RQ system is systems for the 4-engine configurations. For the 2-engine greater than about 1036 m (3400 ft). The 148 passenger most desirable for field lengths or AS system in preference to the RQ system at aircraft should be provided with the GLA ft); lower DOC's than the baseline are achieved at field all field lengths below 991 m (3250 m (2300 to 2600 ft).

lengths below 701 to 792 Similar data are provided in Figures 35 through 37 for DOC-4, 278 km (150 n.mi.); DOC-2, n.mi.). Figures 38 and 39 present percent 926 km (500 n.mi.); and DOC-4, 926 km (500 change in DOC-2 and DOC-4 for the three active control systems as a function of passenger size for the 4-engine configuration at 610 m (2000 ft.) field length. At all sizes benefits are provided by the GLA and AS systems relative to the RQ system. Except at the smallest size the AS system has a lower DOC than the GLA system. Figures 40 and 41 present similar data for the 2-engine configuration at 914 m (3000 ft.) field length and illustrate the passenger sizes above which the GLA and AS systems are more desirable than the RQ system and when the systems provide lower DOC than the baseline aircraft without active controls.

6.1.2 Fuel Consumption Comparisons - The effects on 278 km (150 n.mi.) mission fuel are shown in Figures 42 through 44 for the 44, 100 and 148 of the active control systems airplanes respectively. As expected, the RQ system incurs a penalty while the passenger GLA and AS systems provide fuel savings. The savings at the shortest field length are because of the reduction in airplane size through GLA, and at the longer field lengths because of the increase in aspect ratio provided by the GLA system. An additional increment of saving is provided by the AS system due to the reeuction in horizontal stabilizer area and consequent reduction in airplane size. Note the large increase in mission fuel required by the 2-engine configuration. This is partly due to the speed difference and partly due to the lower wing loading and hence larger airplane required to meet the field performance.

44 RQ ZC, 100 RQ • 18R 00 RQ 4% -1 CHANGE % CHANGE 14 R 148 RQ : IN DOC-2 1I 278 KM e (150 N.MI.)

4-ENGINES 2-ENGINES - 1 0.5M 0.6M - - - -2 -2 1.5 2 2.5 3 3.5 1000 FT 8 10 100M 4 6 FIELD LENGTH Figure 34 Percent Change in DOC-2 (278 km, 150 n.m) Due to Active Controls 44 GLA +2.0 .4 RQ 44 AS / 148 GLA -. .....- / 44 RQ - , -"0 100 GLA 100 AS " * +1 0 RQ -100 148 O"

7 7 /

0 -1.0 1 8 c 0 Z I <r -2.0 /

-- I

I

-4.0 I I 1 2 3 4 1000 FT I I I I 4 6 8 10 1 2 100 M FIELD LENGTH Figure 35 Percent Change DOG-4 (278 kmin, 150 n.m) Due to Active Controls 2-ENGINE 4-ENGINE 0.5 M 0.6 M 2.0 2.

20.... . _;- 2.0 2.0 -ENGINE 2-ENGINENE 2- NE 4-ENGINE 2-ENGINE M 0.5- M 0.6 0.6 M 0.5OM 1.-1.0 -10 Z - Z ---- EQ --- G -- --- G A --- GLA ---- AS g ---- AS ---- AS -2.

-2.0-2.

-3.

-3.

-3.0 44 L -' -4.01,0 3 4 1000 FT 4 1000 FT .- . . *4I0FM 4ISo _ _ _ __ 8 10 12 100 M 10 12 OM 4 6 12 I0M 4 6 8 4 6 B 10 FIELDLENGTH FIELD LENGTH FIELD LENGTH Due to Active Controls (926 km, 500 n.m) Change in DOC-2 Figure 36 Percent 2.0 2.0 4-ENGINE 2-ENGINE 2.0 S 0.5 M 0.6 M - M 0 . M. -EN G IN E 2-ENG IN E S0.5 M 0.6M 4-E NGINE 2-ENGINE 0.5 M 0.6 M S.

--- -- -- - - - - - - - ---- - - - - - - - - - - - - - - - ------- A----- --- GL .A0 G.

-2..

-4G -- 4.

1.__ _4 1000 FT 3_ _4 1000FT 2 3 4 1000 FT I 4 6 8 10 12 OOM 0 10 12 100 M 4 6 . 10 12 100 M 4 6 FIELDLENGTH FIELDLENGTH FIELDLENGTH Percent Change in DOC-4 (926 km, 500 n.m) Due to Active Controls Figure 37 278 KM (150 N.MI.)

4-ENGINES 0.5M 610M (2000 FT.) FIELD LENGTH ft.

-'-- DOC-2 DOC-4 DO- % CHANGE IN DOC 0GLA -1" 100 120 140 40 60 80 NUMBER OF PASSENGERS (2000 ft) F.L.

n.m) vs. Passenger Size - 610 m Change in DOC (278 km, 150 Figure 38 Percent 4-ENG 0.5M 610M (2000 FT) F.L.

926 KM (500 N.MI.)

+2.0 DOC-2 +1.0Q DOC

Z

L)

U 09

0-0 -1.0 60 80 100 120 140 160 NO. OF PAX Figure 39 Percent Change in DOC (926 km, 500 n.m) vs. Passenger Size - 610 m (2000 ft) F.L.

278 KM (150 N.MI.)

0.6M FIELD LENGTH FT (914 M) -3000 2-ENGINES % CHANGE IN DOC DO-4 R i -1 .

80 100 120 140 NUMBER OF PASSENGERS ft) F.L.

Passenger Size - 914 m (3000 DOC (278 km, 150 n.m) vs.

40 Percent Change in Figure 2-ENG 0.6M 914M (3000 FT) F.L.

KM (500 N.MI.)

+2.0 +1.0 RQ DOC-2 \ \ -- -, ....

...

Q DOC-4 U

o

Z 0

P

% NO. OF PAX Figure 41 Percent Change in DOC (926 km, 500 n.m) vs. Passenger Size - 914 m (3000 ft) F.L 44 PAX 1000 LB 2.4 2.4 BASELINE S--- --- RQ I - - - GLA AS - - - - 100 KG 10 -2.2 2.

U.c 2-ENG M 0.6 1.8 Z8 4-ENG

% l0.5M

oo 1.6 1.4 1.2 I I I 4 1000 FT 2 3 12 100M 6 8 FIELD LENGTH Fuel (278 km, 150 n.m) - 44 Pax Figure 42 Effect of Active Controls on Mission 100 PAX 1000 LB 5.0 BASELINE --- RQ SGLA!

---- ASi 100 KG 4.5 20- 4.C U 3.

32-ENG Z1E %0.6 M O 15 % 4-ENG 0.5 M 3.0 N CO 2.

2.0 I I 4 1000 FT 12 100M FIELD LENGTH Figure 43 Effect of Active Controls on Mission Fuel (278 km, 150 n.mi) - 100 Pax 148 PAX 1000 LB 6.0 BASELINE - - -- - -RQ - - - GLA 100 KG --- AS 255.5 5.C 2-ENG 0.6M ,-I U- 4.5 •20 LO4-ENG 0.5 M 4.0 co 3.5 3.0 4 1000 FT 1 2 3 4 6 8 10 12 100M FIELD LENGTH Figure 44 Effect of Active Controls on Mission Fuel (926 km, 500 n.m) - 148 Pax The trends are n.mi.) stage length.

for the 926 km (500 through 47 present data Figures 45 AS systems are provided by the GLA and length but the savings similar to the shorter stage seat km/kg (seat of fuel efficiency, Figure 48 presents of course much greater. As a measure i.e., those with extremes of fuel consumption, for the aircraft with the statute miles/gallon) designs and the poor fuel efficiency of the 2-engine systems. Note again the the RQ and AS lengths. As would be with the shortest field in fuel efficiency associated drastic reduction the AS system capacity, 4-engined airplanes incorporating expected, the largest passenger time provide excellent ride quality.

the best fuel efficiency and at the same have weight as a and 51 compare ramp gross - Figures 49, 50 6.1.3 Weight Comparisons be expected the and 148 passenger sizes. As would of field length for the 44, 100 function design, partly because of the the equivalent 4-engine airplanes are heavier than 2-engined performance requirements because of the engine-out higher cruise speed and partly ratio.

and higher thrust to weight demanding lower wing loading and for all passenger incurs a weight penalty at all field lengths The provision of ride quality or without the relaxed static stability sizes while the adoption of gust load alleviation with wing saving benefits which are highest at the lowest of the system does offer weight portion the in passenger size as.will be seen by comparing loadings and increase with increase as can be seen from Figure is not provided in all cases, This weight saving benefit figures.

of the 4-engine airplane alleviation system actually penalizes the weight where the gust load due m (3500 ft). This weight penalty is and the 2-engine vehicle at 1067 at 914 m (3000 ft) wing exceeding the saving and the weight of the higher aspect ratio to the subsystem weight but at with the larger passenger capacities weights. Note similar trends in fuel and other Figure 53 (2000 ft) field length, 4-engine and lengths. Figure 52 for the 610 m longer field further illustrate the variations (3000 ft) field length, 2-engine configuration for the 914 m of RGW and OWE with passenger capacity.

systems is shown in Figure due to the various active control The percentage change in OWE quality system increases length for each passenger size. The ride 54 as a function of field to 0.4 to 0.75 percent at to 1.5 percent at the shortest field length compared OWE by 1.0 is provided greater improvement in ride quality field lengths. Of course, a much the longest weight is not longest. Since much of the subsystem at the shortest field length than at the savings would be follows that the greatest percentage affected by passenger size, it greatly in the figure. The passenger size; this is confirmed expected to be provided by the largest baseline aircraft in system break-even with the the benefits of GLA or AS airplanes with configuration and around 701 m (2300 ft) for the 4-engine terms of OWE at field lengths of benefits are Below these field lengths OWE 960 m (3150 ft) for the 2-engine configuration.

will be encountered.

provided, while above them OWE penalties 1000 LB 4-ENGINE 2-ENGINE M 0.6 0.5 M 5.0 100 KG 4.5 BASELINE ..

- RQ - - - GLA 4.0 - - - - AS D 3.5 %o 3.0 2.5 12 3 4 1000 FT 12 100 M 6 8 10 FIELD LENGTH 45 Effect of Active Controls on Mission Fuel (926 km, 500 n.m)- 44 Pax Figure 1000 LB 100 KG 4-ENG INEI 2-ENGINE 0.5 M, 0.6M, 7.0 - D 6.5 U- 6.0 25 L5.5 BASELINE - ---

RQ - RQ

- - - GLA - - -

-AS

5.0 4.5 3 4 1000 FT 4 6 8 10 12 100 M FIELD LENGTH Figure 46 Effect of Active Controls on Mission Fuel (926 km, 500 n.m) - 100 Pax 1000 LB 4-ENGINE 2-ENGINE 0.5 M 0.6 M 9.5 KG 9.0

-1 8.5

, 38 8.0 LO BASELINE 34 7.5 RQ % --- GLA AS 7.0 6.5 3 4 1000 FT 4 6 8 10 12 100 M FIELD LENGTH Figure 47 Effect of Active Controls on Mission Fuel (926 km, 500 n.m) - 148 Pax 926 KM (500 N.MI.)

SEAT ST.

MI./GAL., S'EAT SEAT 4-ENGINE.

KM/KG 0.5 M

7-

60 */ / # 0 1 40 / 2-ENGINE0 0.6 M P4 30 - 1 2 '3 4 1000 FT 4 6 8 10 12 100 M FIELD LENGTH Figure 48 Effect of Active Controls on Fuel Usage Per Passenger BASELI NE ......

-RQ - - - GLA 1000 LB 44 PASSENGER AS 1000 KG 30 4-ENGINE 2-ENGINE 0.5 M 0.6 M 25- S S20 I I I I I I I I I II 3 1000 FT 4 1 2 4 6 8 10 100 M 12 FIELD LENGTH on RGW - 44 Pax 49 Effect of Active Controls Figure BASELINE ----- RQ - -- GLA AS - - - - PASSENGER 1000 LB!

KG 55120 4-ENGINE 2-ENGINE.

4 0.5 M 0.6 M.

S 110 1 2 4 6 8 10 12 M BASELINE -------- RQ -- GLA 148 PASSENGER AS 1000 LB 4- ENGINE 2-ENGINE 0.6 M 0.5 M 1000 KG

60 130

2 3 4 1000 FT I I I I I 100 M 10 12 6 8 FIELD LENGTH Figure 51 Effect of Active Controls on RGW - 148 Pax 610M (2000 FT) F.L. 4-ENGINES 1000 KG 1000 LB BASELINE r RQ " .....

---- --- G LA ....

AS 40 0 OWE 40 60 80 100 120 140 NO. OF PASSENGERS Figure 52 RGW and OWE vs. Passenger Capacity - 4 Engines, 610 m (2000 ft) F.L 914M (3000 FT) F.L., 2-ENGINES 0.6M 1000 LB 1000 KG RGW BASELINE RQ - - - - - - - GLA AS - - - - 50- OWE 40- 30 - 10 '20 140 160 80 100 120 40 60 NO. OF PASSENGERS 914 m (3000 ft) F.L and OWE vs. Passenger Capacity - 2 Engines, Figure 53 RGW 44 G LA 100 i AS 0 A

S100

RQ

148 G LA 148 AS 44 AS z GLA S0.5M 2-E4-ENGINE 0.6 M ii,, -4 4 1000 FT 8 10 12 100M FIELD LENGTH J Figure 54 Percent Change in OWE due to Active Controls 1000 LB PAX -- RQ 1000 KG 148 ---- GL AS 2.0 4-ENGINES 2-ENGINES 1.5 1.5 0.5M 0.6M HORIZONTAL STABILIZER WEIGHT 0.56 - 0 0 2 3 w3 3.5 1000 FT 6 8 10 9 11 100M FIELD LENGTH Figure 55 Effect of Active Controls on Horizontal Stabilizer Weight can amount to 6.5 by the AS system which benefits are generally provided The largest OWE of gust load alleviation In addition to the effects shortest field lengths.

to 11 percent at the in a reduction of horizontal the AS system results of the static stability with the relaxation field length and passenger as shown in Figure 55 as a function of stabilizer area and weight shorter than the RQ weights at the weights are lower be noted that the GLA size. It will At GLA at the lower wing loadings.

in airplane size with due to the reduction field lengths in relatively at higher wing aspect ratios results field lengths the optimization the longer chord terms of mean aerodynamic percent c.g. limits in wing chords and larger smaller program reacts stabilizer. The sizing the required size of horizontal which in turn increases higher weights actually being slightly in Figure 55 by the GLA to these effects as shown system weights.

than the RQ Figure 55 by weight is shown clearly in static stability on stabilizer The effect of relaxed application of active controls effects are for the up 386 kg (850 Ib). These reductions of It is trimmable with elevators.

of stabilizer, in this case, retaining the same type while and use of flying tails stabilizer by the in horizontal obtain further reductions possible to not been further reduction has systems. This with active control geared elevators combined quantified in this program.

introduces an controls in the airplanes incorporation of active Airplane Price - The 6.1.4 in some cases, all passenger sizes and field lengths. However, increase in subsystem cost for price airplanes, the total airplane resizing to smaller, more efficient due to weight saving and 60 present initial airplane. Figures 56 through below that of the basic is actually reduced size for airplanes of field length and passenger total aircraft price as a function airframe and in almost all cases the initial price active controls. It can be seen that with and without saving and resizing of control system is not offset by the weight increase due to the active with the GLA and AS that when reoptimizing It must be remembered however the vehicle.

is in efficiency, but it increases with a resulting improvement aspect ratio systems, the wing at the lowest wing loadings with increased weight and cost. Only also generally associated GLA translate into a cost weight saving due to lengths) does the (i.e., shortest field is unchanged by the in this case, the optimum aspect ratio reduction. This occurs because, of GLA.

introduction three cost due to each of the change in airframe 61 through 63 present percent Figures sizes for the 44, 100 and 148 passenger as a function of field length active control systems below the cases does the cost decrease only in special As already mentioned, respectively.

AS systems cost less shortest field lengths the GLA and airframe cost. Note that at the basic BASELINE - - - - - RQ -- -- GLA 4.4 - ---- AS 4-ENGINE, 2-ENGINE 0.5 M' 0.6 M 4.0 TOTAL AIRCRAFT 3.6 , 3.2 LU U AIRFRAME 2.8 2.4~ 2.0 I I I 1 2 3.

i4 1000 FT 10 12 100 M FIELD LENGTH Figure 56 Effect of Active Controls on Airframe and Aircraft Price - 44 Pax BASELINE RQ - - - GLA --- AS 5.8 4-ENGINE 2-ENGINE \ 0.5 M 0.6 M 5.4 TOTAL AIRCRAFT Z 5.0

I o

,4.6 U \ AIRFRAME 4.2 3.8 I 3.I 1 2 3 4 1000 FT 100 M 8 10 12 4 6 FIELD LENGTH Price - 100 Pax Controls on Airframe and Aircraft 57, Effect of Active Figure BASELINE RQ 7.0 - - - GLA ---- AS 4-ENGINE .2-ENGINE 6.5 0.5 M' 0.6 M.

TOTAL AIRCRAFT 6.0 5-.51 4^ U AIRFRAME 5.0 4.5 4.0 I I I I I I I I 1 2 3 4 1000 FT 6 8 10 12 100 M FIELD LENGTH Figure 58 Effect of Active Controls on Airframe and Aircraft Price - 148 Pax 4 ENG 610M (2000 FT) F.L. 0.5M BASELINE ----- RQ /c 6.0 -.-.

AS 5.2 i4.4 3.6 2.8 2.0 m a - a i 40 60 80 100 120 140 160 NO.

OF PASSENGERS Figure 59 Airframe and Aircraft Price vs. Passenger Capacity - 4 Engines, 610 m (2000 ft) F.L 914M (3000 FT) F.L. 0.6M 2 ENG.

BASELINE -- - - - - RQ

o.v iC I

--

GLA

pt.

-...

AS 5.2 ~ 4.4 3.6 2.8 2.0 I I 40 60 80 100 120 140 160 NO. OF PASSENGERS Figure 60 Airframe and Aircraft Price vs. Passenger Capacity - 2 Engines, 914 m (3000 ft) F.L 44 PAX 8 / - - -REL. TO MANUAL RQ 6.

LUE rO POWERED C ::E CONTROLS TROL .

U- 2-ENG S4-ENG 0.5 M 0.6'M I I 0 I 4 1000 FT 2 3 I I I I I 10 12 100M 4 6 8 FIELD LENGTH Figure 61 Percent Change in Airframe Price due to Active Controls - 44 Pax 8 100 PAX 4-ENG 2-ENG < 2 o 0.5 M - 0.6 M 1 2 3 4 1000 FT 4 6 8 10 12 100M FIELD LENGTH in Airframe Price due to Active Controls - 100 Pax Figure 62 Percent Change 148 PAX z - U 4 2 4-ENG 2-ENG 0.5 M 0.6 M 2 3 4 1000 FT 6 8 10 12 100M FIELD LENGTH in Airframe Price due to Active Controls - 148 Pax Figure 63 Percent Change of these systems exceed while at the longer field lengths the cost effects than the RQ system be offset by the improved These increased costs may however that of the RO system.

in lower direct operating costs.

efficiency and result function of passenger capacity for presents percentage airframe cost change as a Figure 64 the percentage change relative to the baseline 4-engine configurations. In all cases the that the percentage change reduces with increase in passenger size. It is noteworthy airplane in field length, but it reduces with RQ system airplanes increases with reduction for the field length for the GLA and AS systems.

reduction in of Concepts 6.2 Turbofan-Powered MF Airplane Characteristics and Comparison m (3000 ft) field length mechanical Ref. 1 study it was determined that the 914 During the turbofans advanced 1.35 fan pressure ratio (FPR) flap (MF) concept powered by two km (500 n.mi.) stage length at 0.7M. The wing loading provided optimum DOC-2 for a 926 Ib./sq. ft.) which in Figure 65, was 287 kg/sq. m. (58.8 of this airplane, which is illustrated discussed in the quality compared to B737 type aircraft. The methodologies gives poor ride active control systems on the turboprop sections for determining the effects of previous 914 m for the 148 passenger, concept have been applied to this turbofan-powered concept (3000 ft) field length configuration.

analyzed with 1972 costs as the basis; the previous study airplanes were configured and The Additionally, a very austere furnishing study has been conducted with 1974 costs.

present costs have been used in the present study for the turboprop standard and reduced crew for short-haul operation. In to obtaining the absolute minimum DOC designs with a view turbofan-powered and a true comparison of the concepts, both the order to present analyzed using consistent ground rules.

aircraft' have been resized and turboprop-powered in Table XIX together with of the resulting airplanes are presented The characteristics (OTW/IBF) powered lift for the over-the-wing/internally blown flap equivalent data in Ref. 1 to have minimum This concept, illustrated in Figure 66, was shown concept.

this 1.35 FPR turbofan engines. The data for DOC-2 at 0.75M when powered by four but since its wing loading is sufficiently high to airplane have been updated to 1974 costs quality it has not been resized with active controls.

provide excellent ride The turboprop configuration is shown in Table XIX to have the lowest fuel consumption, It should be noted that the turboprop engine DOC-2 and DOC-4 of the three concepts.

T-56 engine. It data are based on a rubberized Detroit Diesel Allison performance and cost RELATIVE TO RELATIVE TO MANUAL BASELINE' POWERED BASELINE RQ + RQ --- GLA O GLA 8 ) . AS + F. L. 3000' , 2000' -2 40 60 80 100 120 140 160 NO. OF PASSENGERS in Airframe Price due to Active Controls vs. Passenger Capacity Figure 64 Percent Change 148 PASSENGERS 0.70 MACH OPTIMIZED FOR DOC-2 I MF Vehicle Figure 65 A/C OPTIMIZED FOR DOC-2, 914M (3000 FT.) F.L., 926 KM (500 N.MI.)

D.S.

MF 1.35 FPR TURBOPROP CONCEPT OTW/IBF 1.35 FPR ACTIVE CONTROL NONE NONE RQ GLA AS NONE RQ GLA AS NO.OF ENG. 4 2 2 2 2 4 4 4 4 MACH NO. 0.75 0.70 0.70 0.70 0.70 0.5 0.5 0.5 0.5 OWE- KG 36,510 39,687 39,850 39,189 38,970 34,179 34,303 35,135 34,981 x (LB) 80,490 87,494 87,853 86,396 85,913 75,351 75,623 77,458 77,119 x RGW - KG 56,446 59,848 60,026 59,387 59,110 52,694 52,828 53,278 53,084 (LB) 124,440 131,940 132,332 130,924 130,314 116,168 116,464 117,457 117,028 RATED THRUST- KN 55.33 119.6 119.9 112.8 107.8 41.5 41.59 37.19 37.05 (LB) 12,440 26,890 26,948 25,365 24,231 9,330 9,350 8,355 8,332 MISSION FUEL- KG 4,400 4,790 4,802 4,749 4,708 3,601 3,609 3,335 3,304 0 (LB) 9,700 10,560 10,586 10,470 10,380 7,938 7,956 7,352 7,285 2 W/S T.- KG/SQ.M 554 287 287 287 287 347 347 322 322 (LB/SQ.FT) 113.5 58.8 58.8 58.8 58.8 71.0 71.0 66.0 66.0 12 8 8 10 10 8 - 8 12 12 AR DOC-2 ¢/ASSM 1.911 1.897 1.909 1.884 1.876 1.7866 1.799 1.793 1.788 DOC-4 ¢/ASSM 2.326 2.333 2.347 2.336 2.304 2.117 2.129 2.097 2.090 A/C PRICE $SM 9.103 8.2736 8.3984 8.4015 8.3778 5.5163 5.6248; 5.7253 5.7143 148 PASSENGERS 0.75 MACH OPTIMIZED FOR DOC-2

- - - -- - -- - -

Vehicle Figure 66 OTW/IBF is only possible to achieve these costs at the actual size of the T-56 which would result in a 200 passenger airplane with 4-engines at 914 m (3000 ft.) field length or a 48 passenger design with 2-engines at 0.6 M and 1067 m (3500 ft) field length.

Alternate sizes'are possible for other speeds and field lengths and by using other existing engines.

It is expected that the results would be similar to the data generated using the rubberized T-56.

If a new, advanced turboprop engine having a higher cost and lower fuel consumption (Ref.

1) is used the DOC-2 and DOC-4 values are increased by 11 and 6.5 percent respectively.

The turboprop is then not competitive with the other concepts at DOC-2 which suggests that a new turboprop airplane should be designed to use an existing engine unless fuel costs increase to DOC-4 values or above. II I.I LU IVJ"T VauIoo I aL5V.

The turbofan powered MF with a ride quality system and the OTW/IBF are almost identical in DOC-2 value but the OTW/IBF has the advantage of a 9 percent lower fuel consumption.

The incorporation of either the GLA or AS system improves the MF DOC-2, but it should be noted that a slight improvement in OTW/IBF DOC-2 could also be achieved by incorporating just the relaxed stability portion of the AS system. In order to match the OTW/IBF on DOC-4, the MF must incorporate the AS system.

7.0 CONCLUSIONS AND RECOMMENDATIONS short-haul airplanes with low wing loading can be improved to the The ride quality of of contemporary high wing loading airplanes by the use of active control systems.

standard cost penalty for improved ride quality is 2 percent or less for all cases; The direct operating of gust load alleviation and augmented stability overcomes this penalty and incorporation DOC than aircraft without active controls in all the very low wing loading gives better (1500 ft) field length) and in the 100 and 148 passenger airplanes at 610 m airplanes (457 m ft) and 914 m (3000 ft) field length. For small aircraft (44 passengers) a GLA or AS (2000 for very short field lengths but for field lengths of 610 m (2000 ft) system is recommended and longer the simpler RQ system results in a smaller DOC penalty. For larger airplanes penalties than (100-148 passengers) the GLA and AS systems generally provide smaller DOC less then 914 m (3000 ft). Above this field length the the RQ system for field lengths of to minimize DOC effects except at the longer ranges and higher fuel RO system appears the increased aspect ratio of the GLA and AS systems results in improved fuel prices where active consumption and an advantage in DOC. Fuel savings of 11% were obtained by use of and 347 kg/sq. m (71 controls in a 148 passenger airplane at 914 m (3000 ft) field length lb/sq ft) wing loading.

loadings where Weight savings were obtained with the GLA and AS system at the lower wing field lengths and higher wing reoptimization did not increase wing aspect ratio. At longer controls were obtained at increased loadings the best economics of aircraft with active improved but small weight and cost penalties were aspect ratios. Fuel consumption was Generally, the active control systems increased the incurred compared to baseline aircraft.

the only exception being the largest aircraft at the shorter field initial cost of the airplane; lengths.

consumption and competitive direct operating costs, the Due to the favorable fuel active controls must be considered a major turboprop-powered configuration with density market, particularly for the shorter route contender for the short-haul low/medium due to low speed is negligible. It must be stressed however segments where the time increase turboprop aircraft, to be competitive, must be designed to match existing that these turboprop engines. The increased cost of a new turboprop engine will nullify most of the advantage of this configuration.

s.f.c. of a diesel engine might make consideration of this engine cycle It may be that the low worthwhile. Similarly since the development of a new turboprop engine is questionable it speeds and may be that a new very high bypass ratio fan would be advantageous at these field lengths., The incorporation of active controls in the turbofan MF airplane results in it being equal to the OTW/IBF hybrid in terms of DOC and ride quality. However, the OTW/IBF, because of its higher wing loading, retains its advantage of lower fuel consumption.

This study has been limited to short-haul; it is likely that larger fuel savings are available by the use of active control systems on long haul aircraft which stand to gain so much more from higher aspect ratio wings providing the wing weight increases can be minimized.

Active control systems combining features such as ride quality improvement, gust load alleviation, flutter control and relaxed static stability could result in very efficient high aspect ratio wings. It is recommended that such a program be considered with the final step being the flight demonstration of the wing design.

REFERENCES Short-Haul Aircraft" and Fuel Conservative Advanced Lift Concepts 1. "Evaluation of June 1974.

CR 137525 and 137526, NASA Aircraft Corporation, by Lockheed Lockheed-Georgia for American Airlines" by STOL Transport Proposal 2. "Propeller December, 1970.

Company, ER 10889, Category Aircraft" for Powered Lift Transport Airworthiness Standards 3. "Tentative August, 1970.

Aviation Administration, by Federal Federal Category Airplanes" by Standards: Transport - Airworthiness 4. "Part-25 Aviation Administration.

Transportation" by Lockheed for Short-Haul Turbofan STOL Aircraft 5. "Quiet CR 114612, 14 June, 1973.

Aircraft Corporation, NASA Reaction to Aircraft Motion to Determine Human Simulator Experiments 6. "Flight July, 1974.

of Virginia, ESS-4039-102-74, by University Environments" Ride Functions and Aircraft Error Frequency Response "Development of Tracking 7.

J. W., for Vertical and Lateral Vibration", Rustenberg, Quality Design Criteria Ohio, January, 1971.

AFSC, WPAFB, ASD-TR-70-18, Smoothing" by STOL Transports with Ride "Prospects for Low Wing-Loading 8.

1972.

J., Journal of Aircraft, August, Thompson, G.O., Rohling, W.

Holloway, R.B., Criteria and Analysis on V/STOL Aircraft Handling Qualities 9. "Interim Position April, 1970.

Co., ER-10424, Methods" by Lockheed-Georgia by Innes, Robert C., Holzhauser, Considerations for STOL Aircraft" 10. "Airworthiness January 1970.

Hervey C., NASA TN5594, Curt A., Quigley, December, 1970 Handling", AGARD R 577-70, 11. "V/STOL 83300, for Flying Qualities of Piloted Airplanes", MIL-F 12. "Military Specification August 1969 V/STOL Aircraft", for Flying Qualities of Piloted 13. "Military Specification MIL-F-8785B, December, 1970.